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  • instrument : RADIO SCIENCE SUBSYSTEM for VO1
    INSTRUMENT: RADIO SCIENCE SUBSYSTEM SPACECRAFT: VIKING ORBITER 1 Instrument Information ====================== Instrument Id : RSS Instrument Host Id : VO1 Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Instrument Description ====================== The Radio Science Subsystem (RSS) was used to support the radio science investigation. This investigation utilizes the radio signals to and from the spacecraft to obtain a variety of information. Transponders on the spacecraft send back signals that are coherent with those received from earth so that precision Doppler and ranging measurements can be made. The downlinks from the spacecraft include two coherent frequencies (S band and X band) so that dispersion in the interplanetary medium can be measured and corrected for. Science Objectives ================== From analysis of the radio signals a surprising diversity of informatin can potentially be obtained and the Radio Science Team has identified a large number of scientific objectives, which fall into three categories: dynamical, surface, and internal properties of Mars, atmospheric and ionospheric properties of Mars, and miscellaneuos solar system properties. For more information please refer to the Instrument Information References user view. Operational Considerations ========================== NOT APPLICABLE Calibration Description ======================= NOT APPLICABLE Section 'RSS' ============= Total Fovs : UNK Data Rate : UNK Scan Mode Id : UNK Sample Bits : UNK 'RSS' Detectors --------------- RSSDETEB RSSDETSC 'RSS' Electronics ----------------- RSSELECEB RSSELECSC In modes -------- OPERATING 'RSS' Section FOV Shape 'UNK' ----------------------------- Section Id : RSS Fovs : UNK Horizontal Pixel Fov : UNK Vertical Pixel Fov : UNK Horizontal Fov : UNK Vertical Fov : UNK 'RSS' Section Parameter 'RSSDETEB POWER' ---------------------------------------- RSSDETEB Power is the power measured by the Deep Space Network antenna from the spacecraft antenna for tracking purposes. Instrument Parameter Name : RSSDETEB POWER Sampling Parameter Name : TIME Instrument Parameter Unit : WATTS Minimum Instrument Parameter : UNK Maximum Instrument Parameter : UNK Minimum Sampling Parameter : UNK Maximum Sampling Parameter : UNK Noise Level : UNK Sampling Parameter Interval : UNK Sampling Parameter Resolution : UNK Sampling Parameter Unit : SECOND Instrument Detector 'RSSDETEB' ============================== Detector Type : ANTENNA Detector Aspect Ratio : UNK Minimum Wavelength : UNK Maximum Wavelength : UNK Nominal Operating Temperature : UNK Description ----------- One of the three Deep Space Station Antennae: Goldstone, California; Madrid, Spain; or Canberra, Australia. Sensitivity ----------- UNKNOWN Instrument Detector 'RSSDETSC' ============================== Detector Type : ANTENNA Detector Aspect Ratio : UNK Minimum Wavelength : UNK Maximum Wavelength : UNK Nominal Operating Temperature : UNK Description ----------- See instrument description for information on the spacecraft antenna. Sensitivity ----------- UNKNOWN Instrument Electronics 'RSSELECEB' ================================== Description ----------- See the instrument description for information on the earth-based electronics. Instrument Electronics 'RSSELECSC' ================================== Description ----------- See the instrument description for information on the spacecraft electronics. Instrument Mode 'OPERATING' =========================== Data Path Type : UNK Gain Mode Id : UNK Instrument Power Consumption : UNK In sections ----------- RSS Description ----------- UNKNOWN Mounted On Platform 'UNK' ========================= Cone Offset Angle : UNK Cross Cone Offset Angle : UNK Twist Offset Angle : UNK Description ----------- UNKNOWN
  • instrument : Radio Science Subsystem for Mariner 9
    Instrument Overview =================== The Mariner 9 (MR9) Radio Science investigations utilized instrumentation with elements on the spacecraft and at the NASA Deep Space Network (DSN). Much of this was shared equipment, being used for routine telecommunications as well as for Radio Science. The performance and calibration of both the spacecraft and tracking stations directly affected the radio science data accuracy, and they played a major role in determining the quality of the results. The spacecraft part of the radio science instrument is described immediately below; that is followed by a description of the DSN (ground) part of the instrument. Instrument Specifications - Spacecraft ====================================== The Mariner 9 spacecraft telecommunications subsystem served as part of a radio science subsystem for investigations of Mars. Many details of the subsystem are unknown; its 'launch date' is known to be 1971-05-30. Instrument Id : RSS Instrument Host Id : MR9 Pi Pds User Id : UNK Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : UNK Instrument Mass : UNK Instrument Length : UNK Instrument Width : UNK Instrument Height : UNK Instrument Manufacturer Name : UNK Instrument Overview - Spacecraft ================================ The Doppler shift of the S-band telemetry signal during occultation of the spacecraft by Mars provided the vertical distribution of the index of refraction of the Martian atmosphere. These data yield the vertical distribution of neutral and ionized species.
  • instrument : PLASMA WAVE RECEIVER for VG2
    INSTRUMENT: PLASMA WAVE RECEIVER SPACECRAFT: VOYAGER 2 Instrument Overview =================== Instrument Id : PWS Instrument Host Id : VG2 Principal Investigator : DONALD A. GURNETT PI PDS User Id : DGURNETT Instrument Name : PLASMA WAVE RECEIVER Instrument Type : PLASMA WAVE SPECTROMETER Build Date : 1976-11-28 Instrument Mass : 1.400000 Instrument Length : 0.318000 Instrument Width : 0.185000 Instrument Height : 0.048000 Instrument Serial Number : SN003 Instrument Manufacturer Name : THE UNIVERSITY OF IOWA The Plasma Wave Receiver on Voyager consists of both a 16-channel spectrum analyzer covering the range of 10 Hertz to 56.2 kiloHertz and a wideband waveform receiver which returns the waveform of waves in the frequency range of 40 Hertz to 12 kiloHertz. The spectrum analyzer provides data on a continual basis with a maximum temporal resolution of one spectrum per 4 seconds. The waveform receiver returns 4-bit samples of the electric field measured at a rate of 28,800 samples per second. Because of the very high data rate, the waveform samples must be transmitted in the same manner as the Voyager imaging information. At Jupiter, some 10,000 48-second waveform frames were obtained. At Saturn, Uranus, and Neptune, the number of frames obtained was very small due to the lower telemetry rates available at the greater distances of those planets. Science Objectives ================== The primary science objective of the Voyager plasma wave investigation is to make the first surveys of the plasma wave and low frequency radio wave spectra in the magnetospheres of the outer planets: Jupiter, Saturn, Uranus, and Neptune. Plasma waves participate in a fundamental manner in the dynamics of planetary magnetospheres and in the interactions of that magnetosphere with the external solar wind and internal perturbations such as those induced by satellites interior to the magnetosphere. Plasma waves also provide diagnostic information about the plasma environment near the planets including such parameters as electron density and sometimes temperature. The instrument is also sensitive to low frequency radio emissions and, therefore, acts as a low frequency extension to the Planetary Radio Astronomy investigation. Radio waves are often the only means of remotely observing regions of plasma not accessible to the spacecraft and also lead to remote diagnostics of plasma conditions. The plasma wave receivers are also sensitive to the results of small dust particles impacting on various parts of the spacecraft at high velocities and, hence, provide a direct measure of the rate of impact, the density of the dust, and an estimate of the mass distribution of dust in the vicinity of the large planets, especially those with rings and otherwise dusty environments. Finally, the Plasma Wave Receiver will characterize the plasma wave and radio wave spectrum of the outer heliosphere and perhaps beyond, extending our understanding of solar wind plasma processes and wave-particle interactions to several tens of Astronomical Units. Operational Considerations ========================== The primary operational considerations of the PWS include maintaining the proper operating mode and obtaining waveform samples as often as the spacecraft tape recorder/downlink capabilities allow. The standard instrument mode is with Waveform Power On and Input Gain State Hi. For encounter periods, this corresponds to GS3GAINHI/WFMPWRON. Since there has never been a period when the signal levels were so high as to require the Low input gain state, and it is highly unlikely that such levels will ever be encountered, Low Input Gain State should never be selected. As long as there is power margin available, it is most straightforward to leave the Waveform Receiver Power on. The power consumption is less than 0.5 Watt for this section, hence, the power savings afforded by turning it off is not large. The most involved operational consideration is providing for the transmission of waveform data to the ground. At Jupiter, the majority of the waveform data could be sent directly to the ground via the 115200 bps downlink. This capability disappeared after Jupiter, however, because of the greater distance to the spacecraft, hence, lower telecon rates. Since operating the A/D converter at a rate less than 28800 Hertz would result in aliasing, it is necessary to record the data at the 115200 bps rate on the spacecraft tape recorder using the appropriate data mode and playback the recorded data at a lower rate, commensurate with the link capabilities. Again, a choice of the proper playback mode is required. Since the data modes available on the spacecraft are highly dependent on mission phase, these modes are not described here. Calibration Description ======================= The Voyager plasma wave receiver spectrum analyzers were calibrated by first establishing a relationship between input voltage (of a sine wave at the filter center frequency) and output voltage and second by measuring the effective bandwidth of the filter. The bandwidth is measured by applying a random noise signal of known spectral density and by measuring the output voltage which, by the first part of the calibration, is related to the rms voltage of a sine wave. Dividing the equivalent sine wave voltage squared by the input spectral density gives a bandwidth. This procedure is repeated for each of the frequency channels. A special calibration problem exists for the upper 8 frequency channels (1 kiloHertz and above) due to a failure of a 'tree switch' in the Flight Data System. An in-flight recalibration was attempted by using a Solar type III radio burst observed by both Voyager 1 and 2. The recalibration has known deficiencies, but it has been impossible to date to improve on them. The deficiencies include 'flat-topped' emissions where the emission appears to grow in amplitude up to some plateau level and then stay artificially flat for long periods of time. The background level for each of the channels can vary in step-level fashion based on a number of engineering parameters which utilize the same failed circuitry. Other results of the tree-switch recalibration is that the instrument sensitivity is decreased by some amount which is not well known and the absolute calibration could be off as well. The calibration validity could be a function of frequency since some channels' (mostly the upper 3 channels, 17.8 kHz and above) calibration has been verified with the PRA, but others have not and seem internally inconsistent with the lower frequency, unaffected channels. 'PWS ANTENNA' Detector ====================== Detector Type : DIPOLE ANTENNA Nominal Operating Temperature : 298.000000 The PWS uses a pair of 10 meter antenna elements as a balanced dipole antenna. The two elements are extended from the spacecraft at right angles to each other. (The elements are shared with the Planetary Radio Astronomy instrument, which uses them as a pair of monopoles so that measurements of the degree of right and left hand circular polarization can be made.) The PWS measures the voltage difference between the two elements which, when coupled with the effective length of the antenna system (7.07 m) yields an electric field strength in units of volt/meter. The antenna system has the usual dipole antenna pattern which yields nearly 4*pi steradians in its field of view, although there is a range of fields of view where the detector response drops dramatically as one expects from a dipole pattern. The PWS antenna, used as a balanced dipole with an effective length of 7.07 meters gives a sensitivity to fluctuating (wave) electric fields down to the range of 5.E-6 volt/meter. Even though the antenna elements are extended orthogonally to each other, the antenna pattern is still a dipole since the elements are short with respect to the wavelengths of the waves. The presence of the various parts of the spacecraft in close proximity to the antenna can result in a distorted pattern, but this has not been studied in the frequency range of the PWS. Electronics =========== The PWS electronics system consists of three basic sections. The first is the power supply system which regulates and filters the 28 volt, 2400 Hertz spacecraft power supply and provides DC voltages to the remainder of the instrument electronics. The second section is the spectrum analyzer which consists of two banks of 8 narrowband filters, and two logarithmic detectors, each of which provides an analog voltage proportional to the log of the signal strength delivered to the detector from any of the eight filters it services. The analog outputs from these two compressors, as they are called, are sent to the Flight Data System of the spacecraft for conversion to an 8-bit digital value. The spacecraft steps the inputs to the two compressors periodically (once per 0.5 seconds in GS3 or encounter mode) so that signal strengths in each of the 16 channels is measured over a 4-second interval. The third section consists of a single broadband filter of 40 Hertz to 12 kiloHertz, an automatic gain controlled amplifier, and a 4-bit A/D converter. This section digitizes the electric field waveform at a 28800 Hertz rate. The output amplitude is controlled by the automatic gain control in order to keep the signals within the useful range provided by the 4-bit digitization. Section 'SA' ------------ Total Fovs : 1 Data Rate : 32.000000 Sample Bits : 8 'SA' Detectors -------------- PWS ANTENNAS 'SA' Section FOV Shape 'DIPOLE' ------------------------------- Section Id : SA Fovs : 1 Horizontal Fov : 360.000000 Vertical Fov : 180.000000 'SA' Section Parameter 'WAVE ELECTRIC FIELD INTENSITY' ------------------------------------------------------ A measured parameter equaling the electric field strength in a specific frequency passband (in MKS unit: VOLTS/METER) measured in a single sensor or antenna. Instrument Parameter Name : WAVE ELECTRIC FIELD INTENSITY Sampling Parameter Name : TIME Instrument Parameter Unit : VOLT/METER Minimum Instrument Parameter : 0.000005 Maximum Instrument Parameter : 0.500000 Noise Level : 0.000005 Sampling Parameter Interval : 4.000000 Sampling Parameter Resolution: 4.000000 Sampling Parameter Unit : SECOND Section 'WFRM' -------------- Total Fovs : 1 Data Rate : 115200.000000 Sample Bits : 4 'WFRM' Detectors ---------------- PWS ANTENNA 'WFRM' Section FOV Shape 'DIPOLE' --------------------------------- Section Id : WFRM Fovs : 1 Horizontal Fov : 360.000000 Vertical Fov : 180.000000 'WFRM' Section Parameter 'ELECTRIC FIELD COMPONENT' --------------------------------------------------- A measured parameter equaling the electric field strength (e.g. in milli-Volts per meter) along a particular axis direction. Instrument Parameter Name : ELECTRIC FIELD COMPONENT Sampling Parameter Name : TIME Instrument Parameter Unit : VOLT/METER Minimum Instrument Parameter : 0.000005 Maximum Instrument Parameter : 0.500000 Noise Level : 0.000005 Sampling Parameter Interval : 0.000035 Sampling Parameter Resolution: 0.000035 Sampling Parameter Unit : SECOND Operating Modes 'GS3GAINHI/WFMPWRON' ==================================== Data Path Type : REALTIME Gain Mode Id : HIGH Instrument Power Consumption : 1.600000 The PWS instrument gain is high and the waveform receiver power is on. This is the normal encounter operating mode of the instrument and places it in its most sensitive input gain state with the waveform receiver section turned on. The fact that the waveform receiver power is on does not guarantee that waveform data is available. The spacecraft is in the GS-3 data mode which cycles the plasma wave spectrum analyzer so that a complete spectrum is obtained every 4 seconds. Mounted On Platform 'SPACECRAFT BUS' ==================================== Cone Offset Angle : UNK Cross Cone Offset Angle : UNK Twist Offset Angle : UNK The PWS is mounted on top of the Planetary Radio Astronomy experiment on top of spacecraft bus bays 8 and 9. The two orthogonal antenna elements are attached to the Planetary radio astronomy package.
  • instrument : PLASMA WAVE RECEIVER for VG1
    INSTRUMENT: PLASMA WAVE RECEIVER SPACECRAFT: VOYAGER 1 Instrument Overview =================== Instrument Id : PWS Instrument Host Id : VG1 Principal Investigator : DONALD A. GURNETT PI PDS User Id : DGURNETT Instrument Name : PLASMA WAVE RECEIVER Instrument Type : PLASMA WAVE SPECTROMETER Build Date : UNK Instrument Mass : 1.400000 Instrument Length : 0.318000 Instrument Width : 0.185000 Instrument Height : 0.048000 Instrument Serial Number : SN002 Instrument Manufacturer Name : THE UNIVERSITY OF IOWA The Plasma Wave Receiver on Voyager consists of both a 16-channel spectrum analyzer covering the range of 10 Hertz to 56.2 kiloHertz and a wideband waveform receiver which returns the waveform of waves in the frequency range of 40 Hertz to 12 kiloHertz. The spectrum analyzer provides data on a continual basis with a maximum temporal resolution of one spectrum per 4 seconds. The waveform receiver returns 4-bit samples of the electric field measured at a rate of 28,800 samples per second. Because of the very high data rate, the waveform samples must be transmitted in the same manner as the Voyager imaging information. At Jupiter, some 10,000 48-second waveform frames were obtained. At Saturn, the number of frames obtained was very small due to the lower telemetry rates available at the greater distance of that planet. Science Objectives ================== The primary science objective of the Voyager plasma wave investigation is to make the first surveys of the plasma wave and low frequency radio wave spectra in the magnetospheres of the outer planets: Jupiter and Saturn. Plasma waves participate in a fundamental manner in the dynamics of planetary magnetospheres and in the interactions of that magnetosphere with the external solar wind and internal perturbations such as those induced by satellites interior to the magnetosphere. Plasma waves also provide diagnostic information about the plasma environment near the planets including such parameters as electron density and sometimes temperature. The instrument is also sensitive to low frequency radio emissions and, therefore, acts as a low frequency extension to the Planetary Radio Astronomy investigation. Radio waves are often the only means of remotely observing regions of plasma not accessible to the spacecraft and also lead to remote diagnostics of plasma conditions. The plasma wave receivers are also sensitive to the results of small dust particles impacting on various parts of the spacecraft at high velocities and, hence, provide a direct measure of the rate of impact, the density of the dust, and an estimate of the mass distribution of dust in the vicinity of the large planets, especially those with rings and otherwise dusty environments. Finally, the Plasma Wave Receiver will characterize the plasma wave and radio wave spectrum of the outer heliosphere and perhaps beyond, extending our understanding of solar wind plasma processes and wave-particle interactions to several tens of Astronomical Units. Operational Considerations ========================== The primary operational considerations of the PWS include maintaining the proper operating mode and obtaining waveform samples as often as the spacecraft tape recorder/downlink capabilities allow. The standard instrument mode is with Waveform Power On and Input Gain State Hi. For encounter periods, this corresponds to GS3GAINHI/WFMPWRON. Since there has never been a period when the signal levels were so high as to require the Low input gain state, and it is highly unlikely that such levels will ever be encountered, Low Input Gain State should never be selected. As long as there is power margin available, it is most straightforward to leave the Waveform Receiver Power on. The power consumption is less than 0.5 Watt for this section, hence, the power savings afforded by turning it off is not large. The most involved operational consideration is providing for the transmission of waveform data to the ground. At Jupiter, the majority of the waveform data could be sent directly to the ground via the 115200 bps downlink. This capability disappeared after Jupiter, however, because of the greater distance to the spacecraft, hence, lower telecon rates. Since operating the A/D converter at a rate less than 28800 Hertz would result in aliasing, it is necessary to record the data at the 115200 bps rate on the spacecraft tape recorder using the appropriate data mode and playback the recorded data at a lower rate, commensurate with the link capabilities. Again, a choice of the proper playback mode is required. Since the data modes available on the spacecraft are highly dependent on mission phase, these modes are not described here. Calibration Description ======================= The Voyager plasma wave receiver spectrum analyzers were calibrated by first establishing a relationship between input voltage (of a sine wave at the filter center frequency) and output voltage and second by measuring the effective bandwidth of the filter. The bandwidth is measured by applying a random noise signal of known spectral density and by measuring the output voltage which, by the first part of the calibration, is related to the rms voltage of a sine wave. Dividing the equivalent sine wave voltage squared by the input spectral density gives a bandwidth. This procedure is repeated for each of the frequency channels. 'PWS ANTENNA' Detector ====================== Detector Type : DIPOLE ANTENNA Nominal Operating Temperature : 298.000000 The PWS uses a pair of 10 meter antenna elements as a balanced dipole antenna. The two elements are extended from the spacecraft at right angles to each other. (The elements are shared with the Planetary Radio Astronomy instrument, which uses them as a pair of monopoles so that measurements of the degree of right and left hand circular polarization can be made.) The PWS measures the voltage difference between the two elements which, when coupled with the effective length of the antenna system (7.07 m) yields an electric field strength in units of volt/meter. The antenna system has the usual dipole antenna pattern which yields nearly 4*pi steradians in its field of view, although there is a range of fields of view where the detector response drops dramatically as one expects from a dipole pattern. The PWS antenna, used as a balanced dipole with an effective length of 7.07 meters gives a sensitivity to fluctuating (wave) electric fields down to the range of 5.E-6 volt/meter. Even though the antenna elements are extended orthogonally to each other, the antenna pattern is still a dipole since the elements are short with respect to the wavelengths of the waves. The presence of the various parts of the spacecraft in close proximity to the antenna can result in a distorted pattern, but this has not been studied in the frequency range of the PWS. Electronics =========== The PWS electronics system consists of three basic sections. The first is the power supply system which regulates and filters the 28 volt, 2400 Hertz spacecraft power supply and provides DC voltages to the remainder of the instrument electronics. The second section is the spectrum analyzer which consists of two banks of 8 narrowband filters, and two logarithmic detectors, each of which provides an analog voltage proportional to the log of the signal strength delivered to the detector from any of the eight filters it services. The analog outputs from these two compressors, as they are called, are sent to the Flight Data System of the spacecraft for conversion to an 8-bit digital value. The spacecraft steps the inputs to the two compressors periodically (once per 0.5 seconds in GS3 or encounter mode) so that signal strengths in each of the 16 channels is measured over a 4-second interval. The third section consists of a single broadband filter of 40 Hertz to 12 kiloHertz, an automatic gain controlled amplifier, and a 4-bit A/D converter. This section digitizes the electric field waveform at a 28800 Hertz rate. The output amplitude is controlled by the automatic gain control in order to keep the signals within the useful range provided by the 4-bit digitization. Section 'SA' ------------ Total Fovs : 1 Data Rate : 32.000000 Sample Bits : 8 'SA' Detectors -------------- PWS ANTENNAS 'SA' Section FOV Shape 'DIPOLE' ------------------------------- Section Id : SA Fovs : 1 Horizontal Fov : 360.000000 Vertical Fov : 180.000000 'SA' Section Parameter 'WAVE ELECTRIC FIELD INTENSITY' ------------------------------------------------------ A measured parameter equaling the electric field strength in a specific frequency passband (in MKS unit: VOLTS/METER) measured in a single sensor or antenna. Instrument Parameter Name : WAVE ELECTRIC FIELD INTENSITY Sampling Parameter Name : TIME Instrument Parameter Unit : VOLT/METER Minimum Instrument Parameter : 0.000005 Maximum Instrument Parameter : 0.500000 Noise Level : 0.000005 Sampling Parameter Interval : 4.000000 Sampling Parameter Resolution: 4.000000 Sampling Parameter Unit : SECOND Section 'WFRM' -------------- Total Fovs : 1 Data Rate : 115200.000000 Sample Bits : 4 'WFRM' Detectors ---------------- PWS ANTENNA 'WFRM' Section FOV Shape 'DIPOLE' --------------------------------- Section Id : WFRM Fovs : 1 Horizontal Fov : 360.000000 Vertical Fov : 180.000000 'WFRM' Section Parameter 'ELECTRIC FIELD COMPONENT' --------------------------------------------------- A measured parameter equaling the electric field strength (e.g. in milli-Volts per meter) along a particular axis direction. Instrument Parameter Name : ELECTRIC FIELD COMPONENT Sampling Parameter Name : TIME Instrument Parameter Unit : VOLT/METER Minimum Instrument Parameter : 0.000005 Maximum Instrument Parameter : 0.500000 Noise Level : 0.000005 Sampling Parameter Interval : 0.000035 Sampling Parameter Resolution: 0.000035 Sampling Parameter Unit : SECOND Operating Modes 'GS3GAINHI/WFMPWRON' ==================================== Data Path Type : REALTIME Gain Mode Id : HIGH Instrument Power Consumption : 1.600000 The PWS instrument gain is high and the waveform receiver power is on. This is the normal encounter operating mode of the instrument and places it in its most sensitive input gain state with the waveform receiver section turned on. The fact that the waveform receiver power is on does not guarantee that waveform data is available. The spacecraft is in the GS-3 data mode which cycles the plasma wave spectrum analyzer so that a complete spectrum is obtained every 4 seconds. Mounted On Platform 'SPACECRAFT BUS' ==================================== Cone Offset Angle : UNK Cross Cone Offset Angle : UNK Twist Offset Angle : UNK The PWS is mounted on top of the Planetary Radio Astronomy experiment on top of spacecraft bus bays 8 and 9. The two orthogonal antenna elements are attached to the Planetary radio astronomy package.
  • instrument : PLANETARY RADIO ASTRONOMY RECEIVER for VG1
    INSTRUMENT: PLANETARY RADIO ASTRONOMY RECEIVER SPACECRAFT: VOYAGER 1 Instrument Information ====================== Instrument Id : PRA Instrument Host Id : VG1 Instrument Name : PLANETARY RADIO ASTRONOMY RECEIVER Instrument Type : RADIO SPECTROMETER PI Name : JAMES W. WARWICK Build Date : UNK Instrument Mass : 7.700000 Instrument Height : UNK Instrument Length : UNK Instrument Width : UNK Instrument Manufacturer Name : MARTIN MARIETTA Instrument Serial Number : UNK Science Objectives ================== The Planetary Radio Astronomy (PRA) experiments' primary objective was to locate and explain kilometric, hectometric, and decametric radio emissions from the planets; to measure plasma resonances near the giant planets; and to detect lightning on the giant planets. The instrument was also successful at observing solar radio emissions from the perspective of the outer solar system. Radio emissions can be used to determine the rate of rotation of the inner core of a planet, to determine the existence of a magnetic field, and to search for magnetic anomalies. Radio emissions are often the only remote diagnostic for interactions occurring in the portions of magnetospheres through which a spacecraft does not pass. This is particularly true for the inner magnetosphere, which usually goes unsampled. Further information on the instrument and the investigations performed can be found in [WARWICKETAL1997]. Instrument Description ====================== The Voyager Planetary Radio Astronomy (PRA) instrument consisted of two superheterodyne receivers, one for the range from 1326.0 to 1.2 kHz (center frequency) and a second for the range 40.55 to 1.53 MHz (center frequency). The channels are numbered so that the lowest frequency channel (1.2 kHz) is channel number 198. Channels in the range 1 to 128 are referred to as 'high band', and the remaining channels as 'low band'. In low band, the channel spacing is 19.2 kHz. In high band, the channel spacing is 307.2 kHz. The PRA receiver is driven by two orthogonal antennas mounted on the spacecraft body. Each antenna element is made of BeCu hollow tubes 0.5 inches in diameter and is 10 meters in length. By combining the signals from the two antennas in a 90 degree hybrid, the PRA instrument can distinguish between the opposing states, left hand and right hand, of circular polarization of an incoming wave. The Planetary Radio Astronomy (PRA) receivers were calibrated under environmentally-controlled conditions and over the entire frequency and dynamic range of the instruments. This calibration consisted in application of a known narrow-band signal across the inputs and recording the receiver outputs. The laboratory calibrations provided power levels for each data number (DN) and each frequency in terms of known inputs across the antenna terminals of each of the experiment's two monopoles. Calibrations were carried out over a range of receiver temperatures, but in practice the stability of the receiver as a function of temperature and the stability of the temperature of the receiver as a function of mission phase and the status of the overall spacecraft were such that a single calibration for each DN at each frequency could be used. Receiver output levels were quantized. The minimum value for the wave flux density was frequency dependent varying from 5.E-20 W M**-2 Hz**-1 at frequencies below 1.5 MHz to 5.E-19 at frequencies above 1.5 MHz. The maximum wave flux density was typically 50 dB above the minimum value. The instrument noise level also was frequency dependent. It was about 1.E-19 W M**-2 Hz**-1 below 1.5 MHz. The noise at 10 MHz was still about 1E-19 W M**-2 Hz**-1, increased to about 1.E-17 W M**-2 Hz**-1 at 25 MHz, and then decreased to an intermediate value at 40 MHz. The low-band and high-band operation of the receiver differ. In low-band the receiver operated with a sharply tuned filter only 1 kHz broad at the 3 dB points and in high-band, with a 200 kHz filter. The gain of the receivers was designed in such a way that the output increased discontinuously by 23 dB (corresponding to the 200:1 bandwidth ratio) between the lowest frequency of high-band and the highest frequency of low-band. This caused the instrument output to remain constant across the high-band to low-band transition point if its input was broadband noise. If unpolarized radiation fell orthogonally on each monopole, the total unpolarized flux density for signals below about 5 MHz could be roughly estimated to be S = So (10**(m/1000)), where m was the channel reading in millibels and So is So = 1.5E-21 (W/Hz m**2). No reliable method for estimating the flux density exists for frequencies above 5 MHz due to the increasing effect of antenna resonances. Although the PRA instrument had 14 possible operating modes, in practice the mode called POLLO was used more than 95% of the time. In POLLO, the receiver swept through all 198 channels in sequence from the highest frequency to the lowest. At each frequency step, data were produced every 30 msec, consisting of 25 msec of integration and 5 msec of switching and settling time. Thus, a full sweep through all 200 channels took 6 sec (including 60 msec for two status words). Between steps the 90 degree hybrid was switched such that the receiver was sensitive to the alternate sense of circular polarization. This toggling between left hand and right hand polarization itself alternated with each 6 sec receiver sweep. Thus, for a given frequency, a pair of left hand and right hand measurements were 6 sec apart.
  • instrument : PLANETARY RADIO ASTRONOMY RECEIVER for VG2
    INSTRUMENT: PLANETARY RADIO ASTRONOMY RECEIVER SPACECRAFT: VOYAGER 2 Instrument Information ====================== Instrument Id : PRA Instrument Host Id : VG2 Instrument Name : PLANETARY RADIO ASTRONOMY RECEIVER Instrument Type : RADIO SPECTROMETER PI Name : JAMES W. WARWICK Build Date : UNK Instrument Mass : 7.700000 Instrument Height : UNK Instrument Length : UNK Instrument Width : UNK Instrument Manufacturer Name : MARTIN MARIETTA Instrument Serial Number : UNK Science Objectives ================== The Planetary Radio Astronomy (PRA) experiments' primary objective was to locate and explain kilometric, hectometric, and decametric radio emissions from the planets; to measure plasma resonances near the giant planets; and to detect lightning on the giant planets. The instrument was also successful at observing solar radio emissions from the perspective of the outer solar system. Radio emissions can be used to determine the rate of rotation of the inner core of a planet, to determine the existence of a magnetic field, and to search for magnetic anomalies. Radio emissions are often the only remote diagnostic for interactions occurring in the portions of magnetospheres through which a spacecraft does not pass. This is particularly true for the inner magnetosphere, which usually goes unsampled. Further information on the instrument and the investigations performed can be found in [WARWICKETAL1997]. Instrument Description ====================== The Voyager Planetary Radio Astronomy (PRA) instrument consisted of two superheterodyne receivers, one for the range from 1326.0 to 1.2 kHz (center frequency) and a second for the range 40.55 to 1.53 MHz (center frequency). The channels are numbered so that the lowest frequency channel (1.2 kHz) is channel number 198. Channels in the range 1 to 128 are referred to as 'high band', and the remaining channels as 'low band'. In low band, the channel spacing is 19.2 kHz. In high band, the channel spacing is 307.2 kHz. The PRA receiver is driven by two orthogonal antennas mounted on the spacecraft body. Each antenna element is made of BeCu hollow tubes 0.5 inches in diameter and is 10 meters in length. By combining the signals from the two antennas in a 90 degree hybrid, the PRA instrument can distinguish between the opposing states, left hand and right hand, of circular polarization of an incoming wave. The Planetary Radio Astronomy (PRA) receivers were calibrated under environmentally-controlled conditions and over the entire frequency and dynamic range of the instruments. This calibration consisted in application of a known narrow-band signal across the inputs and recording the receiver outputs. The laboratory calibrations provided power levels for each data number (DN) and each frequency in terms of known inputs across the antenna terminals of each of the experiment's two monopoles. Calibrations were carried out over a range of receiver temperatures, but in practice the stability of the receiver as a function of temperature and the stability of the temperature of the receiver as a function of mission phase and the status of the overall spacecraft were such that a single calibration for each DN at each frequency could be used. Receiver output levels were quantized. The minimum value for the wave flux density was frequency dependent varying from 5.E-20 W M**-2 Hz**-1 at frequencies below 1.5 MHz to 5.E-19 at frequencies above 1.5 MHz. The maximum wave flux density was typically 50 dB above the minimum value. The instrument noise level also was frequency dependent. It was about 1.E-19 W M**-2 Hz**-1 below 1.5 MHz. The noise at 10 MHz was still about 1E-19 W M**-2 Hz**-1, increased to about 1.E-17 W M**-2 Hz**-1 at 25 MHz, and then decreased to an intermediate value at 40 MHz. The low-band and high-band operation of the receiver differ. In low-band the receiver operated with a sharply tuned filter only 1 kHz broad at the 3 dB points and in high-band, with a 200 kHz filter. The gain of the receivers was designed in such a way that the output increased discontinuously by 23 dB (corresponding to the 200:1 bandwidth ratio) between the lowest frequency of high-band and the highest frequency of low-band. This caused the instrument output to remain constant across the high-band to low-band transition point if its input was broadband noise. If unpolarized radiation fell orthogonally on each monopole, the total unpolarized flux density for signals below about 5 MHz could be roughly estimated to be S = So (10**(m/1000)), where m was the channel reading in millibels and So is So = 1.5E-21 (W/Hz m**2). No reliable method for estimating the flux density exists for frequencies above 5 MHz due to the increasing effect of antenna resonances. Although the PRA instrument had 14 possible operating modes, in practice the mode called POLLO was used more than 95% of the time. In POLLO, the receiver swept through all 198 channels in sequence from the highest frequency to the lowest. At each frequency step, data were produced every 30 msec, consisting of 25 msec of integration and 5 msec of switching and settling time. Thus, a full sweep through all 200 channels took 6 sec (including 60 msec for two status words). Between steps the 90 degree hybrid was switched such that the receiver was sensitive to the alternate sense of circular polarization. This toggling between left hand and right hand polarization itself alternated with each 6 sec receiver sweep. Thus, for a given frequency, a pair of left hand and right hand measurements were 6 sec apart.
  • instrument host : MARS RECONNAISSANCE ORBITER
    Instrument Host Overview ======================== Mars Reconnaissance Orbiter Spacecraft -------------------------------------- Mars Reconnaissance Orbiter uses a spacecraft design provided by Lockheed Martin Space Systems that is smarter, more reliable, more agile, and more productive than any previous Mars orbiter. It is the first spacecraft designed from the ground up for aerobraking, a rigorous phase of the mission where the orbiter uses the friction of the martian atmosphere to slow down in order to settle into its final orbit around Mars. When fully assembled and fueled, the spacecraft had to weigh less than 2,180 kilograms (4,806 pounds) so that the Atlas V launch vehicle could lift it into the proper orbit. All subsystems and instruments on board (the so-called 'dry mass') weighed less than 1,031 kilograms (2,273 pounds) to allow room for 1,149 kilograms (2,533 pounds) of propellant for trajectory correction maneuvers that kept the spacecraft on target during the cruise to Mars and for burns that helped capture the spacecraft into orbit around Mars. Spacecraft Configurations ------------------------- During its five-year mission, the spacecraft needed to operate in four distinct mission phases. Launch: During launch, the spacecraft had to fit within the nose cone, or payload fairing, of the launch vehicle, so the large parts like the high-gain antenna and the solar arrays were designed to be folded up. As soon as the launch vehicle placed the spacecraft on a course to leave earth orbit for its journey to Mars, it disconnected itself from the spacecraft. Cruise: As soon as the spacecraft was clear of the launch vehicle, the orbiter deployed its solar arrays to begin producing power. The high-gain antenna was also be deployed at this point. The high-gain antenna moved to track the Earth, while the solar panels remained fixed. Mars Orbit Insertion and Aerobraking: The Mars orbit insertion and aerobraking configuration looked very much like the cruise configuration, except that the high-gain antenna was moved to a position that balanced the solar arrays as it flew through the upper atmosphere of Mars. The heaviest part of the spacecraft (the propellant tank) also made the spacecraft very stable. Due to the large area (37.7 square meters or 405.8 square feet) of the spacecraft in this configuration, each pass through the martian atmosphere during aerobraking caused significant slowing, thus reducing the size of the orbit. Friction from the atmosphere had the additional effect of heating up the spacecraft, so components were designed to withstand this heating. The flight team could further control the heating by changing how deeply the spacecraft dipped into the atmosphere on each orbit. Science Operations: During the primary science phase, the orbiter's job was to point its science instruments at Mars to collect images and other data from targets on the surface of Mars, while ensuring that the high-gain antenna and solar arrays were continuously tracking the Earth and the Sun, respectively. The orbiter typically kept its science instruments pointed to nadir (looking straight down at the surface). A few times per day, and for about fifteen minutes each time, the orbiter pointed side- to-side in order to capture high-priority science targets that did not fall directly beneath the spacecraft. The spacecraft could point off-nadir up to 30 degrees. Major Spacecraft Components --------------------------- Science Payload Instruments: To fulfill the mission science objectives, seven scientific investigations teams were selected by NASA. Four teams (MARCI, MCS, HiRISE, and CRISM) were led by Principal Investigators (PI). Each PI lead team was responsible for the provision and operation of a scientific instrument and the analysis of its data. The PI lead investigations were: Mars Color Imager (MARCI); Mars Climate Sounder, (MCS); High Resolution Imaging Science Experiment, (HiRISE); and Compact Reconnaissance Imaging Spectrometer for Mars, (CRISM). In addition to the PI lead teams, there were two investigation teams that made use of facility instruments. The facility instruments were Context Imager (CTX) and Shallow (Subsurface) Radar (SHARAD). The MARCI PI and Science Team also acted as Team Leader (TL) and Team Members for the CTX facility instrument. The Italian Space Agency (ASI) provided a second facility instrument, SHARAD, for flight on MRO. ASI and NASA both selected members of the SHARAD investigation team with ASI appointing the Team Leader and NASA appointing the Deputy Team Leader. In addition to the instrument investigations, Gravity Science and Atmospheric Structure Facility Investigation Teams used data from the spacecraft telecommunications and accelerometers, respectively, to conduct scientific investigations. The science instruments are summarized below. Instrument: CRISM (Compact Reconnaissance Imaging Spectrometer for Mars) Type: High-Resolution Imaging Spectrometer Measurements: Hyper-spectral Image Cubes, 514 spectral bands, 0.4-4 microns, 7 nm resolution, from 300km; 20 m/pixel, 11 km swath. Science Goals: Regional & local surface composition and morphology. Key Attributes: Moderately high spectral & spatial resolution, targeted and regional survey, very high data rate. Principal Investigator: Scott Murchie, Johns Hopkins University Applied Physics Lab. Instrument: CTX (Context Imager) Type: Mono-chromatic Context Camera Measurements: Panchromatic (minus blue)Images from 300km altitude; 30km swath & 6m/pixel context imaging for HiRISE/CRISM & MRO science. Science Goals: Regional stratigraphy and morphology. Key Attributes: Moderately high resolution with coverage, targeted & regional survey; high data rate. Team Leader: Michael Malin, Malin Space Science Systems (MSSS). Instrument: HiRISE (High Resolution Imaging Science Experiment) Type: High-Resolution Camera (0.5 m aperture) Measurements: Color images, stereo by site revisit, from 300km; < 1m/pixel (ground sampling @ 0.3 m/pixel), 6 km swath in red (broadband), 1.2km swath in blue-green & NIR. Science Goals: Stratigraphy, geologic processes and morphology. Key Attributes: Very high resolution targeted imaging, very high data rate. Principal Investigator: Alfred McEwen, University of Arizona. Instrument: MARCI (Mars Color Imager) Type: Wide-Angle Color Imager Measurements: Coverage of atmospheric clouds, hazes and ozone, and surface albedo in 7 color bands (0.28-0.8 micrometers) (2 UV, 5 Visible). Science Goals: Global weather and surface change. Key Attributes: Daily global coverage daily global mapping, continuous operations dayside; moderate data rate. Principal Investigator: Michael Malin, Malin Space Science Systems (MSSS). Instrument: MCS Type: Atmospheric Sounder Measurements: Atmospheric profiles of water, dust, co2 & temperature, polar radiation balance, 0-80km vertical coverage, vertical resolution ~5km. Science Goals: Atmospheric structure, transport and polar processes. Key Attributes: Global limb sounding; daily, global limb & on-planet mapping; continuous operations day and night; low data rate. Principal Investigator: Daniel J. McCleese, Jet Propulsion Lab (JPL). Instrument: SHARAD Type: Shallow Subsurface Radar Measurements: Ground penetrating radar; transmit split band at 20mhz <1km; 10-20 m vertical resolution 1km x 5km. Science Goals: Regional near-surface ground structure. Key Attributes: Shallow sounding; regional profiling; high data rate. Team Leader: Roberto Seu, University of Rome, Italy. Deputy Team Leader: Roger Phillips, Washington University, St. Louis. Engineering Instruments: Mars Reconnaissance Orbiter carried three instruments that will assist in spacecraft navigation and communications. 1. Electra UHF Communications and Navigation Package: Electra allowed the spacecraft to act as a communications relay between the Earth and landed crafts on Mars that may not have sufficient radio power to communicate directly with Earth by themselves. 2. Optical Navigation Camera: This camera was being tested for improved navigation capability for future missions. Similar cameras placed on orbiters of the future would be able to serve as high-precision interplanetary 'eyes' to guide incoming spacecraft as they near Mars. 3. Ka-band Telecommunications Experiment Package: Mars Reconnaissance Orbiter tested the use of a radio frequency called Ka-band to demonstrate the potential for greater performance in communications using significantly less power. Structures: The structures subsystem is the skeleton around which the spacecraft was assembled. It supported and protected the other engineering subsystems and the science instruments. It was strong enough to survive launch, when the forces can exceed 5 g's. Extremely lightweight but strong materials were used to achieve this strength, including titanium, carbon composites, and aluminum honeycomb. Mechanisms: There were three main mechanisms on board Mars Reconnaissance Orbiter: * one that allowed the high-gain antenna to move in order to point at earth * two that allowed the solar arrays to move to point at the sun Each of these mechanisms, called gimbals, moved about two axes in much the same way that your wrist allows your hand to move in two axes: left/right and up/down. As the spacecraft traveled around Mars each orbit, these gimbals allowed both solar arrays to be always pointed toward the sun, while the high-gain antenna could simultaneously always be pointed at earth. Telecommunications System: The telecommunications subsystem was a two-way radio system used for receiving and transmitting commands and data between the Mars Reconnaissance Orbiter and the Deep Space Network antenna on earth. With its large-dish antenna, powerful amplifier, and fast computer, Mars Reconnaissance Orbiter could transmit data to earth at rates as high as 6 megabits per second, a rate ten times higher than previous Mars orbiters. The orbiter's radio operated in the X-band of the radio spectrum, at a frequency of around 8 Gigahertz. Major components of the telecom subsystem included: * Antennas for transmitting and receiving commands * Amplifiers for boosting the power of radio signals so that they are strong enough to be received at the Deep Space Network antennas * Transponders for translating navigation and other signals from the orbiter Also on board was Electra, a UHF telecommunications package that was one of the engineering instruments providing navigation and communications support to landers and rovers on the surface of Mars. Electra allowed the spacecraft to act as a relay between the earth and landed crafts on Mars, which may not have sufficient radio power to communicate directly with earth. High-gain Antenna: The high-gain antenna is a 3-meter diameter (10-foot) dish antenna for sending and receiving data at high rates. The high-gain antenna was deployed shortly after launch and remained deployed for the remainder of the mission. It served as the primary means of communication to and from the orbiter. The high-gain antenna had to be pointed accurately and was therefore steered using the gimbal mechanism. Low-gain Antennas: Two smaller antennas were present for lower- rate communication during emergencies and special events, such as launch and Mars Orbit Insertion. The data rate of these antennas was lower because they focused the radio beam much more broadly than the high gain antenna, so less of the signal reached earth. But the Deep Space Network station on the earth could detect the signal even when the spacecraft was not pointed at earth, and therefore these antennas were useful for emergencies. The low-gain antennas had the capability to transmit and receive. The two low- gain antennas were mounted on the high-gain antenna dish--one on the front side and one on the back--and were moved with it. Two were needed in that placement so that communication was possible at all times, no matter what the position of the spacecraft might be at a given time. Amplifiers: Located on the backside of the high-gain antenna was the enclosure for the Traveling Wave Tube Amplifiers, which boosted the power of radio signals so that they were strong enough to be received at the Deep Space Network antennas. There were three amplifiers on board: * two for the X-band radio frequency that transmitted radio signals at a power of 100 watts (the second one was for back up to ensure communications in case the first amplifier failed) * one for Ka-band radio frequency that was capable of transmitting at 35 watts. Transponders: Mars Reconnaissance Orbiter carried two transponders, which are special types of radio receiver/transmitters, specially designed for long-range space communications. The second transponder was a backup. The transponders had several functions: * transmit/receive function: translated digital electrical signals into radio signals for sending data to earth, and translated radio signals to digital electrical signals for receiving commands from earth * transponding function: listened for and detected a signal coming in from earth, to which it automatically responded * navigation function: transmitted several types of signals that provided critical navigation clues, enabling navigators on the ground to make precise calculations of the spacecraft speed and distance from earth Propulsion: The propulsion subsystem performed major maneuvers such as trajectory correction maneuvers and Mars orbit insertion. The propulsion subsystem was also used to control the spacecraft's position, as a backup to the reaction wheels. Mars Reconnaissance Orbiter used a monopropellant propulsion system: it carried fuel (hydrazine), but no oxidizer. Thrust was produced by passing the fuel over beds of catalyst material just before it entered the thruster, which caused the hydrazine to combust. The propulsion system included: Propellant Tank The monopropellant hydrazine tank held 1187 kilograms (2617 pounds) of usable propellant. Over 70% of the total propellant was used during Mars orbit insertion. Pressurant Tank Mars Reconnaissance Orbiter fed pressurized helium gas from a separate high- pressure tank, through a regulator, into the propellant tank where it put the hydrazine propellant under pressure. Lines, Valves, and Regulators The pressurized hydrazine flowed through a system of metal tubing to each of the thrusters. Each thruster had a valve so that it could be fired independently. Additional valves in the propellant lines turned on and off the flow to groups of thrusters. Thrusters A total of 20 rocket engine thrusters were onboard: * Six large thrusters, each producing 170 Newtons* (38 pounds force) of thrust for performing the Mars orbit insertion burn. Together, all six produce 1,020 Newtons (230 pounds force) of thrust. * Six medium thrusters, each producing 22 Newtons* (5 pounds force) of thrust for performing trajectory correction maneuvers, and for helping to keep the spacecraft pointing in the right direction during the Mars orbit insertion burn. * Eight small thrusters, each producing 0.9 Newtons* (0.2 pounds force) of thrust for controlling where the orbiter is pointed during normal operations as well as during Mars orbit insertion and trajectory correction maneuvers. Command and Data-Handling Systems: The Command and Data Handling subsystem controlled all spacecraft functions. This system: * managed all forms of data on the spacecraft; * carried out commands sent from earth; * prepared data for transmission to the earth; * managed collection of solar power and charging of the batteries; * collected and processed information about all subsystems and payloads; * kept and distributed the spacecraft time; * calculated its position in orbit around Mars; * carried out commanded maneuvers; and * autonomously monitored and responded to any onboard problems that occurred. The key parts of this system were: * Space Flight Computer (a space-qualified processor based on the 133 MHz PowerPC processor) * Flight Software * Solid State Recorder (total capacity 160 Gigabits) Guidance, Navigation, and Control Systems: The guidance, navigation, and control subsystem was used to control the orientation of the orbiter as it travels through space and to maintain knowledge of where celestial bodies are located (for example, Earth and the sun). This knowledge was critical for the spacecraft to perform the correct maneuvers to get to Mars, to keep its solar arrays pointed toward the sun in order to produce power, and to keep its antenna pointed toward the earth in order to maintain communications. Once in orbit around Mars, this subsystem also maintained constant knowledge of where the spacecraft was in its orbit, and was able to point the science cameras to an accuracy of about 1/20th of one degree. Electrical Power: The electrical power subsystem was responsible for generating, storing, and distributing power to the orbiter systems and included two solar panels and two nickel-hydrogen batteries. Solar panels: The one and only source of power for Mars Reconnaissance Orbiter was sunlight. Mounted on opposite sides of the orbiter and capable of changing position to allow the orbiter to track the sun continuously, each solar panel had an area of approximately 10 square meters (107.6 square feet), and contained 3,744 individual solar cells. The solar cells were able to convert more than 26% of the sun's energy directly into electricity. The solar panels were deployed soon after launch and remained deployed throughout the mission. During aerobraking the solar panels had a special role to play. As the spacecraft skimmed through the upper layers of the martian atmosphere, the large, flat panels acted to slow the spacecraft down and reduce the size of its orbit. The solar arrays were designed to withstand temperatures of almost 200 Celsius. Nickel-hydrogen batteries: Mars Reconnaissance Orbiter used two nickel-hydrogen rechargeable batteries, each with an energy storage capacity of 50 ampere-hours (at 32 volts, 1600 watts per hour). Only about 40% of the battery capacity was intended to be used. The batteries charged during the day side of each two-hour orbit around Mars, using electricity produced by the solar cells, and provided power during the night side of each orbit. Thermal Systems: The thermal subsystem maintained the right temperatures in all parts of the spacecraft. It employed several conduction- and radiation-based techniques for thermal control: * Radiators * Surface coatings * Thermal blankets * Heaters.
  • collection : Dawn Radio Science Tracking and Navigation (TNF) Data Products Collection
    This is the collection of Radio Science Tracking and Navigation (TNF) data products acquired during the Dawn mission. Data were originally stored as chronological records in DSN format TRK-2-34; but they have been sorted according to data type (retaining chronological order within each data type) during migration from PDS3 to PDS4. This is the collection of Radio Science Tracking and Navigation (TNF) data products acquired during the Dawn mission. Data were originally stored as chronological records in DSN format TRK-2-34; but they have been sorted according to data type (retaining chronological order within each data type) during migration from PDS3 to PDS4.
  • collection : MESSENGER Radio Science Tracking and Navigation (TNF) Data Products Collection
    This is the collection of Radio Science Tracking and Navigation (TNF) data products acquired during the MESSENGER mission. Data were originally stored as chronological records in DSN format TRK-2-34; but they have been sorted according to data type (retaining chronological order within each data type) during migration from PDS3 to PDS4. This is the collection of Radio Science Tracking and Navigation (TNF) data products acquired during the MESSENGER mission. Data were originally stored as chronological records in DSN format TRK-2-34; but they have been sorted according to data type (retaining chronological order within each data type) during migration from PDS3 to PDS4.
  • instrument : Radio Science Subsystem for Voyager 1
    Instrument Overview =================== Voyager Radio Science investigations at the giant planets utilized instrumentation with elements both on the spacecraft and at the DSN. Much of this was shared equipment, being used for routine telecommunications as well as for Radio Science. The performance and calibration of both the spacecraft and tracking stations directly affected the radio science data accuracy, and they played a major role in determining the quality of the results. The spacecraft part of the radio science instrument is described immediately below; in most cases, the description applies equally well to both Voyager 1 and Voyager 2 and it applies throughout the Voyager mission. The description of the DSN (ground) part of the instrument may be found in DSN context products. Instrument Specifications ========================= The Voyager spacecraft telecommunications subsystem served as part of a radio science subsystem for investigations of the giant planets. Many details of the subsystem are unknown; its 'build date' is taken to be 1977-09-05, the launch date for Voyager 1. Instrument Id : RSS-VG1S Instrument Host Id : VG1 Pi Pds User Id : UNK Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : 1977-09-05 Instrument Mass : UNK Instrument Length : UNK Instrument Width : UNK Instrument Height : UNK Instrument Manufacturer Name : UNK Instrument Overview =================== The spacecraft radio system was constructed around a redundant pair of transponders. Each transponder was equipped with an S-band receiver (2115 MHz nominal frequency) and transmitters at both S-band (2295 MHz nominal) and X-band (8415 MHz nominal). Compared with S-band, X-band was less sensitive to plasma effects by a factor of about 10; use of both frequencies coherently on the 'downlink' allowed estimation of plasma content along the radio path. Use of X-band also significantly improved the quality of radio tracking data for gravity investigations. The transponder generated downlink signals in either 'coherent' or 'non-coherent' modes, also known as 'two-way' and 'one-way', respectively. When operating in the coherent mode, the transmitted carrier frequency was derived coherently from the received uplink carrier frequency with a 'turn-around ratio' of 240/221 at S-band and (11/3)*240/221 at X-band. In non-coherent mode the transmitted frequency was controlled by an on-board oscillator; the X- and S-band remained coherent in the ratio 11/3. A single Ultra-Stable Oscillator (USO) was used during radio occultations; it provided stabilities several orders of magnitude better than the conventional crystal oscillators, which were part of each transponder. Stability of the Voyager USO was specified in terms of its Allan Deviation -- the fractional frequency deviation from linear drift. Over 10 minute periods, the Allan Deviation ranged from 10**-12 to 4 10**-12 for integrations of 1-10 sec. Long-term fractional drift of the oscillator was about 5 10**-11 per day. Although the oscillator was hardened, there were discontinuities in the drift when the spacecraft passed through the radiation belts of the outer planets. Equivalent X-band microwave frequencies for the Voyager 1 USO during key events were: 8,414,995,272.530 Hz (Titan occultation) 8,414,995,272.376 Hz (Saturn occultation) Multiplying by 3/11 yields the S-band frequency. Traveling wave tube or solid state amplifiers boosted the transponder output. Output powers of 9 and 26 watts could be selected at S-band; the choices at X-band were 12 and 22 watts. The signals were radiated via a 3.66 m diameter parabolic high gain antenna (HGA). The transmit boresight gain of the HGA was 36 dB at S-band and 47 dB at X-band. The half-power half-width of the antenna beam was 0.32 degrees at X-band and 1.1 degrees at S-band. Transmit polarization was right-hand circular at S-band and either right- or left-hand circular at X-band. A Low-Gain Antenna (LGA) was mounted on the feed structure of the HGA and radiated approximately uniformly over the hemisphere into which the HGA pointed. It was used during maneuvers, spacecraft anomalies, and at other times when the HGA was not appropriate. For receiving, the S-band HGA gain was 35 dB at 2115 MHz and the polarization was right-hand circular. The receiving system noise temperature was approximately 2000 K, the carrier tracking loop bandwidth was 18 Hz, and the ranging channel noise bandwidth was 1.5 MHz. Operational Considerations ========================== Descriptions given here are for nominal performance. The spacecraft transponder system comprised redundant units, each with slightly different characteristics. As the transponder units ages, their performance changed slightly. More importantly, the performance for radio science depended on operational factors such as the modulation state for the transmitters, which cannot be predicted in advance. The performance also depended on factors which were not always under the control of the Voyager Project. Spacecraft receivers were designed to lock to the uplink signal. Without locking, Doppler effects -- resulting from relative motion of the spacecraft and ground station -- could result in loss of the radio link as the frequency of the received signal drifted. A series of failures in the Voyager 2 receivers left that transponder unable to track the uplink signal. Beginning in April 1978, Doppler shifts were predicted and the uplink carrier was tuned so that Voyager 2 would see what appeared to be a signal at constant frequency (to an accuracy of 100 Hz. There were no such problems with Voyager 1. During deep occultations by the giant planets, the bending angle resulting from refraction exceeded 10 degrees in some cases -- well beyond the half power beamwidth of the spacecraft antenna. In those cases, the pointing of the HGA was adjusted so that it followed a 'virtual' Earth and maximum signal strength could be sustained. These 'limb-track' maneuvers were critically dependent on accurate timing during the encounter. To protect against Voyager 1 timing errors at Titan (primarily from uncertainties in the radius and position of the satellite), no limb-track was attempted during ingress, and a fixed antenna offset was used during egress. Fortunately, timing was accurate enough that useful data were obtained from each event. Although the spacecraft radioisotope thermoelectric generators were not dependent on solar flux for power, their output decayed as the Voyager spacecraft moved outward through the solar system. During encounters with the outer planets, caution was required in budgeting power and the high-power mode could not be used for the radio transmitters. Calibration Description ======================= Prior to and during some encounter sequences, the spacecraft was commanded to execute a 'mini-ASCAL' maneuver. The HGA was moved slightly above the Earth line then slightly below the Earth line. The procedure was repeated to the left and right of the Earth line so that a 'cross-hair' pattern was mapped out. During the maneuver, the amplitude of the carrier signal was measured carefully. Analysis of the results showed whether the HGA was pointed accurately and, if not, allowed estimation of the error magnitude and direction. Prior to and after encounters, the spacecraft frequency reference was switched to the USO for several hours and the carrier signal was monitored using equipment at the DSN. These 'USO Tests' were used to calibrate the frequency and frequency drift of the USO. USO tests were particularly important before and after the spacecraft entered a severe radiation environment since the radiation typically damaged the crystal and changed its characteristics slightly. Platform Mounting Descriptions ============================== The centerline of the bus was the roll axis of the spacecraft; it also served as the z-axis of the spacecraft coordinate system with the high-gain antenna (HGA) boresight defining the negative z-direction. The HGA boresight was also defined as cone angle 0 degrees and as azimuth 180 degrees, elevation 7 degrees. The Low-Gain Antenna (LGA) was mounted on the feed structure of the HGA and radiated approximately uniformly over the hemisphere into which the HGA pointed. Instrument Section / Operating Mode Descriptions ================================================ The Voyager radio system consisted of two sections, which could be operated in the following modes: Section Mode ------------------------------------------- Oscillator two-way (coherent) one-way (non-coherent) RF output low-gain antenna high-gain antenna Selected parameters describing NASA Standard Transponder (NST) performance are listed below: Oscillator Parameters: S-Band X-Band Two-Way Transponder Turnaround Ratio 240/221 880/221 One-Way Transmit Frequency (MHz) 2296. 8415. Nominal Wavelength (cm) 13.06 3.56 RF Output parameters: S-Band X-Band RF Power Output (w) 9 or 26 12 or 22 Low-Gain Antenna: Half-Power Half Beamwidth (deg) UNK Gain (dBi) UNK EIRP (dBm) UNK Polarization Circular High-Gain Antenna: Half-Power Half-Beamwidth (deg) 1.1 0.32 Gain (dBi) 36 47 Polarization RCP RCP or LCP ACRONYMS AND ABBREVIATIONS ========================== dB decibel dBi dB relative to isotropic dBm dB relative to one milliwatt deg degree DSN Deep Space Network JPL Jet Propulsion Laboratory K Kelvin LCP Left-Circularly Polarized MHz Megahertz RCP Right-Circularly Polarized RF Radio Frequency S-band approximately 2100-2300 MHz sec second SLE Signal Level Estimator UNK unknown X-band approximately 7800-8500 MHz
  • data set : GRAIL MOON LGRS DERIVED GRAVITY SCIENCE DATA PRODUCTS V1.0
    Derived science data originating from the Lunar Gravity Ranging System (LGRS) on each of the two spacecraft comprising the Gravity Recovery and Interior Laboratory (GRAIL) mission.
  • data set : DAWN VESTA GRAVITY SCIENCE DERIVED SCIENCE DATA V1.0
    Dawn Vesta reduced gravity data.
  • instrument : LIGHTNING AND RADIO EMISSION DETECTOR for GP
    Instrument Overview =================== The Lightning and Radio Emission Detector (LRD) instrument will be carried by the Galileo Probe into Jupiter's atmosphere. The LRD will verify the existence of lightning in the atmosphere and will determine the details of many of its basic characteristics. The instrument, operated in its magnetospheric mode at distances of about 5, 4, 3, and 2 planetary radii from Jupiter's center, will also measure the radio frequency (RF) noise spectrum in Jupiter's magnetosphere. The LRD instrument is composed of a ferrite-core radio frequency antenna (~100 Hz to ~100 kHz) and two photodiodes mounted behind individual fisheye lenses. The output of the RF antenna is analyzed both separately and in coincidence with the optical signals from the photodiodes. The RF antenna provides data both in the frequency domain (with three narrow-band channels, primarily for deducing the physical properties of distant lightning) and in the time domain with a priority scheme (primarily for determining from individual RF waveforms the physical properties of closeby-lightning). The LRD instrument has been designed to take into account large uncertainties in the nature of possible Jovian lightning. For example, since Jupiter has no well-defined surface close to the cloud system, there will be no cloud-to-ground discharges, which are the best understood type of lightning on Earth. Lightning in general, and cloud discharges specifically, are very complex physical phenomena and can generate a large variety of RF pulse types and trains: unipolar pulses, bipolar pulses, asymmetric pulses, groups and bursts of pulses. The LRD instrument is designed as a compact and versatile instrument which allows a characterization of these signals with maximum sensitivity and maximum dynamic range. During the design phase, prototype instruments have been intensively tested with Earth lightning during several measuring campaigns. The final instrument characteristics have been set with acceptable margins for the unknown conditions to Jupiter. Modelling of the propagation of RF signals in the frequency range of the LRD instrument in Jupiter's atmosphere shows that direct propagation of signals will occur to distances of order 10^4 km (Rinnert et al., 1979). Hence, it is likely that Jovian atmospheric discharges with the energy of a typical cloud-to-ground discharge on Earth (order 10^8 J) will be detected at 10^4 km or more distance within the atmosphere with the LRD instrument. As noted below, the LRD instrument also includes a 'superbolt' channel, in order to count extremely large events. Hence, in light of all the above, the flown lightning detector instrument must be designed to be as sensitive as possible, limited only by spacecraft noise. The instrument must also cover as large a dynamic range as possible. Principal Investigator ====================== The Principal Investigator for the LRD instrument was Louis Lanzerotti. Scientific Objectives ===================== Radio frequency measurements are made in a reduced mode of operation at altitudes of ~5, 4, 3, 2 planetary radii from the center of Jupiter. These data are stored in the Probe memory and then read out during the atmospheric descent phase of the mission. During the atmospheric descent, the full complement of LRD data are acquired until the loss of the Probe signal by the over-flying Orbiter and/or the demise of the Probe due to atmospheric pressure and heat. The RF data obtained in the magnetosphere will be analyzed also jointly with the Probe Energetic Particle Instrument (EPI) data to gain understanding of magnetospheric particle dynamics. In the magnetosphere, statistics on the characteristics of individual waveforms measured during a sampling interval will be accumulated at the four different altitudes. In addition, noise levels at three different spectral frequencies (3, 5, 90 kHz) will be determined during the measurement intervals. In the atmosphere mode, in addition to statistics on the waveforms and the spectral noise levels at the three narrow-banded frequencies, individual waveforms will be detected, saved, and transmitted to Earth. Such waveforms will provide powerful additional diagnostic capabilities for Jovian RF signals. The LRD instrument, as noted above, has been designed to be as sensitive as possible, limited only by the spacecraft noise, and to be as versatile as possible, limited only by the imposed limitations on power, bit rate, and reliability considerations. It is within these constraints that the scientific objectives will attempt to be achieved. Extensive measurements with Earth lightning have been made and these will be continued in order to gain the maximum understanding of the operational characteristics of the instrument, and therefore the maximum science from the Probe descent through Jupiter's atmosphere. In both the magnetosphere and atmosphere modes the component of the Jovian magnetic field perpendicular to the Probe spin axis will be determined. These data will be used for analyses of EPI data and for determining the spatial distribution of the sources of some of the detected lightning signals. Further, these data will give engineering data on the Probe spin rate. Calibration =========== Because of severe constraints as to weight and power for Probe subsystems, the LRD instrument is very compact. Further, extensive on-board compression of the data is necessary because of the limited available data rate. All sensors and instrument characteristics, of course, have been extensively tested and calibrated. For example, radiation tests were carried out on the sensor electronics and pressure tests were made of the vented electronics box. These latter tests caused a stiffening piece to be added to the microprocessor chip. The calibrations could be verified over long periods because of the delays of the launch of the Galileo spacecraft. A further verification of the instrument parameters is provided by the on-board implemented test generator (ITG). Operational Considerations ========================== Magnetosphere mode. The LRD instrument will operate in the pre-entry phase at distances from the planet's center of about 5, 4, 3, and 2 Rj. The instrument is switched on by the Probe timer at these locations. In this 'magnetosphere mode' the EPI is also in operation. As the Probe is still encapsulated within the heat shield, the MS is less sensitive and the optical sensors are covered. The outputs of the LRD instrument are as noted in the previous section, but without the waveform snapshots. The magnetosphere mode data set at each of the four locations consists of a 64 byte data frame with statistics and the 3 kHz spectral channel subdivided into parallel and perpendicular (to the magnetic field) channels. The data are stored in the Probe memory for transmission during the atmospheric descent phase of the mission. Atmosphere mode. When the LRD instrument is switched on at descent the instrument begins with a test cycle (IFT) and the first data set contains the test pulse data. After that, the instrument runs continuously until the end of the mission and outputs a complete data set every four major frame periods, 256 s. These data sets contain spectral data (the 15 kHz channel being sectored), waveform statistics data, a 1 ms time interval with a selected waveform, optical data and miscellaneous data such as magnetic field component, spin period, and engineering data. The number of complete data sets achieved during descent depends upon the length of time that the Probe survives and/or the length of time that the Probe relay signal is successfully acquired by the over-flying Orbiter. For example, if the total atmosphere data time is ~48 min, then 10 data sets would be sent back (the 11th would be acquired but there would be no time for transmittal). The 10 data sets would contain one test data set and 9 science data sets.
  • collection : MESSENGER Radio Science Delta Differential One-way Ranging (DDOR) Data Products Collection
    DDOR files were generated in the TNF (TRK-2-34) format and contain only TNF data type 10 records. The TNF collection in the MESSENGER archive does not contain data type 10 records. This is the collection of Radio Science Delta Differential One-way Ranging (DDOR) data products acquired during the MESSENGER mission.
  • data set : JUNO DERIVED RADIO SCIENCE GRAVITY DATA V1.0
    Juno reduced gravity data.
  • instrument : GIOTTO RADIOSCIENCE EXPERIMENT for GIO
    Instrument Overview =================== The Giotto spacecraft telecommunications subsystem served as one element of a radio science experiment for investigations of comet Halley. The second element was a set of ground antennas and associated electronics, most of which were in Australia. The spacecraft element of the experiment is specified below. Instrument Id : RSS Instrument Host Id : GIO Instrument Name : GIOTTO RADIOSCIENCE EXPERIMENT Instrument Type : RADIO SCIENCE The Giotto Radio Science Experiment (GRE) utilized instrumentation with elements on the spacecraft and on Earth. Much of this was shared equipment, being used for routine telecommunications as well as for Radio Science. The experiment is described in more detail by [EDENHOFERETAL1986A] and [EDENHOFERETAL1987A]. The spacecraft radio system was constructed around a redundant pair of transponders which received at S-band (2.3 GHz, 13 cm wavelength) and transmitted at both S-band and X-band (8.4 GHz, 3.6 cm wavelength) frequencies. The transmitted frequency during the Halley encounter was controlled by an on-board oscillator; at other times it was controlled by a signal transmitted from the ground. Each transponder included a receiver, command detector, exciter, and low-power amplifier. The transponders provided the usual uplink command and downlink data transmission capabilities. Traveling wave tube amplifiers, driven at saturation, amplified the transponder output before the signals were radiated via a high-gain antenna (HGA). The HGA offset reflector had a diameter of 1.46 m and was despun with respect to the spacecraft body. HGA polarization was right circularly polarized for S-band up/downlink and for X-band downlink. S-band beamwidth was about 5 degrees; X-band beamwidth was about 2 degrees. Ground stations included antennas, associated electronics, and operational systems at two complexes in Australia. The prime system for GRE was the NASA Deep Space Network (DSN) station near Canberra; its 64-m antenna is known as DSS 43. Geodetic coordinates for the 64-m antenna are 148 deg, 58 min, 48 sec E longitude and 35 deg, 24 min, 14 sec S latitude. Raw data from the DSN tracking system included Doppler measurements along the line of sight to the spacecraft. Raw data from the open loop receiver system included digital samples of baseband output. The second system was the 64-m antenna operated by the Australian CSIRO near Parkes; it served as the primary Giotto ground receiving facility for the European Space Agency in addition to its role as a receiving site for the GRE [HALL1986]. Once the radio signal from the spacecraft had been captured by a ground antenna, it was amplified by cryogenic maser amplifiers. Two measurement options were then available: (1) Open-loop Data Acquisition (2) Closed-loop Data Acquisition Open-Loop Data Acquisition -------------------------- Open-loop data acquisition is performed by filtering and then downconverting the received carrier signal to baseband, where it is sampled for subsequent manipulation in digital form. The open-loop receiver is tuned on the basis of frequency predictions that take into account the best estimate of the carrier frequency transmitted by the spacecraft and Doppler corrections based on relative spacecraft-to-ground motion. The quantized samples can then be recorded on magnetic tape. Closed-Loop Data Acquisition ---------------------------- Closed-loop data acquisition is performed with a phase- locked loop receiver; it is usually employed when the spacecraft is operating in its 'coherent' mode. 'Two-way' Doppler shifts are determined by comparing an estimate of the downlink carrier frequency from the phase-locked loop with a reference from the ground station's frequency and timing subsystem. Since the station frequency reference is also used to generate the uplink carrier, the determination can be as accurate as the fundamental station clock -- typically a hydrogen maser frequency standard and, therefore, more stable than the crystal oscillator on board the spacecraft. The Doppler integration time needed to achieve a particular signal-to-noise ratio controls the time interval between successive measurements. Amplitude of the received signal is estimated by sampling the calibrated automatic gain control (AGC) voltage of the phase-locked loop receiver's coherent AGC loop. 'One-way' closed-loop Doppler shifts are determined in much the same way except that the downlink measurement from the phase locked loop must be compared with an estimate of the spacecraft carrier frequency; in this case the accuracy is limited by the relative performance of the reference oscillators on the ground and the spacecraft. Both frequency and amplitude measurements obtained at the DSN station can then be recorded on magnetic tape and/or transmitted electronically to JPL. Files of data specifically intended for navigation and radio science are derived from the raw measurements by the DSN Radio Metric Data Conditioning Team. Operation of the DSN radio science equipment is described in more detail by [ASMAR&HERRERA1993]. The strength of a spacecraft carrier signal, and thus the quality of the radio science data, depends on its modulation state. During the Halley encounter, only the X-band downlink was activated; it was modulated with science data. The link budget for the GRE prime receiver at Canberra, Australia is shown below [EDENHOFERETAL1986B]: S-band S-band X-band uplink downlink downlink ------ -------- -------- Signal frequency (GHz) 2.117 2.299 8.429 TX power (dBm) 73 36.7 43.2 Ground antenna gain (dB) 60.6 61.7 71.9 Propagation loss (dB) -262.5 -263.2 -274.5 S/C antenna gain (dB) 25.3 26.3 39 RF losses (dB) -2.3 -1.9 -1.6 Signal level RX input (dBm) -105.9 -140.4 -122 System noise temperature (dBK) 27.2 13.7 13.8 Boltzmann constant (dBm/HzK) -198.6 -198.6 -198.6 Received S/No (dBHz) 65.5 44.5 62.8 Modulation loss (dB) 0 -3.6 -8.2 PLL bandwidth (dBHz) 15.4 16.8 16.8 Required C/N (dB) 10 15 15 Margin (dB) 40.1 9.1 22.8 Science Objectives ================== The primary objective of the Giotto Radio Science Experiment was determination of the Halley mass fluence resulting from atmospheric drag. A second objective was determining the electron content of the ionosphere of the comet; this part of the experiment was badly degraded when Giotto project management chose to operate only an X-band downlink during the encounter. Mass fluence was determined by measuring the change in spacecraft radial velocity (with respect to receiving stations on the Earth) during the encounter. Velocity changes were inferred from Doppler shifts in the received signal. Doppler shifts not in accord with the expected trajectory could be interpreted as resulting from drag on the spacecraft by dust and gas. Passage of the radio wave through a plasma can also cause Doppler shifts on the received signal. The plasma effects can be readily separated from other effects if measurements are made at two well-separated radio frequencies. The inability to make measurements at S-band meant that only indirect measurements and modeling could be used to infer the density of charged particles in the Halley environment. Calibration Descriptions ======================== Several calibration measurements were carried out; most were directed toward estimating the propagation effects of charged particles along the ray path. Earth Ionosphere Plasma Content ------------------------------- VHF Faraday rotation measurements were made at Canberra using the ETS-2 geostationary satellite. These measurements would provide a measure of the plasma content in the Earth's ionosphere over Canberra. Other Earth Atmospheric Effects ------------------------------- Acquisition of data at two Earth sites (Canberra and Parkes) ensured that local anomalies in propagation conditions over a single antenna (such as changes in the ionosphere) could be calibrated and removed. The results from Canberra and Parkes were very similar [EDENHOFERETAL1987B]. Ray Path Reference ------------------ The Sakigake spacecraft was tracked at S-band by the DSS 42 antenna -- a 34-m antenna that is also part of the Canberra DSN complex. Since Sakigake was in the same part of the sky as Giotto, differences in the properties of the received signals would depend primarily on the propagation conditions in the immediate environment of Halley during the Giotto encounter. Vega Spacecraft Measurements ---------------------------- Vega-1 and Vega-2 used dual-frequency radio methods to obtain estimates of peak electron densities during their encounters with Halley (2500 electrons per cubic centimeter at 8890 km closest approach distance and 1500 el/cc at 8030 km, respectively). Extrapolation of these values to the Giotto flyby conditions gave numbers consistent with simulation [EDENHOFERETAL1986C]. Asymptotic Spacecraft Trajectory Calibration -------------------------------------------- Two-way Doppler and ranging data were acquired both before and after the Halley encounter. These showed that the net change in radial velocity was 23.05 +/- 0.05 cm/sec during the encounter [EDENHOFERETAL1987B]. On-Board Oscillator Drift ------------------------- Measured linear drift of the on-board Giotto oscillator was about +1 Hz per minute [EDENHOFERETAL1987B]. Operational Considerations ========================== Giotto Project management chose to operate only the X-band spacecraft transmitter during the Halley encounter. This made extraction of the electron content in the comet environment extremely difficult, if not impossible [EDENHOFERETAL1987B]. Instrument Section / Operating Mode Descriptions ================================================ The instrument sections and modes for operation during the Halley encounter period are shown below: Section Options Mode ---------------------------------------- Ground Equipment: Antenna 64-m N/A Receiver X-band Open-Loop Closed-Loop Spacecraft Equipment: Antenna High-Gain N/A Transponder Non-Coherent Transmitter X-band N/A The closed-loop and open-loop receivers were operated at their maximum sampling rates (10 Hz and 50 kHz, respectively). During the pre- and post-encounter periods the system was operated in the two-way coherent mode to obtain both Doppler and ranging data. Details on those operations are not available.
  • data set : IHW COMET HALLEY INFRARED FILTER TABLES V1.0
    International Halley Watch Infrared Studies Net observations of 1P/Halley
  • data set : GRAIL MOON RSS RAW DATA V1.0
    Raw radio science data and ancillary files from the Gravity Recovery and Interior Laboratory (GRAIL) mission.
  • instrument : Gravity Science Instrument for Juno
    Instrument Overview =================== The gravity science instrument utilizes the X and Ka-band transponders on-board the Juno spacecraft and Doppler tracking equipment at the Deep Space Network to perform radio science investigations to determine the gravitational field of celestial bodies. The spacecraft part of the radio science instrument is described immediately below; that is followed by a description of the DSN (ground) part of the instrument. For more information about the Juno spacecraft and mission, see [MATOUSEK2006]. Instrument Specifications - Spacecraft ====================================== Instrument Id : RSS Instrument Host Id : JUNO Pi Pds User Id : UNK Instrument Name : GRAVITY SCIENCE INSTRUMENT Instrument Type : RADIO SCIENCE Build Date : UNK Instrument Mass : UNK Instrument Length : UNK Instrument Width : UNK Instrument Height : UNK Instrument Manufacturer Name : UNK Instrument Overview - Spacecraft ================================ The Juno telecommunications system operates at X-band and Ka-band to support the gravity science investigation at Jupiter. The X-band transponder onboard the spacecraft provides the primary communications and telemetry with the ground station. The Ka-band telemetry system is augmented with a Ka-band Translator and downconverter enabling a two-way Ka-band radio science link to the Deep Space Network. The X-band and Ka-band systems can be operated simultaneously for dual X-up/X-down and Ka-up/Ka-down. The ground station uplinks a carrier to the spacecraft which the receiver acquires and tracks. The spacecraft then transmits a signal that is coherent with the uplink signal received. When no uplink signal is present, the downlink signal was referenced to the auxiliary oscillator. Data that are noncoherent contains too much Doppler noise to be useful for gravity science. Science Objectives ================== The radio tracking data are used to improve knowledge of the magnitude and direction of Jupiter's gravity field. The analysis of the interplanetary tracking data (both range data and VLBI) to Juno can be used to improve the modeling of the orbit of Jupiter in future versions of the solar system planetary ephemerides. Gravity Measurements -------------------- Measurement of the gravity field provides significant constraints on inferences about the interior structure of Jupiter. Precise, detailed study of the spacecraft motion in Jovian orbit can yield the mass distribution of the gas giant. Studies of the gravity field emphasize both the global field and local characteristics of the field. The first task is to determine the global field. Doppler and range tracking measurements yield accurate spacecraft trajectory solutions. Simultaneously with reconstruction of the spacecraft orbit, observation equations for field coefficients and a small number of ancillary parameters can be solved. This type of gravity field solution is essential for characterizing tectonic phenomena and can also be used to study localized features. 'Short-arc' line-of-sight Doppler tracking measurements obtained when the Earth-to-spacecraft line-of-sight is within a few degrees of the orbit plane provide the highest resolution of local features. The results from this type of observation typically are presented as contoured acceleration profiles of specific features (e.g., craters, volcanoes, etc.) or line-of-sight acceleration maps of specific regions. Operational Considerations - Spacecraft ======================================= Ka-band measurements are only available when the onboard Ka-band Translator is powered on. Ka-band uplink/downlink is available when the spacecraft is being tracked by the Deep Space Network's DSS-25 in Goldstone, CA because it is the only station in the network with a Ka-band transmitter. During Ka-band tracks not over DSS-25, only non-coherent Ka-band or Ka-band referenced to the X-band uplink is available. During the capture orbit phase, Ka-band checkout passes were conducted to ensure operational status of the Ka-band equipment onboard the spacecraft. Science phase perijoves are all conducted over the DSS-25 antenna in Goldstone, CA. Only X-band data are available for MWR perijoves. Ka-band and X-band data are available for GRAV perijoves. See the MISSION.CAT for details on perijove types and dates. Investigators ============= Folkner, William (Juno Gravity Science Co-I) Asmar, Sami Anderson, John Buccino, Dustin (Juno Gravity Science Instrument Ops) Instrument Overview - DSN ========================= Three Deep Space Communications Complexes (DSCCs) (near Barstow, CA; Canberra, Australia; and Madrid, Spain) comprise the DSN tracking network. Each complex is equipped with several antennas [including at least one each 70-m, 34-m High Efficiency (HEF), and 34-m Beam WaveGuide (BWG)], associated electronics, and operational systems. Primary activity at each complex is radiation of commands to and reception of telemetry data from active spacecraft. Transmission and reception is possible in several radio-frequency bands, the most common being S-band (nominally a frequency of 2100-2300 MHz or a wavelength of 14.2-13.0 cm) and X-band (7100-8500 MHz or 4.2-3.5 cm). Transmitter output powers of up to 400 kW are available. Ground stations have the ability to transmit coded and uncoded waveforms which can be echoed by distant spacecraft. Analysis of the received coding allows navigators to determine the distance to the spacecraft; analysis of Doppler shift on the carrier signal allows estimation of the line-of-sight spacecraft velocity. Range and Doppler measurements are used to calculate the spacecraft trajectory and to infer gravity fields of objects near the spacecraft. Ground stations can record spacecraft signals that have propagated through or been scattered from target media. Measurements of signal parameters after wave interactions with surfaces, atmospheres, rings, and plasmas are used to infer physical and electrical properties of the target. Principal investigators vary from experiment to experiment. See the corresponding section of the spacecraft instrument description or the data set description for specifics. The Deep Space Network is managed by the Jet Propulsion Laboratory of the California Institute of Technology for the U.S. National Aeronautics and Space Administration. Specifications include: Instrument Id : RSS Instrument Host Id : DSN Pi Pds User Id : N/A Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : N/A Instrument Mass : N/A Instrument Length : N/A Instrument Width : N/A Instrument Height : N/A Instrument Manufacturer Name : N/A For more information on the Deep Space Network and its use in radio science see reports by [ASMAR and RENZETTI1993] and [ASMARETAL1995]. For design specifications on DSN subsystems see [DSN810-5]. Subsystems - DSN ================ The Deep Space Communications Complexes (DSCCs) are an integral part of Radio Science instrumentation, along with the spacecraft Radio Frequency Subsystem. Their system performance directly determines the degree of success of Radio Science investigations, and their system calibration determines the degree of accuracy in the results of the experiments. The following paragraphs describe the functions performed by the individual subsystems of a DSCC. This material has been adapted from [ASMARETAL1995]; for additional information, consult [DSN810-5]. Each DSCC includes a set of antennas, a Signal Processing Center (SPC), and communication links to the Jet Propulsion Laboratory (JPL). The general configuration is illustrated below; antennas (Deep Space Stations, or DSS -- a term carried over from earlier times when antennas were individually instrumented) are listed in the table. GOLDSTONE CANBERRA MADRID Antenna SPC 10 SPC 40 SPC 60 -------- --------- -------- -------- 34-m HEF DSS-15 DSS-45 DSS-65 34-m BWG DSS-24 DSS-34 DSS-54 DSS-25 DSS-35 DSS-55 DSS-26 DSS-36 34-m HSB DSS-27 DSS-28 70-m DSS-14 DSS-43 DSS-63 Developmental DSS-13 Subsystem interconnections at each DSCC are shown in the diagram below, and they are described in the sections that follow. The Monitor and Control Subsystem is connected to all other subsystems; the Test Support Subsystem can be. ----------- ------------------ --------- --------- |TRANSMITTER| | | | TRACKING| | COMMAND | | SUBSYSTEM |-| RECEIVER/EXCITER |-|SUBSYSTEM|-|SUBSYSTEM|- ----------- | | --------- --------- | | | SUBSYSTEM | | | | ----------- | | --------------------- | | MICROWAVE | | | | TELEMETRY | | | SUBSYSTEM |-| |-| SUBSYSTEM |- ----------- ------------------ --------------------- | | | ----------- ----------- --------- -------------- | | ANTENNA | | MONITOR | | TEST | | DIGITAL | | | SUBSYSTEM | |AND CONTROL| | SUPPORT | |COMMUNICATIONS|- ----------- | SUBSYSTEM | |SUBSYSTEM| | SUBSYSTEM | ----------- --------- -------------- DSCC Monitor and Control Subsystem ---------------------------------- The DSCC Monitor and Control Subsystem (DMC) is part of the Monitor and Control System (MON) which also includes the ground communications Central Communications Terminal and the Network Operations Control Center (NOCC) Monitor and Control Subsystem. The DMC is the center of activity at a DSCC. The DMC receives and archives most of the information from the NOCC needed by the various DSCC subsystems during their operation. Control of most of the DSCC subsystems, as well as the handling and displaying of any responses to control directives and configuration and status information received from each of the subsystems, is done through the DMC. The effect of this is to centralize the control, display, and archiving functions necessary to operate a DSCC. Communication among the various subsystems is done using a Local Area Network (LAN) hooked up to each subsystem via a network interface unit (NIU). DSCC Antenna Mechanical Subsystem --------------------------------- Multi-mission Radio Science activities require support from the 70-m, 34-m HEF, and 34-m BWG antenna subnets. The antennas at each DSCC function as large-aperture collectors which, by double reflection, cause the incoming radio frequency (RF) energy to enter the feed horns. The large collecting surface of the antenna focuses the incoming energy onto a subreflector, which is adjustable in both axial and angular position. These adjustments are made to correct for gravitational deformation of the antenna as it moves between zenith and the horizon; the deformation can be as large as 5 cm. The subreflector adjustments optimize the channeling of energy from the primary reflector to the subreflector and then to the feed horns. The 70-m and 34-m HEF antennas have 'shaped' primary and secondary reflectors, with forms that are modified paraboloids. This customization allows more uniform illumination of one reflector by another. The BWG reflector shape is ellipsoidal. On the 70-m antennas, the subreflector directs received energy from the antenna onto a dichroic plate, a device which reflects S-band energy to the S-band feed horn and passes X-band energy through to the X-band feed horn. In the 34-m HEF, there is one 'common aperture feed,' which accepts both frequencies without requiring a dichroic plate. In the 34-m BWG, a series of small mirrors (approximately 2.5 meters in diameter) directs microwave energy from the subreflector region to a collection area at the base of the antenna -- typically in a pedestal room. A retractable dichroic reflector separates S- and X-band on some BWG antennas or X- and Ka-band on others. RF energy to be transmitted into space by the horns is focused by the reflectors into narrow cylindrical beams, pointed with high precision (either to the dichroic plate or directly to the subreflector) by a series of drive motors and gear trains that can rotate the movable components and their support structures. The different antennas can be pointed by several means. Two pointing modes commonly used during tracking passes are CONSCAN and 'blind pointing.' With CONSCAN enabled and a closed loop receiver locked to a spacecraft signal, the system tracks the radio source by conically scanning around its position in the sky. Pointing angle adjustments are computed from signal strength information (feedback) supplied by the receiver. In this mode the Antenna Pointing Assembly (APA) generates a circular scan pattern which is sent to the Antenna Control System (ACS). The ACS adds the scan pattern to the corrected pointing angle predicts. Software in the receiver-exciter controller computes the received signal level and sends it to the APA. The correlation of scan position with the received signal level variations allows the APA to compute offset changes which are sent to the ACS. Thus, within the capability of the closed-loop control system, the scan center is pointed precisely at the apparent direction of the spacecraft signal source. An additional function of the APA is to provide antenna position angles and residuals, antenna control mode/status information, and predict-correction parameters to the Area Routing Assembly (ARA) via the LAN, which then sends this information to JPL via the Ground Communications Facility (GCF) for antenna status monitoring. During periods when excessive signal level dynamics or low received signal levels are expected (e.g., during an occultation experiment), CONSCAN should not be used. Under these conditions, blind pointing (CONSCAN OFF) is used, and pointing angle adjustments are based on a predetermined Systematic Error Correction (SEC) model. Independent of CONSCAN state, subreflector motion in at least the z-axis may introduce phase variations into the received Radio Science data. For that reason, during certain experiments, the subreflector in the 70-m and 34-m HEFs may be frozen in the z-axis at a position (often based on elevation angle) selected to minimize phase change and signal degradation. This can be done via Operator Control Inputs (OCIs) from the LMC to the Subreflector Controller (SRC) which resides in the alidade room of the antennas. The SRC passes the commands to motors that drive the subreflector to the desired position. Pointing angles for all antenna types are computed by the NOCC Support System (NSS) from an ephemeris provided by the flight project. These predicts are received and archived by the CMC. Before each track, they are transferred to the APA, which transforms the direction cosines of the predicts into AZ-EL coordinates. The LMC operator then downloads the antenna predict points to the antenna-mounted ACS computer along with a selected SEC model. The pointing predicts consist of time-tagged AZ-EL points at selected time intervals along with polynomial coefficients for interpolation between points. The ACS automatically interpolates the predict points, corrects the pointing predicts for refraction and subreflector position, and adds the proper systematic error correction and any manually entered antenna offsets. The ACS then sends angular position commands for each axis at the rate of one per second. In the 70-m and 34-m HEF, rate commands are generated from the position commands at the servo controller and are subsequently used to steer the antenna. When not using binary predicts (the routine mode for spacecraft tracking), the antennas can be pointed using 'planetary mode' -- a simpler mode which uses right ascension (RA) and declination (DEC) values. These change very slowly with respect to the celestial frame. Values are provided to the station in text form for manual entry. The ACS quadratically interpolates among three RA and DEC points which are on one-day centers. A third pointing mode -- sidereal -- is available for tracking radio sources fixed with respect to the celestial frame. Regardless of the pointing mode being used, a 70-m antenna has a special high-accuracy pointing capability called 'precision' mode. A pointing control loop derives the main AZ-EL pointing servo drive error signals from a two- axis autocollimator mounted on the Intermediate Reference Structure. The autocollimator projects a light beam to a precision mirror mounted on the Master Equatorial drive system, a much smaller structure, independent of the main antenna, which is exactly positioned in HA and DEC with shaft encoders. The autocollimator detects elevation/cross- elevation errors between the two reference surfaces by measuring the angular displacement of the reflected light beam. This error is compensated for in the antenna servo by moving the antenna in the appropriate AZ-EL direction. Pointing accuracies of 0.004 degrees (15 arc seconds) are possible in 'precision' mode. The 'precision' mode is not available on 34-m antennas -- nor is it needed, since their beamwidths are twice as large as on the 70-m antennas. DSCC Antenna Microwave Subsystem -------------------------------- 70-m Antennas: Each 70-m antenna has three feed cones installed in a structure at the center of the main reflector. The feeds are positioned 120 degrees apart on a circle. Selection of the feed is made by rotation of the subreflector. A dichroic mirror assembly, half on the S-band cone and half on the X-band cone, permits simultaneous use of the S- and X-band frequencies. The third cone is devoted to R and D and more specialized work. The Antenna Microwave Subsystem (AMS) accepts the received S- and X-band signals at the feed horn and transmits them through polarizer plates to an orthomode transducer. The polarizer plates are adjusted so that the signals are directed to a pair of redundant amplifiers for each frequency, thus allowing simultaneous reception of signals in two orthogonal polarizations. For S-band these are two Block IVA S-band Traveling Wave Masers (TWMs); for X-band the amplifiers are Block IIA TWMs. 34-m HEF Antennas: The 34-m HEF uses a single feed for both S- and X-band. Simultaneous S- and X-band receive as well as X-band transmit is possible thanks to the presence of an S/X 'combiner' which acts as a diplexer. For S-band, RCP or LCP is user selected through a switch so neither a polarizer nor an orthomode transducer is needed. X-band amplification options include two Block II TWMs or an HEMT Low Noise Amplifier (LNA). S-band amplification is provided by an FET LNA. 34-m BWG Antennas: These antennas use feeds and low-noise amplifiers (LNA) in the pedestal room, which can be switched in and out as needed. Typically the following modes are available: 1. downlink non-diplexed path (RCP or LCP) to LNA-1, with uplink in the opposite circular polarization; 2. downlink non-diplexed path (RCP or LCP) to LNA-2, with uplink in the opposite circular polarization 3. downlink diplexed path (RCP or LCP) to LNA-1, with uplink in the same circular polarization 4. downlink diplexed path (RCP or LCP) to LNA-2, with uplink in the same circular polarization For BWG antennas with dual-band capabilities (e.g., DSS 25) and dual LNAs, each of the above four modes can be used in a single-frequency or dual-frequency configuration. Thus, for antennas with the most complete capabilities, there are sixteen possible ways to receive at a single frequency (2 polarizations, 2 waveguide path choices, 2 LNAs, and 2 bands). DSCC Receiver-Exciter Subsystem ------------------------------- The Receiver-Exciter Subsystem is composed of two groups of equipment: the closed-loop receiver group and the open-loop receiver group. This subsystem is controlled by the Receiver-Exciter Controller (REC) which communicates directly with the DMC for predicts and OCI reception and status reporting. The exciter generates the S-band signal (or X-band for the 34-m HEF only) which is provided to the Transmitter Subsystem for the spacecraft uplink signal. It is tunable under command of the Digitally Controlled Oscillator (DCO) which receives predicts from the Metric Data Assembly (MDA). The diplexer in the signal path between the transmitter and the feed horn for all three antennas (used for simultaneous transmission and reception) may be configured such that it is out of the received signal path (in listen-only or bypass mode) in order to improve the signal-to-noise ratio in the receiver system. Closed Loop Receivers: The Block V receiver-exciter at the 70-m stations allows for two receiver channels, each capable of L-Band (e.g., 1668 MHz frequency or 18 cm wavelength), S-band, or X-band reception, and an S-band exciter for generation of uplink signals through the low-power or high-power transmitter. The closed-loop receivers provide the capability for rapid acquisition of a spacecraft signal and telemetry lockup. In order to accomplish acquisition within a short time, the receivers are predict driven to search for, acquire, and track the downlink automatically. Rapid acquisition precludes manual tuning though that remains as a backup capability. The subsystem utilizes FFT analyzers for rapid acquisition. The predicts are NSS generated, transmitted to the CMC which sends them to the Receiver-Exciter Subsystem where two sets can be stored. The receiver starts acquisition at uplink time plus one round-trip-light-time or at operator specified times. The receivers may also be operated from the LMC without a local operator attending them. The receivers send performance and status data, displays, and event messages to the LMC. Either the exciter synthesizer signal or the simulation (SIM) synthesizer signal is used as the reference for the Doppler extractor in the closed-loop receiver systems, depending on the spacecraft being tracked (and Project guidelines). The SIM synthesizer is not ramped; instead it uses one constant frequency, the Track Synthesizer Frequency (TSF), which is an average frequency for the entire pass. The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. It will be configured such that the expected amplitude changes are accommodated with minimum distortion. The loop bandwidth (2BLo) will be configured such that the expected phase changes can be accommodated while maintaining the best possible loop SNR. Open-Loop Receivers (OLR): The OLR utilized a fixed first Local Oscillator (LO) frequency and a tunable second LO frequency to minimize phase noise and improve frequency stability. The OLR consisted of an RF-to-IF downconverter located at the feed , an IF selection switch (IFS), and a Radio Science Receiver (RSR). The RF-IF downconverters in the 70-m antennas were equipped for four IF channels: S-RCP, S-LCP, X-RCP, and X-LCP. The 34-m HEF stations were equipped with a two-channel RF-IF: S-band and X-band. The IFS switched the IF input among the antennas. DSCC Transmitter Subsystem -------------------------- The Transmitter Subsystem accepts the S-band frequency exciter signal from the Receiver-Exciter Subsystem exciter and amplifies it to the required transmit output level. The amplified signal is routed via the diplexer through the feed horn to the antenna and then focused and beamed to the spacecraft. The Transmitter Subsystem power capabilities range from 18 kW to 400 kW. Power levels above 18 kW are available only at 70-m stations, however, 80 kW transmitters are being installed at the 34-m stations. The Ka-band Transmitter at DSS-25 transmits at 300 W using two combined Traveling Wave Tube Amplifiers (TWTAs). DSCC Tracking Subsystem ----------------------- The Tracking Subsystem primary functions are to acquire and maintain communications with the spacecraft and to generate and format radiometric data containing Doppler and range. The DSCC Tracking Subsystem (DTK) receives the carrier signals and ranging spectra from the Receiver-Exciter Subsystem. The Doppler cycle counts are counted, formatted, and transmitted to JPL in real time. Ranging data are also transmitted to JPL in real time. Also contained in these blocks is the AGC information from the Receiver-Exciter Subsystem. The Radio Metric Data Conditioning Team (RMDCT) at JPL produces a Tracking and Navigation Service File (TNF), which contains Doppler and ranging data. In addition, the Tracking Subsystem receives from the CMC frequency predicts (used to compute frequency residuals and noise estimates), receiver tuning predicts (used to tune the closed-loop receivers), and uplink tuning predicts (used to tune the exciter). From the LMC, it receives configuration and control directives as well as configuration and status information on the transmitter, microwave, and frequency and timing subsystems. The Metric Data Assembly (MDA) controls all of the DTK functions supporting the uplink and downlink activities. The MDA receives uplink predicts and controls the uplink tuning by commanding the DCO. The MDA also controls the Sequential Ranging Assembly (SRA). It formats the Doppler and range measurements and provides them to the GCF for transmission to NOCC. The Sequential Ranging Assembly (SRA) measures the round trip light time (RTLT) of a radio signal traveling from a ground tracking station to a spacecraft and back. From the RTLT, phase, and Doppler data, the spacecraft range can be determined. A coded signal is modulated on an uplink carrier and transmitted to the spacecraft where it is detected and transponded back to the ground station. As a result, the signal received at the tracking station is delayed by its round trip through space and shifted in frequency by the Doppler effect due to the relative motion between the spacecraft and the tracking station on Earth. DSCC Frequency and Timing Subsystem ----------------------------------- The Frequency and Timing Subsystem (FTS) provides all frequency and timing references required by the other DSCC subsystems. It contains four frequency standards of which one is prime and the other three are backups. Selection of the prime standard is done via the CMC. Of these four standards, two are hydrogen masers followed by clean-up loops (CUL) and two are cesium standards. These four standards all feed the Coherent Reference Generator (CRG) which provides the frequency references used by the rest of the complex. It also provides the frequency reference to the Master Clock Assembly (MCA) which in turn provides time to the Time Insertion and Distribution Assembly (TID) which provides UTC and SIM-time to the complex. JPL's ability to monitor the FTS at each DSCC is limited to the MDA calculated Doppler pseudo-residuals, the Doppler noise, the SSI, and to a system which uses the Global Positioning System (GPS). GPS receivers at each DSCC receive a one-pulse-per-second pulse from the station's (hydrogen maser referenced) FTS and a pulse from a GPS satellite at scheduled times. After compensating for the satellite signal delay, the timing offset is reported to JPL where a database is kept. The clock offsets stored in the JPL database are given in microseconds; each entry is a mean reading of measurements from several GPS satellites and a time tag associated with the mean reading. The clock offsets provided include those of SPC 10 relative to UTC (NIST), SPC 40 relative to SPC 10, etc. Detectors - DSN =============== Nominal carrier tracking loop threshold noise bandwidth at X-band is 10 Hz. Coherent (two-way) closed-loop system stability is shown in the table below: integration time Doppler uncertainty (secs) (one sigma, microns/sec) ------ ------------------------ 10 50 60 20 1000 4 For the open-loop subsystem, signal detection is done in software. Calibration - DSN ================= Calibrations of hardware systems are carried out periodically by DSN personnel; these ensure that systems operate at required performance levels -- for example, that antenna patterns, receiver gain, propagation delays, and Doppler uncertainties meet specifications. No information on specific calibration activities is available. Nominal performance specifications are shown in the tables above. Additional information may be available in [DSN810-5]. Prior to each tracking pass, station operators perform a series of calibrations to ensure that systems meet specifications for that operational period. Included in these calibrations is measurement of receiver system temperature in the configuration to be employed during the pass. Results of these calibrations are recorded in (hard copy) Controller's Logs for each pass. The nominal procedure for initializing open-loop receiver attenuator settings is described below. In cases where widely varying signal levels are expected, the procedure may be modified in advance or real-time adjustments may be made to attenuator settings. Operational Considerations - DSN ================================ The DSN is a complex and dynamic 'instrument.' Its performance for Radio Science depends on a number of factors from equipment configuration to meteorological conditions. No specific information on 'operational considerations' can be given here. Operational Modes - DSN ======================= DSCC Antenna Mechanical Subsystem --------------------------------- Pointing of DSCC antennas may be carried out in several ways. For details see the subsection 'DSCC Antenna Mechanical Subsystem' in the 'Subsystem' section. Binary pointing is the preferred mode for tracking spacecraft; pointing predicts are provided, and the antenna simply follows those. With CONSCAN, the antenna scans conically about the optimum pointing direction, using closed-loop receiver signal strength estimates as feedback. In planetary mode, the system interpolates from three (slowly changing) RA-DEC target coordinates; this is 'blind' pointing since there is no feedback from a detected signal. In sidereal mode, the antenna tracks a fixed point on the celestial sphere. In 'precision' mode, the antenna pointing is adjusted using an optical feedback system. It is possible on most antennas to freeze z-axis motion of the subreflector to minimize phase changes in the received signal. DSCC Receiver-Exciter Subsystem ------------------------------- The diplexer in the signal path between the transmitter and the feed horns on all antennas may be configured so that it is out of the received signal path in order to improve the signal-to-noise ratio in the receiver system. This is known as the 'listen-only' or 'bypass' mode. Closed-Loop Receiver AGC Loop ----------------------------- The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. Ordinarily it is configured so that expected signal amplitude changes are accommodated with minimum distortion. The loop bandwidth is ordinarily configured so that expected phase changes can be accommodated while maintaining the best possible loop SNR. Coherent vs. Non-Coherent Operation ----------------------------------- The frequency of the signal transmitted from the spacecraft can generally be controlled in two ways -- by locking to a signal received from a ground station or by locking to an on-board oscillator. These are known as the coherent (or 'two-way') and non-coherent ('one-way') modes, respectively. Mode selection is made at the spacecraft, based on commands received from the ground. When operating in the coherent mode, the transponder carrier frequency is derived from the received uplink carrier frequency with a 'turn-around ratio' typically of 880/749. In the non-coherent mode, the downlink carrier frequency is derived from the spacecraft on-board crystal-controlled oscillator. Either closed-loop or open-loop receivers (or both) can be used with either spacecraft frequency reference mode. Closed-loop reception in two-way mode is usually preferred for routine tracking. Occasionally the spacecraft operates coherently while two ground stations receive the 'downlink' signal; this is sometimes known as the 'three-way' mode. Location - DSN ============== Station locations are documented in [DSN810-5]. Geocentric coordinates are summarized here. Geocentric Geocentric Geocentric Station Radius (km) Latitude (N) Longitude (E) ------------------- ------------ ------------ ------------- Goldstone DSS 13 (34-m R and D) 6372125.096 35.0660180 243.2055410 DSS 14 (70-m) 6371993.267 35.2443523 243.1104618 DSS 15 (34-m HEF) 6371966.511 35.2403129 243.1128049 DSS 24 (34-m BWG) 6371973.601 35.1585346 243.1252056 DSS 25 (34-m BWG) 6371982.537 35.1562591 243.1246368 DSS 26 (34-m BWG) 6371992.264 35.1543409 243.1269835 Canberra DSS 34 (34-m BWG) 6371693.538 -35.2169824 148.9819644 DSS 35 (34-m BWG) 6371697.350 -35.2143052 148.9814558 DSS 43 (70-m) 6371688.998 -35.2209189 148.9812673 DSS 45 (34-m HEF) 6371675.873 -35.2169608 148.9776856 Madrid DSS 54 (34-m BWG) 6370025.490 40.2357726 355.7459032 DSS 55 (34-m BWG) 6370007.988 40.2344478 355.7473667 DSS 63 (70-m) 6370051.198 40.2413554 355.7519915 DSS 65 (34-m HEF) 6370021.709 40.2373555 355.7493011 Measurement Parameters - DSN ============================ Closed-loop data are recorded in Tracking and Navigation Service Files (TNFs), as well as certain other products such as the Orbit Data File (ODF). The TNFs are comprised of SFDUs that have variable-length, variable-format records with mixed typing (i.e., can contain ASCII, integer, and floating-point items in a single record). These files all contain entries that include measurements of Doppler, Range, and signal strength, along with status and uplink frequency information. ACRONYMS AND ABBREVIATIONS - DSN ================================ ACS Antenna Control System ADC Analog-to-Digital Converter AGC Automatic Gain Control AMS Antenna Microwave System APA Antenna Pointing Assembly ARA Area Routing Assembly ATDF Archival Tracking Data File AUX Auxiliary AZ Azimuth BPF Band Pass Filter bps bits per second BWG Beam WaveGuide (antenna) CDU Command Detector Unit CMC Complex Monitor and Control CONSCAN Conical Scanning (antenna pointing mode) CRG Coherent Reference Generator CUL Clean-up Loop DANA a type of frequency synthesizer dB deciBel dBi dB relative to isotropic dBm dB relative to one milliwatt DCO Digitally Controlled Oscillator DDC Digital Down Converter DEC Declination deg degree DIG RSR Digitizer DMC DSCC Monitor and Control Subsystem DOR Differential One-way Ranging DP Data Processor DSCC Deep Space Communications Complex DSN Deep Space Network DSP DSCC Spectrum Processing Subsystem DSS Deep Space Station DTK DSCC Tracking Subsystem E east EIRP Effective Isotropic Radiated Power EL Elevation FET Field Effect Transistor FFT Fast Fourier Transform FIR Finite impulse Response FTS Frequency and Timing Subsystem GCF Ground Communications Facility GHz Gigahertz GPS Global Positioning System HA Hour Angle HEF High-Efficiency (as in 34-m HEF antennas) HEMT High Electron Mobility Transistor (amplifier) HGA High-Gain Antenna HSB High-Speed BWG IF Intermediate Frequency IFS IF Selector Switch IVC IF Selection Switch JPL Jet Propulsion Laboratory K Kelvin Ka-Band approximately 32 GHz KaBLE Ka-Band Link Experiment kbps kilobits per second kHz kilohertz km kilometer kW kilowatt LAN Local Area Network LCP Left-Circularly Polarized LGR Low-Gain Receive (antenna) LGT Low-Gain Transmit (antenna) LMA Lockheed Martin Astronautics LMC Link Monitor and Control LNA Low-Noise Amplifier LO Local Oscillator LPF Low Pass Filter m meters MCA Master Clock Assembly MCCC Mission Control and Computing Center MDA Metric Data Assembly MHz Megahertz MON Monitor and Control System MSA Mission Support Area N north NAR Noise Adding Radiometer NBOC Narrow-Band Occultation Converter NCO Numerically Controlled Oscillator NIST SPC 10 time relative to UTC NIU Network Interface Unit NOCC Network Operations and Control System NRV NOCC Radio Science/VLBI Display Subsystem NSS NOCC Support System OCI Operator Control Input ODF Orbit Data File ODR Original Data Record ODS Original Data Stream OLR Open Loop Receiver OSC Oscillator PDS Planetary Data System POCA Programmable Oscillator Control Assembly PPM Precision Power Monitor RA Right Ascension REC Receiver-Exciter Controller RCP Right-Circularly Polarized RF Radio Frequency RIC RIV Controller RIV Radio Science IF-VF Converter Assembly RMDCT Radio Metric Data Conditioning Team RMS Root Mean Square RSR Radio Science Receiver RSS Radio Science Subsystem RT Real-Time (control computer) RTLT Round-Trip Light Time S-band approximately 2100-2300 MHz sec second SEC System Error Correction SIM Simulation SLE Signal Level Estimator SNR Signal-to-Noise Ratio SNT System Noise Temperature SOE Sequence of Events SPA Spectrum Processing Assembly SPC Signal Processing Center sps samples per second SRA Sequential Ranging Assembly SRC Sub-Reflector Controller SSI Spectral Signal Indicator TID Time Insertion and Distribution Assembly TLM Telemetry TNF Tracking and Navigation File TSF Tracking Synthesizer Frequency TWM Traveling Wave Maser TWNC Two-Way Non-Coherent TWTA Traveling Wave Tube Amplifier UNK unknown USO UltraStable Oscillator UTC Universal Coordinated Time VCO Voltage-Controlled Oscillator VDP VME Data Processor VF Video Frequency X-band approximately 7800-8500 MHz
  • data set : DAWN VESTA GRAVITY SCIENCE DERIVED SCIENCE DATA V2.0
    Dawn Vesta reduced gravity data.
  • data set : IHW COMET HALLEY INFRARED FILTER CURVE MEASUREMENTS V1.0
    International Halley Watch Infrared Studies Net filter curve measurements for 1P/Halley
  • instrument : RADIO SCIENCE SUBSYSTEM for MO
    Instrument Overview =================== The Mars Observer spacecraft telecommunications subsystem was to have served as part of a Radio Science instrument for investigations of Mars. The remainder of the 'instrument' was located at ground stations of the NASA Deep Space Network (DSN). Much of the equipment at both ends was shared, being used for routine telecommunications as well as for Radio Science. Radio data were themselves shared; Doppler and range measurements were used to calculate the spacecraft trajectory and to search for gravity waves. Measurements of signal parameters after waves had propagated through the interplanetary medium were used to infer properties of the medium, characterize small perturbations in spacecraft attitude, and search for gravity waves. Data included in this archive data set are from routine tracking and special tests conducted while Mars Observer was in its Cruise Phase between Earth and Mars. Cruise operations were used to improve operator proficiency in carrying out activities related to radio science observations and to measure the performance of various parts of the radio science and ground data systems. Instrument Overview - Spacecraft ================================ Spacecraft Radio System: The spacecraft radio system was constructed around a redundant pair of Mars Observer Transponders (MOTs) which received at 7.2 GHz and transmitted at 8.4 GHz, both of which are called 'X-band' frequencies. The exact transmitted frequency was controlled by the signal received from a ground station or by an on-board oscillator. The transponders provided the usual uplink command and downlink data transmission capabilities. The MOT also included a ranging 'channel' which could be driven either by signals from the ground for measurement of absolute range or by tones generated on-board to enable measurement of differential range between two ground stations. An MOT generated a downlink signal in either a 'coherent' or a 'non-coherent' mode, also known as the 'two-way' and 'one-way' modes, respectively. When operating in the coherent mode, the MOT behaved as a conventional transponder, its transmitted carrier frequency being derived from the received uplink carrier frequency with a 'turn-around ratio' of 880/749. In the non-coherent mode, the downlink carrier frequency was derived from one of the spacecraft's on-board oscillators. One of the on-board oscillators, the 'ultra-stable oscillator' (USO), had been specially provided to serve as a precision frequency reference for radio occultation experiments. Built by the Applied Physics Laboratory of John Hopkins University, the USO was positioned within the spacecraft so as to minimize its exposure to thermal variations and mechanical vibration. The USO consisted of a quartz crystal resonator contained in a temperature-controlled titanium dewar. Its output frequency was near 19.14 MHz. The X-band carrier for non-coherent downlink transmission was produced by multiplying the USO output frequency by 440. The Mars Observer USO was from a newer generation than those flown on the Voyager and Galileo spacecraft. Its frequency stability, as characterized by the Allan deviation and after accounting for linear drift, was measured to be better than 2E-13 for integration times in the range one to several hundred seconds. This represents more than an order of magnitude improvement in performance over the Voyager and Galileo devices. After a 25 day warm-up period, the USO was expected to have a daily aging rate of 6E-11 and a long term frequency change after five years of significantly less than 1E-6. Laboratory measurements also showed state-of-the-art phase noise and spectral purity as well as amplitude stability performance for a device of this type. During Cruise, the USO drifted 0.2 Hz/day at its X-band output frequency; Allan Deviation for the end-to-end non-coherent system was consistent with the 2E-13 measurements before launch. Each MOT also contained a separate, 'auxiliary' oscillator which could be used to generate the non-coherent downlink carrier. However, the frequency stability of the auxiliary oscillators was about six to seven orders of magnitude poorer than the USO. They were, therefore, not applicable to these radio science investigations. The strength of a spacecraft carrier signal, and thus the quality of the radio occultation data, depends on its modulation state. Mars Observer telemetry data were Manchester encoded and then used to modulate a 1 MHz square wave subcarrier which, in turn, modulated the X-band carrier. Telemetry modulation typically resulted in carrier suppression of either 15.2 dB (simultaneous transmission of science and engineering telemetry) or 9.3 dB (engineering telemetry only). The two-way ranging channel, when activated, suppressed the carrier by about 0.2 dB; the differential range modulation reduced the carrier by another 0.2 dB when in use. Each modulation type -- telemetry, range, and differential range modulation -- could be switched on or off independently. Traveling wave tube amplifiers, driven at saturation, amplified the MOT output before the signals were radiated via (nominally) a steerable 1.5 m diameter parabolic high gain antenna (HGA). Effective isotropic radiated power was about 83 dBm. In contrast to past practice for planetary missions, the HGA was mounted at the end of a 5.6 meter boom so that the phase center of the antenna was about 6 meters from the spacecraft center of mass. Pertinent details of the subsystem are shown below: Instrument Id : RSS Instrument Host Id : MO Pi Pds User Id : LTYLER Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : N/A Instrument Mass : N/A Instrument Length : N/A Instrument Width : N/A Instrument Height : N/A Instrument Manufacturer Name : MOTOROLA Principal Investigator ---------------------- The Team Leader for the Radio Science Team was G. Leonard Tyler of Stanford University. Scientific Objectives ===================== Two different types of radio science experiments were to be conducted with Mars Observer: radio tracking experiments in which the magnitude and direction of the planet's gravity field were derived from the Doppler and ranging measurements, and radio propagation experiments in which modulation on the signal received on Earth could be attributed to properties of an occulting medium, such as the atmosphere of Mars. Gravity Measurements -------------------- Measurement of the gravity field provides significant constraints on inferences about interior structure of Mars. Precise, detailed study of the spacecraft motion in Mars orbit would have yielded the mass distribution of the planet. Topographic data obtained by the Mars Observer Laser Altimeter (MOLA) would have formed a critical adjunct to the measurements of the gravity field since only after the gravitational effects are adjusted for topography can the gravity anomalies be interpreted geophysically. For example, our present understanding of the geometrical flattening is sufficiently uncertain that discussions as to whether or not Mars is in hydrostatic equilibrium cannot be conclusive until altimeter data are available; similarly, the large uncertainty in the precession constant precludes knowledge of individual moments of inertia. Values for these basic quantities derived from Mars Observer would have been useful in studies of Mars' internal stress, radial density, and internal temperature. Studies of the gravity field would have emphasized both the global field and local characteristics of the field. The first task was to determine the global field. Doppler and range tracking measurements spanning periods of several days would have yielded accurate spacecraft trajectory solutions. Simultaneously with reconstruction of the spacecraft orbit, observation equations for approximately 2500 field coefficients and a small number of ancillary parameters would have been derived. This type of gravity field solution is essential for characterizing tectonic phenomena such as the Tharsis bulge and can also be used to study localized features. Results were expected to be available some months after the data were collected, although the final global gravity field would not have been available for one or two years after the end of the Mission. 'Short-arc,' line-of-sight Doppler tracking measurements obtained when the Earth-to-spacecraft line-of-sight was within a few degrees of the orbit plane were expected to provide the highest resolution of local features. The results from this type of observation typically are presented as contoured acceleration profiles of specific features (e.g., craters, volcanoes, etc.) or line-of-sight acceleration maps of specific regions. The high spatial resolution of these products makes them especially useful to geophysicists for study of features in the size range of 300 to 1,000 km. Because of the relative simplicity of the data analysis, results can be available within a few weeks after the data are collected. Radio Occultation Measurements ------------------------------ Atmospheric measurements by the method of radio occultation were expected to contribute to an improved understanding of the circulation, dynamics, and the transport of volatiles and dust in the atmosphere of Mars. These results would have been based on detailed analysis of the radio signal received from Mars Observer as it entered and exited occultation by the planet. Two coordinated investigations of the atmosphere were planned within the Radio Science Team. The first was to focus on obtaining vertical profiles of atmospheric structure with emphasis on investigation of large-scale phenomena, while the second was to concentrate on studies of the small-scale atmospheric structure that causes scintillations in the radio signal. Retrieval of atmospheric profiles requires coherent samples of the radio signal that has propagated through Mars' atmosphere, plus accurate knowledge of the spacecraft trajectory. The latter was to be obtained first from the Mars Observer Navigation Team and later, internally, from Radio Science Team gravity investigators. Initial solutions were expected to provide atmospheric structure -- temperature and pressure vs. absolute radius -- to altitudes of at least 50 km with about 240 m vertical resolution; these 'standard' profiles were expected within 2-4 weeks of data acquisition. After correction for diffractive propagation effects, vertical resolutions of 10-20 m might have been obtained. These 'high-resolution' profiles require an increase in the complexity of retrieval algorithms, which were being revised when the spacecraft was lost. Measures of total columnar content, surface pressure, planet radius at the occultation point, and properties of the ionosphere, if detectable, would have been obtained for each occultation. The spatial and temporal coverage in the radio occultation experiments would have been determined by the geometry of the spacecraft orbit. In the mapping phase of the mission, the Mars Observer spacecraft was to circle Mars in a low-altitude, sun-synchronous, near-polar orbit with a period of about 117 min. As most orbits of the spacecraft included an occultation by Mars, this trajectory provided frequent opportunities for radio occultation measurements of the neutral atmosphere and ionosphere. The occultations generally occurred at high latitudes in both hemispheres. The rotation of the planet between successive orbits would have allowed systematic measurements at regular intervals of about 29 degrees in longitude. Although most occultations were expected to be in polar regions, the view from Earth of the sun-synchronous spacecraft orbit would have evolved slowly in response to the orbital motions of Earth and Mars about the Sun, yielding opportunities for occultation measurements at a variety of latitudes and local times over the course of the 687-day mission. Operational Considerations ========================== Descriptions given here are for nominal performance. The spacecraft transponder system comprised redundant units, each with slightly different characteristics. As transponder units age, their performance changes slightly. More importantly, the performance for Radio Science depended on operational factors such as the modulation state for the transmitters, which cannot be predicted in advance. The performance also depended on factors which were not always under the control of the Mars Observer Project. During radio occultation periods, telemetry modulation would have been turned off; during many of the tests included in the Cruise data set telemetry was also suspended. No special modulation requirements were required during routine tracking for trajectory determination. The steerable 1.5 m diameter parabolic high gain antenna (HGA) was mounted at the end of a 5.6 meter boom so that the phase center of the antenna was about 6 meters from the spacecraft center of mass. An area of concern to the Mars Observer Radio Science and Navigation Teams was the set of 'pseudo-Doppler' effects introduced into the radio tracking and occultation data by motion of the HGA relative to the spacecraft center of mass; it would have been necessary to remove these during data reduction before deriving the gravity field or occultation profiles. Six sources of HGA motion were identified: (i) the spacecraft rotation needed to maintain nadir pointing while in orbit, (ii) articulation of the HGA to maintain Earth point, (iii) HGA boom flexure, (iv) short term migration of the spacecraft center of mass associated with effects such as 'sloshing' of propellants and motions of articulating booms, (v) spacecraft attitude instability, and (vi) spacecraft response to cycling of momentum wheels. The first two error sources are deterministic to the extent that their effects can be taken into account accurately with rather straightforward modeling. Boom flexure was believed to be insignificant. Other motions of the spacecraft center of mass were studied; there were reasons to believe that they would not have proven to be major problems. Item (v) remained a concern at the time the spacecraft was lost. When first analyzed in 1987, tracking error from this source alone was calculated to be nearly an order-of-magnitude greater than that from the next largest source. For the normal 'mapping' configuration, nadir pointing of the spacecraft would have been controlled by referencing spacecraft attitude to the Mars 'horizon' as defined by carbon dioxide emissions from the atmosphere. Atmospheric disturbances can affect determination of the horizon, however, causing the spacecraft attitude to deviate from the true nadir orientation. Doppler error from HGA motion, then, can be coupled to atmospheric features, some of which are stable and associated with surface features (e.g., the semi-permanent waves associated with the Tharsis volcanoes, and the winter polar hood). This coupling would have made such signatures difficult or impossible to distinguish in the data from Doppler effects caused by the spacecraft response to gravity features. After the 1987 analysis, attitude control design engineers at the spacecraft contractor (General Electric, AstroSpace Division) and at the Jet Propulsion Laboratory (JPL) implemented new control algorithms to reduce spacecraft attitude response to local atmospheric features. By mid-1993, analytically predicted Doppler effects resulting from HGA motion relative to the spacecraft center of mass no longer represented the dominant error source for gravity field or atmospheric occultation measurements. Predicted errors associated with the attitude control system remained significant, however, and study of this question was continuing. Item (vi) was identified as a potential source of error only after launch of Mars Observer. Normal rotation of the spacecraft resulted in transfer of angular momentum among the on board momentum wheels. At regular intervals, spaced by 90 degrees of rotation, one set of wheels or another would pass through zero angular momentum and physically stop. Static friction prevented smooth resumption of wheel rotation as the spacecraft continued turning; an attitude transient resulted, which could last for several minutes. Modeling of these 'sticksion' events based on reconstruction of attitude allowed most Doppler effects to be removed from non-coherent Cruise data. But signal-to-noise ratios would have been considerably higher during Mapping, and the accuracy to which this could have been carried out under the best radio science observing conditions remained a concern. Calibration =========== No information is available on calibration of the radio system. Operational Modes ================= The Mars Observer telecommunications system could be viewed as having two sections, which could be operated in the following modes: Section Mode ------------------------------------------- Oscillator two-way (coherent) one-way (non-coherent) RF output low-gain antenna high-gain antenna Selected parameters describing transponder performance are listed below: Oscillator Parameters: X-Band Two-Way Transponder Turnaround Ratio 880/749 One-Way Transmit Frequency (MHz) 8423.200 RF Output parameters: X-Band Power at HGA output 46.5 dBm High-Gain Antenna: Gain (dBi) 39.95 Polarization Right Circular Instrument Overview - DSN ========================= Three Deep Space Communications Complexes (DSCCs) (near Barstow, CA; Canberra, Australia; and Madrid, Spain) comprise the DSN tracking network. Each complex is equipped with several antennas [including at least one each 70-m, 34-m High Efficiency (HEF), and 34-m standard (STD)], associated electronics, and operational systems. Primary activity at each complex is radiation of commands to and reception of telemetry data from active spacecraft. Transmission and reception is possible in several radio-frequency bands, the most common being S-band (nominally a frequency of 2100-2300 MHz or a wavelength of 14.2-13.0 cm) and X-band (7100-8500 MHz or 4.2- 3.5 cm). Transmitter output powers of up to 400 kw are available. Ground stations have the ability to transmit coded and uncoded waveforms which can be echoed by distant spacecraft. Analysis of the received coding allows navigators to determine the distance to the spacecraft; analysis of Doppler shift on the carrier signal allows estimation of the line-of-sight spacecraft velocity. Range and Doppler measurements are used to calculate the spacecraft trajectory and to infer gravity fields of objects near the spacecraft. Ground stations can record spacecraft signals that have propagated through or been scattered from target media. Measurements of signal parameters after wave interactions with surfaces, atmospheres, rings, and plasmas are used to infer physical and electrical properties of the target. Principal investigators vary from experiment to experiment. See the corresponding section of the spacecraft instrument description or the data set description for specifics. The Deep Space Network is managed by the Jet Propulsion Laboratory of the California Institute of Technology for the U.S. National Aeronautics and Space Administration. Specifications include: Instrument Id : RSS Instrument Host Id : DSN Pi Pds User Id : LTYLER Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : N/A Instrument Mass : N/A Instrument Length : N/A Instrument Width : N/A Instrument Height : N/A Instrument Manufacturer Name : N/A For more information on the Deep Space Network and its use in radio science investigations see the reports by [ASMAR&RENZETTI1993] and [ASMAR&HERRERA1993]. For design specifications on DSN subsystems see [DSN810-5]. For an example of use of the DSN for Radio Science see [TYLERETAL1992]. Subsystems - DSN ================ The Deep Space Communications Complexes (DSCCs) are an integral part of the Radio Science instrument, along with other receiving stations and the spacecraft Radio Frequency Subsystem. Their system performance directly determines the degree of success of Radio Science investigations, and their system calibration determines the degree of accuracy in the results of the experiments. The following paragraphs describe the functions performed by the individual subsystems of a DSCC. This material has been adapted from [ASMAR&HERRERA1993]; for additional information, consult [DSN810-5]. Each DSCC includes a set of antennas, a Signal Processing Center (SPC), and communication links to the Jet Propulsion Laboratory (JPL). The general configuration is illustrated below; antennas (Deep Space Stations, or DSS -- a term carried over from earlier times when antennas were individually instrumented) are listed in the table. -------- -------- -------- -------- -------- | DSS 12 | | DSS 18 | | DSS 14 | | DSS 15 | | DSS 16 | |34-m STD| |34-m STD| | 70-m | |34-m HEF| | 26-m | -------- -------- -------- -------- -------- | | | | | | v v | v | --------- | --------- --------->|GOLDSTONE|<---------- |EARTH/ORB| | SPC 10 |<-------------->| LINK | --------- --------- | SPC |<-------------->| 26-M | | COMM | ------>| COMM | --------- | --------- | | | v | v ------ --------- | --------- | NOCC |<--->| JPL |<------- | | ------ | CENTRAL | | GSFC | ------ | COMM | | NASCOMM | | MCCC |<--->| TERMINAL|<-------------->| | ------ --------- --------- ^ ^ | | CANBERRA (SPC 40) <---------------- | | MADRID (SPC 60) <---------------------- GOLDSTONE CANBERRA MADRID Antenna SPC 10 SPC 40 SPC 60 -------- --------- -------- -------- 26-m DSS 16 DSS 46 DSS 66 34-m STD DSS 12 DSS 42 DSS 61 DSS 18 DSS 48 DSS 68 34-m HEF DSS 15 DSS 45 DSS 65 70-m DSS 14 DSS 43 DSS 63 Developmental DSS 13 Subsystem interconnections at each DSCC are shown in the diagram below, and they are described in the sections that follow. The Monitor and Control Subsystem is connected to all other subsystems; the Test Support Subsystem can be. ----------- ------------------ --------- --------- |TRANSMITTER| | | | TRACKING| | COMMAND | | SUBSYSTEM |-| RECEIVER/EXCITER |-|SUBSYSTEM|-|SUBSYSTEM|- ----------- | | --------- --------- | | | SUBSYSTEM | | | | ----------- | | --------------------- | | MICROWAVE | | | | TELEMETRY | | | SUBSYSTEM |-| |-| SUBSYSTEM |- ----------- ------------------ --------------------- | | | ----------- ----------- --------- -------------- | | ANTENNA | | MONITOR | | TEST | | DIGITAL | | | SUBSYSTEM | |AND CONTROL| | SUPPORT | |COMMUNICATIONS|- ----------- | SUBSYSTEM | |SUBSYSTEM| | SUBSYSTEM | ----------- --------- -------------- DSCC Monitor and Control Subsystem ---------------------------------- The DSCC Monitor and Control Subsystem (DMC) is part of the Monitor and Control System (MON) which also includes the ground communications Central Communications Terminal and the Network Operations Control Center (NOCC) Monitor and Control Subsystem. The DMC is the center of activity at a DSCC. The DMC receives and archives most of the information from the NOCC needed by the various DSCC subsystems during their operation. Control of most of the DSCC subsystems, as well as the handling and displaying of any responses to control directives and configuration and status information received from each of the subsystems, is done through the DMC. The effect of this is to centralize the control, display, and archiving functions necessary to operate a DSCC. Communication between the various subsystems is done using a Local Area Network (LAN) hooked up to each subsystem via a network interface unit (NIU). DMC operations are divided into two separate areas: the Complex Monitor and Control (CMC) and the Link Monitor and Control (LMC). The primary purpose of the CMC processor for Radio Science support is to receive and store all predict sets transmitted from NOCC such as Radio Science, antenna pointing, tracking, receiver, and uplink predict sets and then, at a later time, to distribute them to the appropriate subsystems via the LAN. Those predict sets can be stored in the CMC for a maximum of three days under normal conditions. The CMC also receives, processes, and displays event/alarm messages; maintains an operator log; and produces tape labels for the DSP. Assignment and configuration of the LMCs is done through the CMC; to a limited degree the CMC can perform some of the functions performed by the LMC. There are two CMCs (one on-line and one backup) and three LMCs at each DSCC The backup CMC can function as an additional LMC if necessary. The LMC processor provides the operator interface for monitor and control of a link -- a group of equipment required to support a spacecraft pass. For Radio Science, a link might include the DSCC Spectrum Processing Subsystem (DSP) (which, in turn, can control the SSI), or the Tracking Subsystem. The LMC also maintains an operator log which includes operator directives and subsystem responses. One important Radio Science specific function that the LMC performs is receipt and transmission of the system temperature and signal level data from the PPM for display at the LMC console and for inclusion in Monitor blocks. These blocks are recorded on magnetic tape as well as appearing in the Mission Control and Computing Center (MCCC) displays. The LMC is required to operate without interruption for the duration of the Radio Science data acquisition period. The Area Routing Assembly (ARA), which is part of the Digital Communications Subsystem, controls all data communication between the stations and JPL. The ARA receives all required data and status messages from the LMC/CMC and can record them to tape as well as transmit them to JPL via data lines. The ARA also receives predicts and other data from JPL and passes them on to the CMC. DSCC Antenna Mechanical Subsystem --------------------------------- Multi-mission Radio Science activities require support from the 70-m, 34-m HEF, and 34-m STD antenna subnets. The antennas at each DSCC function as large-aperture collectors which, by double reflection, cause the incoming radio frequency (RF) energy to enter the feed horns. The large collecting surface of the antenna focuses the incoming energy onto a subreflector, which is adjustable in both axial and angular position. These adjustments are made to correct for gravitational deformation of the antenna as it moves between zenith and the horizon; the deformation can be as large as 5 cm. The subreflector adjustments optimize the channeling of energy from the primary reflector to the subreflector and then to the feed horns. The 70-m and 34-m HEF antennas have 'shaped' primary and secondary reflectors, with forms that are modified paraboloids. This customization allows more uniform illumination of one reflector by another. The 34-m STD primary reflectors are classical paraboloids, while the subreflectors are standard hyperboloids. On the 70-m and 34-m STD antennas, the subreflector directs received energy from the antenna onto a dichroic plate, a device which reflects S-band energy to the S-band feed horn and passes X-band energy through to the X-band feed horn. In the 34-m HEF, there is one 'common aperture feed,' which accepts both frequencies without requiring a dichroic plate. RF energy to be transmitted into space by the horns is focused by the reflectors into narrow cylindrical beams, pointed with high precision (either to the dichroic plate or directly to the subreflector) by a series of drive motors and gear trains that can rotate the movable components and their support structures. The different antennas can be pointed by several means. Two pointing modes commonly used during tracking passes are CONSCAN and 'blind pointing.' With CONSCAN enabled and a closed loop receiver locked to a spacecraft signal, the system tracks the radio source by conically scanning around its position in the sky. Pointing angle adjustments are computed from signal strength information (feedback) supplied by the receiver. In this mode the Antenna Pointing Assembly (APA) generates a circular scan pattern which is sent to the Antenna Control System (ACS). The ACS adds the scan pattern to the corrected pointing angle predicts. Software in the receiver-exciter controller computes the received signal level and sends it to the APA. The correlation of scan position with the received signal level variations allows the APA to compute offset changes which are sent to the ACS. Thus, within the capability of the closed-loop control system, the scan center is pointed precisely at the apparent direction of the spacecraft signal source. An additional function of the APA is to provide antenna position angles and residuals, antenna control mode/status information, and predict-correction parameters to the Area Routing Assembly (ARA) via the LAN, which then sends this information to JPL via the Ground Communications Facility (GCF) for antenna status monitoring. During periods when excessive signal level dynamics or low received signal levels are expected (e.g., during an occultation experiment), CONSCAN should not be used. Under these conditions, blind pointing (CONSCAN OFF) is used, and pointing angle adjustments are based on a predetermined Systematic Error Correction (SEC) model. Independent of CONSCAN state, subreflector motion in at least the z-axis may introduce phase variations into the received Radio Science data. For that reason, during certain experiments, the subreflector in the 70-m and 34-m HEFs may be frozen in the z-axis at a position (often based on elevation angle) selected to minimize phase change and signal degradation. This can be done via Operator Control Inputs (OCIs) from the LMC to the Subreflector Controller (SRC) which resides in the alidade room of the antennas. The SRC passes the commands to motors that drive the subreflector to the desired position. Unlike the 70-m and 34-m HEFs which have azimuth-elevation (AZ-EL) drives, the 34-m STD antennas use (hour angle-declination) HA-DEC drives. The same positioning of the subreflector on the 34-m STD does not create the same effect as on the 70-m and 34-m HEFs. Pointing angles for all three antenna types are computed by the NOCC Support System (NSS) from an ephemeris provided by the flight project. These predicts are received and archived by the CMC. Before each track, they are transferred to the APA, which transforms the direction cosines of the predicts into AZ-EL coordinates for the 70-m and 34-m HEFs or into HA-DEC coordinates for the 34-m STD antennas. The LMC operator then downloads the antenna AZ-EL or HA-DEC predict points to the antenna-mounted ACS computer along with a selected SEC model. The pointing predicts consist of time-tagged AZ-EL or HA-DEC points at selected time intervals along with polynomial coefficients for interpolation between points. The ACS automatically interpolates the predict points, corrects the pointing predicts for refraction and subreflector position, and adds the proper systematic error correction and any manually entered antenna offsets. The ACS then sends angular position commands for each axis at the rate of one per second. In the 70-m and 34-m HEF, rate commands are generated from the position commands at the servo controller and are subsequently used to steer the antenna. In the 34-m STD antennas motors, rather than servos, are used to steer the antenna; there is no feedback once the 34-m STD has been told where to point. When not using binary predicts (the routine mode for spacecraft tracking), the antennas can be pointed using 'planetary mode' -- a simpler mode which uses right ascension (RA) and declination (DEC) values. These change very slowly with respect to the celestial frame. Values are provided to the station in text form for manual entry. The ACS quadratically interpolates among three RA and DEC points which are on one-day centers. A third pointing mode -- sidereal -- is available for tracking radio sources fixed with respect to the celestial frame. Regardless of the pointing mode being used, a 70-m antenna has a special high-accuracy pointing capability called 'precision' mode. A pointing control loop derives the main AZ-EL pointing servo drive error signals from a two- axis autocollimator mounted on the Intermediate Reference Structure. The autocollimator projects a light beam to a precision mirror mounted on the Master Equatorial drive system, a much smaller structure, independent of the main antenna, which is exactly positioned in HA and DEC with shaft encoders. The autocollimator detects elevation/cross- elevation errors between the two reference surfaces by measuring the angular displacement of the reflected light beam. This error is compensated for in the antenna servo by moving the antenna in the appropriate AZ-EL direction. Pointing accuracies of 0.004 degrees (15 arc seconds) are possible in 'precision' mode. The 'precision' mode is not available on 34-m antennas -- nor is it needed, since their beamwidths are twice as large as on the 70-m antennas. DSCC Antenna Microwave Subsystem -------------------------------- 70-m Antennas: Each 70-m antenna has three feed cones installed in a structure at the center of the main reflector. The feeds are positioned 120 degrees apart on a circle. Selection of the feed is made by rotation of the subreflector. A dichroic mirror assembly, half on the S-band cone and half on the X-band cone, permits simultaneous use of the S- and X-band frequencies. The third cone is devoted to R&D and more specialized work. The Antenna Microwave Subsystem (AMS) accepts the received S- and X-band signals at the feed horn and transmits them through polarizer plates to an orthomode transducer. The polarizer plates are adjusted so that the signals are directed to a pair of redundant amplifiers for each frequency, thus allowing simultaneous reception of signals in two orthogonal polarizations. For S-band these are two Block IVA S-band Traveling Wave Masers (TWMs); for X-band the amplifiers are Block IIA TWMs. 34-m STD Antennas: These antennas have two feed horns, one for S-band signals and one for X-band. The horns are mounted on a cone which is fixed in relation to the subreflector. A dichroic plate mounted above the horns directs energy from the subreflector into the proper horn. The AMS directs the received S- and X-band signals through polarizer plates and on to amplification. There are two Block III S-band TWMs and two Block I X-band TWMs. 34-m HEF Antennas: Unlike the other antennas, the 34-m HEF uses a single feed for both S- and X-band. Simultaneous S- and X-band receive as well as X-band transmit is possible thanks to the presence of an S/X 'combiner' which acts as a diplexer. For S-band, RCP or LCP is user selected through a switch so neither a polarizer nor an orthomode transducer is needed. X-band amplification options include two Block II TWMs or an HEMT Low Noise Amplifier (LNA). S-band amplification is provided by an FET LNA. DSCC Receiver-Exciter Subsystem ------------------------------- The Receiver-Exciter Subsystem is composed of three groups of equipment: the closed-loop receiver group, the open-loop receiver group, and the RF monitor group. This subsystem is controlled by the Receiver-Exciter Controller (REC) which communicates directly with the DMC for predicts and OCI reception and status reporting. The exciter generates the S-band signal (or X-band for the 34-m HEF only) which is provided to the Transmitter Subsystem for the spacecraft uplink signal. It is tunable under command of the Digitally Controlled Oscillator (DCO) which receives predicts from the Metric Data Assembly (MDA). The diplexer in the signal path between the transmitter and the feed horn for all three antennas (used for simultaneous transmission and reception) may be configured such that it is out of the received signal path (in listen-only or bypass mode) in order to improve the signal-to-noise ratio in the receiver system. Closed Loop Receivers: The Block IV receiver-exciter at the 70-m stations allows for two receiver channels, each capable of L-Band (e.g., 1668 MHz frequency or 18 cm wavelength), S-band, or X-band reception, and an S-band exciter for generation of uplink signals through the low-power or high-power transmitter. The Block III receiver-exciter at the 34-m STD stations allows for two receiver channels, each capable of S-band or X-band reception and an exciter used to generate an uplink signal through the low-power transmitter. The receiver-exciter at the 34-m HEF stations allows for one channel only. The closed-loop receivers provide the capability for rapid acquisition of a spacecraft signal and telemetry lockup. In order to accomplish acquisition within a short time, the receivers are predict driven to search for, acquire, and track the downlink automatically. Rapid acquisition precludes manual tuning though that remains as a backup capability. The subsystem utilizes FFT analyzers for rapid acquisition. The predicts are NSS generated, transmitted to the CMC which sends them to the Receiver-Exciter Subsystem where two sets can be stored. The receiver starts acquisition at uplink time plus one round-trip-light-time or at operator specified times. The receivers may also be operated from the LMC without a local operator attending them. The receivers send performance and status data, displays, and event messages to the LMC. Either the exciter synthesizer signal or the simulation (SIM) synthesizer signal is used as the reference for the Doppler extractor in the closed-loop receiver systems, depending on the spacecraft being tracked (and Project guidelines). The SIM synthesizer is not ramped; instead it uses one constant frequency, the Track Synthesizer Frequency (TSF), which is an average frequency for the entire pass. The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. It will be configured such that the expected amplitude changes are accommodated with minimum distortion. The loop bandwidth (2BLo) will be configured such that the expected phase changes can be accommodated while maintaining the best possible loop SNR. Open-Loop Receivers: The Radio Science Open-Loop Receiver (OLR) is a dedicated four channel, narrow-band receiver which provides amplified and downconverted video band signals to the DSCC Spectrum Processing Subsystem (DSP). The OLR utilizes a fixed first Local Oscillator (LO) frequency and a tunable second LO frequency to minimize phase noise and improve frequency stability. The OLR consists of an RF-to-IF downconverter located in the antenna, an IF selection switch (IVC), and a Radio Science IF-VF downconverter (RIV) located in the SPC. The RF-IF downconverters in the 70-m antennas are equipped for four IF channels: S-RCP, S-LCP, X-RCP, and X-LCP. The 34-m HEF stations are equipped with a two-channel RF-IF: S-band and X-band. The IVC switches the IF input between the 70-m and 34-m HEF antennas. The RIV contains the tunable second LO, a set of video bandpass filters, IF attenuators, and a controller (RIC). The LO tuning is done via DSP control of the POCA/PLO combination based on a predict set. The POCA is a Programmable Oscillator Control Assembly and the PLO is a Programmable Local Oscillator (commonly called the DANA synthesizer). The bandpass filters are selectable via the DSP. The RIC provides an interface between the DSP and the RIV. It is controlled from the LMC via the DSP. The RIC selects the filter and attenuator settings and provides monitor data to the DSP. The RIC could also be manually controlled from the front panel in case the electronic interface to the DSP is lost. RF Monitor -- SSI and PPM: The RF monitor group of the Receiver-Exciter Subsystem provides spectral measurements using the Spectral Signal Indicator (SSI) and measurements of the received channel system temperature and spacecraft signal level using the Precision Power Monitor (PPM). The SSI provides a local display of the received signal spectrum at a dedicated terminal at the DSCC and routes these same data to the DSP which routes them to NOCC for remote display at JPL for real-time monitoring and RIV/DSP configuration verification. These displays are used to validate Radio Science Subsystem data at the DSS, NOCC, and Mission Support Areas. The SSI configuration is controlled by the DSP and a duplicate of the SSI spectrum appears on the LMC via the DSP. During real-time operations the SSI data also serve as a quick-look science data type for Radio Science experiments. The PPM measures system noise temperatures (SNT) using a Noise Adding Radiometer (NAR) and downlink signal levels using the Signal Level Estimator (SLE). The PPM accepts its input from the closed-loop receiver. The SNT is measured by injecting known amounts of noise power into the signal path and comparing the total power with the noise injection 'on' against the total power with the noise injection 'off.' That operation is based on the fact that receiver noise power is directly proportional to temperature; thus measuring the relative increase in noise power due to the presence of a calibrated thermal noise source allows direct calculation of SNT. Signal level is measured by calculating an FFT to estimate the SNR between the signal level and the receiver noise floor where the power is known from the SNT measurements. There is one PPM controller at the SPC which is used to control all SNT measurements. The SNT integration time can be selected to represent the time required for a measurement of 30K to have a one-sigma uncertainty of 0.3K or 1%. DSCC Transmitter Subsystem -------------------------- The Transmitter Subsystem accepts the S-band frequency exciter signal from the Block III or Block IV Receiver- Exciter Subsystem exciter and amplifies it to the required transmit output level. The amplified signal is routed via the diplexer through the feed horn to the antenna and then focused and beamed to the spacecraft. The Transmitter Subsystem power capabilities range from 18 kw to 400 kw. Power levels above 18 kw are available only at 70-m stations. DSCC Tracking Subsystem ----------------------- The Tracking Subsystem primary functions are to acquire and maintain communications with the spacecraft and to generate and format radiometric data containing Doppler and range. The DSCC Tracking Subsystem (DTK) receives the carrier signals and ranging spectra from the Receiver-Exciter Subsystem. The Doppler cycle counts are counted, formatted, and transmitted to JPL in real time. Ranging data are also transmitted to JPL in real time. Also contained in these blocks is the AGC information from the Receiver-Exciter Subsystem. The Radio Metric Data Conditioning Team (RMDCT) at JPL produces an Archival Tracking Data File (ATDF) tape which contains Doppler and ranging data. In addition, the Tracking Subsystem receives from the CMC frequency predicts (used to compute frequency residuals and noise estimates), receiver tuning predicts (used to tune the closed-loop receivers), and uplink tuning predicts (used to tune the exciter). From the LMC, it receives configuration and control directives as well as configuration and status information on the transmitter, microwave, and frequency and timing subsystems. The Metric Data Assembly (MDA) controls all of the DTK functions supporting the uplink and downlink activities. The MDA receives uplink predicts and controls the uplink tuning by commanding the DCO. The MDA also controls the Sequential Ranging Assembly (SRA). It formats the Doppler and range measurements and provides them to the GCF for transmission to NOCC. The Sequential Ranging Assembly (SRA) measures the round trip light time (RTLT) of a radio signal traveling from a ground tracking station to a spacecraft and back. From the RTLT, phase, and Doppler data, the spacecraft range can be determined. A coded signal is modulated on an uplink carrier and transmitted to the spacecraft where it is detected and transponded back to the ground station. As a result, the signal received at the tracking station is delayed by its round trip through space and shifted in frequency by the Doppler effect due to the relative motion between the spacecraft and the tracking station on Earth. DSCC Spectrum Processing Subsystem (DSP) ---------------------------------------- The DSCC Spectrum Processing Subsystem (DSP) located at the SPC digitizes and records on magnetic tapes the narrowband output data from the RIV. It consists of a Narrow Band Occultation Converter (NBOC) containing four Analog-to- Digital Converters (ADCs), a ModComp CLASSIC computer processor called the Spectrum Processing Assembly (SPA), and two to six magnetic tape drives. Magnetic tapes are known as Original Data Records (ODRs). Electronic near real-time transmission of data to JPL (an Original Data Stream, or ODS) may be possible in certain circumstances; The DSP is operated through the LMC. Using the SPA-R software, the DSP allows for real-time frequency and time offsets (while in RUN mode) and, if necessary, snap tuning between the two frequency ranges transmitted by the spacecraft: coherent and non-coherent. The DSP receives Radio Science frequency predicts from the CMC, allows for multiple predict set archiving (up to 60 sets) at the SPA, and allows for manual predict generation and editing. It accepts configuration and control data from the LMC, provides display data to the LMC, and transmits the signal spectra from the SSI as well as status information to NOCC and the Project Mission Support Area (MSA) via the GCF data lines. The DSP records the digitized narrowband samples and the supporting header information (i.e., time tags, POCA frequencies, etc.) on 9-track magnetic tapes in 6250 or 1600 bpi GCR format. Through the DSP-RIC interface the DSP controls the RIV filter selection and attenuation levels. It also receives RIV performance monitoring via the RIC. In case of failure of the DSP-RIC interface, the RIV can be controlled manually from the front panel. All the RIV and DSP control parameters and configuration directives are stored in the SPA in a macro-like file called an 'experiment directive' table. A number of default directives exist in the DSP for the major Radio Science experiments. Operators can create their own table entries. Items such as verification of the configuration of the prime open-loop recording subsystem, the selection of the required predict sets, and proper system performance prior to the recording periods will be checked in real-time at JPL via the NOCC displays using primarily the remote SSI display at NOCC and the NRV displays. Because of this, transmission of the DSP/SSI monitor information is enabled prior to the start of recording. The specific run time and tape recording times will be identified in the Sequence of Events (SOE) and/or DSN Keyword File. The DSP can be used to duplicate ODRs. It also has the capability to play back a certain section of the recorded data after conclusion of the recording periods. DSCC Frequency and Timing Subsystem ----------------------------------- The Frequency and Timing Subsystem (FTS) provides all frequency and timing references required by the other DSCC subsystems. It contains four frequency standards of which one is prime and the other three are backups. Selection of the prime standard is done via the CMC. Of these four standards, two are hydrogen masers followed by clean-up loops (CUL) and two are cesium standards. These four standards all feed the Coherent Reference Generator (CRG) which provides the frequency references used by the rest of the complex. It also provides the frequency reference to the Master Clock Assembly (MCA) which in turn provides time to the Time Insertion and Distribution Assembly (TID) which provides UTC and SIM-time to the complex. JPL's ability to monitor the FTS at each DSCC is limited to the MDA calculated Doppler pseudo-residuals, the Doppler noise, the SSI, and to a system which uses the Global Positioning System (GPS). GPS receivers at each DSCC receive a one-pulse-per-second pulse from the station's (hydrogen maser referenced) FTS and a pulse from a GPS satellite at scheduled times. After compensating for the satellite signal delay, the timing offset is reported to JPL where a database is kept. The clock offsets stored in the JPL database are given in microseconds; each entry is a mean reading of measurements from several GPS satellites and a time tag associated with the mean reading. The clock offsets provided include those of SPC 10 relative to UTC (NIST), SPC 40 relative to SPC 10, etc. Optics - DSN ============ Performance of DSN ground stations depends primarily on size of the antenna and capabilities of electronics. These are summarized in the following set of tables. Note that 64-m antennas were upgraded to 70-m between 1986 and 1989. Beamwidth is half-power full angular width. Polarization is circular; L denotes left circular polarization (LCP), and R denotes right circular polarization (RCP). DSS S-Band Characteristics 64-m 70-m 34-m 34-m Transmit STD HEF -------- ----- ----- ----- ----- Frequency (MHz) 2110- 2110- 2025- N/A 2120 2120 2120 Wavelength (m) 0.142 0.142 0.142 N/A Ant Gain (dBi) 62.7 55.2 N/A Beamwidth (deg) 0.119 0.31 N/A Polarization L or R L or R N/A Tx Power (kW) 20-400 20 N/A Receive ------- Frequency (MHz) 2270- 2270- 2270- 2200- 2300 2300 2300 2300 Wavelength (m) 0.131 0.131 0.131 0.131 Ant Gain (dBi) 61.6 63.3 56.2 56.0 Beamwidth (deg) 0.108 0.27 0.24 Polarization L & R L & R L or R L or R System Temp (K) 22 20 22 38 DSS X-Band Characteristics 64-m 70-m 34-m 34-m Transmit STD HEF -------- ----- ----- ----- ----- Frequency (MHz) 8495 8495 N/A 7145- 7190 Wavelength (m) 0.035 0.035 N/A 0.042 Ant Gain (dBi) 74.2 N/A 67 Beamwidth (deg) N/A 0.074 Polarization L or R L or R N/A L or R Tx Power (kW) 360 360 N/A 20 Receive ------- Frequency (MHz) 8400- 8400- 8400- 8400- 8500 8500 8500 8500 Wavelength (m) 0.036 0.036 0.036 0.036 Ant Gain (dBi) 71.7 74.2 66.2 68.3 Beamwidth (deg) 0.031 0.075 0.063 Polarization L & R L & R L & R L & R System Temp (K) 27 20 25 20 Electronics - DSN ================= DSCC Open-Loop Receiver ----------------------- The open loop receiver block diagram shown below is for 70-m and 34-m High-Efficiency (HEF) antenna sites. Based on a tuning prediction file, the POCA controls the DANA synthesizer the output of which (after multiplication) mixes input signals at both S- and X-band to fixed intermediate frequencies for amplification. These signals in turn are down converted and passed through additional filters until they yield baseband output of up to 25 kHz in width. The baseband output is digitally sampled by the DSP and either written to magnetic tape or electronically transferred for further analysis. S-Band X-Band 2295 MHz 8415 MHz Input Input | | v v --- --- --- --- | X |<--|x20|<--100 MHz 100 MHz-->|x81|-->| X | --- --- --- --- | | 295| |315 MHz| |MHz v v --- -- --- --- | X |<--|x3|<------ ------>|x11|-->| X | --- -- |115 33| --- --- | |MHz MHz| | | | | | 50| --- --- |50 MHz| 72 MHz->| X | | X | |MHz v --- --- v --- ^ ^ --- | X |<--60 MHz | 43 43| 60 MHz-->| X | --- |MHz MHz| --- | 9.9 | | 9.9 | | MHz ------------- MHz | | | ^ | | 10| v | v |10 MHz| --- ---------- --- |MHz |------>| X | | DANA | | X |<------| | --- |Synthesizr| --- | | | ---------- | | v v ^ v v ------- ------- | ------- ------- |Filters| |Filters| ---------- |Filters| |Filters| |3,4,5,6| | 1,2 | | POCA | | 1,2 | |3,4,5,6| ------- ------- |Controller| ------- ------- | | ---------- | | 10| |0.1 0.1| |10 MHz| |MHz MHz| |MHz v v v v --- --- --- --- | X |- -| X | | X |- -| X | --- | | --- --- | | --- ^ | | ^ ^ | | ^ | | | | | | | | 10 | | 0.1 0.1 | | 10 MHz | | MHz MHz | | MHz | | | | v v v v Baseband Baseband Output Output Reconstruction of the antenna frequency from the frequency of the signal in the recorded data can be achieved through use of one of the following formulas. Radio Science IF-VF (RIV) Converter Assembly at 70-m and 34-m High-Efficiency (HEF) antennas: FSant=3*[POCA+(79/11)*10^6] + 1.95*10^9 - Fsamp - Frec FXant=11*[POCA-10^7] + 8.050*10^9 - 3*Fsamp + Frec Multi-Mission Receivers at 34-m Standard antennas: FSant=48*POCA + 3*10^8 - 0.75*Fsamp + Frec FXant = (11/3)*[48*POCA + 3*10^8 - 0.75*Fsamp] + Frec where FSant = S-band antenna frequency FXant = X-band antenna frequency POCA = POCA frequency Fsamp = sampling frequency Frec = frequency of recorded signal Filters - DSN ============= DSCC Open-Loop Receiver ----------------------- Nominal filter center frequencies and bandwidths for the Open-Loop Receivers are shown in the table below. Filter Center Frequency 3 dB Bandwidth ------ ---------------- -------------- 1 0.1 MHz 90 Hz 2 0.1 MHz 450 Hz 3 10.0 MHz 2000 Hz 4 10.0 MHz 1700 Hz (S-band) 6250 Hz (X-band) 5 10.0 MHz 45000 Hz 6 10.0 MHz 21000 Hz Detectors - DSN =============== DSCC Open-Loop Receivers ------------------------ Open-loop receiver output is detected in software by the radio science investigator. DSCC Closed-Loop Receivers -------------------------- Nominal carrier tracking loop threshold noise bandwidth at both S- and X-band is 10 Hz. Coherent (two-way) closed-loop system stability is shown in the table below: integration time Doppler uncertainty (secs) (one sigma, microns/sec) ------ ------------------------ 10 50 60 20 1000 4 Calibration - DSN ================= Calibrations of hardware systems are carried out periodically by DSN personnel; these ensure that systems operate at required performance levels -- for example, that antenna patterns, receiver gain, propagation delays, and Doppler uncertainties meet specifications. No information on specific calibration activities is available. Nominal performance specifications are shown in the tables above. Additional information may be available in [DSN810-5]. Prior to each tracking pass, station operators perform a series of calibrations to ensure that systems meet specifications for that operational period. Included in these calibrations is measurement of receiver system temperature in the configuration to be employed during the pass. Results of these calibrations are recorded in (hard copy) Controller's Logs for each pass. The nominal procedure for initializing open-loop receiver attenuator settings is described below. In cases where widely varying signal levels are expected, the procedure may be modified in advance or real-time adjustments may be made to attenuator settings. Open-Loop Receiver Attenuation Calibration ------------------------------------------ The open-loop receiver attenuator calibrations are performed to establish the output of the open-loop receivers at a level that will not saturate the analog-to-digital converters. To achieve this, the calibration is done using a test signal generated by the exciter/translator that is set to the peak predicted signal level for the upcoming pass. Then the output level of the receiver's video band spectrum envelope is adjusted to the level determined by equation (3) below (to five-sigma). Note that the SNR in the equation (2) is in dB while the SNR in equation (3) is linear. Pn = -198.6 + 10*log(SNT) + 10*log(1.2*Fbw) (1) SNR = Ps - Pn (SNR in dB) (2) Vrms = sqrt(SNR + 1)/[1 + 0.283*sqrt(SNR)] (SNR linear) (3) where Fbw = receiver filter bandwidth (Hz) Pn = receiver noise power (dBm) Ps = signal power (dBm) SNT = system noise temperature (K) SNR = predicted signal-to-noise ratio Operational Considerations - DSN ================================ The DSN is a complex and dynamic 'instrument.' Its performance for Radio Science depends on a number of factors from equipment configuration to meteorological conditions. No specific information on 'operational considerations' can be given here. Operational Modes - DSN ======================= DSCC Antenna Mechanical Subsystem --------------------------------- Pointing of DSCC antennas may be carried out in several ways. For details see the subsection 'DSCC Antenna Mechanical Subsystem' in the 'Subsystem' section. Binary pointing is the preferred mode for tracking spacecraft; pointing predicts are provided, and the antenna simply follows those. With CONSCAN, the antenna scans conically about the optimum pointing direction, using closed-loop receiver signal strength estimates as feedback. In planetary mode, the system interpolates from three (slowly changing) RA-DEC target coordinates; this is 'blind' pointing since there is no feedback from a detected signal. In sidereal mode, the antenna tracks a fixed point on the celestial sphere. In 'precision' mode, the antenna pointing is adjusted using an optical feedback system. It is possible on most antennas to freeze z-axis motion of the subreflector to minimize phase changes in the received signal. DSCC Receiver-Exciter Subsystem ------------------------------- The diplexer in the signal path between the transmitter and the feed horns on all three antennas may be configured so that it is out of the received signal path in order to improve the signal-to-noise ratio in the receiver system. This is known as the 'listen-only' or 'bypass' mode. Closed-Loop vs. Open-Loop Reception ----------------------------------- Radio Science data can be collected in two modes: closed- loop, in which a phase-locked loop receiver tracks the spacecraft signal, or open-loop, in which a receiver samples and records a band within which the desired signal presumably resides. Closed-loop data are collected using Closed-Loop Receivers, and open-loop data are collected using Open-Loop Receivers in conjunction with the DSCC Spectrum Processing Subsystem (DSP). See the Subsystems section for further information. Closed-Loop Receiver AGC Loop ----------------------------- The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. Ordinarily it is configured so that expected signal amplitude changes are accommodated with minimum distortion. The loop bandwidth is ordinarily configured so that expected phase changes can be accommodated while maintaining the best possible loop SNR. Coherent vs. Non-Coherent Operation ----------------------------------- The frequency of the signal transmitted from the spacecraft can generally be controlled in two ways -- by locking to a signal received from a ground station or by locking to an on-board oscillator. These are known as the coherent (or 'two-way') and non-coherent ('one-way') modes, respectively. Mode selection is made at the spacecraft, based on commands received from the ground. When operating in the coherent mode, the transponder carrier frequency is derived from the received uplink carrier frequency with a 'turn-around ratio' typically of 240/221. In the non-coherent mode, the downlink carrier frequency is derived from the spacecraft on-board crystal-controlled oscillator. Either closed-loop or open-loop receivers (or both) can be used with either spacecraft frequency reference mode. Closed-loop reception in two-way mode is usually preferred for routine tracking. Occasionally the spacecraft operates coherently while two ground stations receive the 'downlink' signal; this is sometimes known as the 'three-way' mode. DSCC Spectrum Processing Subsystem (DSP) ---------------------------------------- The DSP can operate in four sampling modes with from 1 to 4 input signals. Input channels are assigned to ADC inputs during DSP configuration. Modes and sampling rates are summarized in the tables below: Mode Analog-to-Digital Operation ---- ---------------------------- 1 4 signals, each sampled by a single ADC 2 1 signal, sampled sequentially by 4 ADCs 3 2 signals, each sampled sequentially by 2 ADCs 4 1 signal, sampled sequentially by 3 ADCs 8-bit Samples 12-bit Samples Sampling Rates Sampling Rates (samples/sec per ADC) (samples/sec per ADC) --------------------- --------------------- 50000 31250 25000 15625 12500 10000 10000 6250 5000 5000 4000 3125 2500 2000 1250 1000 1000 500 400 250 200 200 Location - DSN ============== Station locations are documented in [GEO-10REVD]. Geocentric coordinates are summarized here. Geocentric Geocentric Geocentric Station Radius (km) Latitude (N) Longitude (E) --------- ----------- ------------ ------------- Goldstone DSS 12 (34-m STD) 6371.997815 35.1186672 243.1945048 DSS 13 (develop) 6372.117062 35.0665485 243.2051077 DSS 14 (70-m) 6371.992867 35.2443514. 243.1104584 DSS 15 (34-m HEF) 6371.9463 35.2402863 243.1128186 DSS 16 (26-m) 6371.9608 35.1601436 243.1264200 DSS 18 (34-m STD) UNK UNK UNK Canberra DSS 42 (34-m STD) 6371.675607 -35.2191850 148.9812546 DSS 43 (70-m) 6371.688953 -35.2209308 148.9812540 DSS 45 (34-m HEF) 6371.692 -35.21709 148.97757 DSS 46 (26-m) 6371/675 -35.22360 148.98297 DSS 48 (34-m STD) UNK UNK UNK Madrid DSS 61 (34-m STD) 6370.027734 40.2388805 355.7509634 DSS 63 (70-m) 6370.051015 40.2413495 355.7519776 DSS 65 (34-m HEF) 6370.021370 40.2372843 355.7485968 DSS 66 (26-m) 6370.036 40.2400714 355.7485976 DSS 48 (34-m STD) UNK UNK UNK Measurement Parameters - DSN ============================ Open-Loop System ---------------- Output from the Open-Loop Receivers (OLRs), as sampled and recorded by the DSCC Spectrum Processing Subsystem (DSP), is a stream of 8- or 12-bit quantized voltage samples. The nominal input to the Analog-to-Digital Converters (ADCs) is +/-10 volts, but the precise scaling between input voltages and output digitized samples is usually irrelevant for analysis; the digital data are generally referenced to a known noise or signal level within the data stream itself -- for example, the thermal noise output of the radio receivers which has a known system noise temperature (SNT). Raw samples comprise the data block in each DSP record; a header record (presently 83 16-bit words) contains ancillary information such as: time tag for the first sample in the data block RMS values of receiver signal levels and ADC outputs POCA frequency and drift rate Closed-Loop System ------------------ Closed-loop data are recorded in Archival Tracking Data Files (ATDFs), as well as certain secondary products such as the Orbit Data File (ODF). The ATDF Tracking Logical Record contains 117 entries including status information and measurements of ranging, Doppler, and signal strength. ACRONYMS AND ABBREVIATIONS - DSN ================================ ACS Antenna Control System ADC Analog-to-Digital Converter AMS Antenna Microwave System APA Antenna Pointing Assembly ARA Area Routing Assembly ATDF Archival Tracking Data File AZ Azimuth CMC Complex Monitor and Control CONSCAN Conical Scanning (antenna pointing mode) CRG Coherent Reference Generator CUL Clean-up Loop DANA a type of frequency synthesizer dB deciBel dBi dB relative to isotropic DCO Digitally Controlled Oscillator DEC Declination deg degree DMC DSCC Monitor and Control Subsystem DSCC Deep Space Communications Complex DSN Deep Space Network DSP DSCC Spectrum Processing Subsystem DSS Deep Space Station DTK DSCC Tracking Subsystem E east EL Elevation FTS Frequency and Timing Subsystem GCF Ground Communications Facility GPS Global Positioning System HA Hour Angle HEF High-Efficiency (as in 34-m HEF antennas) IF Intermediate Frequency IVC IF Selection Switch JPL Jet Propulsion Laboratory K Kelvin km kilometer kW kilowatt L-band approximately 1668 MHz LAN Local Area Network LCP Left-Circularly Polarized LMC Link Monitor and Control LNA Low-Noise Amplifier LO Local Oscillator m meters MCA Master Clock Assembly MCCC Mission Control and Computing Center MDA Metric Data Assembly MHz Megahertz MON Monitor and Control System MSA Mission Support Area N north NAR Noise Adding Radiometer NBOC Narrow-Band Occultation Converter NIST SPC 10 time relative to UTC NIU Network Interface Unit NOCC Network Operations and Control System NSS NOCC Support System OCI Operator Control Input ODF Orbit Data File ODR Original Data Record ODS Original Data Stream OLR Open Loop Receiver POCA Programmable Oscillator Control Assembly PPM Precision Power Monitor RA Right Ascension REC Receiver-Exciter Controller RCP Right-Circularly Polarized RF Radio Frequency RIC RIV Controller RIV Radio Science IF-VF Converter Assembly RMDCT Radio Metric Data Conditioning Team RTLT Round-Trip Light Time S-band approximately 2100-2300 MHz sec second SEC System Error Correction SIM Simulation SLE Signal Level Estimator SNR Signal-to-Noise Ratio SNT System Noise Temperature SOE Sequence of Events SPA Spectrum Processing Assembly SPC Signal Processing Center SRA Sequential Ranging Assembly SRC Sub-Reflector Controller SSI Spectral Signal Indicator STD Standard (as in 34-m STD antennas) TID Time Insertion and Distribution Assembly TSF Tracking Synthesizer Frequency TWM Traveling Wave Maser UNK unknown UTC Universal Coordinated Time VF Video Frequency X-band approximately 7800-8500 MHz
  • instrument : RADIO SCIENCE SUBSYSTEM for CLEM1
    Instrument Overview =================== The Clementine spacecraft telecommunications subsystem served as part of a Radio Science instrument for investigations of the Moon. The remainder of the 'instrument' was located at ground stations of the NASA Deep Space Network (DSN). Much of the equipment at both ends was shared, being used for routine telecommunications as well as for Radio Science. Radio data were themselves shared; Doppler and range measurements were used to calculate the spacecraft trajectory and to infer the lunar gravity field. Measurements of signal parameters after waves had interacted with the lunar surface were used to infer physical and electrical properties of the surface material. Instrument Overview - Spacecraft ================================ The spacecraft telecommunications subsystem provided simultaneous radio frequency (RF) uplink, downlink, and coherent tracking capabilities with DSN ground stations [REGEONETAL1994]. Clementine used two S-band (2.0-2.3 GHz) omni-directional antennas for low data rate communications and a body-fixed 1.1 m diameter composite parabolic reflector with a deployable feed for high data rate communications. The omni antennas were configured such that commands could be received over a wide range of spacecraft orientations. Two S-Band transponder units provided the transmit and receive functions for the spacecraft. Each transponder contained one exciter and solid-state transmitter and one fixed-tuned phase-locked loop receiver (2093.0542 MHz). In addition, the subsystem contained three diplexers, two 10 dB couplers, and a transfer switch. ---------------------------------------- | | v | -------- | HGA | | |\ ---------- | Switch |-- | \ | |<-| | | | >--| Diplexer | -------- | | / | |- ^ | |/ ---------- | | | | | --------------------------------------- | | | -------- ----- ----- || --------------->| 90 deg | | CMD | | PCM | || OMNI-A | --| Hybrid | ----- ----- ||\ || -------- ^ ^ || \ ---------- || | | | || ><-- ---->| |<- | v v v || / | | | Diplexer | | --------- ------------- ||/ v v | |----->| 10 dB | | RCVR | XMTR |- | -------- ---------- | | Coupler |->| A | A | | | 90 deg | | --------- ------------- | | Hybrid | | --------- ------------- | -------- ---------- ->| 10 dB | | RCVR | XMTR |--\ ^ ^ | |----->| Coupler |->| B | B | \ | | | Diplexer | --------- ------------- ><-- ---->| |--- ^ ^ / ---------- | | |/ / v v OMNI-B \ ----- ----- / | CMD | | PCM | \ ----- ----- | --- /// Both spacecraft receivers were powered on at all times during the mission. Ranging signals, uplinked independently or during commanding, were demodulated by the receiver and passed on to the transmitter for re-transmission to Earth. The receiver's automatic gain control was commanded off to allow tracking of the uplink signal through nulls in the antenna pattern as the spacecraft spun. The exact frequency of the signal transmitted from the spacecraft could be controlled in two ways -- by locking to a signal received from a ground station or by locking to an on-board oscillator. These are known as the coherent (or 'two-way') and non-coherent ('one-way') modes, respectively. When operating in the coherent mode, the transponder carrier frequency was derived from the received uplink carrier frequency with a 'turn-around ratio' of 240/221. In the non-coherent mode, the downlink carrier frequency was derived from the spacecraft on-board crystal-controlled oscillator. After a 3-hour warm-up, the nominal crystal oscillator frequency was 2273.000+/-0.002 MHz. Solid state amplifiers, driven at saturation, amplified the multiplier output before the signals were radiated via either the high-gain parabolic antenna or the low gain omni- directional antennas. The signal transmitted by each of the Clementine antennas was right circularly polarized. Pertinent details of the subsystem are shown below; its 'build date' is taken to be 1994-01-01, which was during the Pre-Launch phase of the Clementine mission. Instrument Id : RSS Instrument Host Id : CLEM1 Pi Pds User Id : N/A Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : 1994-01-01 Instrument Mass : 13.6 Instrument Length : N/A Instrument Width : N/A Instrument Height : N/A Instrument Manufacturer Name : NAVAL RESEARCH LABORATORY Platform Mounting Descriptions ------------------------------ The spacecraft +Z axis vector was in the nominal direction of the instrument sensor panel. The +X axis vector was parallel to the nominal direction of the main thruster nozzle, and the -X axis vector was parallel to the spacecraft high-gain antenna (HGA) boresight. The +Y axis vector formed a right-handed coordinate system and was in the nominal direction of the solar panel rotation axis. The low-gain 'omni-directional' antennas were mounted on the +Z and -Z sides of the spacecraft body. The spacecraft velocity vector was in approximately the -X direction when the spacecraft was oriented for nadir viewing. The nominal HGA S-band polarization was right-hand circular in the -X direction. Cone Offset Angle : 0.00 Cross Cone Offset Angle : 0.00 Twist Offset Angle : 0.00 Principal Investigators ----------------------- The Principal Investigators for the gravity investigations were David E. Smith and Maria T. Zuber. The Principal Investigator for the radio occultation and radio scattering experiments was Christopher L. Lichtenberg. Scientific Objectives ===================== Two different types of radio science experiments were conducted with Clementine: radio tracking experiments in which the magnitude and direction of the planet's gravity field were derived from the Doppler and ranging measurements, and radio propagation experiments in which modulation on the signal received on Earth could be attributed to properties of the intervening medium. Two types of radio propagation experiments were carried out: radio occultations by the lunar limb and scattering from the lunar surface. Gravity Measurements -------------------- Measurement of the gravity field provides significant constraints on inferences about interior structure of the Moon. Precise, detailed study of the spacecraft motion in lunar orbit can yield the mass distribution of the target body. Topographic data obtained by the Clementine laser altimeter forms a critical adjunct to these measurements since only after the gravitational effects are adjusted for topography can the gravity anomalies be interpreted geophysically [ZUBERETAL1994]. Studies of the gravity field emphasize both the global field and local characteristics of the field. The first task is to determine the global field. Doppler and range tracking measurements yield accurate spacecraft trajectory solutions. Simultaneously with reconstruction of the spacecraft orbit, observation equations for field coefficients and a small number of ancillary parameters can be solved [BALMINO1981]. This type of gravity field solution is essential for characterizing large scale phenomena and can also be used to study localized features. 'Short-arc' line-of-sight Doppler tracking measurements obtained when the Earth-to-spacecraft line-of-sight is within a few degrees of the orbit plane provide the highest resolution of local features. The results from this type of observation typically are presented as contoured acceleration profiles of specific features (e.g., craters, ridges, etc.) or line-of-sight acceleration maps of specific regions. The high spatial resolution of these products makes them especially useful to geophysicists for study of features in the size range of 300 to 1,000 km. Because of the relative simplicity of the data analysis, results can be available within a few days after the data are collected [SJOGRENETAL1976]. Radio Occultation Measurements ------------------------------ Radio measurements during occultations can provide information on topography along the target body limb. When combined with accurate reconstructions of the spacecraft trajectory, the time at which occultation occurs provides a precise estimate of the occultation point distance from the target center-of-mass [LINDALETAL1979]. Differences between the ideal diffraction signature expected from a 'knife-edge' opaque screen can also be used to infer details of surface structure in the vicinity of the limb. Bistatic Surface Scattering Measurements ---------------------------------------- When the spacecraft telecommunications antenna is pointed toward the surface of the target, the strength of the scattered signal from the illuminated area may be interpreted in terms of the texture of the surface at that point [SIMPSON1993]. In experiments conducted on 1994-04-10, the Clementine high-gain antenna was aimed toward a single target area -- the lunar south pole (the so-called 'spotlight mode') -- and the scattered signal was recorded at DSN stations in California, Australia, and Spain as the spacecraft moved in its orbit. On 1993-04-23, the spacecraft high-gain antenna was aimed toward the lunar north pole in another spotlight experiment; those signals were received at the Canberra and Madrid DSN sites. On 1994-04-23 the spacecraft antenna was also aimed toward the specular point on the lunar surface for part of a single orbit (the 'quasi- specular' experiment mode). In that case, the mirror-like reflected signal was received at the DSN station near Madrid; the quasi-specular echo is much stronger than the echo obtained in other configurations. Spotlight experiments, conducted on polar targets were designed to detect and measure enhanced backscatter from possible ice deposits in areas permanently shaded from solar illumination. Such enhancements have been seen in Earth-based radar studies of the Galilean satellites [CAMPBELLETAL1978], Mars [MUHLEMANETAL1991], and Mercury [HARMONETAL1994]. Although Earth-based radar studies of the Moon [STACY1993] have had considerably better sensitivity and surface resolution than this experiment, the Clementine geometry uniquely allowed measurements as a function of the bistatic angle -- the separation angle between transmitter and receiver [SIMPSON1993]. The angular variation of any enhancement may be related to the distance the radar signal travels through the ice and, hence, to the thickness and/or clarity of the ice at 13 cm wavelength. Operational Considerations ========================== Descriptions given here are for nominal performance. The spacecraft transponder system comprised redundant units, each with slightly different characteristics. As transponder units age, their performance changes slightly. More importantly, the performance for Radio Science depended on operational factors such as the modulation state for the transmitters, which cannot be predicted in advance. The performance also depended on factors which were not always under the control of the Clementine Project. Clementine telemetry data were sent to Earth at user selectable rates between 125 bits per second and 125 kilobits per second. During radio occultation and radio scattering experiment periods, telemetry modulation was usually turned off; no special modulation requirements were applied during tracking for gravity field determination. Calibration =========== No information is available on calibration of the radio system. Measurements of the high gain antenna radiation pattern were made on an engineering model of the antenna. Operational Modes ================= The Clementine telecommunications system could be viewed as having two sections, which could be operated in the following modes: Section Mode ------------------------------------------- Oscillator two-way (coherent) one-way (non-coherent) RF output low-gain antenna high-gain antenna Selected parameters describing transponder performance are listed below: Oscillator Parameters: S-Band Two-Way Transponder Turnaround Ratio 240/221 One-Way Transmit Frequency (MHz) 2273.000 Frequency Uncertainty (+/- MHz) 0.002 Nominal Wavelength (cm) 13.2 RF Output parameters: S-Band RF Transmitter Power Output (w) 7 Low-Gain Antenna: Half-Power Half Beamwidth (deg) 70 Gain (dBi) 0 EIRP (dBm) 39 Polarization Right Circular High-Gain Antenna: Half-Power Half-Beamwidth (deg) 4.0 Gain (dBi) 26.2 EIRP (dBm) 65.2 Polarization Right Circular Instrument Overview - DSN ========================= Three Deep Space Communications Complexes (DSCCs) (near Barstow, CA; Canberra, Australia; and Madrid, Spain) comprise the DSN tracking network. Each complex is equipped with several antennas [including at least one each 70-m, 34-m High Efficiency (HEF), and 34-m standard (STD)], associated electronics, and operational systems. Primary activity at each complex is radiation of commands to and reception of telemetry data from active spacecraft. Transmission and reception is possible in several radio-frequency bands, the most common being S-band (nominally a frequency of 2100-2300 MHz or a wavelength of 14.2-13.0 cm) and X-band (7100-8500 MHz or 4.2- 3.5 cm). Transmitter output powers of up to 400 kw are available. Ground stations have the ability to transmit coded and uncoded waveforms which can be echoed by distant spacecraft. Analysis of the received coding allows navigators to determine the distance to the spacecraft; analysis of Doppler shift on the carrier signal allows estimation of the line-of-sight spacecraft velocity. Range and Doppler measurements are used to calculate the spacecraft trajectory and to infer gravity fields of objects near the spacecraft. Ground stations can record spacecraft signals that have propagated through or been scattered from target media. Measurements of signal parameters after wave interactions with surfaces, atmospheres, rings, and plasmas are used to infer physical and electrical properties of the target. Principal investigators vary from experiment to experiment. See the corresponding section of the spacecraft instrument description or the data set description for specifics. The Deep Space Network is managed by the Jet Propulsion Laboratory of the California Institute of Technology for the U.S. National Aeronautics and Space Administration. Specifications include: Instrument Id : RSS Instrument Host Id : DSN Pi Pds User Id : N/A Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : N/A Instrument Mass : N/A Instrument Length : N/A Instrument Width : N/A Instrument Height : N/A Instrument Manufacturer Name : N/A For more information on the Deep Space Network and its use in radio science investigations see the reports by [ASMAR&RENZETTI1993] and [ASMAR&HERRERA1993]. For design specifications on DSN subsystems see [DSN810-5]. For an example of use of the DSN for Radio Science see [TYLERETAL1992]. Subsystems - DSN ================ The Deep Space Communications Complexes (DSCCs) are an integral part of the Radio Science instrument, along with other receiving stations and the spacecraft Radio Frequency Subsystem. Their system performance directly determines the degree of success of Radio Science investigations, and their system calibration determines the degree of accuracy in the results of the experiments. The following paragraphs describe the functions performed by the individual subsystems of a DSCC. This material has been adapted from [ASMAR&HERRERA1993]; for additional information, consult [DSN810-5]. Each DSCC includes a set of antennas, a Signal Processing Center (SPC), and communication links to the Jet Propulsion Laboratory (JPL). The general configuration is illustrated below; antennas (Deep Space Stations, or DSS -- a term carried over from earlier times when antennas were individually instrumented) are listed in the table. -------- -------- -------- -------- -------- | DSS 12 | | DSS 18 | | DSS 14 | | DSS 15 | | DSS 16 | |34-m STD| |34-m STD| | 70-m | |34-m HEF| | 26-m | -------- -------- -------- -------- -------- | | | | | | v v | v | --------- | --------- --------->|GOLDSTONE|<---------- |EARTH/ORB| | SPC 10 |<-------------->| LINK | --------- --------- | SPC |<-------------->| 26-M | | COMM | ------>| COMM | --------- | --------- | | | v | v ------ --------- | --------- | NOCC |<--->| JPL |<------- | | ------ | CENTRAL | | GSFC | ------ | COMM | | NASCOMM | | MCCC |<--->| TERMINAL|<-------------->| | ------ --------- --------- ^ ^ | | CANBERRA (SPC 40) <---------------- | | MADRID (SPC 60) <---------------------- GOLDSTONE CANBERRA MADRID Antenna SPC 10 SPC 40 SPC 60 -------- --------- -------- -------- 26-m DSS 16 DSS 46 DSS 66 34-m STD DSS 12 DSS 42 DSS 61 DSS 18 DSS 48 DSS 68 34-m HEF DSS 15 DSS 45 DSS 65 70-m DSS 14 DSS 43 DSS 63 Developmental DSS 13 Subsystem interconnections at each DSCC are shown in the diagram below, and they are described in the sections that follow. The Monitor and Control Subsystem is connected to all other subsystems; the Test Support Subsystem can be. ----------- ------------------ --------- --------- |TRANSMITTER| | | | TRACKING| | COMMAND | | SUBSYSTEM |-| RECEIVER/EXCITER |-|SUBSYSTEM|-|SUBSYSTEM|- ----------- | | --------- --------- | | | SUBSYSTEM | | | | ----------- | | --------------------- | | MICROWAVE | | | | TELEMETRY | | | SUBSYSTEM |-| |-| SUBSYSTEM |- ----------- ------------------ --------------------- | | | ----------- ----------- --------- -------------- | | ANTENNA | | MONITOR | | TEST | | DIGITAL | | | SUBSYSTEM | |AND CONTROL| | SUPPORT | |COMMUNICATIONS|- ----------- | SUBSYSTEM | |SUBSYSTEM| | SUBSYSTEM | ----------- --------- -------------- DSCC Monitor and Control Subsystem ---------------------------------- The DSCC Monitor and Control Subsystem (DMC) is part of the Monitor and Control System (MON) which also includes the ground communications Central Communications Terminal and the Network Operations Control Center (NOCC) Monitor and Control Subsystem. The DMC is the center of activity at a DSCC. The DMC receives and archives most of the information from the NOCC needed by the various DSCC subsystems during their operation. Control of most of the DSCC subsystems, as well as the handling and displaying of any responses to control directives and configuration and status information received from each of the subsystems, is done through the DMC. The effect of this is to centralize the control, display, and archiving functions necessary to operate a DSCC. Communication between the various subsystems is done using a Local Area Network (LAN) hooked up to each subsystem via a network interface unit (NIU). DMC operations are divided into two separate areas: the Complex Monitor and Control (CMC) and the Link Monitor and Control (LMC). The primary purpose of the CMC processor for Radio Science support is to receive and store all predict sets transmitted from NOCC such as Radio Science, antenna pointing, tracking, receiver, and uplink predict sets and then, at a later time, to distribute them to the appropriate subsystems via the LAN. Those predict sets can be stored in the CMC for a maximum of three days under normal conditions. The CMC also receives, processes, and displays event/alarm messages; maintains an operator log; and produces tape labels for the DSP. Assignment and configuration of the LMCs is done through the CMC; to a limited degree the CMC can perform some of the functions performed by the LMC. There are two CMCs (one on-line and one backup) and three LMCs at each DSCC The backup CMC can function as an additional LMC if necessary. The LMC processor provides the operator interface for monitor and control of a link -- a group of equipment required to support a spacecraft pass. For Radio Science, a link might include the DSCC Spectrum Processing Subsystem (DSP) (which, in turn, can control the SSI), or the Tracking Subsystem. The LMC also maintains an operator log which includes operator directives and subsystem responses. One important Radio Science specific function that the LMC performs is receipt and transmission of the system temperature and signal level data from the PPM for display at the LMC console and for inclusion in Monitor blocks. These blocks are recorded on magnetic tape as well as appearing in the Mission Control and Computing Center (MCCC) displays. The LMC is required to operate without interruption for the duration of the Radio Science data acquisition period. The Area Routing Assembly (ARA), which is part of the Digital Communications Subsystem, controls all data communication between the stations and JPL. The ARA receives all required data and status messages from the LMC/CMC and can record them to tape as well as transmit them to JPL via data lines. The ARA also receives predicts and other data from JPL and passes them on to the CMC. DSCC Antenna Mechanical Subsystem --------------------------------- Multi-mission Radio Science activities require support from the 70-m, 34-m HEF, and 34-m STD antenna subnets. The antennas at each DSCC function as large-aperture collectors which, by double reflection, cause the incoming radio frequency (RF) energy to enter the feed horns. The large collecting surface of the antenna focuses the incoming energy onto a subreflector, which is adjustable in both axial and angular position. These adjustments are made to correct for gravitational deformation of the antenna as it moves between zenith and the horizon; the deformation can be as large as 5 cm. The subreflector adjustments optimize the channeling of energy from the primary reflector to the subreflector and then to the feed horns. The 70-m and 34-m HEF antennas have 'shaped' primary and secondary reflectors, with forms that are modified paraboloids. This customization allows more uniform illumination of one reflector by another. The 34-m STD primary reflectors are classical paraboloids, while the subreflectors are standard hyperboloids. On the 70-m and 34-m STD antennas, the subreflector directs received energy from the antenna onto a dichroic plate, a device which reflects S-band energy to the S-band feed horn and passes X-band energy through to the X-band feed horn. In the 34-m HEF, there is one 'common aperture feed,' which accepts both frequencies without requiring a dichroic plate. RF energy to be transmitted into space by the horns is focused by the reflectors into narrow cylindrical beams, pointed with high precision (either to the dichroic plate or directly to the subreflector) by a series of drive motors and gear trains that can rotate the movable components and their support structures. The different antennas can be pointed by several means. Two pointing modes commonly used during tracking passes are CONSCAN and 'blind pointing.' With CONSCAN enabled and a closed loop receiver locked to a spacecraft signal, the system tracks the radio source by conically scanning around its position in the sky. Pointing angle adjustments are computed from signal strength information (feedback) supplied by the receiver. In this mode the Antenna Pointing Assembly (APA) generates a circular scan pattern which is sent to the Antenna Control System (ACS). The ACS adds the scan pattern to the corrected pointing angle predicts. Software in the receiver-exciter controller computes the received signal level and sends it to the APA. The correlation of scan position with the received signal level variations allows the APA to compute offset changes which are sent to the ACS. Thus, within the capability of the closed-loop control system, the scan center is pointed precisely at the apparent direction of the spacecraft signal source. An additional function of the APA is to provide antenna position angles and residuals, antenna control mode/status information, and predict-correction parameters to the Area Routing Assembly (ARA) via the LAN, which then sends this information to JPL via the Ground Communications Facility (GCF) for antenna status monitoring. During periods when excessive signal level dynamics or low received signal levels are expected (e.g., during an occultation experiment), CONSCAN should not be used. Under these conditions, blind pointing (CONSCAN OFF) is used, and pointing angle adjustments are based on a predetermined Systematic Error Correction (SEC) model. Independent of CONSCAN state, subreflector motion in at least the z-axis may introduce phase variations into the received Radio Science data. For that reason, during certain experiments, the subreflector in the 70-m and 34-m HEFs may be frozen in the z-axis at a position (often based on elevation angle) selected to minimize phase change and signal degradation. This can be done via Operator Control Inputs (OCIs) from the LMC to the Subreflector Controller (SRC) which resides in the alidade room of the antennas. The SRC passes the commands to motors that drive the subreflector to the desired position. Unlike the 70-m and 34-m HEFs which have azimuth-elevation (AZ-EL) drives, the 34-m STD antennas use (hour angle-declination) HA-DEC drives. The same positioning of the subreflector on the 34-m STD does not create the same effect as on the 70-m and 34-m HEFs. Pointing angles for all three antenna types are computed by the NOCC Support System (NSS) from an ephemeris provided by the flight project. These predicts are received and archived by the CMC. Before each track, they are transferred to the APA, which transforms the direction cosines of the predicts into AZ-EL coordinates for the 70-m and 34-m HEFs or into HA-DEC coordinates for the 34-m STD antennas. The LMC operator then downloads the antenna AZ-EL or HA-DEC predict points to the antenna-mounted ACS computer along with a selected SEC model. The pointing predicts consist of time-tagged AZ-EL or HA-DEC points at selected time intervals along with polynomial coefficients for interpolation between points. The ACS automatically interpolates the predict points, corrects the pointing predicts for refraction and subreflector position, and adds the proper systematic error correction and any manually entered antenna offsets. The ACS then sends angular position commands for each axis at the rate of one per second. In the 70-m and 34-m HEF, rate commands are generated from the position commands at the servo controller and are subsequently used to steer the antenna. In the 34-m STD antennas motors, rather than servos, are used to steer the antenna; there is no feedback once the 34-m STD has been told where to point. When not using binary predicts (the routine mode for spacecraft tracking), the antennas can be pointed using 'planetary mode' -- a simpler mode which uses right ascension (RA) and declination (DEC) values. These change very slowly with respect to the celestial frame. Values are provided to the station in text form for manual entry. The ACS quadratically interpolates among three RA and DEC points which are on one-day centers. A third pointing mode -- sidereal -- is available for tracking radio sources fixed with respect to the celestial frame. Regardless of the pointing mode being used, a 70-m antenna has a special high-accuracy pointing capability called 'precision' mode. A pointing control loop derives the main AZ-EL pointing servo drive error signals from a two- axis autocollimator mounted on the Intermediate Reference Structure. The autocollimator projects a light beam to a precision mirror mounted on the Master Equatorial drive system, a much smaller structure, independent of the main antenna, which is exactly positioned in HA and DEC with shaft encoders. The autocollimator detects elevation/cross- elevation errors between the two reference surfaces by measuring the angular displacement of the reflected light beam. This error is compensated for in the antenna servo by moving the antenna in the appropriate AZ-EL direction. Pointing accuracies of 0.004 degrees (15 arc seconds) are possible in 'precision' mode. The 'precision' mode is not available on 34-m antennas -- nor is it needed, since their beamwidths are twice as large as on the 70-m antennas. DSCC Antenna Microwave Subsystem -------------------------------- 70-m Antennas: Each 70-m antenna has three feed cones installed in a structure at the center of the main reflector. The feeds are positioned 120 degrees apart on a circle. Selection of the feed is made by rotation of the subreflector. A dichroic mirror assembly, half on the S-band cone and half on the X-band cone, permits simultaneous use of the S- and X-band frequencies. The third cone is devoted to R&D and more specialized work. The Antenna Microwave Subsystem (AMS) accepts the received S- and X-band signals at the feed horn and transmits them through polarizer plates to an orthomode transducer. The polarizer plates are adjusted so that the signals are directed to a pair of redundant amplifiers for each frequency, thus allowing simultaneous reception of signals in two orthogonal polarizations. For S-band these are two Block IVA S-band Traveling Wave Masers (TWMs); for X-band the amplifiers are Block IIA TWMs. 34-m STD Antennas: These antennas have two feed horns, one for S-band signals and one for X-band. The horns are mounted on a cone which is fixed in relation to the subreflector. A dichroic plate mounted above the horns directs energy from the subreflector into the proper horn. The AMS directs the received S- and X-band signals through polarizer plates and on to amplification. There are two Block III S-band TWMs and two Block I X-band TWMs. 34-m HEF Antennas: Unlike the other antennas, the 34-m HEF uses a single feed for both S- and X-band. Simultaneous S- and X-band receive as well as X-band transmit is possible thanks to the presence of an S/X 'combiner' which acts as a diplexer. For S-band, RCP or LCP is user selected through a switch so neither a polarizer nor an orthomode transducer is needed. X-band amplification options include two Block II TWMs or an HEMT Low Noise Amplifier (LNA). S-band amplification is provided by an FET LNA. DSCC Receiver-Exciter Subsystem ------------------------------- The Receiver-Exciter Subsystem is composed of three groups of equipment: the closed-loop receiver group, the open-loop receiver group, and the RF monitor group. This subsystem is controlled by the Receiver-Exciter Controller (REC) which communicates directly with the DMC for predicts and OCI reception and status reporting. The exciter generates the S-band signal (or X-band for the 34-m HEF only) which is provided to the Transmitter Subsystem for the spacecraft uplink signal. It is tunable under command of the Digitally Controlled Oscillator (DCO) which receives predicts from the Metric Data Assembly (MDA). The diplexer in the signal path between the transmitter and the feed horn for all three antennas (used for simultaneous transmission and reception) may be configured such that it is out of the received signal path (in listen-only or bypass mode) in order to improve the signal-to-noise ratio in the receiver system. Closed Loop Receivers: The Block IV receiver-exciter at the 70-m stations allows for two receiver channels, each capable of L-Band (e.g., 1668 MHz frequency or 18 cm wavelength), S-band, or X-band reception, and an S-band exciter for generation of uplink signals through the low-power or high-power transmitter. The Block III receiver-exciter at the 34-m STD stations allows for two receiver channels, each capable of S-band or X-band reception and an exciter used to generate an uplink signal through the low-power transmitter. The receiver-exciter at the 34-m HEF stations allows for one channel only. The closed-loop receivers provide the capability for rapid acquisition of a spacecraft signal and telemetry lockup. In order to accomplish acquisition within a short time, the receivers are predict driven to search for, acquire, and track the downlink automatically. Rapid acquisition precludes manual tuning though that remains as a backup capability. The subsystem utilizes FFT analyzers for rapid acquisition. The predicts are NSS generated, transmitted to the CMC which sends them to the Receiver-Exciter Subsystem where two sets can be stored. The receiver starts acquisition at uplink time plus one round-trip-light-time or at operator specified times. The receivers may also be operated from the LMC without a local operator attending them. The receivers send performance and status data, displays, and event messages to the LMC. Either the exciter synthesizer signal or the simulation (SIM) synthesizer signal is used as the reference for the Doppler extractor in the closed-loop receiver systems, depending on the spacecraft being tracked (and Project guidelines). The SIM synthesizer is not ramped; instead it uses one constant frequency, the Track Synthesizer Frequency (TSF), which is an average frequency for the entire pass. The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. It will be configured such that the expected amplitude changes are accommodated with minimum distortion. The loop bandwidth (2BLo) will be configured such that the expected phase changes can be accommodated while maintaining the best possible loop SNR. Open-Loop Receivers: The Radio Science Open-Loop Receiver (OLR) is a dedicated four channel, narrow-band receiver which provides amplified and downconverted video band signals to the DSCC Spectrum Processing Subsystem (DSP). The OLR utilizes a fixed first Local Oscillator (LO) frequency and a tunable second LO frequency to minimize phase noise and improve frequency stability. The OLR consists of an RF-to-IF downconverter located in the antenna, an IF selection switch (IVC), and a Radio Science IF-VF downconverter (RIV) located in the SPC. The RF-IF downconverters in the 70-m antennas are equipped for four IF channels: S-RCP, S-LCP, X-RCP, and X-LCP. The 34-m HEF stations are equipped with a two-channel RF-IF: S-band and X-band. The IVC switches the IF input between the 70-m and 34-m HEF antennas. The RIV contains the tunable second LO, a set of video bandpass filters, IF attenuators, and a controller (RIC). The LO tuning is done via DSP control of the POCA/PLO combination based on a predict set. The POCA is a Programmable Oscillator Control Assembly and the PLO is a Programmable Local Oscillator (commonly called the DANA synthesizer). The bandpass filters are selectable via the DSP. The RIC provides an interface between the DSP and the RIV. It is controlled from the LMC via the DSP. The RIC selects the filter and attenuator settings and provides monitor data to the DSP. The RIC could also be manually controlled from the front panel in case the electronic interface to the DSP is lost. RF Monitor -- SSI and PPM: The RF monitor group of the Receiver-Exciter Subsystem provides spectral measurements using the Spectral Signal Indicator (SSI) and measurements of the received channel system temperature and spacecraft signal level using the Precision Power Monitor (PPM). The SSI provides a local display of the received signal spectrum at a dedicated terminal at the DSCC and routes these same data to the DSP which routes them to NOCC for remote display at JPL for real-time monitoring and RIV/DSP configuration verification. These displays are used to validate Radio Science Subsystem data at the DSS, NOCC, and Mission Support Areas. The SSI configuration is controlled by the DSP and a duplicate of the SSI spectrum appears on the LMC via the DSP. During real-time operations the SSI data also serve as a quick-look science data type for Radio Science experiments. The PPM measures system noise temperatures (SNT) using a Noise Adding Radiometer (NAR) and downlink signal levels using the Signal Level Estimator (SLE). The PPM accepts its input from the closed-loop receiver. The SNT is measured by injecting known amounts of noise power into the signal path and comparing the total power with the noise injection 'on' against the total power with the noise injection 'off.' That operation is based on the fact that receiver noise power is directly proportional to temperature; thus measuring the relative increase in noise power due to the presence of a calibrated thermal noise source allows direct calculation of SNT. Signal level is measured by calculating an FFT to estimate the SNR between the signal level and the receiver noise floor where the power is known from the SNT measurements. There is one PPM controller at the SPC which is used to control all SNT measurements. The SNT integration time can be selected to represent the time required for a measurement of 30K to have a one-sigma uncertainty of 0.3K or 1%. DSCC Transmitter Subsystem -------------------------- The Transmitter Subsystem accepts the S-band frequency exciter signal from the Block III or Block IV Receiver- Exciter Subsystem exciter and amplifies it to the required transmit output level. The amplified signal is routed via the diplexer through the feed horn to the antenna and then focused and beamed to the spacecraft. The Transmitter Subsystem power capabilities range from 18 kw to 400 kw. Power levels above 18 kw are available only at 70-m stations. DSCC Tracking Subsystem ----------------------- The Tracking Subsystem primary functions are to acquire and maintain communications with the spacecraft and to generate and format radiometric data containing Doppler and range. The DSCC Tracking Subsystem (DTK) receives the carrier signals and ranging spectra from the Receiver-Exciter Subsystem. The Doppler cycle counts are counted, formatted, and transmitted to JPL in real time. Ranging data are also transmitted to JPL in real time. Also contained in these blocks is the AGC information from the Receiver-Exciter Subsystem. The Radio Metric Data Conditioning Team (RMDCT) at JPL produces an Archival Tracking Data File (ATDF) tape which contains Doppler and ranging data. In addition, the Tracking Subsystem receives from the CMC frequency predicts (used to compute frequency residuals and noise estimates), receiver tuning predicts (used to tune the closed-loop receivers), and uplink tuning predicts (used to tune the exciter). From the LMC, it receives configuration and control directives as well as configuration and status information on the transmitter, microwave, and frequency and timing subsystems. The Metric Data Assembly (MDA) controls all of the DTK functions supporting the uplink and downlink activities. The MDA receives uplink predicts and controls the uplink tuning by commanding the DCO. The MDA also controls the Sequential Ranging Assembly (SRA). It formats the Doppler and range measurements and provides them to the GCF for transmission to NOCC. The Sequential Ranging Assembly (SRA) measures the round trip light time (RTLT) of a radio signal traveling from a ground tracking station to a spacecraft and back. From the RTLT, phase, and Doppler data, the spacecraft range can be determined. A coded signal is modulated on an uplink carrier and transmitted to the spacecraft where it is detected and transponded back to the ground station. As a result, the signal received at the tracking station is delayed by its round trip through space and shifted in frequency by the Doppler effect due to the relative motion between the spacecraft and the tracking station on Earth. DSCC Spectrum Processing Subsystem (DSP) ---------------------------------------- The DSCC Spectrum Processing Subsystem (DSP) located at the SPC digitizes and records on magnetic tapes the narrowband output data from the RIV. It consists of a Narrow Band Occultation Converter (NBOC) containing four Analog-to- Digital Converters (ADCs), a ModComp CLASSIC computer processor called the Spectrum Processing Assembly (SPA), and two to six magnetic tape drives. Magnetic tapes are known as Original Data Records (ODRs). Electronic near real-time transmission of data to JPL (an Original Data Stream, or ODS) may be possible in certain circumstances; The DSP is operated through the LMC. Using the SPA-R software, the DSP allows for real-time frequency and time offsets (while in RUN mode) and, if necessary, snap tuning between the two frequency ranges transmitted by the spacecraft: coherent and non-coherent. The DSP receives Radio Science frequency predicts from the CMC, allows for multiple predict set archiving (up to 60 sets) at the SPA, and allows for manual predict generation and editing. It accepts configuration and control data from the LMC, provides display data to the LMC, and transmits the signal spectra from the SSI as well as status information to NOCC and the Project Mission Support Area (MSA) via the GCF data lines. The DSP records the digitized narrowband samples and the supporting header information (i.e., time tags, POCA frequencies, etc.) on 9-track magnetic tapes in 6250 or 1600 bpi GCR format. Through the DSP-RIC interface the DSP controls the RIV filter selection and attenuation levels. It also receives RIV performance monitoring via the RIC. In case of failure of the DSP-RIC interface, the RIV can be controlled manually from the front panel. All the RIV and DSP control parameters and configuration directives are stored in the SPA in a macro-like file called an 'experiment directive' table. A number of default directives exist in the DSP for the major Radio Science experiments. Operators can create their own table entries. Items such as verification of the configuration of the prime open-loop recording subsystem, the selection of the required predict sets, and proper system performance prior to the recording periods will be checked in real-time at JPL via the NOCC displays using primarily the remote SSI display at NOCC and the NRV displays. Because of this, transmission of the DSP/SSI monitor information is enabled prior to the start of recording. The specific run time and tape recording times will be identified in the Sequence of Events (SOE) and/or DSN Keyword File. The DSP can be used to duplicate ODRs. It also has the capability to play back a certain section of the recorded data after conclusion of the recording periods. DSCC Frequency and Timing Subsystem ----------------------------------- The Frequency and Timing Subsystem (FTS) provides all frequency and timing references required by the other DSCC subsystems. It contains four frequency standards of which one is prime and the other three are backups. Selection of the prime standard is done via the CMC. Of these four standards, two are hydrogen masers followed by clean-up loops (CUL) and two are cesium standards. These four standards all feed the Coherent Reference Generator (CRG) which provides the frequency references used by the rest of the complex. It also provides the frequency reference to the Master Clock Assembly (MCA) which in turn provides time to the Time Insertion and Distribution Assembly (TID) which provides UTC and SIM-time to the complex. JPL's ability to monitor the FTS at each DSCC is limited to the MDA calculated Doppler pseudo-residuals, the Doppler noise, the SSI, and to a system which uses the Global Positioning System (GPS). GPS receivers at each DSCC receive a one-pulse-per-second pulse from the station's (hydrogen maser referenced) FTS and a pulse from a GPS satellite at scheduled times. After compensating for the satellite signal delay, the timing offset is reported to JPL where a database is kept. The clock offsets stored in the JPL database are given in microseconds; each entry is a mean reading of measurements from several GPS satellites and a time tag associated with the mean reading. The clock offsets provided include those of SPC 10 relative to UTC (NIST), SPC 40 relative to SPC 10, etc. Optics - DSN ============ Performance of DSN ground stations depends primarily on size of the antenna and capabilities of electronics. These are summarized in the following set of tables. Note that 64-m antennas were upgraded to 70-m between 1986 and 1989. Beamwidth is half-power full angular width. Polarization is circular; L denotes left circular polarization (LCP), and R denotes right circular polarization (RCP). DSS S-Band Characteristics 64-m 70-m 34-m 34-m Transmit STD HEF -------- ----- ----- ----- ----- Frequency (MHz) 2110- 2110- 2025- N/A 2120 2120 2120 Wavelength (m) 0.142 0.142 0.142 N/A Ant Gain (dBi) 62.7 55.2 N/A Beamwidth (deg) 0.119 0.31 N/A Polarization L or R L or R N/A Tx Power (kW) 20-400 20 N/A Receive ------- Frequency (MHz) 2270- 2270- 2270- 2200- 2300 2300 2300 2300 Wavelength (m) 0.131 0.131 0.131 0.131 Ant Gain (dBi) 61.6 63.3 56.2 56.0 Beamwidth (deg) 0.108 0.27 0.24 Polarization L & R L & R L or R L or R System Temp (K) 22 20 22 38 DSS X-Band Characteristics 64-m 70-m 34-m 34-m Transmit STD HEF -------- ----- ----- ----- ----- Frequency (MHz) 8495 8495 N/A 7145- 7190 Wavelength (m) 0.035 0.035 N/A 0.042 Ant Gain (dBi) 74.2 N/A 67 Beamwidth (deg) N/A 0.074 Polarization L or R L or R N/A L or R Tx Power (kW) 360 360 N/A 20 Receive ------- Frequency (MHz) 8400- 8400- 8400- 8400- 8500 8500 8500 8500 Wavelength (m) 0.036 0.036 0.036 0.036 Ant Gain (dBi) 71.7 74.2 66.2 68.3 Beamwidth (deg) 0.031 0.075 0.063 Polarization L & R L & R L & R L & R System Temp (K) 27 20 25 20 Electronics - DSN ================= DSCC Open-Loop Receiver ----------------------- The open loop receiver block diagram shown below is for 70-m antenna sites as used for Clementine observations. This was a non-standard configuration required to capture the Clementine signal, which was about 22 MHz below normal NASA transmitter frequencies [DEVEREAUX1994]. The 2273 MHz Clementine S-band signals are first downconverted to approximately 273 MHz. Then a second mixer upconverts the signals to 318 MHz, where they can be fed into the normal X-band receiver chain. Conversion loss through the added mixer is about 6 dB. Based on a tuning prediction file, the POCA controls the DANA synthesizer the output of which (after multiplication) mixes the input signals at 318 MHz to fixed intermediate frequencies for subsequent amplification. These signals in turn are down converted and passed through additional filters until they yield baseband output of up to 25 kHz in width. The baseband output is digitally sampled by the DSP and either written to magnetic tape or electronically transferred for further analysis. S-Band X-Band 2273 MHz 8415 MHz Input 45 Input | MHz | v | v --- --- | --- --- | X |<--|x20|<--100 MHz | 100 MHz-->|x81|-->| X | --- --- V --- --- 273| --- MHz ----------------------| X |--------------------- 318 --- |MHz v --- -- 115 MHz 33.4545 MHz --- --- | X |<--|x3|<------ ------>|x11|-->| X | --- -- | | --- --- | | | | | | | | 50| 72.8181 --- --- |50 MHz| MHz->| X | | X |<-10 MHz |MHz v --- --- v --- ^ ^ --- | X |<--60 MHz | | 60 MHz-->| X | --- | | --- | 9.9 | 43.4545 MHz | 9.9 | | MHz ------------- MHz | | | ^ | | 10| v | v |10 MHz| --- ----------- --- |MHz |------>| X | | DANA | | X |<------| | --- |Synthesizer| --- | | | ----------- | | v v ^ v v ------- ------- | ------- ------- |Filters| |Filters| ---------- |Filters| |Filters| |3,4,5,6| | 1,2 | | POCA | | 1,2 | |3,4,5,6| ------- ------- |Controller| ------- ------- | | ---------- | | 10| |0.1 0.1| |10 MHz| |MHz MHz| |MHz v v v v --- --- --- --- | X |- -| X | | X |- -| X | --- | | --- --- | | --- ^ | | ^ ^ | | ^ | | | | | | | | 10 | | 0.1 0.1 | | 10 MHz | | MHz MHz | | MHz | | | | v v v v Baseband Baseband Output Output Reconstruction of the antenna frequency from the frequency of the signal in the recorded data can be achieved through use of one of the following formula. Radio Science IF-VF (RIV) Converter Assembly at 70-m antenna for Clementine: FSant = 11*[POCA - 10^7] + 1.905*10^9 - 3*Fsamp + Frec where FSant = S-band antenna frequency POCA = POCA frequency Fsamp = sampling frequency Frec = frequency of recorded signal Filters - DSN ============= DSCC Open-Loop Receiver ----------------------- Nominal filter center frequencies and bandwidths for the Open-Loop Receivers are shown in the table below. Filter Center Frequency 3 dB Bandwidth ------ ---------------- -------------- 1 0.1 MHz 90 Hz 2 0.1 MHz 450 Hz 3 10.0 MHz 2000 Hz 4 10.0 MHz 1700 Hz (S-band) 6250 Hz (X-band) 5 10.0 MHz 45000 Hz 6 10.0 MHz 21000 Hz Detectors - DSN =============== DSCC Open-Loop Receivers ------------------------ Open-loop receiver output is detected in software by the radio science investigator. DSCC Closed-Loop Receivers -------------------------- Nominal carrier tracking loop threshold noise bandwidth at both S- and X-band is 10 Hz. Coherent (two-way) closed-loop system stability is shown in the table below: integration time Doppler uncertainty (secs) (one sigma, microns/sec) ------ ------------------------ 10 50 60 20 1000 4 Calibration - DSN ================= Calibrations of hardware systems are carried out periodically by DSN personnel; these ensure that systems operate at required performance levels -- for example, that antenna patterns, receiver gain, propagation delays, and Doppler uncertainties meet specifications. No information on specific calibration activities is available. Nominal performance specifications are shown in the tables above. Additional information may be available in [DSN810-5]. Prior to each tracking pass, station operators perform a series of calibrations to ensure that systems meet specifications for that operational period. Included in these calibrations is measurement of receiver system temperature in the configuration to be employed during the pass. Results of these calibrations are recorded in (hard copy) Controller's Logs for each pass. The nominal procedure for initializing open-loop receiver attenuator settings is described below. In cases where widely varying signal levels are expected, the procedure may be modified in advance or real-time adjustments may be made to attenuator settings. Open-Loop Receiver Attenuation Calibration ------------------------------------------ The open-loop receiver attenuator calibrations are performed to establish the output of the open-loop receivers at a level that will not saturate the analog-to-digital converters. To achieve this, the calibration is done using a test signal generated by the exciter/translator that is set to the peak predicted signal level for the upcoming pass. Then the output level of the receiver's video band spectrum envelope is adjusted to the level determined by equation (3) below (to five-sigma). Note that the SNR in the equation (2) is in dB while the SNR in equation (3) is linear. Pn = -198.6 + 10*log(SNT) + 10*log(1.2*Fbw) (1) SNR = Ps - Pn (SNR in dB) (2) Vrms = sqrt(SNR + 1)/[1 + 0.283*sqrt(SNR)] (SNR linear) (3) where Fbw = receiver filter bandwidth (Hz) Pn = receiver noise power (dBm) Ps = signal power (dBm) SNT = system noise temperature (K) SNR = predicted signal-to-noise ratio Operational Considerations - DSN ================================ The DSN is a complex and dynamic 'instrument.' Its performance for Radio Science depends on a number of factors from equipment configuration to meteorological conditions. No specific information on 'operational considerations' can be given here. Operational Modes - DSN ======================= DSCC Antenna Mechanical Subsystem --------------------------------- Pointing of DSCC antennas may be carried out in several ways. For details see the subsection 'DSCC Antenna Mechanical Subsystem' in the 'Subsystem' section. Binary pointing is the preferred mode for tracking spacecraft; pointing predicts are provided, and the antenna simply follows those. With CONSCAN, the antenna scans conically about the optimum pointing direction, using closed-loop receiver signal strength estimates as feedback. In planetary mode, the system interpolates from three (slowly changing) RA-DEC target coordinates; this is 'blind' pointing since there is no feedback from a detected signal. In sidereal mode, the antenna tracks a fixed point on the celestial sphere. In 'precision' mode, the antenna pointing is adjusted using an optical feedback system. It is possible on most antennas to freeze z-axis motion of the subreflector to minimize phase changes in the received signal. DSCC Receiver-Exciter Subsystem ------------------------------- The diplexer in the signal path between the transmitter and the feed horns on all three antennas may be configured so that it is out of the received signal path in order to improve the signal-to-noise ratio in the receiver system. This is known as the 'listen-only' or 'bypass' mode. Closed-Loop vs. Open-Loop Reception ----------------------------------- Radio Science data can be collected in two modes: closed- loop, in which a phase-locked loop receiver tracks the spacecraft signal, or open-loop, in which a receiver samples and records a band within which the desired signal presumably resides. Closed-loop data are collected using Closed-Loop Receivers, and open-loop data are collected using Open-Loop Receivers in conjunction with the DSCC Spectrum Processing Subsystem (DSP). See the Subsystems section for further information. Closed-Loop Receiver AGC Loop ----------------------------- The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. Ordinarily it is configured so that expected signal amplitude changes are accommodated with minimum distortion. The loop bandwidth is ordinarily configured so that expected phase changes can be accommodated while maintaining the best possible loop SNR. Coherent vs. Non-Coherent Operation ----------------------------------- The frequency of the signal transmitted from the spacecraft can generally be controlled in two ways -- by locking to a signal received from a ground station or by locking to an on-board oscillator. These are known as the coherent (or 'two-way') and non-coherent ('one-way') modes, respectively. Mode selection is made at the spacecraft, based on commands received from the ground. When operating in the coherent mode, the transponder carrier frequency is derived from the received uplink carrier frequency with a 'turn-around ratio' typically of 240/221. In the non-coherent mode, the downlink carrier frequency is derived from the spacecraft on-board crystal-controlled oscillator. Either closed-loop or open-loop receivers (or both) can be used with either spacecraft frequency reference mode. Closed-loop reception in two-way mode is usually preferred for routine tracking. Occasionally the spacecraft operates coherently while two ground stations receive the 'downlink' signal; this is sometimes known as the 'three-way' mode. DSCC Spectrum Processing Subsystem (DSP) ---------------------------------------- The DSP can operate in four sampling modes with from 1 to 4 input signals. Input channels are assigned to ADC inputs during DSP configuration. Modes and sampling rates are summarized in the tables below: Mode Analog-to-Digital Operation ---- ---------------------------- 1 4 signals, each sampled by a single ADC 2 1 signal, sampled sequentially by 4 ADCs 3 2 signals, each sampled sequentially by 2 ADCs 4 1 signal, sampled sequentially by 3 ADCs 8-bit Samples 12-bit Samples Sampling Rates Sampling Rates (samples/sec per ADC) (samples/sec per ADC) --------------------- --------------------- 50000 31250 25000 15625 12500 10000 10000 6250 5000 5000 4000 3125 2500 2000 1250 1000 1000 500 400 250 200 200 Location - DSN ============== Station locations are documented in [GEO-10REVD]. Geocentric coordinates are summarized here. Geocentric Geocentric Geocentric Station Radius (km) Latitude (N) Longitude (E) --------- ----------- ------------ ------------- Goldstone DSS 12 (34-m STD) 6371.997815 35.1186672 243.1945048 DSS 13 (develop) 6372.117062 35.0665485 243.2051077 DSS 14 (70-m) 6371.992867 35.2443514. 243.1104584 DSS 15 (34-m HEF) 6371.9463 35.2402863 243.1128186 DSS 16 (26-m) 6371.9608 35.1601436 243.1264200 DSS 18 (34-m STD) UNK UNK UNK Canberra DSS 42 (34-m STD) 6371.675607 -35.2191850 148.9812546 DSS 43 (70-m) 6371.688953 -35.2209308 148.9812540 DSS 45 (34-m HEF) 6371.692 -35.21709 148.97757 DSS 46 (26-m) 6371/675 -35.22360 148.98297 DSS 48 (34-m STD) UNK UNK UNK Madrid DSS 61 (34-m STD) 6370.027734 40.2388805 355.7509634 DSS 63 (70-m) 6370.051015 40.2413495 355.7519776 DSS 65 (34-m HEF) 6370.021370 40.2372843 355.7485968 DSS 66 (26-m) 6370.036 40.2400714 355.7485976 DSS 48 (34-m STD) UNK UNK UNK Measurement Parameters - DSN ============================ Open-Loop System ---------------- Output from the Open-Loop Receivers (OLRs), as sampled and recorded by the DSCC Spectrum Processing Subsystem (DSP), is a stream of 8- or 12-bit quantized voltage samples. The nominal input to the Analog-to-Digital Converters (ADCs) is +/-10 volts, but the precise scaling between input voltages and output digitized samples is usually irrelevant for analysis; the digital data are generally referenced to a known noise or signal level within the data stream itself -- for example, the thermal noise output of the radio receivers which has a known system noise temperature (SNT). Raw samples comprise the data block in each DSP record; a header record (presently 83 16-bit words) contains ancillary information such as: time tag for the first sample in the data block RMS values of receiver signal levels and ADC outputs POCA frequency and drift rate Closed-Loop System ------------------ Closed-loop data are recorded in Archival Tracking Data Files (ATDFs), as well as certain secondary products such as the Orbit Data File (ODF). The ATDF Tracking Logical Record contains 117 entries including status information and measurements of ranging, Doppler, and signal strength. ACRONYMS AND ABBREVIATIONS - DSN ================================ ACS Antenna Control System ADC Analog-to-Digital Converter AMS Antenna Microwave System APA Antenna Pointing Assembly ARA Area Routing Assembly ATDF Archival Tracking Data File AZ Azimuth CMC Complex Monitor and Control CONSCAN Conical Scanning (antenna pointing mode) CRG Coherent Reference Generator CUL Clean-up Loop DANA a type of frequency synthesizer dB deciBel dBi dB relative to isotropic DCO Digitally Controlled Oscillator DEC Declination deg degree DMC DSCC Monitor and Control Subsystem DSCC Deep Space Communications Complex DSN Deep Space Network DSP DSCC Spectrum Processing Subsystem DSS Deep Space Station DTK DSCC Tracking Subsystem E east EL Elevation FTS Frequency and Timing Subsystem GCF Ground Communications Facility GPS Global Positioning System HA Hour Angle HEF High-Efficiency (as in 34-m HEF antennas) IF Intermediate Frequency IVC IF Selection Switch JPL Jet Propulsion Laboratory K Kelvin km kilometer kW kilowatt L-band approximately 1668 MHz LAN Local Area Network LCP Left-Circularly Polarized LMC Link Monitor and Control LNA Low-Noise Amplifier LO Local Oscillator m meters MCA Master Clock Assembly MCCC Mission Control and Computing Center MDA Metric Data Assembly MHz Megahertz MON Monitor and Control System MSA Mission Support Area N north NAR Noise Adding Radiometer NBOC Narrow-Band Occultation Converter NIST SPC 10 time relative to UTC NIU Network Interface Unit NOCC Network Operations and Control System NSS NOCC Support System OCI Operator Control Input ODF Orbit Data File ODR Original Data Record ODS Original Data Stream OLR Open Loop Receiver POCA Programmable Oscillator Control Assembly PPM Precision Power Monitor RA Right Ascension REC Receiver-Exciter Controller RCP Right-Circularly Polarized RF Radio Frequency RIC RIV Controller RIV Radio Science IF-VF Converter Assembly RMDCT Radio Metric Data Conditioning Team RTLT Round-Trip Light Time S-band approximately 2100-2300 MHz sec second SEC System Error Correction SIM Simulation SLE Signal Level Estimator SNR Signal-to-Noise Ratio SNT System Noise Temperature SOE Sequence of Events SPA Spectrum Processing Assembly SPC Signal Processing Center SRA Sequential Ranging Assembly SRC Sub-Reflector Controller SSI Spectral Signal Indicator STD Standard (as in 34-m STD antennas) TID Time Insertion and Distribution Assembly TSF Tracking Synthesizer Frequency TWM Traveling Wave Maser UNK unknown UTC Universal Coordinated Time VF Video Frequency X-band approximately 7800-8500 MHz
  • data set : LRO MOON RADIO SCIENCE RAW TRACKING DATA V1.0
    Raw radio science data and ancillary files from the Lunar Reconnaissance Orbiter mission.
  • investigation : MAGELLAN
    Mission Overview ================ The Magellan spacecraft was launched from the Kennedy Space Center on 4 May 1989. The spacecraft was deployed from the Shuttle cargo bay after the Shuttle achieved parking orbit. Magellan, using an inertial upper stage rocket, was then placed into a Type IV transfer orbit to Venus where it carried out radar mapping and gravity studies starting in August 1990. The Mission has been described in many papers including two special issues of the Journal of Geophysical Research [VRMPP1983; SAUNDERSETAL1990; JGRMGN1992]. The radar system is also described in [JOHNSON1990]. Magellan was powered by single degree of freedom, sun-tracking, solar panels. The spacecraft was 3-axis stabilized by reaction wheels using gyros and a star sensor for attitude reference. The spacecraft carried a solid rocket motor for Venus orbit insertion. A small hydrazine system was used for trajectory corrections and certain attitude control functions. Earth communication with the Deep Space Network (DSN) was by means of S- and X-band channels. The high-gain antenna also functioned as the SAR mapping antenna during orbital operations. The interplanetary cruise phase lasted until 10 August 1990. During the cruise phase there were small trajectory correction maneuvers to ensure proper approach geometry. Using the solid rocket motor, the spacecraft was placed into an elliptical orbit around the planet, with a periapsis latitude of approximately 10 degrees north, a periapsis altitude of 295 km, a period of 3.263 hours, and an apoapsis altitude of approximately 7762 km. After orbit insertion, the radar system acquired test data. Then, unexpectedly, the signal from the spacecraft was lost twice. Following an intense recovery process, commands were sent to avoid further communication interruptions, and the spacecraft resumed mapping operations on 15 September 1990. Each mapping cycle lasted 243 days, which was the time required for Venus to make one rotation under the spacecraft orbit. The first mapping cycle ended on 15 May 1991. Typical activities during a single mapping pass on Cycle 1 were as follows. As the spacecraft neared periapsis, it was oriented so the high-gain antenna pointed slightly to the side of the ground track. At a true anomaly of -59 degrees, the radar was commanded on. The radar continued to take data to a true anomaly of 80 degrees and then the radar was commanded off. On the next pass the swath started at -80 degrees and went to 59 degrees. Alternating north and south swaths were repeated throughout Cycle 1. The range of latitudes covered by the synthetic aperture radar (SAR) during Cycle 1 was 67 degrees S to 90 degrees N. The range of SAR incidence angles was from just under 20 to just over 40 degrees. The SAR data were taken at a data rate of 750 kilobits/second and were stored in the spacecraft tape recorder. Altimeter and radiometer data were also taken when SAR data were acquired. The altimeter data were taken using a small fan beam antenna at a data rate of 30 kb/s. As the spacecraft moved away from the planet toward apoapsis, the spacecraft reoriented the high-gain antenna towards Earth and the stored radar data were transmitted to DSN stations. This data taking- and transmitting-cycle was repeated for every orbit. By 15 May 1991, the planet had been completely mapped except for the area near the South Pole and a few regions which had been missed because of temporary equipment failures. Cycle 2 observations focused on filling the gaps in Cycle 1 coverage (including the south pole area), acquiring SAR data at a constant incidence angle (25 degrees), and conducting a suite of ad hoc experiments, including high resolution imaging and radar stereo. To observe the south pole the spacecraft was rotated 180 degrees about its nadir-pointing axis so as to conduct right-looking SAR observations. Gaps in the Cycle 1 coverage were filled by rotating the spacecraft back to its initial left-looking direction. The orbit plane was adjusted slightly at the beginning of Cycle 2 so that altimetry tracks would be offset by about 10 km at the equator, bisecting the orbit-to-orbit offset of altimetry tracks in Cycle 1. The spacecraft was rotated 90 deg about the HGA boresight on orbits 3716-3719 to obtain SAR and radiometry data with VV polarization. Radio occultation measurements were made on orbits 3212-3214. The principal objective of Cycle 3 was to perform radar stereo mapping of the Venusian surface. About 30 percent of the Cycle 1 coverage was remapped in this cycle with a different, left-looking incidence angle on the surface. Gravity data were collected over Artemis Chasma. In addition, high resolution altimetry data were collected by pointing the high gain antenna straight down during orbits 4919 to 4921. Transmission of acquired radar data to Earth became nearly impossible after spacecraft equipment failures late in Cycle 3, and the radar was not used for science purposes after that. Cycle 4 was used for full (360 degree) longitudinal collection of gravity data because of favorable planetary and spacecraft geometry. The cycle was extended by about ten days to compensate for passage of the radio ray through the Venus atmosphere during the first ten days. To improve sensitivity to gravity features, orbit periapsis was lowered on orbit 5752. Radio occultation measurements were made on orbits 6369, 6370, 6471, and 6472. The aerobraking phase of the mission was designed to change the Magellan orbit from eccentric to nearly circular. This was accomplished by dropping periapsis to less than 150 km above the surface and using atmospheric drag to reduce the energy in the orbit. Aerobraking ended on 3 August 1993, and periapsis was boosted above the atmosphere leaving the spacecraft in an orbit that was 540 km above the surface at apoapsis and 197 km above the surface at periapsis. The orbit period was 94 minutes. The spacecraft remained on its medium-gain antenna in this orbit until Cycle 5 began officially on 16 August 1993. During Cycles 5 and 6 the orbit was low and approximately circular. The emphasis was on collecting high-resolution gravity data. Two bistatic surface scattering experiments were conducted, one on 6 October 1993 (orbits 9331, 9335, and 9336) and the second on 9 November 1993 (orbits 9846-9848). Mission Phases ============== Mission phases were defined for significant spacecraft activity periods. During orbital operations a 'cycle' was approximately the time required for Venus to rotate once under the spacecraft (about 243 days). But there were orbit adjustments and other activities that made some mapping cycles not strictly contiguous and slightly longer or shorter than the rotation period. PRELAUNCH --------- The prelaunch phase extended from delivery of the spacecraft to Kennedy Space Center until the start of the launch countdown. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1988-09-01 Mission Phase Stop Time : 1989-05-04 Spacecraft Operations Type : ORBITER LAUNCH ------ The launch phase extended from the start of launch countdown until completion of the injection into the Earth-Venus trajectory. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1989-05-04 Mission Phase Stop Time : 1989-05-04 Spacecraft Operations Type : ORBITER CRUISE ------ The cruise phase extended from injection into the Earth-Venus trajectory until 10 days before Venus orbit insertion. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1989-05-04 Mission Phase Stop Time : 1990-08-01 Spacecraft Operations Type : ORBITER ORBIT INSERTION --------------- The Venus orbit insertion phase extended from 10 days before Venus orbit insertion until burnout of the solid rocket injection motor. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1990-08-01 Mission Phase Stop Time : 1990-08-10 Spacecraft Operations Type : ORBITER ORBIT CHECKOUT -------------- The orbit trim and checkout phase extended from burnout of the solid rocket injection motor until the beginning of radar mapping. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1990-08-10 Mission Phase Stop Time : 1990-09-15 Spacecraft Operations Type : ORBITER MAPPING CYCLE 1 --------------- The first mapping cycle extended from completion of the orbit trim and checkout phase until completion of one cycle of radar mapping (approximately 243 days). Mapping orbits included in the first cycle were 373 through 2165. Orbits 2159-2171 were used for an interferometry test, and orbits 2172-2175 were used to conduct an orbit trim maneuver (OTM). Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1990-09-15 Mission Phase Stop Time : 1991-05-15 Spacecraft Operations Type : ORBITER MAPPING CYCLE 2 --------------- The second mapping cycle extended from completion of the first mapping cycle through an additional cycle of mapping. Acquisition of 'right-looking' SAR data was emphasized. Orbits included in the second cycle were 2176 through 3976. Radio occultation measurements were first carried out on orbits 3212-3214. A period of battery reconditioning followed completion of Cycle 2. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1991-05-16 Mission Phase Stop Time : 1992-01-17 Spacecraft Operations Type : ORBITER MAPPING CYCLE 3 --------------- The third mapping cycle extended from completion of battery reconditioning through an additional cycle of mapping (approximately 243 days). Acquisition of 'stereo' SAR data was emphasized. Orbits included in the third cycle were 4031 through 5747. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1992-01-24 Mission Phase Stop Time : 1992-09-14 Spacecraft Operations Type : ORBITER MAPPING CYCLE 4 --------------- The fourth mapping cycle extended from completion of the third mapping cycle through an additional cycle of mapping. Acquisition of radio tracking data for gravity studies was emphasized. Radio occultation measurements were carried out on orbits 6369, 6370, 6471, and 6472. Because of poor observing geometry for gravity data collection at the beginning of the cycle, this cycle was extended 10 days beyond the nominal 243 days. Orbits included within the fourth cycle were 5748 through 7626. Periapsis was lowered on orbit 5752 to improve sensitivity to gravity features in Cycle 4. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1992-09-14 Mission Phase Stop Time : 1993-05-25 Spacecraft Operations Type : ORBITER AEROBRAKING ----------- The aerobraking phase extended from completion of the fourth mapping cycle through achievement of a near-circular orbit. Circularization was achieved more quickly than expected; the first gravity data collection in the circular orbit was not scheduled until 11 days later. Orbits included within the aerobraking phase were 7627 through 8392. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1993-05-26 Mission Phase Stop Time : 1993-08-05 Spacecraft Operations Type : ORBITER MAPPING CYCLE 5 --------------- The fifth mapping cycle extended from completion of the aerobraking phase through an additional cycle of mapping (approximately 243 days). Acquisition of radio tracking data for gravity studies was emphasized. The first orbit in the fifth cycle was orbit 8393, and the last was orbit 12248. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1993-08-16 Mission Phase Stop Time : 1994-04-15 Spacecraft Operations Type : ORBITER MAPPING CYCLE 6 --------------- The sixth mapping cycle extended from completion of the fifth mapping cycle through an additional cycle of mapping (approximately 180 days). Acquisition of radio tracking data for gravity studies was emphasized. The first orbit in the sixth cycle was orbit 12249, and the last was orbit 15032. The sixth cycle ended when radio contact was lost as the spacecraft entered the atmosphere and was destroyed in a 'terminal windmill' experiment. Spacecraft Id : MGN Target Name : VENUS Mission Phase Start Time : 1994-04-16 Mission Phase Stop Time : 1994-10-12 Spacecraft Operations Type : ORBITER
  • collection : DART Radio Science Tracking and Navigation Files (TRK-2-34) Data Collection
    The DART mission receives Tracking and Navigation Files (TRK-2-34) from the Deep Space Network (DSN) collected primarily from signals emitted by the high gain antenna (HGA) onboard the DART spacecraft. These are binary files of table data whose fields and format are described by the TRK-2-34 DSN Tracking System Data Archival Format document referenced in the DART Radio Science SIS. These files have been sorted by data record type, of which there are approximately 18. Both mission navigators and those working on radio science investigations use these data.
  • data set : CASSINI RSS RAW DATA SET - TIGR1 V1.0
    Cassini RSS TIGR1 raw data.
  • collection : LICIACube Radio Science Tracking and Navigation Files (TRK-2-34) Data Collection
    The LICIACube mission receives Tracking and Navigation Files (TRK-2-34) from the Deep Space Network (DSN) collected primarily from signals emitted by the high gain antenna (HGA) onboard the LICIACube spacecraft. These are binary files of table data whose fields and format are described by the TRK-2-34 DSN Tracking System Data Archival Format document referenced in the LICIACube Radio Science SIS. These files have been sorted by data record type, of which there are approximately 18. Both mission navigators and those working on radio science investigations use these data.
  • instrument : RADAR SYSTEM for MGN
    Instrument Specifications ========================= The radar was manufactured by Hughes Aircraft Company and the 'build date' is taken to be 1989-01-01. The radar dimensions were 0.304 by 1.35 by 0.902 (height by length by width in meters) and the mass was 126.1 kg. Instrument Id : RDRS Instrument Host Id : MGN Pi Pds User Id : GPETTENGILL Instrument Name : RADAR SYSTEM Instrument Type : RADAR Build Date : 1989-01-01 Instrument Mass : 126.100000 Instrument Length : 1.350000 Instrument Width : 0.902000 Instrument Height : 0.304000 Instrument Manufacturer Name : HUGHES AIRCRAFT For more information on the radar system see the papers by [JOHNSON1990] and [SAUNDERSETAL1990]. Instrument Overview =================== The Magellan radar system included a 3.7 m diameter high gain antenna (HGA) for SAR and radiometry and a smaller fan-beam antenna (ALTA) for altimetry. The system operated at 12.6 cm wavelength. Common electronics were used in SAR, altimetry, and radiometry modes. The SAR operated in a burst mode; altimetry and radiometry observations were interleaved with the SAR bursts. Between SAR bursts (typically several times a second) groups of altimeter pulses were transmitted from a dedicated fan-beam altimeter antenna directed toward the spacecraft's nadir. The altimeter pulses were identical in waveform and bandwidth to the SAR pulses, resulting in a range accuracy of better than 15 m. The pulse-repetition rate and burst duration differed between the two modes. Radiometry data were obtained by spending a portion of the time between SAR bursts and after altimeter operation in a passive (receive-only) mode, with the HGA antenna capturing the microwave thermal emission from the planet. Noise power within the 10-MHz receiver bandwidth was detected and accumulated for 50 ms. To reduce the sensitivity to receiver gain changes in this mode, the receiver was connected on alternate bursts first to a comparison dummy load at a known physical temperature and then to the HGA. The short-term temperature resolution was about 2 K; the long-term absolute accuracy after calibration was about 20 K. Science Objectives ================== See MISSION_OBJECTIVES_SUMMARY under MISSION. Operational Considerations ========================== The Magellan radar system was used to acquire radar back-scatter (SAR) images, altimetry, and radiometry when the spacecraft was close to the planet. Nominal operation extended from about 20 minutes before periapsis until about 20 minutes after periapsis. In the SAR mode output from the radar receiver was sampled, blocks of samples were quantized using an adaptive procedure, and the results were stored on tape. In the altimetry mode samples were recorded directly, without quantization. Radiometry measurements were stored in the radar header records. During most of the remainder of each orbit, the HGA was pointed toward Earth and the contents of the tape recorder were transmitted to a station of the DSN at approximately 270 kilobits/second. SAR, altimetry, and radiometry data were then processed using ground software into images, altimetry profiles, estimates of backscatter coefficient, emissivity, and other quantities. Calibration Description ======================= The radar was calibrated before flight using an active electronic target simulator [CUEVAS1989]. Platform Mounting Descriptions ============================== The spacecraft +Z axis vector was in the nominal direction of the HGA boresight. The +X axis vector was parallel to the nominal rotation axis of the solar panels. The +Y axis vector formed a right-handed coordinate system and was in the nominal direction of the star scanner boresight. The spacecraft velocity vector was in approximately the -Y direction when the spacecraft was oriented for left-looking SAR operation. The nominal HGA polarization was linear in the y-direction. Cone Offset Angle : 0.00 Cross Cone Offset Angle : 0.00 Twist Offset Angle : 0.00 The altimetry antenna boresight was in the x-z plane 25 degrees from the +Z direction and 65 degrees from the +X direction. The altimetry antenna was aimed approximately toward nadir during nominal radar operation. The altimetry antenna polarization was linear in the y-direction. The medium gain antenna boresight was 70 degrees from the +Z direction and 20 degrees from the -Y direction. The low gain antenna was mounted on the back of the HGA feed; it's boresight was in the +Z direction and it had a hemispherical radiation pattern. Principal Investigator ====================== The Principal Investigator for the radar instrument was Gordon H. Pettengill. Instrument Section / Operating Mode Descriptions ================================================ The Magellan radar system consisted of the following sections, each of which operated in the following modes: Section Mode ------------------------------------------- SAR Synthetic Aperture Radar (SAR) ALT Altimetry RAD Radiometry (1) SAR Characteristics ----------------------- In the Synthetic Aperture Radar mode, the radar transmitted bursts of phase-modulated pulses through its high gain antenna. Echo signals were captured by the antenna, sampled at the receiver output, and stored on tape after being quantized to reduce data volume. Pulse repetition rate and incidence angle were chosen to meet a minimum signal-to-noise ratio requirement (8 dB) for image pixels after ground processing. Multiple looks were used in processing to reduce speckle noise. Incidence angles varied from about 13 degrees at the pole to about 44 degrees at periapsis during normal mapping operations (e.g., Cycle 1); but other 'look angle profiles' were used during the mission. Peak transmit power : 350 watts Transmitted pulse length : 26.5 microsecs Pulse repetition frequency : 4400-5800 per sec Time bandwidth product : 60 Inverse baud width : 2.26 MHz Data quantization (I and Q) : 2 bits each Recorded data rate : 750 kilobits/sec Polarization (nominal) : linear horizontal HGA half-power full beam width : 2.2 deg (azimuth) : 2.5 deg (elev) one-way gain (from SAR RF port) : 35.7 dBi System temperature (viewing Venus) : 1250 K Surface resolution (range) : 120-360 m (along track) : 120-150 m Number of looks : 4 or more Swath width : 25 km (approx) Antenna look angle : 13-47 deg Incidence angle on surface : 18-50 deg Data Path Type : RECORDED DATA PLAYBACK Instrument Power Consumption : UNK (2) ALT Characteristics ----------------------- After SAR bursts (typically several times a second) groups of altimeter pulses were transmitted from a dedicated fan beam altimeter antenna (ALTA) directed toward the spacecraft's nadir. Output from the radar receiver was sampled, and the samples were stored on tape for transmission to Earth. During nominal left-looking SAR operation the ALTA pointed approximately 20 deg to the left of the spacecraft ground track at periapsis and about 10 deg to the right of the ground track near the north and south pole. Data quantization (I and Q) : 4 bits each Recorded data rate : 35 kbs Polarization : linear ALTA half-power full beamwidth (along track): 11 deg (cross track): 31 deg one-way gain referenced to ALT RF port : 18.9 dBi ALTA offset from HGA : 25 deg Burst interval : 0.5-1.0 sec duration : 1.0 millisec Dynamic range : 30 dB (or more) Data Path Type : RECORDED DATA PLAYBACK Instrument Power Consumption : UNK (3) RAD Characteristics ----------------------- Radiometry measurements were made by the radar receiver and HGA in a receive-only mode that was activated after the altimetry mode to record the level of microwave radiothermal emission from the planet. Noise power within the 10-MHz receiver bandwidth was detected and accumulated for 50 ms. To reduce the sensitivity to receiver gain changes in this mode, the receiver was connected on alternate bursts first to a comparison dummy load at a known physical temperature and then to the HGA. The short-term temperature resolution was about 2 K; the long-term absolute accuracy after calibration was about 20 K. At several times during the mission, radiometry measurements were carried out using known cosmic radio sources. Receiver Bandwidth : 10 MHz Integration Time : 50 millisecs Polarization (nominal) : linear horizontal Data Quantization : 12 bits Data Rate : 10-48 bits/sec HGA half-power full beam width : 2.2 deg System temperature (viewing Venus) : 1250 K Antenna look angle : 13-47 deg Incidence angle on surface : 18-50 deg Surface resolution (along track) : 15-120 km (cross track) : 20-125 km Data Path Type : RECORDED DATA PLAYBACK Instrument Power Consumption : UNK
  • data set : NEAR EROS RADIO SCIENCE DERIVED PRODUCTS - GRAVITY V1.0
    NEAR Eros gravity science derived data products from JPL.
  • instrument : MARS EXPRESS ORBITER RADIO SCIENCE for MEX
    Instrument Overview =================== Mars Express (MEX) Radio Science investigations utilized instrumentation with elements on both the spacecraft and ground (Earth). Much of this was shared equipment, being used for routine telecommunications as well as for Radio Science. Ground systems were provided by the European Space Agency (ESA) at New Norcia, Australia, and by the U.S. National Aeronautics and Space Administration (NASA) Deep Space Network (DSN) at sites in Australia, Spain, and the United States. Performance and calibration of both the spacecraft and ground systems directly affected the radio science data accuracy and played a major role in determining the quality of the results. The spacecraft was able to receive and transmit signals at both S-band (approximately 13 cm wavelength) and X-band (approximately 3.5 cm). The spacecraft transmissions could use either an onboard oscillator for the frequency reference ('one- way' mode) or a signal transmitted from the ground ('two-way' mode); in the latter case, either an S- or X-band signal from the ground could be used as the reference. Science Objectives ================== Two different types of radio science measurements were carried out with Mars Express: Radiometric Measurements: Ground stations have the ability to transmit coded and uncoded waveforms which can be echoed by distant spacecraft. Analysis of the received coding allows navigators to determine the distance to the spacecraft; analysis of Doppler shift on the carrier signal allows estimation of the line-of-sight spacecraft velocity. Range and Doppler measurements are used to calculate the spacecraft trajectory and to infer gravity fields of objects near the spacecraft. NB: Doppler measurements can be made in one-way but are usually more accurate if carried out in two- way mode. Radio Propagation Measurements: Ground stations can record spacecraft signals that have propagated through or been scattered from target media. Measurements of signal parameters after wave interactions with surfaces, atmospheres, rings, and plasmas are used to infer physical and electrical properties of the target. Radio propagation measurements can be conducted in either one-way or two-way mode. These measurements were applied - separately and together - to Mars science objectives such as inference of local gravity field anomalies, mass of Phobos, temperature and pressure of the atmosphere, electron density in the ionosphere, scattering properties of the surface, and structure of the solar wind. Gravity Measurements -------------------- Measurement of the gravity field provides significant constraints on inferences about the interior structure of Mars. Precise, detailed study of the spacecraft motion in Mars orbit can yield the mass distribution of the planet. Topographic data, such as those obtained by the Mars Global Surveyor (MGS) Mars Orbiting Laser Altimeter (MOLA), form a critical adjunct to these measurements since only after the gravitational effects are adjusted for topography can the gravity anomalies be interpreted geophysically. Mars Express studies of the gravity field emphasized the local and time varying characteristics of the field; but the first task was to determine the global field. Doppler and range tracking measurements yield accurate spacecraft trajectory solutions. Simultaneously with reconstruction of the spacecraft orbit, observation equations for field coefficients and a small number of ancillary parameters can be solved. This type of gravity field solution is essential for characterizing tectonic phenomena and can also be used to study localized features. Differences in the solution can be used to infer variation of low degree and order coefficients on time scales of months to years - such as might be expected from seasonal mass exchange between polar cap deposits and the atmosphere. These kind of global gravity measurements were typically conducted around apocenter. Gravity models based on MGS data have been published by [LEMOINEETAL2001] and [YUANETAL2001]. Early results from studies of time variability in the MGS results have been presented by [ZUBER&SMITH2002]. Global gravity measurements were typically done around Mars Express Apocenter. 'Short-arc' line-of-sight Doppler tracking measurements obtained when the Earth-to-spacecraft line-of-sight is within a few degrees of the orbit plane provide the highest resolution of local features. The results from this type of observation typically are presented as contoured acceleration profiles of specific features (e.g., craters, volcanoes, etc.) or line-of-sight acceleration maps of specific regions. The high spatial resolution of these products makes them especially useful to geophysicists for study of features in the size range of 300 to 1000 km. These kind of measurements were typically conducted during Mars Express pericenter over interesting geophysical structures like: Tempe Fossae and Olympus Mons. An early example of such analysis was conducted on Viking Orbiter 2 data and published by [SJOGREN1979]. A possible by-product of the gravity field analysis is information on the density structure of the upper atmosphere [TRACADISETAL2001]. Phobos flyby ------------ During the Mars Express Mission several close flybys at Phobos occurred. When the distance between orbiter and Phobos is < 500 km Mars Radio Science will be able to derive the mass and density of the moon and to determine the Phobos orbit to great accuracy. Radio Occultation Measurements ------------------------------ Atmospheric measurements by the method of radio occultation contribute to an improved understanding of structure, circulation, dynamics, and transport in the atmosphere of Mars. These results are based on detailed analysis of the radio signal phase as the ray path enters and exits occultation by the planet, leading to profiles of temperature and pressure in the neutral atmosphere and profiles of electron density in the ionosphere. Retrieval of atmospheric profiles requires coherent samples with a sample rate of at least 10 per second of the radio signal that has propagated through the atmosphere, plus accurate knowledge of the spacecraft trajectory. The latter was obtained from the MEX Flight Dynamics Team. Solutions from MEX occultations provided neutral atmospheric structure to about 50 km from the surface and electron densities over a range of about 50 km centered on the altitude of the ionization peak. Spatial and temporal coverage in radio occultation experiments are determined by the geometry of the spacecraft orbit and the dates and times at which occultation data are acquired. Since MEX radio occultation experiments were conducted on a regular basis using a polar orbit, there was extensive occultation coverage at high northern and southern latitudes (e.g., beyond 60 degrees). As the orbit appeared to drift from edge-on to nearly broadside (as viewed from Earth), occultation points moved toward the equator and the entry/exit angle approached grazing. During the first Mars year of MEX operations, there were three occultation 'seasons' between which the spacecraft was not occulted for several months at a time. In the year 2004 only Occultation Ingress measurements were performed. Bistatic Surface Scattering Measurements ------------------------------------------ The spacecraft high-gain antenna (HGA) could also be pointed toward the surface of the planet. The strength of the signal scattered from the illuminated area could be measured and the results interpreted in terms of the dielectric constant of the surface material. The model for interpretation assumes Fresnel reflection at the specular angle. Under certain circumstances, the dispersion of the echo (its spectral width) could be interpreted in terms of the surface roughness on scales comparable to the wavelength. One such MGS bistatic radar was conducted over the Mars Polar Lander/Deep Space 2 site in May 2000 [SIMPSON&TYLER2001]. For a few seconds before and after geometrical occultation the HGA illuminated a small strip of surface as well as the atmosphere. In some cases, an echo could be observed from the surface. The interpretation of these transient echoes is more difficult than for the case above, possibly involving diffraction and surface waves in addition to Fresnel reflection. On Mars Express this operation was done in ONED mode. That is no uplink but with X- and S-Band downlink. The HGA was pointed to Mars. Pointing was inertial. That is no slew was performed during the measurement. Solar Scintillation and Faraday Rotation Experiments ---------------------------------------------------- Solar scintillation and Faraday rotation experiments were conducted to improve understanding of the structure and dynamics of the solar corona and wind. Because Mars orbits the Sun, spacecraft like MEX are transported behind the solar disk, as seen from Earth. Radio waves propagating between MEX and Earth stations are refracted and scattered by the solar plasma [WOO1993]. Intensity fluctuations can be related to fluctuations in electron density along the path, while Doppler or phase scintillations can be related to both electron density fluctuations and also the speed of the solar wind. Many plasma effects decrease as the square of the radio frequency; scintillations are about an order of magnitude stronger at S-band than X-band. The first solar conjunction observations with MEX were conducted during the solar conjunction season of 2004: 16.8.2004 - 22.10.2004. Measurements during solar conjunction should be typically been done in TWOD-S configuration. That is in two-way mode with S-Band uplink and coherent and simultaneous in X- and S- Band. However, due to problems to lock S-Band in the 2004 conjunction season. The TWOD-X configuration was used instead. That is in two-way mode with X-Band uplink and coherent and simultaneous in X- and S-Band Investigators and Other Key Personnel ===================================== Martin Paetzold University of Cologne Principal Investigator; solar physics Bernd Hausler Universitaet der Experiment Manager Bundeswehr Munich Richard Simpson Stanford University Data Manager; surface scattering Joerg Selle Universitaet der Operations Manager Bundeswehr Munich Sami Asmar Jet Propulsion JPL/DSN operations Laboratory G. Leonard Tyler Stanford University radio propagation David Hinson Stanford University atmosphere, ionosphere, radio occultation Jean-Pierre Centre National gravity Barriot d'Etudes Spatiale Toulouse Veronique Dehant Observatoire Royale gravity Brussels Instrument Specification - Spacecraft ===================================== The Mars Express spacecraft telecommunications subsystem served as part of a radio science subsystem for investigations of Mars. Many details of the subsystem are unknown; but they are not of importance for understanding the science. The spacecraft 'build date' is taken to be 2003-06-01, shortly before launch. Instrument Id : MRS Instrument Host Id : MEX Pi Pds User Id : MPAETZOLD Instrument Name : MARS EXPRESS ORBITER RADIO SCIENCE EXPERIMENT Instrument Type : RADIO SCIENCE Build Date : 2003-06-01 Instrument Mass : UNK Instrument Length : UNK Instrument Width : UNK Instrument Height : UNK Instrument Manufacturer Name : UNK Subsystems ---------- SWITCH TRANSPONDER 1 -------- ------ ----- -------------------- \ | |---| TWTA |---|\ /|<--------|X-Band Transmitter| \ | | ------ | \ / | | | HGA >--| | | X | ------|S-Band Transmitter| / | | ------ | / \ | | | | / | RFDU |---| TWTA |---|/ \| | --->|X-band Receiver | | | ------ ----- | | | | LGA >---| |<---------------------- | ->|S-Band Receiver | | | | | -------------------- | |------------------------- | | |--------------------------- TRANSPONDER 2 | | -------------------- | |---< LGA <---|X-Band Transmitter| -------- | | <---|S-Band Transmitter| TRANSPONDERS 1 and 2 were | | connected to provide fully --->|X-band Receiver | redundant, switchable | | functions. --->|S-band Receiver | -------------------- The Mars Express radio subsystem comprised several components (shown above), configured to provide redundant functions should any single component fail (except the high-gain antenna). The high-gain antenna (HGA) was a body-fixed 1.60 m diameter parabolic dish which allowed transmission and reception at both S- and X-band. The HGA boresight was in the -X direction of the spacecraft coordinate system, offset 5 degrees in the +Z direction. Its gain was 29.56 dB and 41.43 dB at S- and X-band, respectively. Two low-gain antennas (LGA) were mounted on the front and rear of the spacecraft; they operated only at S-band. The HGA was the main antenna for receiving telecommands from and transmitting telemetry signals to the ground. The LGAs were used during the commissioning phase after launch and for emergency operations. The Radio Frequency Distribution Unit (RFDU) switched the onboard radio frequency hardware among the three antennas. Switchable Traveling Wave Tube Amplifiers (TWTA) provided 60 watts of X-band transmitter power to the RFDU; their inputs could come from either Transponder 1 or Transponder 2. The S-band transmitter power was 5 watts, which was generated within the transponder units. The S-band uplink was received via the LGA or HGA. In the coherent two-way mode the received frequency was used to derive the downlink frequencies by using the constant transponder ratios 880/221 and 240/221 for X-band and S-band downlink, respectively. The X-band uplink was received via the HGA only. In the coherent two-way mode the received frequency was used to derive the downlink frequencies by using the constant transponder ratios 880/749 and 240/749 for X-band and S-band downlink, respectively. An X-band uplink generally enhanced the performance of the radio link because X-band is less sensitive to the interplanetary plasma along the propagation path. The X-band and S-band frequencies were related by a factor of 11/3. If an uplink existed, the downlinks were also coherent with the uplink by their respective transponding ratios. The dual-frequency downlink allowed separation of the classical Doppler shift, due to relative motion of the spacecraft and the ground station, from the dispersive media effects, due to the propagation of the radio waves through the ionosphere and interplanetary medium. In one-way mode, the downlink transmitter frequency was derived from an onboard Temperature Controlled Crystal Oscillator (TCXO). The one-way mode could be selected by command from the ground. If the spacecraft receiver could not detect an uplink signal from the ground, the TCXO was selected by default. TCXO stability was several orders of magnitude less than the uplink reference, so the one-way mode was used only when no uplink was available (such as during bistatic radar experiments, when the HGA was pointed toward Mars) or when signal conditions were expected to be very dynamic and the transponder might not be able to lock to the uplink (such as during egress occultation observations). The redundant transponders each consisted of an S-band and X- band receiver and transmitter. The spacecraft was capable of receiving one uplink signal at S-band (2100 MHz) via the LGAs, or at either X-band (7100 MHz) or S-band via the HGA. The spacecraft could transmit a downlink signal at S-band (2300 MHz) and (simultaneously) a downlink signal at X-band (8400 MHz) using the HGA; or it could transmit one downlink signal at S-band via the LGAs. Operational Considerations -------------------------- Radio science observations often required operation of the spacecraft in orientations and configurations that were not compatible with spacecraft constraints, telecommunications, and requirements for other instruments. There were also limitations within the Radio Science Team, which resulted in a prioritization of radio science observations. The following list is representative but not complete. Spacecraft transmissions were very limited while in eclipse to conserve battery power. During the first occultation season, no egress occultations were observed because those were always in eclipse. Spacecraft cooling panels could not be exposed to direct sunlight. Since those were located on the side opposite the HGA, no bistatic radar experiments could be conducted which required pointing of the HGA more than about 90 degrees from the Sun. Bistatic radar experiments during the first half of 2004 were conducted only with fixed inertial pointing. That is, the HGA pointing was fixed in inertial space and the target was allowed to drift through the beam. Immediately after turn-on, output power from the S-band transmitter was variable. To ensure stability, a warm-up period of about 60 minutes was scheduled before each use of the S-band transmitter. Solar observations had highest priority from 30 days before solar conjunction to 30 days after. Otherwise, radio occultations had highest priority. Bistatic radar experiments had third priority; they required use of a 70-m DSN antenna, so were difficult to schedule on short notice. Gravity observations were most interesting when two-way dual-frequency data could be collected as the spacecraft passed through pericenter. But pericenter time was highly contested with several other instruments, which also sought those opportunities to acquire surface data with the highest resolution. Phobos encounters were rare and were scheduled separately. Competition among instruments for those times was extremely fierce. Calibration ----------- For many experiments, calibration data were collected in conjunction with the scientific observations. For example, carrier power and frequency could be determined before and/or after bistatic radar and radio occultation experiments when the antenna was pointed toward Earth. The gain, beam patterns, and pointing of the HGA were calibrated during post-launch tests. The half-power points were about 2.6 and 0.8 degrees from the boresight at S- and X-band, respectively. For radio tracking data, error sources in two-way mode are shown below, where the tabulated error values are given in terms of equivalent spacecraft velocity error. These values were based on pre-launch tests. |======================================================| | Error Source | Equivalent Velocity | | | Error (mm/s) | | |---------------------| | | S-band | X-Band | |================================+==========+==========| |Total phase error (thermal and | 1.0 | 0.3 | |ground station contributions) | | | |--------------------------------+----------+----------| |Transponder quantization error | 0.4 | 0.1 | |in frequency | | | |--------------------------------+----------+----------| |Transponder quantization error | 0.01 | 0.004 | |in phase | | | |================================+==========+==========| |Total error (coherent mode) | 1.1 | 0.32 | |======================================================| Platform Mounting ----------------- The MEX High Gain Antenna was rigidly attached to the -X side of the spacecraft bus. Therefore, the MEX HGA frame (MEX_HGA) was defined as a fixed offset frame with its orientation given relative to the MEX_SPACECRAFT frame: +Z axis of the HGA frame was in the antenna boresight direction (nominally 5 degrees off the spacecraft -X axis toward the spacecraft +Z axis); +Y axis of the HGA frame was in the direction of the spacecraft +Y axis; +X completed the right hand frame; The origin of the HGA frame was located at the geometric center of the HGA dish outer rim circle. ^+Zhga | | | +Xhga | +Yhga _____o-------> \ / .________________. .__`._____.'__. .________________. | \ | | / | | \ | ___ | / | | | | .' ` +Ysc | | | |o=| | o------->o| | | | | `_|+Zsc | | | | / | | | \ | ._________________/ .______|______. \_________________. -Y Solar Array | +Y Solar Array V +Xsc Nominally a single rotation of -85 degrees about the +Y axis was needed to align the spacecraft frame with the HGA frame. Operating Modes --------------- A two-way dual-frequency radio link was used for occultations, gravity observations, and solar corona investigations. Such a radio link benefited from the superior frequency stability of the ground station. The dual-frequency downlink at X-band and S-band was used to separate classical and dispersive Doppler shifts, allowing correction of the observed frequency shift by any plasma contribution. For some observations (e.g., solar corona) an S-band uplink was used to increase sensitivity to plasma effects along the path. In the above experiments, operation was usually preferred with full power in the carrier (no telemetry or other modulation on the downlink) to maximize signal-to-noise ratio. The dual-frequency one-way radio link at S- and X-band was used for bistatic radar experiments. In these experiments, the HGA was pointed toward Mars and could not be used to capture an uplink signal, receive commands, or transmit telemetry. The dual-frequency one-way radio link was also used for egress occultation experiments because there was no time to establish a two-way link. Stability of the one-way link was not sufficient to allow scientifically useful probing of the neutral atmosphere on egress; but the ionospheric analysis could be carried out using the differential phase/frequency effects at S- and X-band which were proportional to each other. Instrument Specification - New Norcia ===================================== ESA completed construction of a 35 m ground station at its complex near New Norcia, Australia, in the year before launch of Mars Express. The station provided uplink at either S- or X-band and simultaneous dual-frequency downlink at both bands. Specifications are given below. the 'build date' is taken arbitrarily to be 1 January 2003. Instrument Id : RSS Instrument Host Id : NNO Pi Pds User Id : MPAETZOLD Instrument Name : UNK Instrument Type : RADIO SCIENCE Build Date : 2003-01-01 Instrument Mass : UNK Instrument Length : UNK Instrument Width : UNK Instrument Height : UNK Instrument Manufacturer Name : UNK The IFMS (Intermediate Frequency Modulation System) at NNO is a piece of equipment which mainly provides: - generation of the uplink IF carrier, possibly modulated with a TC signal (from an external source) and a Ranging signal (internally generated) - reception of the downlink IF signal - diversity combination estimation - demodulation (remnant and suppressed carrier demodulators and Ranging) demodulator) and generation of bit stream for the telemetry decoding system - collection of Doppler, Meteo and Ranging measurements into data sets, later available for local display and remote retrieval via DDS - telemetry decoding is provided by the integrated TCDS (Telemetry Channel Decoding System) functional unit (the presence of the TCDS may be optional) System overview --------------- Antenna o \ /|\ / \ / | \ / --v-- / \ / \ / \ |----------------------------------------------------| | Front-End | |----------------------------------------------------| ^ ------- | | |Meteo | | | |sensors| | | ------- | | | | | v ||----| |------|---------------------------|------------| |---|| TC | |IFMS | v | |TM || | | v |-------------------------| | | || | | |---------| | Common Front End/ | | | || |------->| Up-link | ..>| Diversity Combiner | | | || | | |Modulator| . |-------------------------| | | || |<------>|---------| . | | | | | || |Uplink| ^ . v v v | | || |hand | . . ----- ------- -------- | | || |shake | . . | OLP | |RG Dmod| |R/S carr|--->| ||----| | . . ----- ------- | Demod || | | | . . ^ ^ -------- | | | | . . . . | | | | | . . . . v | | | | . . . . ------- | | | | . . . . | TCDS |---->| | | . . . . ------- | | | | . . . . ^ | |---| | . . . . . | | . . . . . | | v v v v v | | |----------------------------------------| | | | System Monitoring & Control | | | | (UCPU software) | | | |----------------------------------------| | |-----------------------------------------------| | | | ----- ------ ----- | DCP | |STC II| | OCC | ----- ------ ----- IFMS software ------------- The IFMS software is mainly in charge of the following functions: - handle the Digital Signal Processing (DSP) units (Uplink Modulator, Common Front End, Diversity Combiner, Ranging Demodulator, Remnant Carrier Demodulator, Suppressed Carrier Demodulator, Meteo system and TCDS) - execute data acquisition requests and collect independently Doppler, AGC, Meteorological, Ranging and Open-Loop data - allow the Control Centre to retrieve the collected data - provide Monitoring & Control access to the Station Computer (STC) - provide Monitoring & Control access to an operator via the Development Control Position (DCP) for both local control and engineering purpose Subsystems ---------- Of primary interest to radio science are the three Intermediate Frequency and Modem Systems (IFMS) at New Norcia which controlled both the uplink and downlink. The IFMS baseband processor operated on a 17.5 Msps 24-bit complex sample stream (12 bit words each for the I and Q channels) which resulted from filtering and decimating the 280 Msps 8-bit stream output by the Common Front End (CFE) analog-to-digital converter. These channels were provided for both the right circular and left-circular polarizations (RCP and LCP, respectively). The Radio Science raw data could be directly transferred to a mass storage device and/or processed by a Fast Fourier routine. Data transfer rates from the digital signal processor to data storage (disk) were limited to 10 samples/s. Data transfer to the European Space Operations Center (ESOC) could be done at a rate of 2 ksps. Subsystems overview ------------------- The IFMS is constituted of a 19'' crate containing the UNIX CPU (UCPU), the Time Code Reader (TCR) and the DSP units (Uplink Modulator (ULM), CFE units, General DSP units (GDSP), except the Meteo unit which is external to the system). - Internal network and IP Processor (IPP): The Internet Protocol suite is used to interface most of the IFMS elements on an internal IP network. For this the GDSP units are equipped with an IPP (IP Processor), in charge of managing data communication with the UNIX-CPU (based on IP) and the DSP board controller (based on serial interface). - The Time Code Reader (TCR) receives: - the Time Reference (IRIG-B on 1 kHz or 5 MHz carrier) - the Frequency Reference (5 MHz or 10 MHz) It distributes Time to the other units to be used for measurement time-tagging. - Meteo Unit: The Meteo Unit includes outdoor sensors providing analogue data for humidity, pressure, temperature and indoor electronic equipment (located in fact outside of the IFMS rack). It provides ASCII formatted numerical measurement of humidity, pressure and temperature to the IFMS management processor. - Uplink Modulator (ULM): The ULM unit generates internally the ranging signal (Tone and Code) in digital form. It receives the telecommand signal in either digital or analogue form from external equipment. It outputs an IF signal (230 MHz or 70 MHz) modulated by the uplink ranging signal and/or the telecommand signal. - Common Front End Unit: The CFE unit receives (from the down-converters) the 70 MHz down- link IF signals modulated by telemetry and possibly ranging signals and digitises them for further processing. The digital data is propagated on the rack back-plane. Note: A second CFE can be present in the IFMS. - Diversity combiner: The DCE unit makes estimates of: - the depolarisation angle between the LH and RH channels - the phase error between the LH and the RH channels It then provides qualification information on the rack back-plane for further use by the demodulators. - Ranging receiver and demodulator: The RGD unit receives, from the Common Front End and Diversity Combiner units, the digital demodulated 70 MHz signal and qualification information. It demodulates the down-link signal and extracts Doppler measurement. It generates a replica Ranging signal and performs the Tone PLL and the Ambiguity Resolution in order to measure signal round-trip delay, modulo the maximum code length. It provides Doppler and Ranging measurement. - Remnant and Suppress Carrier demodulators: The RCD and SCD units receive, from the Common Front End and Diversity Combiner units, the digital demodulated 70 MHz signal and qualifier. They provide demodulated telemetry data and Doppler measurement. - Telemetry Channel Decoder System: The TCDS unit receives, from the demodulator units, the telemetry bit stream and performs Viterbi and Reed-Solomon decoding and frame synchronisation. It provides decoded and synchronised telemetry data emitted via a UDP/IP protocol. - Open-Loop Processor: The OLP unit receives, from the Common Front End and Diversity Combiner units, the digital demodulated 70 MHz signal and qualifier. It provides Open-Loop measurements. Operational Considerations -------------------------- By agreement between the Mars Express Radio Science (MaRS) Team and the European Space Operations Centre (ESOC), the three IFMS units at New Norcia were configured as follows: IFMS 1: Controlled uplink, including choice of band (usually X-band, but S-band for solar conjunction). Two channels of closed-loop downlink were possible; these could be any two of the four X-RCP, X-LCP, S- RCP, and S-LCP combinations. If X-RCP and X-LCP were selected, then the IFMS computed polarization. IFMS 2: Backup for IFMS 1; MaRS could specify its configuration if it was not assigned otherwise. IFMS 3: MaRS could always specify the configuration Platform Mounting ----------------- In the IRTF2000 reference system at epoch 2002-07-24 12:00:00, the Cartesian coordinates of the intersection of the azimuth and elevation axes of the New Norcia antenna were (meters): X = -2414067.051 Y = 4907869.387 Z = -3270605.276 Using the WGS84 reference ellipsoid with equatorial radius 6378137 m and inverse flattening 298.257223563, the geodetic latitude, longitude, and height were geodetic latitude = -31.04822306 degrees north longitude = 116.19150227 degrees east height = 252.224 meters Calibration ----------- See Calibration section for spacecraft. Modes ----- See Modes section for spacecraft. In addition, there were two mode choices on the ground. Closed-loop data acquisition was done with a phase-locked loop receiver at the ground station. The downlink signal arriving at the station could be either one-way or two-way. Two-way Doppler shifts were extracted by comparing each measurement of the downlink carrier frequency from the phase- locked loop with a reference from the ground station frequency reference source -- e.g., a hydrogen maser with a frequency stability on the order of 1E-15 to 1E-16. Because this frequency reference source was also used for generation of the uplink carrier, the accuracy of the frequency determination was as good as the reference source. The Doppler integration time needed to achieve a certain signal to noise ratio determined the time between successive frequency determinations. The amplitude of the radio signal was estimated by the Automatic Gain Control (AGC). Open-loop data recording was done by filtering and down- converting the received radio carrier signal to baseband where it was digitally sampled and stored for subsequent analysis. The open-loop receiver was tuned by a local oscillator. The frequency of the local oscillator was given by the best available estimate of the carrier frequency transmitted by the spacecraft and applying Doppler corrections due to the relative spacecraft-to-Earth motion. Measurement Parameters ---------------------- Each IFMS generated up to four types of data records: Doppler, gain, range, and/or meteorology. Each included a header with the following information: station identifier; spacecraft identifier; time tag of the first and last samples; sample period; total number of samples; and several flags or other markers to identify the data. Doppler samples could be taken at 1000, 100, 10, 1, or 0.1 per second; the data records contain: sample number and time; unwrapped phase and accumulated phase with respect to a reference. Gain records contain: sample number and time; carrier level and polarization angle. Range data could be taken every 1-120 seconds (in user selectable increments of 1 second); range records contain: sample number and time; round trip delay modulo the ranging code; current code number and several flags and status words. Meteorological records contain: sample number and time; humidity; pressure and temperature. Instrument Specification - DSN ============================== Three Deep Space Communications Complexes (DSCCs) (near Barstow, CA; Canberra, Australia; and Madrid, Spain) comprised the DSN tracking network. Each complex was equipped with several antennas [including at least one each 70-m, 34-m High Efficiency (HEF), and 34-m Beam WaveGuide (BWG)], associated electronics, and operational systems. Primary activity at each complex was radiation of commands to and reception of telemetry from active spacecraft. Transmission and reception was possible in several radio- frequency bands, the most common being S-band (nominally a frequency of 2100-2300 MHz or a wavelength of 14.2-13.0 cm) and X-band (7100-8500 MHz or 4.2-3.5 cm). Transmitter output powers of up to 400 kW were available. The Deep Space Network was managed by the Jet Propulsion Laboratory of the California Institute of Technology for the U.S. National Aeronautics and Space Administration. Specifications included: Instrument Id : RSS Instrument Host Id : DSN Pi Pds User Id : MPAETZOLD Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : UNK Instrument Mass : UNK Instrument Length : UNK Instrument Width : UNK Instrument Height : UNK Instrument Manufacturer Name : UNK So far as radio science was concerned, the DSN was an evolving 'instrument;' the paragraphs which follow describe its capabilities during the first year of Mars Express orbital operations. For more information on the Deep Space Network and its use in radio science see reports by [ASMAR&RENZETTI1993], [ASMAR&HERRERA1993], and [ASMARETAL1995]. For design specifications on DSN subsystems see [DSN810-5]. For DSN use with MGS Radio Science see [TYLERETAL1992], [TYLERETAL2001], and [JPLD-14027]. Subsystems - DSN ---------------- The Deep Space Communications Complexes (DSCCs) were an integral part of Radio Science instrumentation. Their system performance directly determined the degree of success of Radio Science investigations, and their system calibration determined the degree of accuracy in the results of the experiments. The following paragraphs describe the functions performed by the individual subsystems of a DSCC. This material has been adapted from [ASMAR&HERRERA1993] and [DSN871-049-041]; for additional information, consult [DSN810-5], [DSN821-110], and [DSN821-104]. Each DSCC included a set of antennas, a Signal Processing Center (SPC), and communication links to the Jet Propulsion Laboratory (JPL). The general configuration is illustrated below; not all antennas are shown. -------- -------- -------- -------- -------- | DSS 25 | | DSS 27 | | DSS 14 | | DSS 15 | | DSS 16 | |34-m BWG| |34-m HSB| | 70-m | |34-m HEF| | 26-m | -------- -------- -------- -------- -------- | | | | | | v v | v | --------- | --------- --------->|GOLDSTONE|<---------- |EARTH/ORB| | SPC 10 |<-------------->| LINK | |---------| |---------| | SPC |<-------------->| 26-M | | COMM | ------>| COMM | --------- | --------- | | | v | v ------ --------- | --------- | NOCC |<--->| JPL |<------- | | ------ | CENTRAL | | GSFC | ------ | COMM | | NASCOMM | |AMMOS |<--->| TERMINAL|<-------------->| | ------ --------- --------- ^ ^ | | CANBERRA (SPC 40) <---------------- | | MADRID (SPC 60) <---------------------- The following table lists the DSN antennas (Deep Space Stations, or DSS's -- a term carried over from earlier times when antennas were individually instrumented) available for Mars Express. Not all antennas were actually used for MEX; their capabilities varied and some were more suitable for MEX Radio Science than others. GOLDSTONE CANBERRA MADRID Antenna SPC 10 SPC 40 SPC 60 -------- --------- -------- -------- 26-m DSS 16 DSS 46 DSS 66 34-m HEF DSS 15 DSS 45 DSS 65 34-m BWG DSS 24 DSS 34 DSS 54 DSS 25 DSS 55 DSS 26 34-m HSB DSS 27 DSS 28 70-m DSS 14 DSS 43 DSS 63 Developmental DSS 13 Subsystem interconnections at each DSCC are shown in the diagram below, and are described in the sections that follow. The Monitor and Control Subsystem was connected to all other subsystems; and the Test Support Subsystem could have been. ----------- ------------------ --------------------- |TRANSMITTER|_| UPLINK |_| COMMAND |_ | SUBSYSTEM | | SUBSYSTEM | | SUBSYSTEM | | ----------- ------------------ --------------------- | | | ----------- ------------------ --------------------- | | MICROWAVE |_| DOWNLINK |_| TELEMETRY |_| | SUBSYSTEM | | SUBSYSTEM | | SUBSYSTEM | | ----------- ------------------ --------------------- | | | ----------- ----------- --------- -------------- | | ANTENNA | | MONITOR | | TEST | | DIGITAL |_| | SUBSYSTEM | |AND CONTROL| | SUPPORT | |COMMUNICATIONS| ----------- | SUBSYSTEM | |SUBSYSTEM| | SUBSYSTEM | ----------- --------- -------------- DSCC Monitor and Control Subsystem ---------------------------------- The DSCC Monitor and Control Subsystem (DMC) was part of the Monitor and Control System (MON) which also included the ground communications Central Communications Terminal (CCT) and the Network Operations Control Center (NOCC) Monitor and Control Subsystem. The DMC was the center of activity at a DSCC. The DMC received and archived most of the information from the NOCC needed by the various DSCC subsystems during their operation. Control of most of the DSCC subsystems, as well as the handling and displaying of any responses to control directives and configuration and status information received from each of the subsystems, was done through the DMC. The effect of this was to centralize the control, display, and short-term archiving functions necessary to operate a DSCC. Communication among the various subsystems was done using a Local Area Network (LAN) hooked up to each subsystem via a network interface unit (NIU). DMC operations were divided into two separate areas: the Complex Monitor and Control (CMC) and the Network Monitor and Control (NMC). The primary purpose of the CMC processor for Radio Science support was to receive and store all predict sets transmitted from NOCC -- such as antenna pointing, tracking, receiver, and uplink predict sets -- and then, at a later time, to distribute them to the appropriate subsystems via the LAN. Those predict sets could be stored in the CMC for a maximum of three days under normal conditions. The CMC also received, processed, and displayed event/alarm messages, and maintained an operator log. Assignment and configuration of the NMCs was done through the CMC; to a limited degree the CMC could perform some of the functions performed by the NMC. There were two CMCs (one on-line and one backup) and three NMCs at each DSCC. The backup CMC could function as an additional NMC if necessary. The NMC processor provided the operator interface for monitor and control of a link -- a group of equipment required to support a spacecraft pass. For Radio Science, a link might include one or more Radio Science Receivers (RSRs), the DSCC Uplink Subsystem (UPL), and one or more DSCC Downlink Tracking and Telemetry Subsystems (DTTs). The NMC also maintained an operator log which included all operator directives and subsystem responses. One important Radio Science-specific function that the NMC performed was receipt and transmission of the system temperature and signal level data from the PPM, for display at the NMC console, and for inclusion in Monitor blocks. These blocks were recorded on magnetic tape as well as appearing in the NOCC displays. The NMC was required to operate without interruption for the duration of the Radio Science data acquisition period. The Area Routing Assembly (ARA), which was part of the Digital Communications Subsystem, controlled all data communication between the stations and JPL. The ARA received all required data and status messages from the NMC/CMC, and could record them to tape as well as transmit them to JPL via data lines. The ARA also received predicts and other data from JPL, and passed them on to the CMC. DSCC Antenna Mechanical Subsystem --------------------------------- Radio Science activities generally required support from the 70-m, 34-m HEF, and 34-m BWG antenna subnets. The antennas at each DSCC functioned as large-aperture collectors which, by double reflection, caused the incoming radio frequency (RF) energy to enter the feed horns. The large collecting surface of the antenna focused the incoming energy onto a subreflector, which was adjustable in both axial and angular position. These adjustments were made to correct for gravitational deformation of the antenna as it moved between zenith and the horizon; the deformation could be as large as 7 cm. The subreflector adjustments optimized the channeling of energy from the primary reflector to the subreflector, and then to the feed horns. The 70-m and 34-m HEF antennas had 'shaped' primary and secondary reflectors, with forms that were modified paraboloids. This customization allowed more uniform illumination of one reflector by another. The BWG reflector shape was ellipsoidal. On the 70-m antennas, the subreflector directed received energy from the antenna onto a dichroic plate, a device which reflected S-band energy to the S-band feed horn and passed X-band energy through to the X-band feed horn. In the 34-m HEF, there was one 'common aperture feed', which accepted both frequencies without requiring a dichroic plate. In the 34-m BWG, a series of small mirrors (approximately 2.5 meters in diameter) directed microwave energy from the subreflector region to a collection area at the base of the antenna -- typically in a pedestal room. A retractable dichroic reflector separated the S and X bands on some BWG antennas, or the X and Ka bands on others. RF energy to be transmitted into space by the horns was focused by the reflectors into narrow cylindrical beams, pointed with high precision (either to the dichroic plate or directly to the subreflector) by a series of drive motors and gear trains that could rotate the movable components and their support structures. The different antennas could be pointed by several means. Two pointing modes commonly used during tracking passes were CONSCAN and 'blind pointing.' With CONSCAN enabled and a closed-loop receiver locked to a spacecraft signal, the system tracked the radio source by conically scanning around its position in the sky. Pointing angle adjustments were computed from signal strength information (feedback) supplied by the receiver. In this mode the Antenna Pointing Assembly (APA) generated a circular scan pattern which was sent to the Antenna Control System (ACS). The ACS added the scan pattern to the corrected pointing angle predicts. Software in the receiver-exciter controller computed the received signal level and sent it to the APA. The correlation of scan position with the received signal level variations allowed the APA to compute offset changes which were sent to the ACS. Thus, within the capability of the closed-loop control system, the scan center was pointed precisely at the apparent direction of the spacecraft signal source. An additional function of the APA was to provide antenna position angles and residuals, antenna control mode/status information, and predict-correction parameters to the Area Routing Assembly (ARA) via the LAN, which then sent this information to JPL via the Ground Communications Facility (GCF) for antenna status monitoring. During periods when excessive signal level dynamics or low received signal levels were expected (e.g., during an occultation experiment), CONSCAN could not be used. Under these conditions, blind pointing (CONSCAN OFF) was used, and pointing angle adjustments were based on a predetermined Systematic Error Correction (SEC) model. Independent of CONSCAN state, subreflector motion in at least the z-axis could introduce phase variations into the received Radio Science data. For that reason, during certain experiments, the subreflector in the 70-m and 34-m HEFs was frozen in the z-axis at a position (often based on elevation angle) selected to minimize phase change and signal degradation. This could be done via Operator Control Inputs (OCIs) from the NMC to the Subreflector Controller (SRC) which resided in the alidade room of the antennas. The SRC passed the commands to motors that drove the subreflector to the desired position. Pointing angles for all antenna types were computed by the NOCC Support System (NSS) from an ephemeris provided by the flight project. These predicts were received and archived by the CMC. Before each track, they were transferred to the APA, which transformed the direction cosines of the predicts into AZ-EL coordinates. The LMC operator then downloaded the antenna predict points to the antenna-mounted ACS computer along with a selected SEC model. The pointing predicts consisted of time-tagged AZ-EL points at selected time intervals along with polynomial coefficients for interpolation between points. The ACS automatically interpolated the predict points, corrected the pointing predicts for refraction and subreflector position, and added the proper systematic error correction and any manually entered antenna offsets. The ACS then sent angular position commands for each axis at the rate of one per second. In the 70-m and 34-m HEF, rate commands were generated from the position commands at the servo controller and were subsequently used to steer the antenna. When not using binary predicts (the routine mode for spacecraft tracking), the antennas could be pointed using 'planetary' mode -- a simpler mode which used right ascension (RA) and declination (DEC) values. These changed very slowly with respect to the celestial frame. Values were provided to the station in text form for manual entry. The ACS quadratically interpolated among three RA and DEC points which were on one-day centers. A third pointing mode -- sidereal -- was available for tracking radio sources fixed with respect to the celestial frame. Regardless of the pointing mode being used, a 70-m antenna had a special high-accuracy pointing capability called 'precision' mode. A pointing control loop derived the main AZ-EL pointing servo drive error signals from a two- axis autocollimator mounted on the Intermediate Reference Structure. The autocollimator projected a light beam to a precision mirror mounted on the Master Equatorial drive system, a much smaller structure, independent of the main antenna, which was exactly positioned in HA and DEC with shaft encoders. The autocollimator detected elevation/cross- elevation errors between the two reference surfaces by measuring the angular displacement of the reflected light beam. This error was compensated for in the antenna servo by moving the antenna in the appropriate AZ-EL direction. Pointing accuracies of 0.004 degrees (15 arc seconds) were possible in 'precision' mode. The 'precision' mode was not available on 34-m antennas -- nor was it needed, since their beamwidths were twice as large as on the 70-m antennas. DSCC Antenna Microwave Subsystem -------------------------------- 70-m Antennas: Each 70-m antenna had three feed cones installed in a structure at the center of the main reflector. The feeds were positioned 120 degrees apart on a circle. Selection of the feed was made by rotation of the subreflector. A dichroic mirror assembly, half on the S-band cone and half on the X-band cone, permitted simultaneous use of the S- and X-band frequencies. The third cone was devoted to R&D and more specialized work. The Antenna Microwave Subsystem (AMS) accepted the received S- and X-band signals at the feed horn and transmitted them through polarizer plates to an orthomode transducer. The polarizer plates were adjusted so that the signals were directed to a pair of redundant amplifiers for each frequency, thus facilitating the simultaneous reception of signals in two orthogonal polarizations. For S-band these were two Block IVA S-band Traveling Wave Masers (TWMs); for X-band the amplifiers were Block IIA TWMs. 34-m HEF Antennas: The 34-m HEF used a single feed for both S- and X-band. Simultaneous S- and X-band receive as well as X-band transmit was possible thanks to the presence of an S/X 'combiner' which acted as a diplexer. For S-band, RCP or LCP was user selected through a switch, so neither a polarizer nor an orthomode transducer was needed. The X-band amplification options included two Block II TWMs or a High Electron Mobility Transistor (HEMT) Low Noise Amplifier (LNA), while the S-band amplification was provided by a Field Effect Transistor (FET) LNA. 34-m BWG Antennas: These antennas used feeds and low-noise amplifiers (LNA) in the pedestal room, which could be switched in and out as needed. Typically the following modes were available: 1. downlink non-diplexed path (RCP or LCP) to LNA-1, with uplink in the opposite circular polarization; 2. downlink non-diplexed path (RCP or LCP) to LNA-2, with uplink in the opposite circular polarization; 3. downlink diplexed path (RCP or LCP) to LNA-1, with uplink in the same circular polarization; 4. downlink diplexed path (RCP or LCP) to LNA-2, with uplink in the same circular polarization. For BWG antennas with dual-band capabilities and dual LNAs, each of the above four modes could be used in a single- or dual-frequency configuration. Thus, for antennas with the most complete capabilities, there were sixteen possible ways to receive (2 polarizations, 2 waveguide path choices, 2 LNAs, and 2 bands). DSCC Uplink Subsystem --------------------- The Uplink Subsystem (UPL) comprised the Exciter, the Command Modulation, Uplink Controller, and Uplink Ranging assemblies. The UPL was based around the Block V Exciter (BVE) equipment. The BVEs generated uplink carrier and uplink range phase data, and delivered these data directly to the Project. The exciter generated a sky-level signal which was provided to the Transmitter Subsystem for the spacecraft uplink signal. Based on predicts from the CMC, the BVE provided a sky-level uplink signal to either the low-power or the high-power transmitter. It was tunable under command of the DCO (Digitally Controlled Oscillator). The diplexer in the signal path between the transmitter and the feed horn for all antenna types (used for simultaneous transmission and reception) could be configured such that it was out of the received signal path (in listen-only or bypass mode) in order to improve the signal-to-noise ratio in the receiver system. DSCC Downlink Subsystem ----------------------- The Downlink Subsystem consisted of three groups of equipment: the closed-loop receiver group, the open-loop receiver group, and the RF monitor group. Closed-Loop Receivers: The current closed-loop group, called the Downlink Tracking and Telemetry Subsystem (DTK), consisted of the Downlink Controller, the Receiver and Ranging Processor (RRP), and the Telemetry Processor (TLP) assemblies. The DTT was currently based around the Block V Receiver (BVR) equipment. The BVRs generated downlink carrier and downlink range phase data (not Doppler counts and ranging units, as had been the case before early 2003), and delivered these (phase) data directly to the Project. The DTT could simultaneously support as many downlink channels as could be assigned by the NMC, up to the total number of RRPs available at a given complex (allowing for the reception of several different frequencies/wavelengths/bands, or different polarizations of the same downlink band). The only other constraint was that any selected downlink band/bands had to be supported by that antenna. The closed-loop receivers provided the capability for rapid acquisition of a spacecraft signal, and telemetry lock-up. In order to accomplish signal acquisition within a short time, the receivers were predict driven to search for, acquire, and track the downlink automatically. Rapid acquisition precluded manual tuning, though that remained as a backup capability. The BVRs utilized FFT analyzers for rapid lock-up. The downlink predicts were generated by the NSS and then transmitted to the CMC, which sent them to the Receiver Subsystem where two sets could be stored. The receiver started acquisition at the beginning of a track (pass), or at an operator-specified time. The BVRs could also be operated from the NMC without local operators attending them. The receivers also sent performance and status data, displays, and event messages to the NMC. With the BVRs, the simulation (SIM) synthesizer signal was used as the reference for the Doppler extractor. The synthesizer was adjusted before the beginning of the pass to a frequency that was appropriate for the channel (i.e., within the band) of the incoming signal; and would generally remain constant during the pass. The closed-loop receiver AGC loop could be configured to one of three settings: narrow, medium, or wide. It was configured such that the expected amplitude changes were accommodated with minimum distortion. The loop bandwidth (2BLo) was configured such that the expected phase changes could be accommodated while maintaining the best possible loop SNR. Open-Loop Receivers: The open-loop Radio Science Receiver (RSR) was a dedicated receiver that got a downconverted signal (about 300 MHz), filtered the signal to limit its bandwidth (to 265-375 MHz, centered at 320 MHz), and then further downconverted (to a center frequency of 64 MHz) and digitized the signal. The RSR filters were specified by their bandwidths, desired resolution, and offset from the predicted sky frequency. The open-loop receivers operated in both a link-assigned and a stand-alone mode. In the link-assigned mode, the NMC received monitor data from the RSR for incorporation into the data set for tracking support, and provided a workstation from which the RSR could be operated. RSRs that were not assigned to a link could be operated in a stand-alone mode without interference to any activities in progress at the complex. Monitor data were not sent to the NMC by RSRs operating in the stand-alone mode. DSCC Transmitter Subsystem -------------------------- The Transmitter (TXR) Subsystem accepted a sky-level frequency exciter signal from the Uplink (Exciter) Subsystem exciter. This signal was routed via the diplexer through the feed horn to the antenna, where it was then focused and beamed to the spacecraft. The Transmitter Subsystem power capabilities ranged from 18 kW to 400 kW, for S- and X-band uplink. Power levels above 20 kW were available only at 70-m stations. DSCC Tracking Subsystem ----------------------- Beginning in early 2003, all the Tracking Subsystem functions were incorporated within the Uplink Subsystem (UPL) and the Downlink Tracking and Telemetry Subsystem (DTT) -- the DTK was eliminated. DSCC Frequency and Timing Subsystem ----------------------------------- The Frequency and Timing Subsystem (FTS) provided all of the frequency and timing references required by the other DSCC subsystems. It contained four frequency standards, of which one was prime and the other three were backups. Selection of the prime standard was done via the CMC. Of these four standards, two were hydrogen masers followed by clean-up loops (CUL) and two were cesium standards. These four standards all fed the Coherent Reference Generator (CRG), which provided the frequency references used by the rest of the complex. FTS also provided the frequency reference to the Master Clock Assembly (MCA), which in turn provided time to the Time Insertion and Distribution Assembly (TID), which provided UTC and SIM-time to the complex. JPL's ability to monitor the FTS at each DSCC was limited to the station-calculated Doppler pseudo-residuals, the Doppler noise, the RSR, the SSI, and to a system that used the Global Positioning System (GPS). GPS receivers at each DSCC received a one-pulse-per-second signal from the station's (hydrogen- maser-referenced) FTS and a pulse from a GPS satellite at scheduled times. After compensating for the satellite signal delay, the timing offset was reported to JPL, where a database was kept. The clock offsets stored in the JPL database were given in microseconds; each entry was a mean reading of the measurements from several GPS satellites, and a time tag associated with the mean reading. The clock offsets that were provided included those of SPC 10 relative to UTC (NIST), SPC 40 relative to SPC 10, etc. Optics - DSN ============ Performance of the DSN ground stations depended primarily on size of the antenna and capabilities of the electronics. These are summarized in the following set of tables. Beamwidth is half-power full angular width. Polarization is circular; L denotes left circular polarization (LCP), and R denotes right circular polarization (RCP). DSS S-Band Characteristics 70-m 34-m 34-m Transmit BWG HEF -------- ----- ----- ----- Frequency (MHz) 2110- 2025- N/A 2120 2120 Wavelength (m) 0.142 0.142 N/A Ant Gain (dBi) 62.7 56.1 N/A Beamwidth (deg) 0.119 N/A N/A Polarization L or R L or R N/A Tx Power (kW) 20-100 20 N/A Receive ------- Frequency (MHz) 2270- 2270- 2200- 2300 2300 2300 Wavelength (m) 0.131 0.131 0.131 Ant Gain (dBi) 63.3 56.7 56.0 Beamwidth (deg) 0.108 N/A 0.24 Polarization L & R L or R L or R System Temp (K) 20 31 38 DSS X-Band Characteristics 70-m 34-m 34-m Transmit BWG HEF -------- ----- ----- ----- Frequency (MHz) 8495 7145- 7145- 7190 7190 Wavelength (m) 0.035 0.042 0.042 Ant Gain (dBi) 74.2 66.9 67 Beamwidth (deg) N/A 0.074 Polarization L or R L or R L or R Tx Power (kW) 20 20 20 Receive ------- Frequency (MHz) 8400- 8400- 8400- 8500 8500 8500 Wavelength (m) 0.036 0.036 0.036 Ant Gain (dBi) 74.2 68.1 68.3 Beamwidth (deg) 0.031 N/A 0.063 Polarization L & R L & R L & R System Temp (K) 20 30 20 NB: The X-band 70-m transmitting parameters are given at 8495 MHz, the frequency used by the Goldstone planetary radar system. For telecommunications, the transmitting frequency was in the range 7145-7190 MHz, the power would typically be 20 kW, and the gain would be about 72.6 dB (70-m antenna). When ground transmitters were used in spacecraft radio science experiments, the details of transmitter and antenna performance rarely impacted the results. Calibration - DSN ================= Calibrations of hardware systems were carried out periodically by DSN personnel; these ensured that systems operated at required performance levels -- for example, that antenna patterns, receiver gain, propagation delays, and Doppler uncertainties met specifications. No information on specific calibration activities was available. Nominal performance specifications are shown in the tables above. Additional information may be available in [DSN810-5]. Prior to each tracking pass, station operators performed a series of calibrations to ensure that systems met specifications for that operational period. Included in these calibrations was measurement of receiver system temperature in the configuration to be employed during the pass. Results of these calibrations were recorded in (hard copy) Controller's Logs for each pass. Operational Considerations - DSN ================================ The DSN was a complex and dynamic 'instrument.' Its performance for Radio Science depended on a number of factors from equipment configuration to meteorological conditions. No specific information on 'operational considerations' can be given here. Operational Modes - DSN ======================= DSCC Antenna Mechanical Subsystem --------------------------------- Pointing of DSCC antennas could be carried out in several ways. For details see the subsection 'DSCC Antenna Mechanical Subsystem' in the 'Subsystem' section. Binary pointing was the preferred mode for tracking spacecraft; pointing predicts were provided, and the antenna simply followed those. With CONSCAN, the antenna scanned conically about the optimum pointing direction, using closed-loop receiver signal strength estimates as feedback. In planetary mode, the system interpolated from three (slowly changing) RA-DEC target coordinates; this was 'blind' pointing since there was no feedback from a detected signal. In sidereal mode, the antenna tracked a fixed point on the celestial sphere. In 'precision' mode, the antenna pointing was adjusted using an optical feedback system. In addition, it was possible on most antennas to freeze the z-axis motion of the subreflector to minimize phase changes in the received signal. DSCC Downlink Tracking and Telemetry Subsystem ---------------------------------------------- The diplexer in the signal path between the transmitter and the feed horns on all antennas could be configured so that it was out of the received signal path in order to improve the signal-to-noise ratio in the receiver system. This was known as the 'listen-only' or 'bypass' mode. Closed-Loop vs. Open-Loop Reception ----------------------------------- Radio Science data could be collected in two modes: closed- loop, in which a phase-locked loop receiver tracked the spacecraft signal, or open-loop, in which a receiver sampled and recorded a band within which the desired signal presumably resided. Closed-loop data were collected using Closed-Loop Receivers, and open-loop data were collected using Open-Loop Receivers in conjunction with the Full Spectrum Processing Subsystem (FSP). See the Subsystems section for further information. Closed-Loop Receiver AGC Loop ----------------------------- The closed-loop receiver AGC loop could be configured to one of three settings: narrow, medium, or wide. In general, it was configured so that expected signal amplitude changes were accommodated with minimum distortion. The loop bandwidth was typically configured so that expected phase changes could be accommodated while maintaining the best possible loop SNR. Coherent vs. Non-Coherent Operation ----------------------------------- The frequency of the signal transmitted from the spacecraft could generally be controlled in two ways -- by locking to a signal received from a ground station or by locking to an on-board oscillator. These were known as the coherent (or 'two-way') and non-coherent ('one-way') modes, respectively. Mode selection was made at the spacecraft, based on commands received from the ground. When operating in the coherent mode, the transponder carrier frequency was derived from the received uplink carrier frequency with a 'turn-around ratio' as expressed in the table below: Uplink Downlink Turn-Around Band Band Ratio --------------------------------- X X 880/749 X S 240/749 In the non-coherent mode, the downlink carrier frequency was derived from the spacecraft's on-board, crystal-controlled oscillator. Either closed-loop or open-loop receivers (or both) could be used with either spacecraft frequency reference mode. Closed-loop reception in the two-way mode was usually preferred for routine tracking/navigation. Occasionally the spacecraft operated coherently such that one ground station did the transmitting, and a second/different ground station received the 'downlink' signal -- this was referred to as the 'three-way' mode. Media Calibration System ------------------------ The Earth's atmosphere contributes phase and amplitude noise to the spacecraft radio signal received at a ground station. Each DSCC had a GPS receiver subsystem to calibrate for both the ionosphere and troposphere (both wet and dry components), along the zenith direction. This subsystem also measured the temperature, pressure, humidity, wind speed and direction, and Faraday rotation. Location - DSN ============== Accurate spacecraft navigation using radio metric data required knowledge of the locations of the DSN tracking stations. The coordinate system in which the locations of the tracking stations were expressed needed to be consistent with the reference frame definitions used to provide Earth orientation calibrations. The International Earth Rotation Service (IERS) had established a terrestrial reference frame for use with Earth orientation measurements. The IERS issued a new realization of the terrestrial reference frame each year. The definition of the coordinate system changed slowly as the data improved, and as ideas about how best to define the coordinate system developed. The overall changes from year to year were at the level of as few centimeters. The 1993 version of the IERS Terrestrial Reference Frame (IRTF1993) was most used for DSN station locations. The DSN station locations were determined by use of VLBI measurements, and by conventional and GPS surveying. Tables of station locations were available in either Cartesian or geodetic coordinates. The geodetic coordinates were referred to a geoid with an equatorial radius of 6378136.3 m, and a flattening factor f=298.257, as described in IERS Technical Note 13. The DSN Station Locations in ITRF1993 Cartesian reference frame at epoch 1993.0 (assuming subreflector-fixed configuration) were as follows: Antenna x(m) y(m) z(m) ------------------------------------------------ DSS 12 -2350443.812 -4651980.837 +3665630.988 DSS 13 -2351112.491 -4655530.714 +3660912.787 DSS 14 -2353621.251 -4641341.542 +3677052.370 DSS 15 -2353538.790 -4641649.507 +3676670.043 DSS 16 -2354763.158 -4646787.462 +3669387.069 DSS 17 -2354730.357 -4646751.776 +3669440.659 DSS 23 -2354757.567 -4646934.675 +3669207.824 DSS 24 -2354906.528 -4646840.114 +3669242.295 DSS 25 -2355021.795 -4646953.325 +3669040.628 DSS 26 -2354890.967 -4647166.925 +3668872.212 DSS 27 -2349915.260 -4656756.484 +3660096.529 DSS 28 -2350101.849 -4656673.447 +3660103.577 DSS 33 -4461083.514 +2682281.745 -3674570.392 DSS 34 -4461146.720 +2682439.296 -3674393.517 DSS 42 -4460981.016 +2682413.525 -3674582.072 DSS 43 -4460894.585 +2682361.554 -3674748.580 DSS 45 -4460935.250 +2682765.710 -3674381.402 DSS 46 -4460828.619 +2682129.556 -3674975.508 DSS 49 -4554231.843 +2816758.983 -3454036.065 (Parkes) DSS 53 +4849330.129 -0360338.092 +4114758.766 DSS 54 +4849434.496 -0360724.062 +4114618.570 DSS 55 +4849525.318 -0360606.299 +4114494.905 DSS 61 +4849245.211 -0360278.166 +4114884.445 DSS 63 +4849092.647 -0360180.569 +4115109.113 DSS 65 +4849336.730 -0360488.859 +4114748.775 DSS 66 +4849148.543 -0360474.842 +4114995.021 The DSN Station Locations in ITRF1993 Geodetic reference frame at epoch 1993.0 (assuming subreflector-fixed configuration) were as follows: latitude longitude height Antenna deg min sec deg min sec (m) ---------------------------------------------------------- DSS 12 35 17 59.77577 243 11 40.24697 962.87517 DSS 13 35 14 49.79342 243 12 19.95493 1071.17855 DSS 14 35 25 33.24518 243 6 37.66967 1002.11430 DSS 15 35 25 18.67390 243 6 46.10495 973.94523 DSS 16 35 20 29.54391 243 7 34.86823 944.71108 DSS 17 35 20 31.83778 243 7 35.38803 937.65000 DSS 23 35 20 22.38335 243 7 37.70043 946.08556 DSS 24 35 20 23.61492 243 7 30.74701 952.14515 DSS 25 35 20 15.40494 243 7 28.70236 960.38138 DSS 26 35 20 8.48213 243 7 37.14557 970.15911 DSS 27 35 14 17.78052 243 13 24.06569 1053.20312 DSS 28 35 14 17.78136 243 13 15.99911 1065.38171 DSS 33 -35 24 1.76138 148 58 59.12204 684.83864 DSS 34 -35 23 54.53984 148 58 55.06236 692.71119 DSS 42 -35 24 2.44494 148 58 52.55396 675.35557 DSS 43 -35 24 8.74388 148 58 52.55394 689.60780 DSS 45 -35 23 54.46400 148 58 39.65992 675.08630 DSS 46 -35 24 18.05462 148 58 59.08571 677.55141 DSS 49 -32 59 54.25297 148 15 48.64683 415.52885 DSS 53 40 25 38.48036 355 45 1.24307 827.50081 DSS 54 40 25 32.23152 355 44 45.24459 837.60097 DSS 55 40 25 27.45965 355 44 50.51161 819.70966 DSS 61 40 25 43.45508 355 45 3.51113 841.15897 DSS 63 40 25 52.34908 355 45 7.16030 865.54412 DSS 65 40 25 37.86055 355 44 54.88622 834.53926 DSS 66 40 25 47.90367 355 44 54.88739 850.58213 Measurement Parameters - DSN ============================ Open-Loop System ---------------- Output from the Open-Loop Receivers (OLRs), as sampled and recorded by the Radio Science Receiver (RSR), was a stream of 1-, 2-, 4-, 8-, or 16-bit I (In-Phase) and Q (Quadrature- Phase) samples. The spacecraft transmitted an RF signal to an antenna, where the signal was downconverted to IF. The RSR selected an IF signal for a particular frequency band and passed it through a digitizer (where it was attenuated and then mixed with timing information). The signal was then decimated, filtered (to I&Q samples), and then multiplied by the signal from a numerically controlled oscillator. Finally, the RSR reduced the bandwidth and sample rate of the samples, and truncated the results (thus creating an offset of -0.5 in the output data). The samples of data were packed into SFDU blocks (nominally containing a single second's worth of data), and a header was attached to provide the following associated data for the record: - time tag for the first sample in the data block - data source identification (DSS, RSR, and sub-channel), and frequency band - data sample resolution (bits per sample) and rate (samples per second) - filter gain, ADC RMS amplitude, and attenuation - frequency and phase polynomial coefficients Closed-Loop System ------------------ Since mid 2003, closed-loop data were recorded and provided in Tracking and Navigation Files (TNFs). The TNFs were comprised of SFDUs that had variable-length, variable-format records with mixed typing (i.e., can contain ASCII, integer, and floating-point items in a single record). These files all contained entries that included measurements of Doppler, range, and signal strength, along with status and uplink frequency information. Acronyms and Abbreviations ========================== 1PPS One Pulse per Second ACS Antenna Control System ADEV Allan Deviation AGC Automatic Gain Control APA Antenna Pointing Assembly BPF Band Pass Filter BWG Beam Wave Guide CFE Common Front End CONSCAN Conical Scanning D/L downlink dBi decibel relative to isotropic DCP Development Control Position DDC Digital Down Converter DDS Data Distribution System DSCC Deep Space Communications Complex DSN Deep Space Network DSP Digital Signal Processing DSS Deep Space Network Station ESA European Space Agency ESOC European Space Operations Centre FTS Frequency and Timing subsystem HEF High Efficiency HGA High Gain Antenna HSB High-Speed BWG IFMS Intermediate Frequency Modulation System IVC Intermediate Frequency Selection Switch JPL Jet Propulsion Laboratory LCP Left Circular Polarization LGA Low Gain Antenna LNA Low Noise Amplifier LPF Low Pass filter MaRS Mars Express orbiter Radio Science Experiment MB Medium band Mbit Mega bit MEXX Mars Express MOLA Mars Orbiting Laser Altimeter MRS Mars Express Radio Science N/A not applicable NASA National Aeronautics and Space Administration NMC Network Monitor and Control NNO New Norcia OCC Operation Control Centre ODF Orbit Data File OLR Open Loop Receiver PDS Planetary Data System PI Principal Investigator pwr power rcvrs receivers RCP Right Circular Polarization RF Radio Frequency RFDU Radio Frequency Distribution Unit RIV Radio Science IF-VF Downconverter rms root mean square RSR Radio- Science Receiver RSS Radio Science Subsystem SIM SNR Signal-Noise-Ratio SNT System Noise Temperature SPC Signal Processing Center STC Station Computer sps samples per second STAT Science Time Analysis Tool TCDS Telemetry Channel Decoding System TCXO Temperature Controlled Crystal Oscillator TID Time Insertion and Distribution Assembly TNF Tracking and Navigation File TWOD Two-way dual-frequency mode TWOS Two-way single-frequency mode TWTA Traveling wave tube amplifier Tx Transmitter U/L uplink UNK unknown UTC Coordinated Universal Time VDP VME Data Processor VF Video Frequency VME Versa Module Eurocard (standard bus) w watt
  • instrument : Comet Nucleus Sounding Experiment by Radiowave Transmission for RL
    INSTRUMENT OVERVIEW =================== THE COMPLETE CONSERT EXPERIMENT IS COMPOSED OF: - ONE ORBITER PART (ELECTRONICS, ANTENNA, HARNESS) - ONE LANDER PART (ELECTRONICS, ANTENNAS,HARNESS) SCIENTIFIC OBJECTIVES ===================== Our experiment concerns the rough tomography of the comet nucleus performed by the CONSERT instrument. It works as a time domain transponder between one module landed on the comet surface (Lander) and another flying around the comet (Orbiter). The CONSERT signal consists in a 90 MHz sinusoidal waveform which is phase modulated by a pseudorandom code or PSK (phase shift keying) coding. Such frequency, in the radio range, is expected to minimize the losses during the propagation inside the comet material and the generated pulse code maximizes the signal to noise ratio. In these experimental conditions CONSERT will perform a measurement of the mean dielectric properties and on the detection of large size embedded structures or small irregularities within the comet nucleus. CALIBRATION =========== Due to the design of the instrument, there is no special calibration operation mode. The science data are composed of the signal characterising the propagation channel of the comet nucleus as a function of time: - The propagation time is the main data to be inverted and its accuracy is waranted by the Consert clock absolute accuracy and stability. - The signal amplitude can also provide information about the nucleus structure but there is no internal calibration channel to increase the link budget accuracy. OPERATIONAL CONSIDERATIONS ========================== The Lander makes a coherent addition and a detection of the position (delay) of the correlation principal peak. A clean coded signal is finally back emitted to the Orbiter with the found delay. The orbiter accumulates the received signal and sends it to the Earth (via the satellite interface). OPERATIONAL MODES ================= Each scientific measurement sequence (called scanning sequence) involves the Orbiter part and the Lander. The duration of this scan sequence is related to the duration of the orbit of the Rosetta spacecraft relatively to the Lander on the rotating comet nucleus. This duration is typically of the order of 10 hours during one revolution around the nucleus. The individual duration of each sounding is less than one second. This duration of scanning sequence should correspond to the time when the Lander and the Orbiter are separated by the comet. During the direct line of sight periods, where Philae and Rosetta are in visibility of each others, the synchronization occurs. This mandatory and preliminary phase is also called the CONSERT 'tuning'. This means that the duration of the data recording does not correspond to the total time of one revolution, but only for the part where Philae and Rosetta are in occultation by the comet nucleus. The number of samples which theoretically should be taken is given by the following formula: 2 * pi * radius of comet / ( lambda / 2 ) where lambda is wavelength. During the scanning sequence for a hypothetical circular comet with a 750 m radius, about 3000 individual measurements, called soundings should be taken. The general structure of the Consert operational scenario does not depend on the comet type that is explored during the Rosetta mission, however a certain amount of the parameters depends on the shape and size of the comet nucleus and on the orbit of the spacecraft and nucleus rotation. MEASURED PARAMETERS =================== While acquiring the signal which propagates through the comet nucleus, the CONSERT intrument measures the following physical parameters: - Propagation time of one to three peaks (in case of multi-path propagation). - Attenuation of these signal peaks along the propagation channel.
  • data set : NEW HORIZONS REX KEMCRUISE1 CALIBRATED V2.0
    Calibrated data taken by New Horizons Radio Science Experiment instrument during the CRUISE TO FIRST KBO ENCOUNTER mission phase. This is VERSION 2.0 of this data set.
  • data set : DAWN CERES GRAVITY SCIENCE DERIVED SCIENCE DATA V4.0
    Dawn Ceres reduced gravity data.
  • data set : NEW HORIZONS REX KEMCRUISE1 RAW V2.0
    Raw data taken by New Horizons Radio Science Experiment instrument during the CRUISE TO FIRST KBO ENCOUNTER mission phase. This is VERSION 2.0 of this data set.
  • data set : MRO RADIO SCIENCE DERIVED GRAVITY SCIENCE DATA PRODUCTS V1.0
    Mars Reconnaissance Orbiter reduced gravity data.
  • instrument : RADIO SCIENCE SUBSYSTEM for MRO
    Radio science investigations utilize instrumentation with elements both on a spacecraft and at ground stations -- in this case, at the NASA Deep Space Network (DSN). The spacecraft part of the radio science instrument is described immediately below; that is followed by a description of the DSN (ground) part of the instrument.
  • data set : CASSINI RSS RAW DATA SET - SROC1 V1.0
    Cassini RSS SROC1 raw data.
  • data set : CASSINI RSS RAW DATA SET - SROC2 V1.0
    Cassini RSS SROC2 raw data.
  • instrument : RADIO SCIENCE SUBSYSTEM for MER2
    The Mars Exploration Rover (MER) mission includes two spacecraft, Spirit (MER-2) and Opportunity (MER-1). The MER Radio Science data consist of measurements of the Doppler shift of the rover radio signal as measured by the NASA Deep Space Network (DSN) and by the Mars Odyssey orbiter (ODY). The primary purpose of all equipment was collection of telemetry from the rovers, with Doppler measurements made for rover position determination as required and on a best efforts basis to support Radio Science. The performance and calibration of both the spacecraft and tracking stations directly affected the radio science data accuracy, and they played a major role in determining the quality of the results. The MER part of the radio science instrument is described immediately below; that is followed by a description of the relevant ODY relay radio system and a description of the DSN (ground) part of the instrument.
  • instrument : RADIO SCIENCE SUBSYSTEM for GO
    Instrument Overview =================== Galileo Radio Science investigations utilized instrumentation with elements on the spacecraft and at the Deep Space Network (DSN). Much of this was shared equipment, being used for routine telecommunications as well as for Radio Science. The performance and calibration of both the spacecraft and tracking stations directly affected the radio science data accuracy, and they played a major role in determining the quality of the results. The spacecraft part of the radio science instrument is described immediately below; that is followed by a description of the DSN (ground) part of the instrument. Radio Science investigations were carried out by two teams. The Celestial Mechanics Team, under Team Leader John Anderson, conducted experimental tests of general relativity (including searching for gravitational waves), made measurements to improve solar system ephemerides, and sought to improve gravitational models for Jupiter and its satellites [ANDERSONETAL1992]. The Radio Propagation Team, under Team Leader Tay Howard, investigated the solar corona and carried out various studies in the Jovian system primarily concerning atmospheres and ionospheres [HOWARDETAL1992]. Instrument Specifications - Spacecraft ====================================== The Galileo spacecraft telecommunications subsystem served as part of a radio science subsystem for investigations primarily of Jupiter and its satellites, but also including Venus, the Earth-Moon system, and the Sun. Many details of the subsystem are unknown; its 'build date' is taken to be 1989-01-01, which was during the prelaunch phase of the Galileo mission. Instrument Id : RSS Instrument Host Id : GO Pi Pds User Id : UNK Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : 1989-01-01 Instrument Mass : UNK Instrument Length : UNK Instrument Width : UNK Instrument Height : UNK Instrument Manufacturer Name : UNK Instrument Overview - Spacecraft ================================ The spacecraft radio system was constructed around a redundant pair of transponders which received and transmitted at both S-band (2.3 GHz, 13 cm wavelength) and X-band (8.4 GHz, 3.6 cm wavelength) frequencies; the following combinations of uplink/downlink were supported by the design: S/S, X/X, S/X and S. The exact frequency transmitted from the spacecraft was controlled by the signal received from a ground station ('two-way' or 'coherent' mode) or by an on-board oscillator ('one-way' or 'non-coherent' mode). In some circumstances an uplink signal was transmitted from one ground station while two ground stations participated in reception; this was known as the 'three-way' mode. In the absence of an uplink signal, the spacecraft system switched automatically to the one-way mode. The on-board frequency reference could be either of two redundant 'auxiliary' crystal oscillators or a single ultra-stable oscillator (USO) provided specifically to support radio science observations. Each transponder included a receiver, command detector, exciter, and low-power amplifier. The transponders provided the usual uplink command and downlink data transmission capabilities. The following modulation states could be commanded: telemetry alone, ranging alone, telemetry and ranging, or carrier only. Each transponder could be operated through one of two low-gain antennas at S-band only; a furlable high-gain antenna (HGA) never deployed properly during Cruise, resulting in a serious degradation of radio science measurements, including loss of X-band capability. The HGA was aligned with the spin axis of the rotor part of the spacecraft. Low-Gain Antenna 1 (LGA-1) was located at the end of the HGA feed, so it is also aligned with the spin axis. LGA-2 was at the end of a boom, 3.52 m from the spin axis. When operating in the coherent mode, the transponder downlink frequency was related to the uplink frequency by the 'turn-around ratio' of 240/221 at S-band. At X-band it would have been 880/749. An X-band downlink controlled by an S-band uplink would have had a turn-around ratio of (240/221)*(11/3). Science Objectives ================== Two different types of radio science measurements were conducted with the Galileo Orbiter: radio tracking in which the magnitude and direction of gravitational forces could be derived from 'closed-loop' Doppler (and, sometimes, ranging) measurements, and radio propagation experiments in which modulation on the signal received at Earth stations could be attributed to properties of the intervening medium. The radio science measurements were analyzed by two investigation teams; the Celestial Mechanics Team was primarily interested in characterizing variations in gravitational forces, and the Radio Propagation Team was primarily interested in the atmospheres of the Sun, Jupiter, and Jupiter's satellites. Gravity Measurements -------------------- Measurement of the gravity field provides significant constraints on inferences about interior structure of Jupiter and its satellites. Precise, detailed study of spacecraft motion in Jupiter orbit and during satellite flybys can yield a mass distribution of each body and higher-order field terms if the measurements are sensitive enough. Compared with determinations from previous missions, improvements in the gravity field of Jupiter itself were not expected from tracking the Galileo Orbiter, but second-order gravity harmonics were expected from flyby encounters with satellites. One equatorial and one polar flyby at Ganymede were sought to determine independently the rotational and tidal response of the body assuming hydrostatic equilibrium. Departures from hydrostatic equilibrium were expected to confuse that issue at Europa, though the measurements were expected to be useful, while the relatively weak response to rotation and tides at Callisto made the experiment most marginal there [HUBBARD&ANDERSON1978]. Differences in principal moments of inertia to an accuracy of one percent or better were sought at Io [ANDERSONETAL1996]. Tests of General Relativity --------------------------- There has been continuing interest in testing the theory of general relativity by bouncing radar signals from hard planetary surfaces and using two-way ranging data from spacecraft anchored to other planetary bodies. No hard surface exists at Jupiter and no previous spacecraft had orbited the planet, so Galileo represented a unique opportunity to investigate this question. Two years of ranging to Galileo were expected to fix the range to Jupiter to an accuracy of about 150 m, with the limit set by orbit determination error along the Earth-Jupiter line and not by limitations of the radio 'instrument'. In combination with results from the Pioneer and Voyager spacecraft, these measurements were expected to lead to an improved ephemeris for Jupiter. As Jupiter (and Galileo) appear to pass behind the Sun when viewed from Earth, solar gravity should retard the radio signal propagating between the spacecraft and Earth. One set of time delay measurements to/from the Viking Orbiters and Landers agreed to within 0.1 percent of the General Relativity prediction. Measurements with Galileo were expected to be a factor of 5 worse, but the next best measurements were only to 2 percent of the General Relativity prediction. Not only would another set of measurements at the sub-one percent level be good experimental practice, but Galileo measurements could also verify the agreement over a range of directions in inertial space [WILL1981]. The red shift of the signal in Jupiter's gravitational field could be measured to an accuracy of about +/-1 percent after radiation hardening of the USO crystal in Jupiter's charged particle environment. Search for Gravitational Radiation ---------------------------------- Matter undergoing asymmetrical motion (theoretically) radiates gravitational waves which propagate at the velocity of light. Observed acceleration of the mean orbital motion of binary pulsar PSR 1913+16 is consistent with predictions [TAYLOR&WEISBERG1989]; other evidence is more ambiguous, and gravity waves themselves had not been detected with certainty before Galileo. For several extended periods during Galileo's cruise to Jupiter, when other spacecraft activity was at a minimum and when the spacecraft was near opposition, its radio link with Earth was monitored carefully for signs of passing, cosmicly generated, long period gravitational waves. Similar observations were conducted simultaneously with the Mars Observer and Ulysses spacecraft so that detections could be confirmed and direction of propagation of the gravitational waves inferred from time differences along other paths. Previous searches have been conducted using Viking, Voyager, and Pioneers 10 and 11 [ARMSTRONG1989]. Solar Corona Observations ------------------------- For several weeks around each of four superior conjunctions Galileo's radio link passed through the solar corona. Signals were scattered and refracted as they propagated through the turbulent plasma; the resulting modulation could be analyzed to obtain estimates of coronal structure and dynamics [WOO1993]. Specific objectives of the Galileo solar corona experiments included better understanding of: (1) three-dimensional electron density distribution and its relation to the photospheric magnetic field configuration, solar cycle, distance from the surface, and solar latitude; (2) structural differences among coronal 'holes', active regions, and the 'quiet' Sun; (3) characteristics of the acceleration regions of the solar wind in coronal holes, streamers, and other parts of the corona; (4) energy sources responsible for creation of coronal materials with temperatures over 1000000K; (5) resonant solar oscillations on the dynamical characteristics of the tenuous solar atmosphere; (6) excitation and propagation conditions for magnetoacoustic, Alfven, and other waves; and (7) form and evolution of disturbances near the Sun and their relationship to white light coronal mass ejections. Jupiter Occultations -------------------- Radio occultation measurements can contribute to an improved understanding of structure, circulation, dynamics, and transport in the atmosphere of Jupiter. Results from Galileo were based on detailed analysis of the radio signal as it entered and exited occultation by the planet. Three phases of the atmospheric investigation may be defined. The first is to obtain vertical profiles of electron content in the ionosphere; second is to extract large scale structure in the neutral atmosphere; third is to detect and interpret fine scale structure in both the ionospheric and neutral atmosphere profiles and to measure absorption in the neutral atmosphere. The Galileo tour permitted radio occultations on approximately half of the planned orbits at a number of latitudes. Pioneers 10 and 11 had earlier shown sharp, multiple, dense, low-lying ionospheric layers [FJELDBOETAL1976]. The vertical extent of the ionized layers, their time histories, and detailed structure were sought as keys to both the composition and chemistry of the upper atmosphere. With precise pointing of the HGA, Galileo was expected to penetrate below the condensation level for ammonia in the neutral atmosphere, providing global measures of ammonia concentration in well-mixed regions where Voyager had produced only one [LINDALETAL1981]. Measurements between 15N and 15S latitudes were expected to provide snapshots of vertical structure of waves propagating in the atmosphere; ingress and egress measurements from the same occultation could provide strong constraints on zonal wavenumber and meridional structure [HINSON&MAGALHAES1991]. Satellite Occultations ---------------------- Radio data acquired during occultation by a satellite could be used to determine its diameter to accuracies on the order of 1 km and, possibly, properties of any satellite atmosphere or ionosphere. In the case of Io a substantial ionosphere had been detected by Pioneer 10 [KLIOREETAL1975]; repeated occultations by Io were intended to improve understanding of spatial and temporal variability of the charged particles and their interaction with Jupiter's magnetic field. Occultations by the Io torus would provide a measure of the total number of free electrons along the propagation path, a useful constraint of the spatial structure of the torus. Jupiter's Magnetic Field ------------------------ Galileo was the first spacecraft equipped to measure both Faraday rotation of propagating waves and differential phase retardation between S- and X-band. Faraday rotation measurements were planned during each occultation by Jupiter and were to be used to investigate the characteristics of the magnetic field in the planet's ionosphere. Different models of the magnetic field yield differences in the predicted Faraday rotation on the order of 0.3 radians; the Faraday rotation experiment designed for Galileo exceeded this threshold by a factor of 10. Bistatic Scattering from Icy Galilean Moons ------------------------------------------- Monostatic radar echoes from Europa, Ganymede, and Callisto were found to be anomalously diffuse, strong, and polarized [CAMPBELLETAL1978]. By using the Galileo spacecraft as a microwave signal source during encounters with each of these bodies, the bistatic scattering as a function of angle could be determined, providing constraints on both the models for the anomalous scattering process and also the properties of the ice that presumably is responsible. Operational Considerations - Spacecraft ======================================= Because the HGA never deployed and only right-circularly polarized signals at S-band were available from LGA-1, the Faraday and dual-frequency measurements were never realized. For the Celestial Mechanics Team, the single frequency meant that signal dispersion resulting from passage through the solar wind, Earth's ionosphere, and other media could not be removed easily from data. For the Radio Propagation Team, the loss of antenna gain meant that only observations with the strongest signals could be made. Penetration below the ionosphere during Jupiter occultations and sensing charged and neutral particle environments of satellites became very difficult, and the bistatic surface experiments were dropped. Because Faraday Rotation experiments required linearly transmitted polarizations (available only from the HGA), those were also dropped. Calibration Description - Spacecraft ==================================== No information available. Platform Mounting Descriptions - Spacecraft =========================================== The HGA and LGA-1 antennas were mounted facing in the negative Zr direction; see the GO_SPACECRAFT_DESC_INST.CAT file for more information. Principal Investigators ======================= The Team Leader for the Celestial Mechanics Team was John D. Anderson of the Jet Propulsion Laboratory. Team members were (all from JPL): J.W. Armstrong J.K. Campbell F.B. Estabrook T.P. Krisher E.L. Lau The Team Leader for the Radio Propagation Team was H. Taylor Howard of Stanford University. Team members and affiliations were: V.R. Eshleman Stanford University D.P. Hinson Stanford University A.J. Kliore Jet Propulsion Laboratory G.F. Lindal Jet Propulsion Laboratory R. Woo Jet Propulsion Laboratory M.K. Bird University of Bonn, Germany H. Volland University of Bonn, Germany P. Edenhofer University of Bochum, Germany M. Paetzold DFLR, Germany H. Porsche DFLR, Germany Experiment Representative at JPL for both teams was Randy Herrera. Instrument Section / Operating Mode Descriptions - Spacecraft ============================================================= The Galileo radio system consisted of two sections, which could be operated in the following modes: Section Mode ------------------------------------------- Oscillator two-way (coherent) one-way (non-coherent) RF output low-gain antenna (choice from two) high-gain antenna (failed to deploy properly) Details for the radio system, as designed, are given in the table below: Transmitting Parameters: Frequency (MHz) 8415 2295 Transmit Power (w) 12 or 21 9 or 27 HGA Gain (dBi) 50 38 HGA Half-Power Beamwidth (deg) 0.6 1.5 Polarization LCP or RCP Linear Axial Ratio (dB) 2 32 Receiving Parameters: Frequency (MHz) 7167 2115 HGA Gain (dBi) 46 36 Polarization LCP or RCP Linear Noise Temperature (K) 270 1000 Instrument Overview - DSN ========================= Three Deep Space Communications Complexes (DSCCs) (near Barstow, CA; Canberra, Australia; and Madrid, Spain) comprise the DSN tracking network. Each complex is equipped with several antennas [including at least one each 70-m, 34-m High Efficiency (HEF), and 34-m standard (STD)], associated electronics, and operational systems. Primary activity at each complex is radiation of commands to and reception of telemetry data from active spacecraft. Transmission and reception is possible in several radio-frequency bands, the most common being S-band (nominally a frequency of 2100-2300 MHz or a wavelength of 14.2-13.0 cm) and X-band (7100-8500 MHz or 4.2- 3.5 cm). Transmitter output powers of up to 400 kw are available. Ground stations have the ability to transmit coded and uncoded waveforms which can be echoed by distant spacecraft. Analysis of the received coding allows navigators to determine the distance to the spacecraft; analysis of Doppler shift on the carrier signal allows estimation of the line-of-sight spacecraft velocity. Range and Doppler measurements are used to calculate the spacecraft trajectory and to infer gravity fields of objects near the spacecraft. Ground stations can record spacecraft signals that have propagated through or been scattered from target media. Measurements of signal parameters after wave interactions with surfaces, atmospheres, rings, and plasmas are used to infer physical and electrical properties of the target. Principal investigators vary from experiment to experiment. See the corresponding section of the spacecraft instrument description or the data set description for specifics. The Deep Space Network is managed by the Jet Propulsion Laboratory of the California Institute of Technology for the U.S. National Aeronautics and Space Administration. Specifications include: Instrument Id : RSS Instrument Host Id : DSN Pi Pds User Id : N/A Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : N/A Instrument Mass : N/A Instrument Length : N/A Instrument Width : N/A Instrument Height : N/A Instrument Manufacturer Name : N/A For more information on the Deep Space Network and its use in radio science investigations see the reports by [ASMAR&RENZETTI1993] and [ASMAR&HERRERA1993]. For design specifications on DSN subsystems see [DSN810-5]. For an example of use of the DSN for Radio Science see [TYLERETAL1992]. Subsystems - DSN ================ The Deep Space Communications Complexes (DSCCs) are an integral part of the Radio Science instrument, along with other receiving stations and the spacecraft Radio Frequency Subsystem. Their system performance directly determines the degree of success of Radio Science investigations, and their system calibration determines the degree of accuracy in the results of the experiments. The following paragraphs describe the functions performed by the individual subsystems of a DSCC. This material has been adapted from [ASMAR&HERRERA1993]; for additional information, consult [DSN810-5]. Each DSCC includes a set of antennas, a Signal Processing Center (SPC), and communication links to the Jet Propulsion Laboratory (JPL). The general configuration is illustrated below; antennas (Deep Space Stations, or DSS -- a term carried over from earlier times when antennas were individually instrumented) are listed in the table. -------- -------- -------- -------- -------- | DSS 12 | | DSS 18 | | DSS 14 | | DSS 15 | | DSS 16 | |34-m STD| |34-m STD| | 70-m | |34-m HEF| | 26-m | -------- -------- -------- -------- -------- | | | | | | v v | v | --------- | --------- --------->|GOLDSTONE|<---------- |EARTH/ORB| | SPC 10 |<-------------->| LINK | --------- --------- | SPC |<-------------->| 26-M | | COMM | ------>| COMM | --------- | --------- | | | v | v ------ --------- | --------- | NOCC |<--->| JPL |<------- | | ------ | CENTRAL | | GSFC | ------ | COMM | | NASCOM | | MCCC |<--->| TERMINAL|<-------------->| | ------ --------- --------- ^ ^ | | CANBERRA (SPC 40) <---------------- | | MADRID (SPC 60) <---------------------- GOLDSTONE CANBERRA MADRID Antenna SPC 10 SPC 40 SPC 60 -------- --------- -------- -------- 26-m DSS 16 DSS 46 DSS 66 34-m STD DSS 12 DSS 42 DSS 61 DSS 18 DSS 48 DSS 68 34-m HEF DSS 15 DSS 45 DSS 65 70-m DSS 14 DSS 43 DSS 63 Developmental DSS 13 Subsystem interconnections at each DSCC are shown in the diagram below, and they are described in the sections that follow. The Monitor and Control Subsystem is connected to all other subsystems; the Test Support Subsystem can be. ----------- ------------------ --------- --------- |TRANSMITTER| | | | TRACKING| | COMMAND | | SUBSYSTEM |-| RECEIVER/EXCITER |-|SUBSYSTEM|-|SUBSYSTEM|- ----------- | | --------- --------- | | | SUBSYSTEM | | | | ----------- | | --------------------- | | MICROWAVE | | | | TELEMETRY | | | SUBSYSTEM |-| |-| SUBSYSTEM |- ----------- ------------------ --------------------- | | | ----------- ----------- --------- -------------- | | ANTENNA | | MONITOR | | TEST | | DIGITAL | | | SUBSYSTEM | |AND CONTROL| | SUPPORT | |COMMUNICATIONS|- ----------- | SUBSYSTEM | |SUBSYSTEM| | SUBSYSTEM | ----------- --------- -------------- DSCC Monitor and Control Subsystem ---------------------------------- The DSCC Monitor and Control Subsystem (DMC) is part of the Monitor and Control System (MON) which also includes the ground communications Central Communications Terminal and the Network Operations Control Center (NOCC) Monitor and Control Subsystem. The DMC is the center of activity at a DSCC. The DMC receives and archives most of the information from the NOCC needed by the various DSCC subsystems during their operation. Control of most of the DSCC subsystems, as well as the handling and displaying of any responses to control directives and configuration and status information received from each of the subsystems, is done through the DMC. The effect of this is to centralize the control, display, and archiving functions necessary to operate a DSCC. Communication between the various subsystems is done using a Local Area Network (LAN) hooked up to each subsystem via a network interface unit (NIU). DMC operations are divided into two separate areas: the Complex Monitor and Control (CMC) and the Link Monitor and Control (LMC). The primary purpose of the CMC processor for Radio Science support is to receive and store all predict sets transmitted from NOCC such as Radio Science, antenna pointing, tracking, receiver, and uplink predict sets and then, at a later time, to distribute them to the appropriate subsystems via the LAN. Those predict sets can be stored in the CMC for a maximum of three days under normal conditions. The CMC also receives, processes, and displays event/alarm messages; maintains an operator log; and produces tape labels for the DSP. Assignment and configuration of the LMCs is done through the CMC; to a limited degree the CMC can perform some of the functions performed by the LMC. There are two CMCs (one on-line and one backup) and three LMCs at each DSCC The backup CMC can function as an additional LMC if necessary. The LMC processor provides the operator interface for monitor and control of a link -- a group of equipment required to support a spacecraft pass. For Radio Science, a link might include the DSCC Spectrum Processing Subsystem (DSP) (which, in turn, can control the SSI), or the Tracking Subsystem. The LMC also maintains an operator log which includes operator directives and subsystem responses. One important Radio Science specific function that the LMC performs is receipt and transmission of the system temperature and signal level data from the PPM for display at the LMC console and for inclusion in Monitor blocks. These blocks are recorded on magnetic tape as well as appearing in the Mission Control and Computing Center (MCCC) displays. The LMC is required to operate without interruption for the duration of the Radio Science data acquisition period. The Area Routing Assembly (ARA), which is part of the Digital Communications Subsystem, controls all data communication between the stations and JPL. The ARA receives all required data and status messages from the LMC/CMC and can record them to tape as well as transmit them to JPL via data lines. The ARA also receives predicts and other data from JPL and passes them on to the CMC. DSCC Antenna Mechanical Subsystem --------------------------------- Multi-mission Radio Science activities require support from the 70-m, 34-m HEF, and 34-m STD antenna subnets. The antennas at each DSCC function as large-aperture collectors which, by double reflection, cause the incoming radio frequency (RF) energy to enter the feed horns. The large collecting surface of the antenna focuses the incoming energy onto a subreflector, which is adjustable in both axial and angular position. These adjustments are made to correct for gravitational deformation of the antenna as it moves between zenith and the horizon; the deformation can be as large as 5 cm. The subreflector adjustments optimize the channeling of energy from the primary reflector to the subreflector and then to the feed horns. The 70-m and 34-m HEF antennas have 'shaped' primary and secondary reflectors, with forms that are modified paraboloids. This customization allows more uniform illumination of one reflector by another. The 34-m STD primary reflectors are classical paraboloids, while the subreflectors are standard hyperboloids. On the 70-m and 34-m STD antennas, the subreflector directs received energy from the antenna onto a dichroic plate, a device which reflects S-band energy to the S-band feed horn and passes X-band energy through to the X-band feed horn. In the 34-m HEF, there is one 'common aperture feed,' which accepts both frequencies without requiring a dichroic plate. RF energy to be transmitted into space by the horns is focused by the reflectors into narrow cylindrical beams, pointed with high precision (either to the dichroic plate or directly to the subreflector) by a series of drive motors and gear trains that can rotate the movable components and their support structures. The different antennas can be pointed by several means. Two pointing modes commonly used during tracking passes are CONSCAN and 'blind pointing.' With CONSCAN enabled and a closed loop receiver locked to a spacecraft signal, the system tracks the radio source by conically scanning around its position in the sky. Pointing angle adjustments are computed from signal strength information (feedback) supplied by the receiver. In this mode the Antenna Pointing Assembly (APA) generates a circular scan pattern which is sent to the Antenna Control System (ACS). The ACS adds the scan pattern to the corrected pointing angle predicts. Software in the receiver-exciter controller computes the received signal level and sends it to the APA. The correlation of scan position with the received signal level variations allows the APA to compute offset changes which are sent to the ACS. Thus, within the capability of the closed-loop control system, the scan center is pointed precisely at the apparent direction of the spacecraft signal source. An additional function of the APA is to provide antenna position angles and residuals, antenna control mode/status information, and predict-correction parameters to the Area Routing Assembly (ARA) via the LAN, which then sends this information to JPL via the Ground Communications Facility (GCF) for antenna status monitoring. During periods when excessive signal level dynamics or low received signal levels are expected (e.g., during an occultation experiment), CONSCAN should not be used. Under these conditions, blind pointing (CONSCAN OFF) is used, and pointing angle adjustments are based on a predetermined Systematic Error Correction (SEC) model. Independent of CONSCAN state, subreflector motion in at least the z-axis may introduce phase variations into the received Radio Science data. For that reason, during certain experiments, the subreflector in the 70-m and 34-m HEFs may be frozen in the z-axis at a position (often based on elevation angle) selected to minimize phase change and signal degradation. This can be done via Operator Control Inputs (OCIs) from the LMC to the Subreflector Controller (SRC) which resides in the alidade room of the antennas. The SRC passes the commands to motors that drive the subreflector to the desired position. Unlike the 70-m and 34-m HEFs which have azimuth-elevation (AZ-EL) drives, the 34-m STD antennas use (hour angle-declination) HA-DEC drives. The same positioning of the subreflector on the 34-m STD does not create the same effect as on the 70-m and 34-m HEFs. Pointing angles for all three antenna types are computed by the NOCC Support System (NSS) from an ephemeris provided by the flight project. These predicts are received and archived by the CMC. Before each track, they are transferred to the APA, which transforms the direction cosines of the predicts into AZ-EL coordinates for the 70-m and 34-m HEFs or into HA-DEC coordinates for the 34-m STD antennas. The LMC operator then downloads the antenna AZ-EL or HA-DEC predict points to the antenna-mounted ACS computer along with a selected SEC model. The pointing predicts consist of time-tagged AZ-EL or HA-DEC points at selected time intervals along with polynomial coefficients for interpolation between points. The ACS automatically interpolates the predict points, corrects the pointing predicts for refraction and subreflector position, and adds the proper systematic error correction and any manually entered antenna offsets. The ACS then sends angular position commands for each axis at the rate of one per second. In the 70-m and 34-m HEF, rate commands are generated from the position commands at the servo controller and are subsequently used to steer the antenna. In the 34-m STD antennas motors, rather than servos, are used to steer the antenna; there is no feedback once the 34-m STD has been told where to point. When not using binary predicts (the routine mode for spacecraft tracking), the antennas can be pointed using 'planetary mode' -- a simpler mode which uses right ascension (RA) and declination (DEC) values. These change very slowly with respect to the celestial frame. Values are provided to the station in text form for manual entry. The ACS quadratically interpolates among three RA and DEC points which are on one-day centers. A third pointing mode -- sidereal -- is available for tracking radio sources fixed with respect to the celestial frame. Regardless of the pointing mode being used, a 70-m antenna has a special high-accuracy pointing capability called 'precision' mode. A pointing control loop derives the main AZ-EL pointing servo drive error signals from a two- axis autocollimator mounted on the Intermediate Reference Structure. The autocollimator projects a light beam to a precision mirror mounted on the Master Equatorial drive system, a much smaller structure, independent of the main antenna, which is exactly positioned in HA and DEC with shaft encoders. The autocollimator detects elevation/cross- elevation errors between the two reference surfaces by measuring the angular displacement of the reflected light beam. This error is compensated for in the antenna servo by moving the antenna in the appropriate AZ-EL direction. Pointing accuracies of 0.004 degrees (15 arc seconds) are possible in 'precision' mode. The 'precision' mode is not available on 34-m antennas -- nor is it needed, since their beamwidths are twice as large as on the 70-m antennas. DSCC Antenna Microwave Subsystem -------------------------------- 70-m Antennas: Each 70-m antenna has three feed cones installed in a structure at the center of the main reflector. The feeds are positioned 120 degrees apart on a circle. Selection of the feed is made by rotation of the subreflector. A dichroic mirror assembly, half on the S-band cone and half on the X-band cone, permits simultaneous use of the S- and X-band frequencies. The third cone is devoted to R&D and more specialized work. The Antenna Microwave Subsystem (AMS) accepts the received S- and X-band signals at the feed horn and transmits them through polarizer plates to an orthomode transducer. The polarizer plates are adjusted so that the signals are directed to a pair of redundant amplifiers for each frequency, thus allowing simultaneous reception of signals in two orthogonal polarizations. For S-band these are two Block IVA S-band Traveling Wave Masers (TWMs); for X-band the amplifiers are Block IIA TWMs. 34-m STD Antennas: These antennas have two feed horns, one for S-band signals and one for X-band. The horns are mounted on a cone which is fixed in relation to the subreflector. A dichroic plate mounted above the horns directs energy from the subreflector into the proper horn. The AMS directs the received S- and X-band signals through polarizer plates and on to amplification. There are two Block III S-band TWMs and two Block I X-band TWMs. 34-m HEF Antennas: Unlike the other antennas, the 34-m HEF uses a single feed for both S- and X-band. Simultaneous S- and X-band receive as well as X-band transmit is possible thanks to the presence of an S/X 'combiner' which acts as a diplexer. For S-band, RCP or LCP is user selected through a switch so neither a polarizer nor an orthomode transducer is needed. X-band amplification options include two Block II TWMs or an HEMT Low Noise Amplifier (LNA). S-band amplification is provided by an FET LNA. DSCC Receiver-Exciter Subsystem ------------------------------- The Receiver-Exciter Subsystem is composed of three groups of equipment: the closed-loop receiver group, the open-loop receiver group, and the RF monitor group. This subsystem is controlled by the Receiver-Exciter Controller (REC) which communicates directly with the DMC for predicts and OCI reception and status reporting. The exciter generates the S-band signal (or X-band for the 34-m HEF only) which is provided to the Transmitter Subsystem for the spacecraft uplink signal. It is tunable under command of the Digitally Controlled Oscillator (DCO) which receives predicts from the Metric Data Assembly (MDA). The diplexer in the signal path between the transmitter and the feed horn for all three antennas (used for simultaneous transmission and reception) may be configured such that it is out of the received signal path (in listen-only or bypass mode) in order to improve the signal-to-noise ratio in the receiver system. Closed Loop Receivers: The Block IV receiver-exciter at the 70-m stations allows for two receiver channels, each capable of L-Band (e.g., 1668 MHz frequency or 18 cm wavelength), S-band, or X-band reception, and an S-band exciter for generation of uplink signals through the low-power or high-power transmitter. The Block III receiver-exciter at the 34-m STD stations allows for two receiver channels, each capable of S-band or X-band reception and an exciter used to generate an uplink signal through the low-power transmitter. The receiver-exciter at the 34-m HEF stations allows for one channel only. The closed-loop receivers provide the capability for rapid acquisition of a spacecraft signal and telemetry lockup. In order to accomplish acquisition within a short time, the receivers are predict driven to search for, acquire, and track the downlink automatically. Rapid acquisition precludes manual tuning though that remains as a backup capability. The subsystem utilizes FFT analyzers for rapid acquisition. The predicts are NSS generated, transmitted to the CMC which sends them to the Receiver-Exciter Subsystem where two sets can be stored. The receiver starts acquisition at uplink time plus one round-trip-light-time or at operator specified times. The receivers may also be operated from the LMC without a local operator attending them. The receivers send performance and status data, displays, and event messages to the LMC. Either the exciter synthesizer signal or the simulation (SIM) synthesizer signal is used as the reference for the Doppler extractor in the closed-loop receiver systems, depending on the spacecraft being tracked (and Project guidelines). The SIM synthesizer is not ramped; instead it uses one constant frequency, the Track Synthesizer Frequency (TSF), which is an average frequency for the entire pass. The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. It will be configured such that the expected amplitude changes are accommodated with minimum distortion. The loop bandwidth (2BLo) will be configured such that the expected phase changes can be accommodated while maintaining the best possible loop SNR. Open-Loop Receivers: The Radio Science Open-Loop Receiver (OLR) is a dedicated four channel, narrow-band receiver which provides amplified and downconverted video band signals to the DSCC Spectrum Processing Subsystem (DSP). The OLR utilizes a fixed first Local Oscillator (LO) frequency and a tunable second LO frequency to minimize phase noise and improve frequency stability. The OLR consists of an RF-to-IF downconverter located in the antenna, an IF selection switch (IVC), and a Radio Science IF-VF downconverter (RIV) located in the SPC. The RF-IF downconverters in the 70-m antennas are equipped for four IF channels: S-RCP, S-LCP, X-RCP, and X-LCP. The 34-m HEF stations are equipped with a two-channel RF-IF: S-band and X-band. The IVC switches the IF input between the 70-m and 34-m HEF antennas. The RIV contains the tunable second LO, a set of video bandpass filters, IF attenuators, and a controller (RIC). The LO tuning is done via DSP control of the POCA/PLO combination based on a predict set. The POCA is a Programmable Oscillator Control Assembly and the PLO is a Programmable Local Oscillator (commonly called the DANA synthesizer). The bandpass filters are selectable via the DSP. The RIC provides an interface between the DSP and the RIV. It is controlled from the LMC via the DSP. The RIC selects the filter and attenuator settings and provides monitor data to the DSP. The RIC could also be manually controlled from the front panel in case the electronic interface to the DSP is lost. RF Monitor -- SSI and PPM: The RF monitor group of the Receiver-Exciter Subsystem provides spectral measurements using the Spectral Signal Indicator (SSI) and measurements of the received channel system temperature and spacecraft signal level using the Precision Power Monitor (PPM). The SSI provides a local display of the received signal spectrum at a dedicated terminal at the DSCC and routes these same data to the DSP which routes them to NOCC for remote display at JPL for real-time monitoring and RIV/DSP configuration verification. These displays are used to validate Radio Science Subsystem data at the DSS, NOCC, and Mission Support Areas. The SSI configuration is controlled by the DSP and a duplicate of the SSI spectrum appears on the LMC via the DSP. During real-time operations the SSI data also serve as a quick-look science data type for Radio Science experiments. The PPM measures system noise temperatures (SNT) using a Noise Adding Radiometer (NAR) and downlink signal levels using the Signal Level Estimator (SLE). The PPM accepts its input from the closed-loop receiver. The SNT is measured by injecting known amounts of noise power into the signal path and comparing the total power with the noise injection 'on' against the total power with the noise injection 'off.' That operation is based on the fact that receiver noise power is directly proportional to temperature; thus measuring the relative increase in noise power due to the presence of a calibrated thermal noise source allows direct calculation of SNT. Signal level is measured by calculating an FFT to estimate the SNR between the signal level and the receiver noise floor where the power is known from the SNT measurements. There is one PPM controller at the SPC which is used to control all SNT measurements. The SNT integration time can be selected to represent the time required for a measurement of 30K to have a one-sigma uncertainty of 0.3K or 1%. DSCC Transmitter Subsystem -------------------------- The Transmitter Subsystem accepts the S-band frequency exciter signal from the Block III or Block IV Receiver- Exciter Subsystem exciter and amplifies it to the required transmit output level. The amplified signal is routed via the diplexer through the feed horn to the antenna and then focused and beamed to the spacecraft. The Transmitter Subsystem power capabilities range from 18 kw to 400 kw. Power levels above 18 kw are available only at 70-m stations. DSCC Tracking Subsystem ----------------------- The Tracking Subsystem primary functions are to acquire and maintain communications with the spacecraft and to generate and format radiometric data containing Doppler and range. The DSCC Tracking Subsystem (DTK) receives the carrier signals and ranging spectra from the Receiver-Exciter Subsystem. The Doppler cycle counts are counted, formatted, and transmitted to JPL in real time. Ranging data are also transmitted to JPL in real time. Also contained in these blocks is the AGC information from the Receiver-Exciter Subsystem. The Radio Metric Data Conditioning Team (RMDCT) at JPL produces an Archival Tracking Data File (ATDF) tape which contains Doppler and ranging data. In addition, the Tracking Subsystem receives from the CMC frequency predicts (used to compute frequency residuals and noise estimates), receiver tuning predicts (used to tune the closed-loop receivers), and uplink tuning predicts (used to tune the exciter). From the LMC, it receives configuration and control directives as well as configuration and status information on the transmitter, microwave, and frequency and timing subsystems. The Metric Data Assembly (MDA) controls all of the DTK functions supporting the uplink and downlink activities. The MDA receives uplink predicts and controls the uplink tuning by commanding the DCO. The MDA also controls the Sequential Ranging Assembly (SRA). It formats the Doppler and range measurements and provides them to the GCF for transmission to NOCC. The Sequential Ranging Assembly (SRA) measures the round trip light time (RTLT) of a radio signal traveling from a ground tracking station to a spacecraft and back. From the RTLT, phase, and Doppler data, the spacecraft range can be determined. A coded signal is modulated on an uplink carrier and transmitted to the spacecraft where it is detected and transponded back to the ground station. As a result, the signal received at the tracking station is delayed by its round trip through space and shifted in frequency by the Doppler effect due to the relative motion between the spacecraft and the tracking station on Earth. DSCC Spectrum Processing Subsystem (DSP) ---------------------------------------- The DSCC Spectrum Processing Subsystem (DSP) located at the SPC digitizes and records on magnetic tapes the narrowband output data from the RIV. It consists of a Narrow Band Occultation Converter (NBOC) containing four Analog-to- Digital Converters (ADCs), a ModComp CLASSIC computer processor called the Spectrum Processing Assembly (SPA), and two to six magnetic tape drives. Magnetic tapes are known as Original Data Records (ODRs). Electronic near real-time transmission of data to JPL (an Original Data Stream, or ODS) may be possible in certain circumstances; The DSP is operated through the LMC. Using the SPA-R software, the DSP allows for real-time frequency and time offsets (while in RUN mode) and, if necessary, snap tuning between the two frequency ranges transmitted by the spacecraft: coherent and non-coherent. The DSP receives Radio Science frequency predicts from the CMC, allows for multiple predict set archiving (up to 60 sets) at the SPA, and allows for manual predict generation and editing. It accepts configuration and control data from the LMC, provides display data to the LMC, and transmits the signal spectra from the SSI as well as status information to NOCC and the Project Mission Support Area (MSA) via the GCF data lines. The DSP records the digitized narrowband samples and the supporting header information (i.e., time tags, POCA frequencies, etc.) on 9-track magnetic tapes in 6250 or 1600 bpi GCR format. Through the DSP-RIC interface the DSP controls the RIV filter selection and attenuation levels. It also receives RIV performance monitoring via the RIC. In case of failure of the DSP-RIC interface, the RIV can be controlled manually from the front panel. All the RIV and DSP control parameters and configuration directives are stored in the SPA in a macro-like file called an 'experiment directive' table. A number of default directives exist in the DSP for the major Radio Science experiments. Operators can create their own table entries. Items such as verification of the configuration of the prime open-loop recording subsystem, the selection of the required predict sets, and proper system performance prior to the recording periods will be checked in real-time at JPL via the NOCC displays using primarily the remote SSI display at NOCC and the NRV displays. Because of this, transmission of the DSP/SSI monitor information is enabled prior to the start of recording. The specific run time and tape recording times will be identified in the Sequence of Events (SOE) and/or DSN Keyword File. The DSP can be used to duplicate ODRs. It also has the capability to play back a certain section of the recorded data after conclusion of the recording periods. DSCC Frequency and Timing Subsystem ----------------------------------- The Frequency and Timing Subsystem (FTS) provides all frequency and timing references required by the other DSCC subsystems. It contains four frequency standards of which one is prime and the other three are backups. Selection of the prime standard is done via the CMC. Of these four standards, two are hydrogen masers followed by clean-up loops (CUL) and two are cesium standards. These four standards all feed the Coherent Reference Generator (CRG) which provides the frequency references used by the rest of the complex. It also provides the frequency reference to the Master Clock Assembly (MCA) which in turn provides time to the Time Insertion and Distribution Assembly (TID) which provides UTC and SIM-time to the complex. JPL's ability to monitor the FTS at each DSCC is limited to the MDA calculated Doppler pseudo-residuals, the Doppler noise, the SSI, and to a system which uses the Global Positioning System (GPS). GPS receivers at each DSCC receive a one-pulse-per-second pulse from the station's (hydrogen maser referenced) FTS and a pulse from a GPS satellite at scheduled times. After compensating for the satellite signal delay, the timing offset is reported to JPL where a database is kept. The clock offsets stored in the JPL database are given in microseconds; each entry is a mean reading of measurements from several GPS satellites and a time tag associated with the mean reading. The clock offsets provided include those of SPC 10 relative to UTC (NIST), SPC 40 relative to SPC 10, etc. Optics - DSN ============ Performance of DSN ground stations depends primarily on size of the antenna and capabilities of electronics. These are summarized in the following set of tables. Note that 64-m antennas were upgraded to 70-m between 1986 and 1989. Beamwidth is half-power full angular width. Polarization is circular; L denotes left circular polarization (LCP), and R denotes right circular polarization (RCP). DSS S-Band Characteristics 64-m 70-m 34-m 34-m Transmit STD HEF -------- ----- ----- ----- ----- Frequency (MHz) 2110- 2110- 2025- N/A 2120 2120 2120 Wavelength (m) 0.142 0.142 0.142 N/A Ant Gain (dBi) 62.7 55.2 N/A Beamwidth (deg) 0.119 0.31 N/A Polarization L or R L or R N/A Tx Power (kW) 20-400 20 N/A Receive ------- Frequency (MHz) 2270- 2270- 2270- 2200- 2300 2300 2300 2300 Wavelength (m) 0.131 0.131 0.131 0.131 Ant Gain (dBi) 61.6 63.3 56.2 56.0 Beamwidth (deg) 0.108 0.27 0.24 Polarization L & R L & R L or R L or R System Temp (K) 22 20 22 38 DSS X-Band Characteristics (N/A for Galileo) 64-m 70-m 34-m 34-m Transmit STD HEF -------- ----- ----- ----- ----- Frequency (MHz) 8495 8495 N/A 7145- 7190 Wavelength (m) 0.035 0.035 N/A 0.042 Ant Gain (dBi) 74.2 N/A 67 Beamwidth (deg) N/A 0.074 Polarization L or R L or R N/A L or R Tx Power (kW) 360 360 N/A 20 Receive ------- Frequency (MHz) 8400- 8400- 8400- 8400- 8500 8500 8500 8500 Wavelength (m) 0.036 0.036 0.036 0.036 Ant Gain (dBi) 71.7 74.2 66.2 68.3 Beamwidth (deg) 0.031 0.075 0.063 Polarization L & R L & R L & R L & R System Temp (K) 27 20 25 20 Electronics - DSN ================= DSCC Open-Loop Receiver ----------------------- The open loop receiver block diagram shown below is for 70-m and 34-m High-Efficiency (HEF) antenna sites. Based on a tuning prediction file, the POCA controls the DANA synthesizer the output of which (after multiplication) mixes input signals at both S- and X-band to fixed intermediate frequencies for amplification. These signals in turn are down converted and passed through additional filters until they yield baseband output of up to 25 kHz in width. The baseband output is digitally sampled by the DSP and either written to magnetic tape or electronically transferred for further analysis. S-Band X-Band 2295 MHz 8415 MHz Input Input | | v v --- --- --- --- | X |<--|x20|<--100 MHz 100 MHz-->|x81|-->| X | --- --- --- --- | | 295| |315 MHz| |MHz v v --- -- 33.1818 --- --- | X |<--|x3|<------ MHz ------>|x11|-->| X | --- -- |115 | --- --- | |MHz | | | | | | 50| 71.8181 --- --- |50 MHz| MHz->| X | | X |<-10 MHz |MHz v --- --- v --- ^ ^ --- | X |<--60 MHz | | 60 MHz-->| X | --- | | --- | 9.9 | 43.1818 MHz | 9.9 | | MHz ------------- MHz | | | ^ | | 10| v | v |10 MHz| --- ---------- --- |MHz |------>| X | | DANA | | X |<------| | --- |Synthesizr| --- | | | ---------- | | v v ^ v v ------- ------- | ------- ------- |Filters| |Filters| ---------- |Filters| |Filters| |3,4,5,6| | 1,2 | | POCA | | 1,2 | |3,4,5,6| ------- ------- |Controller| ------- ------- | | ---------- | | 10| |0.1 0.1| |10 MHz| |MHz MHz| |MHz v v v v --- --- --- --- | X |- -| X | | X |- -| X | --- | | --- --- | | --- ^ | | ^ ^ | | ^ | | | | | | | | 10 | | 0.1 0.1 | | 10 MHz | | MHz MHz | | MHz | | | | v v v v Baseband Baseband Output Output Reconstruction of the antenna frequency from the frequency of the signal in the recorded data can be achieved through use of one of the following formulas. Radio Science IF-VF (RIV) Converter Assembly at 70-m and 34-m High-Efficiency (HEF) antennas: FSant=3*[POCA+(790/11)*10^6] + 1.95*10^9 - Fsamp - Frec FXant=11*[POCA-10^7] + 8.050*10^9 - 3*Fsamp + Frec Multi-Mission Receivers at 34-m Standard antennas (DSS 42 and 61; the diagram above does not apply): FSant=48*POCA + 3*10^8 - 0.75*Fsamp + Frec FXant = (11/3)*[48*POCA + 3*10^8 - 0.75*Fsamp] + Frec where FSant = S-band antenna frequency FXant = X-band antenna frequency POCA = POCA frequency Fsamp = sampling frequency Frec = frequency of recorded signal Filters - DSN ============= DSCC Open-Loop Receiver ----------------------- Nominal filter center frequencies and bandwidths for the Open-Loop Receivers are shown in the table below. Filter Center Frequency 3 dB Bandwidth ------ ---------------- -------------- 1 0.1 MHz 90 Hz 2 0.1 MHz 450 Hz 3 10.0 MHz 2000 Hz 4 10.0 MHz 1700 Hz (S-band) 6250 Hz (X-band) 5 10.0 MHz 45000 Hz 6 10.0 MHz 21000 Hz MMR filters (DSS 42 and 61) include the following: Filter Center Frequency 3 dB Bandwidth ------ ---------------- -------------- 5 Unknown 2045 Hz (S-band) 7500 Hz (X-band) Detectors - DSN =============== DSCC Open-Loop Receivers ------------------------ Open-loop receiver output is detected in software by the radio science investigator. DSCC Closed-Loop Receivers -------------------------- Nominal carrier tracking loop threshold noise bandwidth at both S- and X-band is 10 Hz. Coherent (two-way) closed-loop system stability is shown in the table below: integration time Doppler uncertainty (secs) (one sigma, microns/sec) ------ ------------------------ 10 50 60 20 1000 4 Calibration - DSN ================= Calibrations of hardware systems are carried out periodically by DSN personnel; these ensure that systems operate at required performance levels -- for example, that antenna patterns, receiver gain, propagation delays, and Doppler uncertainties meet specifications. No information on specific calibration activities is available. Nominal performance specifications are shown in the tables above. Additional information may be available in [DSN810-5]. Prior to each tracking pass, station operators perform a series of calibrations to ensure that systems meet specifications for that operational period. Included in these calibrations is measurement of receiver system temperature in the configuration to be employed during the pass. Results of these calibrations are recorded in (hard copy) Controller's Logs for each pass. The nominal procedure for initializing open-loop receiver attenuator settings is described below. In cases where widely varying signal levels are expected, the procedure may be modified in advance or real-time adjustments may be made to attenuator settings. Open-Loop Receiver Attenuation Calibration ------------------------------------------ The open-loop receiver attenuator calibrations are performed to establish the output of the open-loop receivers at a level that will not saturate the analog-to-digital converters. To achieve this, the calibration is done using a test signal generated by the exciter/translator that is set to the peak predicted signal level for the upcoming pass. Then the output level of the receiver's video band spectrum envelope is adjusted to the level determined by equation (3) below (to five-sigma). Note that the SNR in the equation (2) is in dB while the SNR in equation (3) is linear. Pn = -198.6 + 10*log(SNT) + 10*log(1.2*Fbw) (1) SNR = Ps - Pn (SNR in dB) (2) Vrms = sqrt(SNR + 1)/[1 + 0.283*sqrt(SNR)] (SNR linear)(3) where Fbw = receiver filter bandwidth (Hz) Pn = receiver noise power (dBm) Ps = signal power (dBm) SNT = system noise temperature (K) SNR = predicted signal-to-noise ratio Operational Considerations - DSN ================================ The DSN is a complex and dynamic 'instrument.' Its performance for Radio Science depends on a number of factors from equipment configuration to meteorological conditions. No specific information on 'operational considerations' can be given here. Operational Modes - DSN ======================= DSCC Antenna Mechanical Subsystem --------------------------------- Pointing of DSCC antennas may be carried out in several ways. For details see the subsection 'DSCC Antenna Mechanical Subsystem' in the 'Subsystem' section. Binary pointing is the preferred mode for tracking spacecraft; pointing predicts are provided, and the antenna simply follows those. With CONSCAN, the antenna scans conically about the optimum pointing direction, using closed-loop receiver signal strength estimates as feedback. In planetary mode, the system interpolates from three (slowly changing) RA-DEC target coordinates; this is 'blind' pointing since there is no feedback from a detected signal. In sidereal mode, the antenna tracks a fixed point on the celestial sphere. In 'precision' mode, the antenna pointing is adjusted using an optical feedback system. It is possible on most antennas to freeze z-axis motion of the subreflector to minimize phase changes in the received signal. DSCC Receiver-Exciter Subsystem ------------------------------- The diplexer in the signal path between the transmitter and the feed horns on all three antennas may be configured so that it is out of the received signal path in order to improve the signal-to-noise ratio in the receiver system. This is known as the 'listen-only' or 'bypass' mode. Closed-Loop vs. Open-Loop Reception ----------------------------------- Radio Science data can be collected in two modes: closed- loop, in which a phase-locked loop receiver tracks the spacecraft signal, or open-loop, in which a receiver samples and records a band within which the desired signal presumably resides. Closed-loop data are collected using Closed-Loop Receivers, and open-loop data are collected using Open-Loop Receivers in conjunction with the DSCC Spectrum Processing Subsystem (DSP). See the Subsystems section for further information. Closed-Loop Receiver AGC Loop ----------------------------- The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. Ordinarily it is configured so that expected signal amplitude changes are accommodated with minimum distortion. The loop bandwidth is ordinarily configured so that expected phase changes can be accommodated while maintaining the best possible loop SNR. Coherent vs. Non-Coherent Operation ----------------------------------- The frequency of the signal transmitted from the spacecraft can generally be controlled in two ways -- by locking to a signal received from a ground station or by locking to an on-board oscillator. These are known as the coherent (or 'two-way') and non-coherent ('one-way') modes, respectively. Mode selection is made at the spacecraft, based on commands received from the ground. When operating in the coherent mode, the transponder carrier frequency is derived from the received uplink carrier frequency with a 'turn-around ratio' typically of 240/221. In the non-coherent mode, the downlink carrier frequency is derived from the spacecraft on-board crystal-controlled oscillator. Either closed-loop or open-loop receivers (or both) can be used with either spacecraft frequency reference mode. Closed-loop reception in two-way mode is usually preferred for routine tracking. Occasionally the spacecraft operates coherently while two ground stations receive the 'downlink' signal; this is sometimes known as the 'three-way' mode. DSCC Spectrum Processing Subsystem (DSP) ---------------------------------------- The DSP can operate in four sampling modes with from 1 to 4 input signals. Input channels are assigned to ADC inputs during DSP configuration. Modes and sampling rates are summarized in the tables below: Mode Analog-to-Digital Operation ---- ---------------------------- 1 4 signals, each sampled by a single ADC 2 1 signal, sampled sequentially by 4 ADCs 3 2 signals, each sampled sequentially by 2 ADCs 4 2 signals, the first sampled by ADC #1 and the second sampled sequentially at 3 times the rate by ADCs #2-4 8-bit Samples 12-bit Samples Sampling Rates Sampling Rates (samples/sec per ADC) (samples/sec per ADC) --------------------- --------------------- 50000 31250 25000 15625 12500 10000 10000 6250 5000 5000 4000 3125 2500 2000 1250 1000 1000 500 400 250 200 200 Input to each ADC is identified in header records by a Signal Channel Number (J1 - J4). Nominal channel assignments are shown below. Signal Channel Number Receiver (70-m or HEF) (34-m STD) --------------------- ------------- ---------- J1 X-RCP not used J2 S-RCP not used J3 X-LCP X-RCP J4 S-LCP S-RCP Location - DSN ============== Station locations are documented in [GEO-10REVD]. Geocentric coordinates are summarized here. Geocentric Geocentric Geocentric Station Radius (km) Latitude (N) Longitude (E) --------- ----------- ------------ ------------- Goldstone DSS 12 (34-m STD) 6371.997815 35.1186672 243.1945048 DSS 13 (develop) 6372.117062 35.0665485 243.2051077 DSS 14 (70-m) 6371.992867 35.2443514 243.1104584 DSS 15 (34-m HEF) 6371.9463 35.2402863 243.1128186 DSS 16 (26-m) 6371.9608 35.1601436 243.1264200 DSS 18 (34-m STD) UNK UNK UNK Canberra DSS 42 (34-m STD) 6371.675607 -35.2191850 148.9812546 DSS 43 (70-m) 6371.688953 -35.2209308 148.9812540 DSS 45 (34-m HEF) 6371.692 -35.21709 148.97757 DSS 46 (26-m) 6371.675 -35.22360 148.98297 DSS 48 (34-m STD) UNK UNK UNK Madrid DSS 61 (34-m STD) 6370.027734 40.2388805 355.7509634 DSS 63 (70-m) 6370.051015 40.2413495 355.7519776 DSS 65 (34-m HEF) 6370.021370 40.2372843 355.7485968 DSS 66 (26-m) 6370.036 40.2400714 355.7485976 DSS 48 (34-m STD) UNK UNK UNK Measurement Parameters - DSN ============================ Open-Loop System ---------------- Output from the Open-Loop Receivers (OLRs), as sampled and recorded by the DSCC Spectrum Processing Subsystem (DSP), is a stream of 8- or 12-bit quantized voltage samples. The nominal input to the Analog-to-Digital Converters (ADCs) is +/-10 volts, but the precise scaling between input voltages and output digitized samples is usually irrelevant for analysis; the digital data are generally referenced to a known noise or signal level within the data stream itself -- for example, the thermal noise output of the radio receivers which has a known system noise temperature (SNT). Raw samples comprise the data block in each DSP record; a header record (presently 83 16-bit words) contains ancillary information such as: time tag for the first sample in the data block RMS values of receiver signal levels and ADC outputs POCA frequency and drift rate Closed-Loop System ------------------ Closed-loop data are recorded in Archival Tracking Data Files (ATDFs), as well as certain secondary products such as the Orbit Data File (ODF). The ATDF Tracking Logical Record contains 117 entries including status information and measurements of ranging, Doppler, and signal strength. ACRONYMS AND ABBREVIATIONS - DSN ================================ ACS Antenna Control System ADC Analog-to-Digital Converter AGC Automatic Gain Control AMS Antenna Microwave System APA Antenna Pointing Assembly ARA Area Routing Assembly ATDF Archival Tracking Data File AZ Azimuth CMC Complex Monitor and Control CONSCAN Conical Scanning (antenna pointing mode) CRG Coherent Reference Generator CUL Clean-up Loop DANA a type of frequency synthesizer dB deciBel dBi dB relative to isotropic DCO Digitally Controlled Oscillator DEC Declination deg degree DFLR Deutsche Forschungsanstalt fur Luft- und Raumfahrt DMC DSCC Monitor and Control Subsystem DSCC Deep Space Communications Complex DSN Deep Space Network DSP DSCC Spectrum Processing Subsystem DSS Deep Space Station DTK DSCC Tracking Subsystem E east EL Elevation FET Field Effect Transistor FFT Fast Fourier Transform FTS Frequency and Timing Subsystem GCF Ground Communications Facility GCR Group Coded Recording GHz gigahertz GPS Global Positioning System GSFC Goddard Space Flight Center HA Hour Angle HEF High-Efficiency (as in 34-m HEF antennas) HEMT HGA High Gain Antenna IF Intermediate Frequency IVC IF Selection Switch JPL Jet Propulsion Laboratory K Kelvin km kilometer kW kilowatt L-band approximately 1668 MHz LAN Local Area Network LCP Left-Circularly Polarized LGA Low Gain Antenna LMC Link Monitor and Control LNA Low-Noise Amplifier LO Local Oscillator m meters MCA Master Clock Assembly MCCC Mission Control and Computing Center MDA Metric Data Assembly MHz Megahertz MMR Multi-Mission Radio (Science) MON Monitor and Control System MSA Mission Support Area N north NAR Noise Adding Radiometer NASA National Aeronautics and Space Administration NASCOM NASA Communications NBOC Narrow-Band Occultation Converter NIST SPC 10 time relative to UTC NIU Network Interface Unit NOCC Network Operations and Control System NRV NOCC Radio Science/VLBI Display Subsystem NSS NOCC Support System OCI Operator Control Input ODF Orbit Data File ODR Original Data Record ODS Original Data Stream OLR Open Loop Receiver PLO Programmable Local Oscillator POCA Programmable Oscillator Control Assembly PPM Precision Power Monitor RA Right Ascension REC Receiver-Exciter Controller RCP Right-Circularly Polarized RF Radio Frequency RIC RIV Controller RIV Radio Science IF-VF Converter Assembly RMDCT Radio Metric Data Conditioning Team RTLT Round-Trip Light Time S-band approximately 2100-2300 MHz sec second SEC System Error Correction SIM Simulation SLE Signal Level Estimator SNR Signal-to-Noise Ratio SNT System Noise Temperature SOE Sequence of Events SPA Spectrum Processing Assembly SPC Signal Processing Center SRA Sequential Ranging Assembly SRC Sub-Reflector Controller SSI Spectral Signal Indicator STD Standard (as in 34-m STD antennas) TID Time Insertion and Distribution Assembly TSF Tracking Synthesizer Frequency TWM Traveling Wave Maser Tx Transmitter UNK unknown UTC Universal Coordinated Time VF Video Frequency X-band approximately 7800-8500 MHz
  • instrument : MOON MINERALOGY MAPPER for CH1-ORB
    The following text summarizes the NASA public website for the Moon Mineralogy Mapper, http://m3.jpl.nasa.gov/. Instrument Overview =================== The Moon Mineralogy Mapper (M3) was a near-infrared imaging spectrometer of the pushbroom type on-board the Indian Space Research Organization's (ISRO) Chandrayaan-1 spacecraft that launched on 22 October 2008. M3 measured the spectral range from 430 to 3000 nm, had a 40-km-wide field of view from a 100-km orbit, and operated in a high resolution (70 m/pixel) target mode or a lower resolution (140 m/pixel) global mode. Originally scheduled to make target and global observations from a 100-km circular polar obit for two years, the first M3 imaging data were acquired on 18 and 19 November 2008 using both the global and target imaging modes. Several days later on 22 November, M3 acquired its first scientific images of Oriental Basin. Data acquisition from the 100-km primary science orbit continued into May 2009. On 19 May, the Chandrayaan-1 spacecraft was raised to a 200-km to compensate for star tracker failures and orbiter temperatures that were higher than expected. On 28 August 2009 ISRO abruptly lost radio contact with the Chandrayaan-1 spacecraft and officially terminated the mission three days later. After ten months in operation, M3 had obtained over 4.6 billion spectra of the lunar surface, and its low resolution spatial coverage encompassed nearly the entire Moon. The instrument was funded by NASA's Discovery Mission of Opportunity program and was designed, built, and tested by a partnership between Jet Propulsion Laboratory and Brown University, USA. Preliminary results are presented by Pieters, et al. (2009b) [PIETERSETAL2009B] and Pieters, et al. (2009c) [PIETERSETAL2009C]. Instrument Characteristics ========================== M3 was a near-infrared imaging spectrometer with a compact system of optics in the Offner design which produces little or not optical distortion, either spatially or spectrally. The instrument included a robust on-board calibrator and was cooled by a multi-stage passive cooler. M3 was designed to operate in two distinct modes. The target mode was used for full resolution imaging of specific science targets, and the global mode was used for low resolution global imaging. M3 used a pushbroom method in which the field of view was passively swept through the lunar scene below it, along a line perpendicular to the slit, simultaneously exposing all 600 spatial pixels in an entire row using 260 spectral channels (in target mode). The characteristics of the instrument are summarized here: Telescope : f/2.7, all-aluminum optics Grating : Single dual-blaze electron-beam grating Detector : Single two-dimensional HgCdTe array Mass : 8 kg Volume : 25 x 18 x 21 cm Power : 15 W Spectral Range : 430 - 3000 nm Field of View : 40 km (both modes) from a 100-km orbit Imaging Modes : Target : 600 pixel crosstrack 70 m/pix spatial at 100-km orbit 260 bands (10-nm width) ~12-deg latitudinal swaths : Global : 300 pixel crosstrack 140 m/pix spatial at 100-km orbit 85 bands (20- & 40-nm width selected) ~145-deg latitudinal swaths Spectral Response : FWHM < 12.5 nm (target mode in lab) Spatial Response : FWHM < 1.0 mrad (target mode in lab) Spatial Sampling : 0.7 mrad (target mode in lab) Signal-to-Noise Ratio : >400 at the equator (0-deg solar zenith) >100 at polar regions (80-deg solar zenith) For more information about the M3 instrument, see Green, et al. (2008a) [GREENETAL2008A], Green, et al. (2008b) [GREENETAL2008B], Green, et al. (2009) [GREENETAL2009], and Petro, et al. (2008) [PETROETAL2008], and Pieters, et al. (2009a) [PIETERSETAL2009A]. Scientific Objectives ===================== The specific scientific objectives of the M3 instrument were to: - Evaluate the primary components of the lunar crust and their distribution across the highlands, - Characterize the diversity and extent of the different types of basaltic volcanism, - Identify and access deposits containing volatiles including water, - Map fresh craters to access the properties of impacts in the recent past, and - Identify and evaluate concentrations of unusual and unexpected minerals.
  • investigation : NEAR EARTH ASTEROID RENDEZVOUS
    Mission Overview ================ The Near Earth Asteroid Rendezvous (NEAR) mission inaugurated NASA's Discovery Program. It was the first mission to orbit an asteroid and made the first comprehensive scientific measurements of an asteroid's surface composition, geology, physical properties, and internal structure. NEAR was launched successfully on 17 February 1996 aboard a Delta II-7925. It made the first reconnaissance of a C-type asteroid during its flyby of the main-belt asteroid 253 Mathilde in June 1997. It became the first spacecraft to enter orbit around an asteroid, doing so at the large near-Earth asteroid 433 Eros in February 2000. The spacecraft, renamed NEAR Shoemaker, landed on Eros at 37.2 South by 278.4 West, ending its mission on February 12, 2001 with another spacecraft first. NEAR obtained new information on the nature and evolution of asteroids, improved our understanding of planetary formation processes in the early solar system, and clarified the relationships between asteroids and meteorites. The NEAR Mission Operations Center and Science Data Center were both located at APL. The latter maintained the entire NEAR data set on-line and made data from all instruments accessible over the Internet to every member of the NEAR science team. For a detailed description of the mission see [CHENGETAL1998]. Introduction ============ Of the more than 7000 asteroids that have been named, most are found in the main asteroid belt between the orbits of Mars and Jupiter, but those that come within 1.3 AU of the Sun are known as near-Earth asteroids. The orbits of these dynamically young bodies have evolved on 100-million-year timescales because of collisions and gravitational interactions with planets. The present-day orbits of such asteroids do not necessarily indicate where they formed. Some are already in Earth-crossing orbits, and those that are not are highly likely to evolve into one. More than 250 near-Earth asteroids are known, and they appear to typify a broad sample of the main-belt population. Before NEAR, knowledge of the nature of asteroids came from three sources: Earth-based remote sensing, data from the Galileo spacecraft flybys of the two main-belt asteroids 951 Gaspra and 243 Ida, and laboratory analyses of meteorites. Most meteorites are believed to be collisional fragments of asteroids, but they may represent a biased and incomplete sampling of the materials actually found in near-Earth asteroids. Firm links between meteorite types and asteroid types have been difficult to establish [GAFFEYETAL1993A]. The uncommon eucrite (a basaltic achondrite) meteorites have been linked by visible and near-infrared reflectance measurements to the relatively rare V-type asteroids [MCCORDETAL1970], [BINZEL&XU1993]. However, a major controversy has been whether and how the most common meteorite types (the ordinary chondrites) may be linked to the most common asteroid types (the S-type or stony asteroids) in the inner part of the asteroid belt [BELLETAL1989], [GAFFEYETAL1993B]. Galileo and NEAR targets 951 Gaspra, 243 Ida, and the 433 Eros are all S-type asteroids.) The S-type asteroids are a diverse class of objects known to contain the silicate minerals olivine and pyroxene plus an admixture of iron/nickel metal. Some appear to be fragments of bodies that underwent substantial melting and differentiation. Others may consist of primitive materials like ordinary chondrites that never underwent melting and that may preserve characteristics of the solid material from which the inner planets accreted. The Galileo flybys provided the very first high-resolution images of S asteroids, revealing complex surfaces covered by craters, fractures, grooves, and subtle color variations [BELTONETAL1992], [OSTROETAL1990]. Galileo also discovered a satellite at Ida, which is a member of the Koronis family (Eros is not an asteroid family member). The near-infrared spectrum of Gaspra indicates a high olivine abundance such that it is inferred to be a fragment of a differentiated body. Conversely, Ida and Eros display infrared spectra that may be consistent with a silicate mineralogy like that in ordinary chondrites [CHAPMAN1996], [MURCHIE&PIETERS1996]. The Galileo instrument complement did not include any capability to measure elemental composition, and debate continues about whether ordinary chondrites are related to S-type asteroids. The NEAR mission spent about a year in orbit around Eros, entering 14 February 2000 and landing at the asteroid surface 12 February 2001 from when spacecraft operations were continued until 28 February 2001. It acquired the first comprehensive, spatially resolved measurements of the geomorphology, reflectance spectral properties, and shape of an asteroid, and X-ray and gamma-ray spectral measurements of elemental abundances from orbit and the surface. The ambient magnetic field in the vicinity of the asteroid was also measured. NEAR orbited Eros at low altitude, as close as about 1 body radius above the surface, for several months so as to allow NEAR's instruments to to acquire their highest spatial resolution measurements. The NEAR data, especially when combined with those from the Galileo flybys, greatly advanced our understanding of S-type asteroids and their possible relationships to meteorites and other small bodies of the solar system. NEAR also conducted a thorough search for satellites. Spacecraft Design ================= NEAR was a solar-powered, three-axis-stabilized spacecraft [SANTOETAL1995] with a launch mass, including propellant, of 805 kg and a dry mass of 468 kg. The spacecraft was simple and highly redundant. It used X-band telemetry to the NASA deep space network; data rates at Eros were selectable in the range of 2.9 to 8.8 kbps using a 34-m high-efficiency antenna. With a 70-m antenna, the data rates from Eros ranged from 17.6 to 26.5 kbps. The command and telemetry systems were fully redundant. Two solid-state recorders were accommodated with a combined memory capacity of 1.6 Gbit. Spacecraft attitude was determined using a star camera, a fully redundant inertial measurement unit, and redundant digital Sun sensors. The propulsion sub-system was dual mode (hydrazine was used as fuel for both the monopropellant and bipropellant systems) and included one 450-N bipropellant thruster for large maneuvers, four 21-N thrusters, and seven 3.5-N thrusters for fine velocity control and momentum dumping. Attitude was controlled by a redundant set of four reaction wheels or by the thruster complement to within 1.7 mrad. NEAR's line-of-sight pointing stability was within 20 microrad 1 s, and postprocessing attitude knowledge was within 130 microrad. Forward and aft aluminum honeycomb decks were connected with eight aluminum honeycomb side panels. Mounted on the outside of the forward deck were a fixed, 1.5-m-dia. X-band high-gain antenna (HGA), four fixed solar panels, and the X-ray solar monitor system. When the solar panels were fully illuminated, the Sun was in the center of the solar monitor field of view (FOV). No booms were accommodated on the spacecraft. The electronics were mounted on the inside of the forward and aft decks. NEAR contained six scientific instruments, which are detailed in the next section. 1. Multispectral Imager (MSI) 2. Near-Infrared Spectrograph (NIS) 3. X-Ray Spectrometer (XRS) 4. Gamma-Ray Spectrometer (GRS) 5. NEAR Laser Rangefinder (NLR) 6. Magnetometer (MAG) The MAG was mounted on top of the HGA feed, where it was exposed to the minimum level of spacecraft-generated magnetic fields. The remaining instruments (MSI, NIS, XRS, GRS, and NLR) were all mounted on the outside of the aft deck. They were on fixed mounts and were co-aligned to view a common boresight direction. The NIS had a scan mirror that allowed it to look 30 degrees forward and 110 degrees aft from the common boresight. Key properties of the mission design permitted the use of this fixed spacecraft geometry. Throughout most of the orbital rendezvous with Eros, the angle between the Sun and the Earth, as seen from the spacecraft, remained less than about 30 degrees. In addition, the mission aphelion was reached during cruise. Hence, if the solar panels were sufficiently large to sustain NEAR at aphelion, there was sufficient power margin at Eros for the spacecraft to pull its solar panels over 30 deg off full illumination to point the HGA at Earth. Moreover, the rendezvous orbit plane was maintained so that the orbit normal pointed approximately at the Sun. In this case, as NEAR orbited Eros, it was usually able to roll around the HGA axis so as to keep the instruments pointed at the asteroid while maintaining adequate solar panel illumination. The instruments were usually pointed away from the asteroid when the HGA was used to downlink to Earth. This mode of operation motivated the requirement for on-board data storage. With on-board image compression, NEAR could store more than 1000 images and downlink them within 10 hrs at its maximum data rate of 26.5 kbps. The spacecraft was designed using a distributed architecture, partitioned so that subsystems generally did not share common hardware or software. One major benefit of this approach was that careful design of interfaces allowed development, test, and integration of sub-systems in parallel. In addition, this architecture had a natural advantage of built-in contingencies and design margins. Truly parallel subsystem development required independence at the subsystem interface, through careful partitioning of functional requirements and ample design margins at subsystem inter-faces. On NEAR, subsystems were interfaced through a MIL-STD-1553 data bus, chosen because it was compatible with many off-the-shelf industry components. The data bus had additional attractive features: fewer interconnecting cables; built-in redundancy and cross-strapping; simplification of interface definition; a fault-tolerant, transformer-coupled interface; a common data architecture for sharing information among subsystems; and a flexible software-defined interface instead of a rigid hardware-defined interface. When it was launched, NEAR was the lowest-cost U.S. planetary mission ever. The spacecraft's 27-month development schedule was unusually rapid. The distributed architecture and the selection of the 1553 data bus were key to developing NEAR on time and under budget. Previous planetary missions have not used a distributed architecture because they have been optimized for performance, i.e., to return maximum science within available technology. The distributed architecture approach comes with a mass penalty, and therefore a performance penalty: some hardware that can be combined at the system level is duplicated at the subsystem level. The distributed architecture approach for NEAR features interface margin and testability, optimizing the spacecraft for low cost and rapid schedule. Nevertheless, the performance penalty is minuscule, and the mass penalty for using the distributed architecture approach is only about 10 kg. Instrument Tasks ================ Details on the many science objectives of the NEAR instruments can be found elsewhere [VEVERKAETAL1997A], [TROMBKAETAL1997], [ACUNAETAL1997], [ZUBERETAL1997], [YEOMANSETAL1997] and [CHENGETAL1997]. A brief summary of instrument characteristics is given in this section. (Full descriptions of each science investigation and instrument appeared in a special issue of Space Science Reviews, vol. 82, 1997.) Detailed instrument descriptions and results of ground and in-flight calibrations appear in the companion articles of this issue of the Technical Digest. Multispectral Imager -------------------- The main goals of the MSI were to determine the shape of Eros and to map the mineralogy and morphology of features on its surface at high spatial resolution. MSI was a 537 x 244 pixel charge-coupled device camera with five-element radiation-hardened refractive optics. It covered the spectral range from 0.4 to 1.1 microns, and it had an eight-position filter wheel. Seven narrow-band filters were chosen to discriminate the major iron-bearing silicates present (olivine and pyroxene); the eight, broad-band filter was for fast exposures and high sensitivity, including optical navigation. occur on Eros. The camera had an FOV of 2.93 x 2.26 degrees and a pixel resolution of 96 x 162 microrad. It had a maximum framing rate of 1 per second with images digitized to 12 bits and a dedicated digital processing unit with an image buffer in addition to both lossless and lossy on-board image compression. Near-Infrared Spectrograph -------------------------- NIS measured the spectrum of sunlight reflected from Eros in the near-infrared range from 0.8 to 2.5 microns to determine the distribution and abundance of surface minerals like olivine and pyroxene. This grating spectrometer dispersed the light from the slit FOV (0.38 x 0.76 degrees in its narrow position and 0.76 x 0.76 degrees in the wide position) across a pair of passively cooled one-dimensional array detectors. A 32-channel germanium array covering the lower wavelengths, with channel centers at 0.82 to 1.49 microns with a 0.022 micron spacing between channels. A 32-channel indium/gallium- arsenide array covering longer wavelengths, with channel centers at 1.37 to 2.71 microns with a 0.043 micron spacing between channels. Due to configuration of the optics and the sensitivity of this array, useful measurements were acquired by it over the wavelength range 1.5 to 2.5 microns. The slit could be closed for dark current measurements, which were routinely interleaved with measurements of the asteroid. NIS had a scan mirror that enabled it to step across the range from 30 degrees forward of the common boresight to 110 degrees aft, in 0.4 degree steps. Spectral images were built up by a combination of scan mirror and spacecraft motions. In addition, the NIS had a gold calibration target that viewed at the forward limit of the mirror's scan ranges. It scattered sunlight into the instrument and provided a quantitative, in-flight calibration of instrument stability. X-Ray Spectrometer ------------------ The XRS was an X-ray resonance fluorescence spectrometer that detected the characteristic X-ray line emissions excited by solar X-rays from major elements in the asteroid's surface. It covered X-rays in the energy range from 1 to 10 keV using three gas proportional counters. The balanced, differential filter technique was used to separate the closely spaced Mg, Al, and Si lines lying below 2 keV. The gas proportional counters directly resolved higher energy line emissions from Ca and Fe. A mechanical collimator gave the XRS a 5 degree FOV, with which it mapped the chemical composition of the asteroid at spatial resolutions as fine as 2 km in the low orbits. It also included a separate solar monitor system to measure continuously the incident spectrum of solar X-rays, using both a gas proportional counter and a high-spectral-resolution silicon X-ray detector. The XRS performed in-flight calibration using a calibration rod with Fe-55 sources that could be rotated into or out of the detector FOV. Gamma-Ray Spectrometer ---------------------- The GRS detected characteristic gamma rays in the 0.3- to 10-MeV range emitted from specific elements in the asteroid surface. Some of these emissions were excited by cosmic rays and some arose from natural radioactivity in the asteroid. The GRS used a body-mounted, passively cooled NaI scintillator detector with a bismuth germanate anticoincidence shield that defined a 45 degree FOV. Abundances of several important elements such as K, Si, and Fe were measured. NEAR Laser Rangefinder ---------------------- The NLR was a laser altimeter that measured the distance from the spacecraft to the asteroid surface by sending out a short burst of laser light and then recording the time required for the signal to return from the asteroid. It used a chromium-doped neodymium/yttrium-aluminum-garnet (Cr-Nd-YAG) solid-state laser and a compact reflecting telescope. It sent a small portion of each emitted laser pulse through an optical fiber of known length and into the receiver, providing a continuous in-flight calibration of the timing circuit. The ranging data were used to construct a global shape model and a global topographic map of Eros with horizontal resolution of about 300 m. The NLR also measured detailed topographic profiles of surface features on Eros with a best spatial resolution of under 5 m. These topographic profiles enhanced and complemented the study of surface morphology from imaging. Magnetometer ------------ The fluxgate magnetometer used ring core sensors made of highly magnetically permeable material. MAG searched for any intrinsic magnetic fields of Eros. The recent Galileo flybys of the S-type asteroids Gaspra and Ida yielded evidence that both of these bodies are magnetic, although this evidence is ambiguous [KIVELSONETAL1993]. Discovery of an intrinsic magnetic field at Eros would have been the first definitive detection of magnetism at an asteroid and would have yielded important insights about its thermal and geological history. Radio Science ------------- In addition to the six major instruments, a coherent X-band transponder was used to conduct a radio science investigation by measuring the Doppler shift from the spacecraft's radial velocity component relative to the Earth. Accurate measurements of the Doppler shift and the range to Earth as the spacecraft orbited Eros allowed mapping of the asteroid's gravity field. In conjunction with MSI/NIS and NLR data, gravity determinations were combined with global shape and rotation data to constrain the internal density structure of Eros and search for heterogeneity. Mission Profile =============== The NEAR spacecraft was successfully launched in February 1996, taking advantage of the unique alignment of Earth and Eros that occurs only once every 7 years [FARQUHARETAL1995]. A Delta-II 7925 rocket placed NEAR into a 2-year DV (trajectory correction maneuver)/Earth gravity-assist trajectory (DVEGA). This trajectory represents a new application of the DVEGA technique: Instead of using an Earth swingby maneuver to increase the aphelion of the spacecraft trajectory, the maneuver actually decreased the aphelion distance while increasing the inclination from 0 to about 10 deg. The circuitous 3-year flight path to Eros was the result of a Discovery Program requirement to use an inexpensive, but less capable, launch vehicle. With a larger launch vehicle such as an Atlas or Titan, a 1-year direct trajectory could have been used, but the total mission cost would have increased by at least $50 million. The Mathilde encounter occurred 1 week before the deep space maneuver on 3 July 1997. The Earth swingby occurred on 23 January 1998. Rendezvous operations at Eros were scheduled to begin on 20 December 1998, but a main rocket engine abort occurred. A flyby of Eros was accomplished on 23 December 1998, and the rendezvous was rescheduled for 14 February 2000, when orbit insertion occurred. On 12 February 2001, NEAR accomplished a soft landing on Eros. Mathilde Flyby -------------- Asteroid 253 Mathilde was discovered on 12 November 1885 by Johann Palisa in Vienna, Austria. The name was suggested by V. A. Lebeuf (1859-1929), a staff member of the Paris Observatory, who first computed an orbit for the new asteroid. The name is thought to honor the wife of astronomer Moritz Loewy (1833-1907), then the vice director of the Paris Observatory. Although Mathilde's existence has been known since 1885, it was only following the announcement of NEAR's possible flyby that extensive physical observations were carried out using telescopes on Earth. These showed that Mathilde was an unusual object, especially because of its rotation, which is at least an order of magnitude slower than typical main-belt asteroids. Using a series of observations of this asteroid made in the first half of 1995, Stefano Mottola and his colleagues [MOTTOLAETAL1995] determined that Mathilde's rotation period is an extremely long 17.4 days. Only two asteroids, 288 Glauke and 1220 Clocus, have longer periods (48 and 31 days, respectively), and there is no obvious mechanism that can account for these extremely long asteroid 'days.' The only previous spacecraft encounters with asteroids, as noted earlier, had been the Galileo flybys of 951 Gaspra in October 1991 and 243 Ida in August 1993. Both of these objects, as well as Eros, are S-type asteroids. However, the most common type of asteroid in the outer asteroid belt, the dark and primitive C-type objects, had not yet been investigated. Spectral observations of Mathilde showed that its spectrum was consistent with those of C-type asteroids and that it was similar to those of the large carbonaceous asteroids 1 Ceres and 2 Pallas (the two largest asteroids). (Mathilde is about twice the size of Ida and four times the size of Gaspra.) Before the NEAR spacecraft executed its flyby of Mathilde on 27 June 1997, these additional facts were known about the asteroid: estimated diameter, 61 km; H magnitude (a measure of absolute visual brightness), 10.30; perihelion, 1.94 AU; aphelion, 3.35 AU; and orbital inclination, 6.71 degrees. Prior to the NEAR spacecraft encounter with Mathilde, on 27 June 1997 Mission Operations sent a command to the NEAR spacecraft that had the effect of advancing the Mission Elapsed Time (MET) clock by 10 seconds. This command was issued in order to correct for a timing error in the Mathilde fly-by observing sequence due to ephemeris uncertainties which existed at the time the sequence was generated and loaded to the spacecraft. After analysis of the final optical navigation data, the navigation team determined an additional shift decrementing the MET clock by 1 second was necessary. Mission Operations sent the additional command to the spacecraft; thus collectively these commands had the effect of incrementing the MET clock on board the NEAR spacecraft by 9 seconds. The NEAR spacecraft fly-by of Mathilde was then successfully executed. Following the Mathilde fly-by Mission Operations commanded the spacecraft to restore the MET clock. NEAR's encounter with Mathilde occurred at about 2 AU from the Sun, where available power from the solar panels was reduced to about 25% of its maximum mission level. Furthermore, a requirement to point the solar panels about 50 deg away from the optimal solar direction during the encounter reduced the available power by another 36%. Because of this power constraint, the only science instrument operated during the encounter period was MSI [LANDSHOF&CHENG1995]. However, spacecraft tracking data for the radio science experiment were obtained for an asteroid mass determination [CHENGETAL1994]. The imaging experiment during the flyby had three major objectives: 1. Most importantly, to obtain at least one image of Mathilde near closest approach to provide the highest-spatial-resolution view of the surface 2. To obtain an image of the complete illuminated portion of the asteroid visible during the flyby 3. To acquire images of the sky around the asteroid to search for possible satellites The entire imaging sequence was accomplished in about 25 min around closest approach (1200 km) at a speed of 9.93 km/s (Sun distance, 1.99 AU; Earth distance, 2.19 AU). A total of 534 images (24 high phase angle, 144 high-resolution, 188 global color imaging, 178 satellite search) were obtained during this interval. The whole illuminated portion of the asteroid was imaged in color at about a 500 m/pixel at a phase angle near 40 degrees. The best partial views were at 200 to 350 m/pixel. Mathilde's mass was determined by accurately tracking NEAR before and after the encounter. Apart from an interval of 1 to 2 h during the closest approach period, when imaging experiments were conducted, continuous tracking of the spacecraft was conducted for 3 days on either side of closest approach. During the flyby, Mathilde exerted a slight gravitational tug on NEAR. The corresponding gravitational tugs on the Galileo spacecraft at Gaspra and Ida were too small to allow mass determinations. However, because Mathilde's mass is so much larger than either Gaspra's or Ida's, its effects on NEAR's path were detectable in the spacecraft's radio tracking data. Earth Swingby ------------- The next critical phase of NEAR's flight profile was scheduled for 23 January 1998, when the spacecraft would pass by the Earth at an altitude of only 532 km. This maneuver was expected to drastically alter NEAR's heliocentric trajectory, changing the inclination from 0.52 to 10.04 deg, and reducing the aphelion distance from 2.18 to 1.77 AU and perihelion distance from 0.95 to 0.98 AU. An interesting consequence of the Earth flyby was that the post-swingby trajectory remained over the Earth's south polar region for a considerable time. During the encounter MSI and NIS observations of both Earth and the Moon were acquired from 23 January through 26 January, to test instrument performance during extended operations like at Eros, and to perform inflight radiance and alignment calibrations. Eros Encounter -------------- The NEAR mission target, 433 Eros, is the second largest asteroid and is intermediate in size between Gaspra and Ida. Eros is one of only three near-Earth asteroids with maximum diameter above 10 km, and it is the only large one whose heliocentric orbit is accessible enough to permit a rendezvous mission using the Delta II launch vehicle. The mean diameter of Eros, about 17 km, is an order of magnitude larger than that of typical known near-Earth asteroids. Eros was discovered in 1898. It was the subject of a worldwide ground-based observing campaign in 1975 when it passed within 0.15 AU of Earth. Visible, infrared, and radar observations determined the approximate size, shape, rotation rate, and pole position of Eros (Table 1) and showed that a regolith (fragmentary material produced by impacts) was present on its surface. 433 Eros is presently in a Mars-crossing (but not Earth-crossing) orbit; however, numerical simulations suggest that it may evolve into an Earth crosser within 2 million years. [MICHELETAL1996] Spectroscopic analyses have found the visible and near-infrared spectra of Eros to be consistent with a silicate mineralogy like that found in ordinary chondrite meteorites. These measurements were extended to higher spatial resolution by NEAR. Rendezvous operations at Eros were scheduled to begin on 20 December 1998, culminating in orbit insertion on 10 January. During the first of four main rocket engine firings to match velocity with Eros, on 20 December, an abort occurred and NEAR flew by Eros on 23 December at a relative velocity of 1 km/s. At this time a contingency sequence was executed during which data were collected by MSI, NIS, and MAG. The whole illuminated portion of the asteroid was imaged in color at about 500 m/pixel before and after closest approach at phase angles of 80 to 110 degrees. The best partial views were at about 400 m/pixel. Eros Operations =============== Beginning in January 2000, a sequence of small maneuvers decreased the relative velocity between NEAR and Eros to only 5 m/s. On 13 Feb 2000, NEAR performed a flyby of Eros on its sunward side at a distance as of 200 km. In addition to gathering NIS spectra at an optimal illumination geometry, this first pass provided improved estimates of the asteroid's physical parameters, such as a mass determination to 1% accuracy, identification of surface landmarks, and an improved estimate of Eros's spin vector. Orbit insertion occurred 14 Feb. As the spacecraft orbiter altitude was subsequently lowered, the mass, moments of inertia, gravity harmonics, spin state, and landmark locations were determined with increasing precision. NEAR operated in a series of orbits that came as close as 3 km to the asteroid's surface, culminating with a soft landing on 12 February 2001. The evolution of low-altitude orbits around Eros was strongly influenced by its irregular gravity field. In unstable orbits, the spacecraft could crash into Eros in a matter of days. Safe operation of NEAR during its 11-month prime science phase required close coordination between the science, mission design, navigation, and mission operations teams. [LANDSHOF&CHENG1995] To simplify science operations, the rendezvous was divided into distinct phases [CHENGETAL1994]. During each mission phase, particular aspects of the science were emphasized for science planning, so the highest priority investigation controlled instrument pointing for the majority of the observing time. The highest-priority science varied by mission phase, because of the changing orbital geometry. While in orbits at 100 km or more from the center of Eros, the highest priority science was global mapping by MSI. In orbits at 50 km or lower, the highest priority science was compositional measurement by XRS/GRS. A two-week period was allocated to altimetry by NLR at the start of the 50 km polar orbits. NEAR spent more than 150 days in orbits at 50 km or less from the center of Eros, plus two additional weeks on the surface acquiring GRS data. Data Flow --------- All data from the NEAR mission were down-linked to the NASA Deep Space Network and then forwarded to the Mission Operations Center (MOC) at APL. Doppler and ranging data from the spacecraft were analyzed primarily by the NEAR navigation team at the Jet Propulsion Laboratory (JPL) and processed to determine the spacecraft ephemeris as well as to perform radio science investigations. The entire spacecraft telemetry stream, including spacecraft and instrument housekeeping data and all science data, was forwarded to the APL MOC together with the radiometric Doppler and range data. Navigation data including spacecraft Ephemeris were forwarded to MOC in the form of SPICE kernels. (SPICE is an information system developed by the Navigation Ancillary Information Facility at JPL. It consists of data files and software for managing navigation-related data including spacecraft and planetary ephemerides, spacecraft pointing, timekeeping, gravity data, etc.) From the APL MOC the spacecraft telemetry stream were passed to the Science Data Center (SDC), the project facility responsible for low-level processing of spacecraft telemetry, data distribution, and data archiving. As such, the SDC supported the activities of the science team in data analysis and mission planning. The SDC created and maintained an archive, which was the central project repository for science data products such as images, asteroid models, and asteroid maps. The SDC enabled easy access to mission data sets by members of the science team and by others, and it collected observing requests and science priorities from the science team. It maintained a telemetry archive, a record of instrument and spacecraft commands as executed, and records of science sequences as requested and as executed. It provided ancillary data (spacecraft and planetary ephemerides, spacecraft and planetary attitudes, shape and gravity files, and spacecraft clock files) in the form of SPICE kernels to the science team. Conclusion ========== NEAR substantially increased our knowledge of primitive bodies in the solar system by providing a long, up-close look at the S-type asteroid 433 Eros and the first resolved images of the C-type asteroid 253 Mathilde. NEAR was the first mission to a near-Earth asteroid and a C-type asteroid, and it was the first spacecraft to flyby, orbit, and land on a small body.
  • data set : ULY JUP URAP RADIO ASTRONOMY REC AVERAGE E-FIELD 144 SEC
    ULY JUP URAP RADIO ASTRONOMY REC AVERAGE E-FIELD 144 SEC
  • instrument : RADIO SCIENCE SUBSYSTEM for MGN
    Instrument Overview =================== The Magellan Radio Science investigations utilized instrumentation with elements on the spacecraft and at the DSN. Much of this is shared equipment, being used for routine telecommunications as well as for Radio Science. The performance and calibration of both the spacecraft and tracking stations directly affect the radio science data accuracy, and they play a major role in determining the quality of the results. The spacecraft part of the radio science instrument is described immediately below; that is followed by a description of the DSN (ground) part of the instrument. Instrument Specifications - Spacecraft ====================================== The Magellan spacecraft telecommunications subsystem served as part of a radio science subsystem for investigations of Venus. Many details of the subsystem are unknown; its 'build date' is taken to be 1989-01-01, which was during the prelaunch phase of the Magellan mission. Instrument Id : RSS Instrument Host Id : MGN Pi Pds User Id : UNK Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : 1989-01-01 Instrument Mass : 102.2 KG Instrument Length : UNK Instrument Width : UNK Instrument Height : UNK Instrument Manufacturer Name : UNK Instrument Overview - Spacecraft ================================ The spacecraft radio system was constructed around a redundant pair of NASA S-band Standard Transponders (NSTs). In front of the S-band transponders was a single cross-strapped X-band down-converter, which allowed the NST to receive and transmit at both S-band (2.3 GHz, 13 cm wavelength) and X-band (8.4 GHz, 3.6 cm wavelength) frequencies. The X-band capability reduced plasma effects on radio signals by a factor of 10 and significantly improved the quality of radio tracking data for gravity investigations. The exact transmitted frequency was controlled by the signal received from a ground station or by an on-board oscillator. Each transponder included a receiver, command detector, exciter, and low-power amplifier. The transponders provided the usual uplink command and downlink data transmission capabilities. The spacecraft was capable of either S- or X-band uplink for commanding and simultaneous S- and X-band downlink for telemetry. The NST generated a downlink signal in either a 'coherent' or a 'non-coherent' mode, also known as the 'two- way' and 'one-way' modes, respectively. When operating in the coherent mode, the NST behaved as a conventional transponder; its transmitted carrier frequency was derived coherently from the received uplink carrier frequency with a 'turn-around ratio' of 880/749 at X-band and 240/221 at S-band. When the X-band downlink was controlled by an S-band uplink, the turn-around ratio was (240/221)*(11/3). In the non-coherent mode, the downlink carrier frequency was derived from one of the spacecraft's on-board crystal-controlled oscillators. After a 3-hour warm-up, the 'local' oscillator frequency was estimated to be 8425.864 +/- 0.002 MHz at X-band; the S-band signal was lower by a factor 3/11. Auxiliary Oscillator A was judged to have slightly better stability. During bistatic radar experiments in May-June 1994 it served as the frequency reference. After a 1.5 hour warm-up (first orbit, T=41C) its X-band frequency was estimated to be 8425.831 +/- 0.002 MHz; during subsequent orbits (T=43.3C) its frequency was estimated to be 8425.827+/-0.002 MHz. These frequencies depended primarily on the thermal environment, which could range over 43C < T < 45C after initial warm-up. Sensitivity of Auxiliary Oscillator A to thermal changes was -1.9 kHz per degree C. The strength of a spacecraft carrier signal, and thus the quality of the radio science data, depends on its modulation state. Magellan radar data were sent to Earth at a nominal rate of 268.8 kilobits per second at X-band; S-band was used for transmission of engineering data at 1.2 kilobits per second. Backup data rates of 115.2 kilobits for radar data and 40 bits per second for engineering were available for special contingencies and were required frequently during later phases of the mission. Traveling wave tube amplifiers, driven at saturation, amplified the NST output before the signals were radiated via (nominally) a 3.7 m diameter parabolic high gain antenna (HGA). The same antenna was used for the on-board Synthetic Aperture Radar (SAR) system. Medium-gain and low-gain antennas (MGA and LGA, respectively) were also provided. The S-band signal transmitted by the Magellan HGA was linearly polarized. For radar operation, the nominal S-band polarization was linear in the y-direction of the spacecraft coordinate system (HH on the surface). For telecommunications (and radio science) the S-band polarization was linear in the x-direction. The X-band signal transmitted by the HGA was circularly polarized, with the sense of polarization depending on the transponder connected. More information can be found in the Magellan Spacecraft Final Report [MGN-MA-011-1995]. Science Objectives ================== Two different types of radio science experiments were conducted with Magellan: radio tracking experiments in which the magnitude and direction of the planet's gravity field were derived from the Doppler (and, sometimes, ranging) measurements, and radio propagation experiments in which signal modulation detected on Earth could be attributed to properties of the medium. Several variations of the radio propagation experiments were carried out including radio occultations by the atmosphere of Venus and scattering from its surface. Polarization and scintillation measurements were also obtained when the radio wave passed through the solar corona. Gravity Measurements -------------------- Measurement of the gravity field provides significant constraints on inferences about the interior structure of Venus. Precise, detailed study of the spacecraft motion in Venus orbit can yield constraints on the mass distribution of the planet. Topographic data obtained by the Magellan Altimeter (ALT) forms a critical adjunct to these measurements since only after the gravitational effects are adjusted for topography can the gravity anomalies be interpreted geophysically. Studies of the gravity field emphasize both the global field and local characteristics of the field. The first task is to determine the global field. Doppler and range tracking measurements yield accurate spacecraft trajectory solutions. Simultaneously with reconstruction of the spacecraft orbit, observation equations for field coefficients and a small number of ancillary parameters can be solved. This type of gravity field solution is essential for characterizing tectonic phenomena and can also be used to study localized features. 'Short-arc' line-of-sight Doppler tracking measurements obtained when the Earth-to-spacecraft line-of-sight is within a few degrees of the orbit plane provide the highest resolution of local features. The results from this type of observation typically are presented as contoured acceleration profiles of specific features (e.g., craters, volcanoes, etc.) or line-of-sight acceleration maps of specific regions. The high spatial resolution of these products makes them especially useful to geophysicists for study of features in the size range of 100 to 1,000 km. Because of the relative simplicity of the data analysis, results can be available within a few weeks after the data are collected. The gravity investigation has been described in more detail by [KONOPLIV&SJOGREN1996] and [KONOPLIVETAL1999]. Radio Occultation Measurements ------------------------------ Atmospheric measurements by the method of radio occultation contribute to an improved understanding of structure, circulation, dynamics, and transport in the atmosphere of Venus. These results are based on detailed analysis of the radio signal received from Magellan as it enters and exits occultation by the planet. Two phases of the atmospheric investigation may be defined. The first is to obtain vertical profiles of atmospheric structure with emphasis on investigation of large-scale phenomena. The second is to concentrate on studies of the absorption at various levels in the atmosphere (absorption correlates with concentration of sulfuric acid vapor). Retrieval of atmospheric profiles requires coherent samples of the radio signal that has propagated through the atmosphere, plus accurate knowledge of the antenna pointing and the spacecraft trajectory. The latter is obtained from the Magellan Navigation Team. Initial solutions from Magellan occultations provided atmospheric structure -- temperature and pressure vs. absolute radius -- to altitudes as low as about 35 km from the surface. The atmosphere becomes critically refracting at levels only a short distance lower, so these results represent very nearly the maximum depth achievable using radio occultation probing. The spatial and temporal coverage in the radio occultation experiments are determined by the geometry of the spacecraft orbit and the dates and times at which occultation data were acquired. Since radio occultation experiments were conducted on an ad hoc basis, the Magellan coverage is sparse. Bistatic Surface Scattering Measurements ---------------------------------------- The spacecraft telecommunications antenna can also be pointed toward the surface of the planet. The strength of the scattered signal from the illuminated area may then be interpreted in terms of the texture of the surface at that point. In an experiment conducted on 6 October 1993, the Magellan high-gain antenna was aimed toward the summit of Gula Mons and the scattered signal was measured at DSN stations in California and Australia. On 9 November 1993, the high-gain antenna was aimed toward the (moving) point on the surface which would give mirror-like reflection toward Earth; those signals were received at the DSN station in Spain. In May and June 1994 experiments were conducted over Maxwell Montes; those data were processed to give the polarization of the reflected signal, which can lead to estimates of complex dielectric constant of the surface material. Results were published by [PETTENGILLETAL1996]. Solar Scintillation and Faraday Rotation Experiments ---------------------------------------------------- Solar scintillation and Faraday rotation experiments are conducted to improve our understanding of the structure and dynamics of the solar corona and wind. Because Venus orbits the Sun, spacecraft like Magellan are transported behind the solar disk, as seen from Earth. Radio waves propagating between Magellan and Earth stations are refracted and scattered (scintillation) by the solar plasma [WOO1993]. Intensity fluctuations can be related to fluctuations in electron density along the path, while Doppler or phase scintillations can be related to both electron density fluctuations and also the speed of the solar wind. The plane of linearly polarized waves can be rotated if there is an accompanying magnetic field (Faraday rotation). Many plasma effects decrease as the square of the radio frequency; scintillations are about an order of magnitude stronger at S-band than X-band. But only Magellan's S-band radio system used linear polarization; the circularly polarized X-band system was essentially immune to Faraday rotation. Operational Considerations - Spacecraft ======================================= Descriptions given here are for nominal performance. The spacecraft transponder system comprised redundant units, each with slightly different characteristics. As transponder units age, their performance changes slightly. More importantly, the performance for radio science depended on operational factors such as the modulation state for the transmitters, which cannot be predicted in advance. The performance also depended on factors which were not always under the control of the Magellan Project. The sample design control table below is adapted from [BUROW1986] for the X-band link carrying 268.8 kbps telemetry. Mean and variance denote expectations after accounting for best and worst case variations on design values. Parameter Design Mean Variance Value ----------------------------------------------------------- Transmit Parameters: Total transmit power (dBm) 43.42 43.26 0.0139 Transmit circuit loss (dB) -0.80 -0.85 0.0176 Transmit antenna gain (dB) 48.75 48.75 0.0104 Transmit antenna pointing (dB) 1.03* -0.65* 0.1484 Path Parameters: Space loss (dB) -279.09 -279.09 -- Atmospheric loss (dB) -0.31 -0.31 0.0000 Receive Parameters: Polarization loss (dB) -0.04 -0.04 0.0005 Receive antenna gain (dB) 73.42 73.42 0.1200 Receive antenna pointing (dB) -0.10 -0.07 0.0006 Receive circuit loss (dB) 0.00 0.00 0.0000 Total Power Summary: Net loss (dB) -158.84 0.2975 Total receiver power (dBm) -115.58 0.3114 Receiver noise (dBm/Hz) -179.64 -179.76 0.0416 Carrier Performance: Carrier modulation loss (dB) -13.83 -14.32 2.1302 Polarization loss (dB) 0.00 0.00 0.0000 Receiver carrier power (dBm) -129.90 2.4416 Carrier noise bandwidth (dB/Hz) 20.00 19.98 0.2111 Carrier threshold SNR (dB) 10.00 10.00 0.0000 Carrier threshold power (dBm) -149.78 0.0627 Performance margin (dB) 19.98 2.5042 * Archivist's Note: The sign difference between design value and mean is carried forward from the original [BUROW1986]. We were unable to contact the author and confirm his intent. A plausible explanation is that the design value should be -1.03, reflecting performance loss from antenna pointing errors, while the -0.65 mean represents a more favorable expectation after further design study. The specific values are not crucial to interpretation of these radio science data. On 4 January 1992 at 15:39 UTC, after a star calibration on orbit 3880, the X-band telemetry and high-rate science data did not reappear in the Magellan downlink. After several configuration changes and diagnostic tests, the Spacecraft Team reached the preliminary conclusion that a summing amplifier in Transponder-A had failed. This summing amplifier combined the X-band engineering and science telemetry signals before phase-modulation, amplification, and transmission of the signal to Earth via the High-Gain Antenna. As part of the testing on 4 January, Transponder-B was turned on and substituted for Transponder-A for about 25 minutes. Transponder-B had been turned off on 13 March 1991 when it developed a spurious sideband that masked the data subcarriers. The spurious signal appeared several MHz above the carrier frequency at X-band, sometimes containing more power than the carrier. During the short test on 4 January, Transponder-B started out at 28-deg C and transmitted good data to Earth. Later it warmed to 34-deg C and the downlink signal degraded. The probable failure mechanism of Transponder-B was determined to be a cracked, failed capacitor. Informal Spacecraft Team contacts with Motorola on 6 January indicated that Transponder-B might be returned to use if it were operated with little or no thermal cycling. That strategy was adopted for most of the remainder of the mission. Use of the 115.2 kbs science data rate allowed reception of radar data on Earth through most of that time despite presence of the spur, but the mapping time on each orbit was reduced commensurate with the playback capabilities. Both Transponder-A and -B could support gravity observations in their degraded states. Both transponders had strong X-band carriers in spite of their failures. The existence of the spur also affected radio propagation observations to the extent that carrier power was reduced to between a tenth and a half of its nominal level when Transponder-B was in use. Over time periods of an orbit, the frequency and strength of the spur did not change appreciably; so a calibration of carrier level at one point in the orbit was sufficient to establish carrier level during other parts of the orbit. For the bistatic surface scattering observations in October and November 1993, estimates of transmitted power and spur parameters were derived from spacecraft engineering telemetry and ground measurements. The total power at S-band was 4.188+/-0.025 watts; the total power at X-band was 17.298+/-0.100 watts. The X-band spur was approximately 9.3 MHz above and at least 0.5 dB stronger than the carrier. For experiments conducted in May and June of 1994, Transponder-A was used so that full power was available in each carrier (no X-band spur). Calibration Description - Spacecraft ==================================== No information available. Platform Mounting Descriptions - Spacecraft =========================================== The spacecraft +Z axis vector was in the nominal direction of the HGA boresight. The +X axis vector was parallel to the nominal rotation axis of the solar panels. The +Y axis vector formed a right-handed coordinate system and was in the nominal direction of the star scanner boresight. The spacecraft velocity vector was in approximately the -Y direction when the spacecraft was oriented for left-looking SAR operation. The nominal HGA S-band polarization for radar operation was linear in the y-direction; the nominal HGA S-band polarization for telecommunications (and radio science) was linear in the x-direction. The nominal X-band polarization was circular, with the sense depending on the transmitter attached. Cone Offset Angle : 0.00 Cross Cone Offset Angle : 0.00 Twist Offset Angle : 0.00 The medium gain antenna boresight was 70 degrees from the +Z direction and 20 degrees from the -Y direction. The low gain antenna was mounted on the back of the HGA feed; its boresight was in the +Z direction and it had a hemispherical radiation pattern. Principal Investigators ======================= The Principal Investigators for the gravity investigations were William L. Sjogren, Michel Lefebvre, and Georges Balmino. The Principal Investigator for the radio occultation experiments was G. Leonard Tyler. The Principal Investigator for the surface bistatic radar experiments was Gordon H. Pettengill. Investigators for the solar scintillation and Faraday rotation observations were Richard Woo and Michael Bird, respectively. Instrument Section / Operating Mode Descriptions - Spacecraft ============================================================= The Magellan radio system consisted of two sections, which could be operated in the following modes: Section Mode ------------------------------------------- Oscillator two-way (coherent) one-way (non-coherent) RF output low-gain antenna (no information available) medium-gain antenna high-gain antenna Selected parameters describing NST performance are listed below: Oscillator Parameters: S-Band X-Band Two-Way Transponder Turnaround Ratio 880/749 880/749 One-Way Transmit Frequency (MHz) 2297.963 8425.864 Frequency Uncertainty (+/- MHz) 0.0005 0.002 Nominal Wavelength (cm) 12.97 3.54 RF Output parameters: S-Band X-Band RF Power Output (w) 5.6 17.8 Medium-Gain Antenna: Half-Power Half Beamwidth (deg) 9.2 Gain (dBi) 19. EIRP (dBm) 56.5 Polarization Circular High-Gain Antenna: Half-Power Half-Beamwidth (deg) 1.1 0.32 Gain (dBi) 35.9 48.3 EIRP (dBm) 71.4 90.5 Polarization Linear Circular Instrument Overview - DSN ========================= Three Deep Space Communications Complexes (DSCCs) (near Barstow, CA; Canberra, Australia; and Madrid, Spain) comprise the DSN tracking network. Each complex is equipped with several antennas [including at least one each 70-m, 34-m High Efficiency (HEF), and 34-m standard (STD)], associated electronics, and operational systems. Primary activity at each complex is radiation of commands to and reception of telemetry data from active spacecraft. Transmission and reception is possible in several radio-frequency bands, the most common being S-band (nominally a frequency of 2100-2300 MHz or a wavelength of 14.2-13.0 cm) and X-band (7100-8500 MHz or 4.2- 3.5 cm). Transmitter output powers of up to 400 kw are available. Ground stations have the ability to transmit coded and uncoded waveforms which can be echoed by distant spacecraft. Analysis of the received coding allows navigators to determine the distance to the spacecraft; analysis of Doppler shift on the carrier signal allows estimation of the line-of-sight spacecraft velocity. Range and Doppler measurements are used to calculate the spacecraft trajectory and to infer gravity fields of objects near the spacecraft. Ground stations can record spacecraft signals that have propagated through or been scattered from target media. Measurements of signal parameters after wave interactions with surfaces, atmospheres, rings, and plasmas are used to infer physical and electrical properties of the target. Principal investigators vary from experiment to experiment. See the corresponding section of the spacecraft instrument description or the data set description for specifics. The Deep Space Network is managed by the Jet Propulsion Laboratory of the California Institute of Technology for the U.S. National Aeronautics and Space Administration. Specifications include: Instrument Id : RSS Instrument Host Id : DSN Pi Pds User Id : N/A Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : N/A Instrument Mass : N/A Instrument Length : N/A Instrument Width : N/A Instrument Height : N/A Instrument Manufacturer Name : N/A For more information on the Deep Space Network and its use in radio science investigations see the reports by [ASMAR&RENZETTI1993], [ASMAR&HERRERA1993], and [ASMARETAL1995]. For design specifications on DSN subsystems see [DSN810-5]. For an example of use of the DSN for Radio Science see [TYLERETAL1992]. Subsystems - DSN ================ The Deep Space Communications Complexes (DSCCs) are an integral part of the Radio Science instrument, along with other receiving stations and the spacecraft Radio Frequency Subsystem. Their system performance directly determines the degree of success of Radio Science investigations, and their system calibration determines the degree of accuracy in the results of the experiments. The following paragraphs describe the functions performed by the individual subsystems of a DSCC. This material has been adapted from [ASMAR&HERRERA1993]; for additional information, consult [DSN810-5]. Each DSCC includes a set of antennas, a Signal Processing Center (SPC), and communication links to the Jet Propulsion Laboratory (JPL). The general configuration is illustrated below; antennas (Deep Space Stations, or DSS -- a term carried over from earlier times when antennas were individually instrumented) are listed in the table. -------- -------- -------- -------- -------- | DSS 12 | | DSS 18 | | DSS 14 | | DSS 15 | | DSS 16 | |34-m STD| |34-m STD| | 70-m | |34-m HEF| | 26-m | -------- -------- -------- -------- -------- | | | | | | v v | v | --------- | --------- --------->|GOLDSTONE|<---------- |EARTH/ORB| | SPC 10 |<-------------->| LINK | --------- --------- | SPC |<-------------->| 26-M | | COMM | ------>| COMM | --------- | --------- | | | v | v ------ --------- | --------- | NOCC |<--->| JPL |<------- | | ------ | CENTRAL | | GSFC | ------ | COMM | | NASCOMM | | MCCC |<--->| TERMINAL|<-------------->| | ------ --------- --------- ^ ^ | | CANBERRA (SPC 40) <---------------- | | MADRID (SPC 60) <---------------------- GOLDSTONE CANBERRA MADRID Antenna SPC 10 SPC 40 SPC 60 -------- --------- -------- -------- 26-m DSS 16 DSS 46 DSS 66 34-m STD DSS 12 DSS 42 DSS 61 DSS 18 DSS 48 DSS 68 34-m HEF DSS 15 DSS 45 DSS 65 70-m DSS 14 DSS 43 DSS 63 Developmental DSS 13 Subsystem interconnections at each DSCC are shown in the diagram below, and they are described in the sections that follow. The Monitor and Control Subsystem is connected to all other subsystems; the Test Support Subsystem can be. ----------- ------------------ --------- --------- |TRANSMITTER| | | | TRACKING| | COMMAND | | SUBSYSTEM |-| RECEIVER/EXCITER |-|SUBSYSTEM|-|SUBSYSTEM|- ----------- | | --------- --------- | | | SUBSYSTEM | | | | ----------- | | --------------------- | | MICROWAVE | | | | TELEMETRY | | | SUBSYSTEM |-| |-| SUBSYSTEM |- ----------- ------------------ --------------------- | | | ----------- ----------- --------- -------------- | | ANTENNA | | MONITOR | | TEST | | DIGITAL | | | SUBSYSTEM | |AND CONTROL| | SUPPORT | |COMMUNICATIONS|- ----------- | SUBSYSTEM | |SUBSYSTEM| | SUBSYSTEM | ----------- --------- -------------- DSCC Monitor and Control Subsystem ---------------------------------- The DSCC Monitor and Control Subsystem (DMC) is part of the Monitor and Control System (MON) which also includes the ground communications Central Communications Terminal and the Network Operations Control Center (NOCC) Monitor and Control Subsystem. The DMC is the center of activity at a DSCC. The DMC receives and archives most of the information from the NOCC needed by the various DSCC subsystems during their operation. Control of most of the DSCC subsystems, as well as the handling and displaying of any responses to control directives and configuration and status information received from each of the subsystems, is done through the DMC. The effect of this is to centralize the control, display, and archiving functions necessary to operate a DSCC. Communication among the various subsystems is done using a Local Area Network (LAN) hooked up to each subsystem via a network interface unit (NIU). DMC operations are divided into two separate areas: the Complex Monitor and Control (CMC) and the Link Monitor and Control (LMC). The primary purpose of the CMC processor for Radio Science support is to receive and store all predict sets transmitted from NOCC such as Radio Science, antenna pointing, tracking, receiver, and uplink predict sets and then, at a later time, to distribute them to the appropriate subsystems via the LAN. Those predict sets can be stored in the CMC for a maximum of three days under normal conditions. The CMC also receives, processes, and displays event/alarm messages; maintains an operator log; and produces tape labels for the DSP. Assignment and configuration of the LMCs is done through the CMC; to a limited degree the CMC can perform some of the functions performed by the LMC. There are two CMCs (one on-line and one backup) and three LMCs at each DSCC The backup CMC can function as an additional LMC if necessary. The LMC processor provides the operator interface for monitor and control of a link -- a group of equipment required to support a spacecraft pass. For Radio Science, a link might include the DSCC Spectrum Processing Subsystem (DSP) (which, in turn, can control the SSI), or the Tracking Subsystem. The LMC also maintains an operator log which includes operator directives and subsystem responses. One important Radio Science specific function that the LMC performs is receipt and transmission of the system temperature and signal level data from the PPM for display at the LMC console and for inclusion in Monitor blocks. These blocks are recorded on magnetic tape as well as appearing in the Mission Control and Computing Center (MCCC) displays. The LMC is required to operate without interruption for the duration of the Radio Science data acquisition period. The Area Routing Assembly (ARA), which is part of the Digital Communications Subsystem, controls all data communication between the stations and JPL. The ARA receives all required data and status messages from the LMC/CMC and can record them to tape as well as transmit them to JPL via data lines. The ARA also receives predicts and other data from JPL and passes them on to the CMC. DSCC Antenna Mechanical Subsystem --------------------------------- Multi-mission Radio Science activities require support from the 70-m, 34-m HEF, and 34-m STD antenna subnets. The antennas at each DSCC function as large-aperture collectors which, by double reflection, cause the incoming radio frequency (RF) energy to enter the feed horns. The large collecting surface of the antenna focuses the incoming energy onto a subreflector, which is adjustable in both axial and angular position. These adjustments are made to correct for gravitational deformation of the antenna as it moves between zenith and the horizon; the deformation can be as large as 5 cm. The subreflector adjustments optimize the channeling of energy from the primary reflector to the subreflector and then to the feed horns. The 70-m and 34-m HEF antennas have 'shaped' primary and secondary reflectors, with forms that are modified paraboloids. This customization allows more uniform illumination of one reflector by another. The 34-m STD primary reflectors are classical paraboloids, while the subreflectors are standard hyperboloids. On the 70-m and 34-m STD antennas, the subreflector directs received energy from the antenna onto a dichroic plate, a device which reflects S-band energy to the S-band feed horn and passes X-band energy through to the X-band feed horn. In the 34-m HEF, there is one 'common aperture feed,' which accepts both frequencies without requiring a dichroic plate. RF energy to be transmitted into space by the horns is focused by the reflectors into narrow cylindrical beams, pointed with high precision (either to the dichroic plate or directly to the subreflector) by a series of drive motors and gear trains that can rotate the movable components and their support structures. The different antennas can be pointed by several means. Two pointing modes commonly used during tracking passes are CONSCAN and 'blind pointing.' With CONSCAN enabled and a closed loop receiver locked to a spacecraft signal, the system tracks the radio source by conically scanning around its position in the sky. Pointing angle adjustments are computed from signal strength information (feedback) supplied by the receiver. In this mode the Antenna Pointing Assembly (APA) generates a circular scan pattern which is sent to the Antenna Control System (ACS). The ACS adds the scan pattern to the corrected pointing angle predicts. Software in the receiver-exciter controller computes the received signal level and sends it to the APA. The correlation of scan position with the received signal level variations allows the APA to compute offset changes which are sent to the ACS. Thus, within the capability of the closed-loop control system, the scan center is pointed precisely at the apparent direction of the spacecraft signal source. An additional function of the APA is to provide antenna position angles and residuals, antenna control mode/status information, and predict-correction parameters to the Area Routing Assembly (ARA) via the LAN, which then sends this information to JPL via the Ground Communications Facility (GCF) for antenna status monitoring. During periods when excessive signal level dynamics or low received signal levels are expected (e.g., during an occultation experiment), CONSCAN should not be used. Under these conditions, blind pointing (CONSCAN OFF) is used, and pointing angle adjustments are based on a predetermined Systematic Error Correction (SEC) model. Independent of CONSCAN state, subreflector motion in at least the z-axis may introduce phase variations into the received Radio Science data. For that reason, during certain experiments, the subreflector in the 70-m and 34-m HEFs may be frozen in the z-axis at a position (often based on elevation angle) selected to minimize phase change and signal degradation. This can be done via Operator Control Inputs (OCIs) from the LMC to the Subreflector Controller (SRC) which resides in the alidade room of the antennas. The SRC passes the commands to motors that drive the subreflector to the desired position. Unlike the 70-m and 34-m HEFs which have azimuth-elevation (AZ-EL) drives, the 34-m STD antennas use (hour angle-declination) HA-DEC drives. The same positioning of the subreflector on the 34-m STD does not create the same effect as on the 70-m and 34-m HEFs. Pointing angles for all three antenna types are computed by the NOCC Support System (NSS) from an ephemeris provided by the flight project. These predicts are received and archived by the CMC. Before each track, they are transferred to the APA, which transforms the direction cosines of the predicts into AZ-EL coordinates for the 70-m and 34-m HEFs or into HA-DEC coordinates for the 34-m STD antennas. The LMC operator then downloads the antenna AZ-EL or HA-DEC predict points to the antenna-mounted ACS computer along with a selected SEC model. The pointing predicts consist of time-tagged AZ-EL or HA-DEC points at selected time intervals along with polynomial coefficients for interpolation between points. The ACS automatically interpolates the predict points, corrects the pointing predicts for refraction and subreflector position, and adds the proper systematic error correction and any manually entered antenna offsets. The ACS then sends angular position commands for each axis at the rate of one per second. In the 70-m and 34-m HEF, rate commands are generated from the position commands at the servo controller and are subsequently used to steer the antenna. In the 34-m STD antennas motors, rather than servos, are used to steer the antenna; there is no feedback once the 34-m STD has been told where to point. When not using binary predicts (the routine mode for spacecraft tracking), the antennas can be pointed using 'planetary mode' -- a simpler mode which uses right ascension (RA) and declination (DEC) values. These change very slowly with respect to the celestial frame. Values are provided to the station in text form for manual entry. The ACS quadratically interpolates among three RA and DEC points which are on one-day centers. A third pointing mode -- sidereal -- is available for tracking radio sources fixed with respect to the celestial frame. Regardless of the pointing mode being used, a 70-m antenna has a special high-accuracy pointing capability called 'precision' mode. A pointing control loop derives the main AZ-EL pointing servo drive error signals from a two- axis autocollimator mounted on the Intermediate Reference Structure. The autocollimator projects a light beam to a precision mirror mounted on the Master Equatorial drive system, a much smaller structure, independent of the main antenna, which is exactly positioned in HA and DEC with shaft encoders. The autocollimator detects elevation/cross- elevation errors between the two reference surfaces by measuring the angular displacement of the reflected light beam. This error is compensated for in the antenna servo by moving the antenna in the appropriate AZ-EL direction. Pointing accuracies of 0.004 degrees (15 arc seconds) are possible in 'precision' mode. The 'precision' mode is not available on 34-m antennas -- nor is it needed, since their beamwidths are twice as large as on the 70-m antennas. DSCC Antenna Microwave Subsystem -------------------------------- 70-m Antennas: Each 70-m antenna has three feed cones installed in a structure at the center of the main reflector. The feeds are positioned 120 degrees apart on a circle. Selection of the feed is made by rotation of the subreflector. A dichroic mirror assembly, half on the S-band cone and half on the X-band cone, permits simultaneous use of the S- and X-band frequencies. The third cone is devoted to R&D and more specialized work. The Antenna Microwave Subsystem (AMS) accepts the received S- and X-band signals at the feed horn and transmits them through polarizer plates to an orthomode transducer. The polarizer plates are adjusted so that the signals are directed to a pair of redundant amplifiers for each frequency, thus allowing simultaneous reception of signals in two orthogonal polarizations. For S-band these are two Block IVA S-band Traveling Wave Masers (TWMs); for X-band the amplifiers are Block IIA TWMs. 34-m STD Antennas: These antennas have two feed horns, one for S-band signals and one for X-band. The horns are mounted on a cone which is fixed in relation to the subreflector. A dichroic plate mounted above the horns directs energy from the subreflector into the proper horn. The AMS directs the received S- and X-band signals through polarizer plates and on to amplification. There are two Block III S-band TWMs and two Block I X-band TWMs. 34-m HEF Antennas: Unlike the other antennas, the 34-m HEF uses a single feed for both S- and X-band. Simultaneous S- and X-band receive as well as X-band transmit is possible thanks to the presence of an S/X 'combiner' which acts as a diplexer. For S-band, RCP or LCP is user selected through a switch so neither a polarizer nor an orthomode transducer is needed. X-band amplification options include two Block II TWMs or an HEMT Low Noise Amplifier (LNA). S-band amplification is provided by an FET LNA. DSCC Receiver-Exciter Subsystem ------------------------------- The Receiver-Exciter Subsystem is composed of three groups of equipment: the closed-loop receiver group, the open-loop receiver group, and the RF monitor group. This subsystem is controlled by the Receiver-Exciter Controller (REC) which communicates directly with the DMC for predicts and OCI reception and status reporting. The exciter generates the S-band signal (or X-band for the 34-m HEF only) which is provided to the Transmitter Subsystem for the spacecraft uplink signal. It is tunable under command of the Digitally Controlled Oscillator (DCO) which receives predicts from the Metric Data Assembly (MDA). The diplexer in the signal path between the transmitter and the feed horn for all three antennas (used for simultaneous transmission and reception) may be configured such that it is out of the received signal path (in listen-only or bypass mode) in order to improve the signal-to-noise ratio in the receiver system. Closed Loop Receivers: The Block IV receiver-exciter at the 70-m stations allows for two receiver channels, each capable of L-Band (e.g., 1668 MHz frequency or 18 cm wavelength), S-band, or X-band reception, and an S-band exciter for generation of uplink signals through the low-power or high-power transmitter. The Block III receiver-exciter at the 34-m STD stations allows for two receiver channels, each capable of S-band or X-band reception and an exciter used to generate an uplink signal through the low-power transmitter. The receiver-exciter at the 34-m HEF stations allows for one channel only. The closed-loop receivers provide the capability for rapid acquisition of a spacecraft signal and telemetry lockup. In order to accomplish acquisition within a short time, the receivers are predict driven to search for, acquire, and track the downlink automatically. Rapid acquisition precludes manual tuning though that remains as a backup capability. The subsystem utilizes FFT analyzers for rapid acquisition. The predicts are NSS generated, transmitted to the CMC which sends them to the Receiver-Exciter Subsystem where two sets can be stored. The receiver starts acquisition at uplink time plus one round-trip-light-time or at operator specified times. The receivers may also be operated from the LMC without a local operator attending them. The receivers send performance and status data, displays, and event messages to the LMC. Either the exciter synthesizer signal or the simulation (SIM) synthesizer signal is used as the reference for the Doppler extractor in the closed-loop receiver systems, depending on the spacecraft being tracked (and Project guidelines). The SIM synthesizer is not ramped; instead it uses one constant frequency, the Track Synthesizer Frequency (TSF), which is an average frequency for the entire pass. The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. It will be configured such that the expected amplitude changes are accommodated with minimum distortion. The loop bandwidth (2BLo) will be configured such that the expected phase changes can be accommodated while maintaining the best possible loop SNR. Open-Loop Receivers: There are two types of Radio Science Open-Loop Receivers (OLR) in use. At 70-m and 34-m HEF stations the OLR is a a dedicated four channel, narrow-band receiver which provides amplified and downconverted video band signals to the DSCC Spectrum Processing Subsystem (DSP); it sometimes goes by the designation 'RIV'. At 34-m STD stations (DSS 42 and DSS 61) the OLR is an older system, the Multi-Mission Receiver (MMR), which provides two channels of narrow-band receiver output. Both OLR systems are described in detail below under 'Electronics - DSN'; here the overview continues only for the RIV system. The 70-m and 34-m HEF OLR utilizes a fixed first Local Oscillator (LO) frequency and a tunable second LO frequency to minimize phase noise and improve frequency stability. The OLR consists of an RF-to-IF downconverter located in the antenna, an IF selection switch (IVC), and a Radio Science IF-VF downconverter (RIV) located in the SPC. The RF-IF downconverters in the 70-m antennas are equipped for four IF channels: S-RCP, S-LCP, X-RCP, and X-LCP. The 34-m HEF stations are equipped with a two-channel RF-IF: S-band and X-band. The IVC switches the IF input between the 70-m and 34-m HEF antennas. The RIV contains the tunable second LO, a set of video bandpass filters, IF attenuators, and a controller (RIC). The LO tuning is done via DSP control of the POCA/PLO combination based on a predict set. The POCA is a Programmable Oscillator Control Assembly and the PLO is a Programmable Local Oscillator (commonly called the DANA synthesizer). The bandpass filters are selectable via the DSP. The RIC provides an interface between the DSP and the RIV. It is controlled from the LMC via the DSP. The RIC selects the filter and attenuator settings and provides monitor data to the DSP. The RIC could also be manually controlled from the front panel in case the electronic interface to the DSP is lost. RF Monitor -- SSI and PPM: The RF monitor group of the Receiver-Exciter Subsystem provides spectral measurements using the Spectral Signal Indicator (SSI) and measurements of the received channel system temperature and spacecraft signal level using the Precision Power Monitor (PPM). The SSI provides a local display of the received signal spectrum at a dedicated terminal at the DSCC and routes these same data to the DSP which routes them to NOCC for remote display at JPL for real-time monitoring and RIV/DSP configuration verification. These displays are used to validate Radio Science Subsystem data at the DSS, NOCC, and Mission Support Areas. The SSI configuration is controlled by the DSP and a duplicate of the SSI spectrum appears on the LMC via the DSP. During real-time operations the SSI data also serve as a quick-look science data type for Radio Science experiments. The PPM measures system noise temperatures (SNT) using a Noise Adding Radiometer (NAR) and downlink signal levels using the Signal Level Estimator (SLE). The PPM accepts its input from the closed-loop receiver. The SNT is measured by injecting known amounts of noise power into the signal path and comparing the total power with the noise injection 'on' against the total power with the noise injection 'off.' That operation is based on the fact that receiver noise power is directly proportional to temperature; thus measuring the relative increase in noise power due to the presence of a calibrated thermal noise source allows direct calculation of SNT. Signal level is measured by calculating an FFT to estimate the SNR between the signal level and the receiver noise floor where the power is known from the SNT measurements. There is one PPM controller at the SPC which is used to control all SNT measurements. The SNT integration time can be selected to represent the time required for a measurement of 30K to have a one-sigma uncertainty of 0.3K or 1%. DSCC Transmitter Subsystem -------------------------- The Transmitter Subsystem accepts the S-band frequency exciter signal from the Block III or Block IV Receiver- Exciter Subsystem exciter and amplifies it to the required transmit output level. The amplified signal is routed via the diplexer through the feed horn to the antenna and then focused and beamed to the spacecraft. The Transmitter Subsystem power capabilities range from 18 kw to 400 kw. Power levels above 18 kw are available only at 70-m stations. DSCC Tracking Subsystem ----------------------- The Tracking Subsystem primary functions are to acquire and maintain communications with the spacecraft and to generate and format radiometric data containing Doppler and range. The DSCC Tracking Subsystem (DTK) receives the carrier signals and ranging spectra from the Receiver-Exciter Subsystem. The Doppler cycle counts are counted, formatted, and transmitted to JPL in real time. Ranging data are also transmitted to JPL in real time. Also contained in these blocks is the AGC information from the Receiver-Exciter Subsystem. The Radio Metric Data Conditioning Team (RMDCT) at JPL produces an Archival Tracking Data File (ATDF) which contains Doppler and ranging data. In addition, the Tracking Subsystem receives from the CMC frequency predicts (used to compute frequency residuals and noise estimates), receiver tuning predicts (used to tune the closed-loop receivers), and uplink tuning predicts (used to tune the exciter). From the LMC, it receives configuration and control directives as well as configuration and status information on the transmitter, microwave, and frequency and timing subsystems. The Metric Data Assembly (MDA) controls all of the DTK functions supporting the uplink and downlink activities. The MDA receives uplink predicts and controls the uplink tuning by commanding the DCO. The MDA also controls the Sequential Ranging Assembly (SRA). It formats the Doppler and range measurements and provides them to the GCF for transmission to NOCC. The Sequential Ranging Assembly (SRA) measures the round trip light time (RTLT) of a radio signal traveling from a ground tracking station to a spacecraft and back. From the RTLT, phase, and Doppler data, the spacecraft range can be determined. A coded signal is modulated on an uplink carrier and transmitted to the spacecraft where it is detected and transponded back to the ground station. As a result, the signal received at the tracking station is delayed by its round trip through space and shifted in frequency by the Doppler effect due to the relative motion between the spacecraft and the tracking station on Earth. DSCC Spectrum Processing Subsystem (DSP) ---------------------------------------- The DSCC Spectrum Processing Subsystem (DSP) located at the SPC digitizes and records the narrowband output data from the RIV. It consists of a Narrow Band Occultation Converter (NBOC) containing four Analog-to-Digital Converters (ADCs), a ModComp CLASSIC computer processor called the Spectrum Processing Assembly (SPA), and several magnetic tape drives. Magnetic tapes containing DSP output are known as Original Data Records (ODRs). Electronic near real-time data transmission (known as an Original Data Stream, or ODS) may be possible in certain circumstances. The DSP is operated through the LMC. Using the SPA-Radioscience (SPA-R) software, the DSP allows for real-time frequency and time offsets (while in RUN mode) and, if necessary, snap tuning between the two frequency ranges transmitted by the spacecraft: coherent and non-coherent. The DSP receives Radio Science frequency predicts from the CMC, allows for multiple predict set archiving (up to 60 sets) at the SPA, and allows for manual predict generation and editing. It accepts configuration and control data from the LMC, provides display data to the LMC, and transmits the signal spectra from the SSI as well as status information to NOCC and the Project Mission Support Area (MSA) via the GCF data lines. The DSP records the digitized narrowband samples and the supporting header information (i.e., time tags, POCA frequencies, etc.) on 9-track magnetic tapes in 6250 or 1600 bpi GCR format. Through the DSP-RIC interface the DSP controls the RIV filter selection and attenuation levels. It also receives RIV performance monitoring via the RIC. In case of failure of the DSP-RIC interface, the RIV can be controlled manually from the front panel. All the RIV and DSP control parameters and configuration directives are stored in the SPA in a macro-like file called an 'experiment directive' table. A number of default directives exist in the DSP for the major Radio Science experiments. Operators can create their own table entries. Items such as verification of the configuration of the prime open-loop recording subsystem, the selection of the required predict sets, and proper system performance prior to the recording periods will be checked in real-time at JPL via the NOCC displays using primarily the remote SSI display at NOCC and the NRV displays. Because of this, transmission of the DSP/SSI monitor information is enabled prior to the start of recording. The specific run time and tape recording times will be identified in the Sequence of Events (SOE) and/or DSN Keyword File. The DSP can be used to duplicate ODRs. It also has the capability to play back a certain section of the recorded data after conclusion of the recording periods. DSCC Frequency and Timing Subsystem ----------------------------------- The Frequency and Timing Subsystem (FTS) provides all frequency and timing references required by the other DSCC subsystems. It contains four frequency standards of which one is prime and the other three are backups. Selection of the prime standard is done via the CMC. Of these four standards, two are hydrogen masers followed by clean-up loops (CUL) and two are cesium standards. These four standards all feed the Coherent Reference Generator (CRG) which provides the frequency references used by the rest of the complex. It also provides the frequency reference to the Master Clock Assembly (MCA) which in turn provides time to the Time Insertion and Distribution Assembly (TID) which provides UTC and SIM-time to the complex. JPL's ability to monitor the FTS at each DSCC is limited to the MDA calculated Doppler pseudo-residuals, the Doppler noise, the SSI, and to a system which uses the Global Positioning System (GPS). GPS receivers at each DSCC receive a one-pulse-per-second pulse from the station's (hydrogen maser referenced) FTS and a pulse from a GPS satellite at scheduled times. After compensating for the satellite signal delay, the timing offset is reported to JPL where a database is kept. The clock offsets stored in the JPL database are given in microseconds; each entry is a mean reading of measurements from several GPS satellites and a time tag associated with the mean reading. The clock offsets provided include those of SPC 10 relative to UTC (NIST), SPC 40 relative to SPC 10, etc. Optics - DSN ============ Performance of DSN ground stations depends primarily on size of the antenna and capabilities of electronics. These are summarized in the following set of tables. Note that 64-m antennas were upgraded to 70-m between 1986 and 1989. Beamwidth is half-power full angular width. Polarization is circular; L denotes left circular polarization (LCP), and R denotes right circular polarization (RCP). DSS S-Band Characteristics 64-m 70-m 34-m 34-m Transmit STD HEF -------- ----- ----- ----- ----- Frequency (MHz) 2110- 2110- 2025- N/A 2120 2120 2120 Wavelength (m) 0.142 0.142 0.142 N/A Ant Gain (dBi) 62.7 55.2 N/A Beamwidth (deg) 0.119 0.31 N/A Polarization L or R L or R N/A Tx Power (kW) 20-400 20 N/A Receive ------- Frequency (MHz) 2270- 2270- 2270- 2200- 2300 2300 2300 2300 Wavelength (m) 0.131 0.131 0.131 0.131 Ant Gain (dBi) 61.6 63.3 56.2 56.0 Beamwidth (deg) 0.108 0.27 0.24 Polarization L & R L & R L or R L or R System Temp (K) 22 20 22 38 DSS X-Band Characteristics 64-m 70-m 34-m 34-m Transmit STD HEF -------- ----- ----- ----- ----- Frequency (MHz) 8495 8495 N/A 7145- 7190 Wavelength (m) 0.035 0.035 N/A 0.042 Ant Gain (dBi) 74.2 N/A 67 Beamwidth (deg) N/A 0.074 Polarization L or R L or R N/A L or R Tx Power (kW) 360 360 N/A 20 Receive ------- Frequency (MHz) 8400- 8400- 8400- 8400- 8500 8500 8500 8500 Wavelength (m) 0.036 0.036 0.036 0.036 Ant Gain (dBi) 71.7 74.2 66.2 68.3 Beamwidth (deg) 0.031 0.075 0.063 Polarization L & R L & R L & R L & R System Temp (K) 27 20 25 20 NB: X-band 64-m and 70-m transmitting parameters are given at 8495 MHz, the frequency used by the Goldstone planetary radar system. For telecommunications, the transmitting frequency would be in the range 7145-7190 MHz, the power would typically be 20 kW, and the gain would be about 72.6 dB (70-m antenna). When ground transmitters are used in spacecraft radio science experiments, the details of transmitter and antenna performance rarely impact the results. Electronics - DSN ================= DSCC Open-Loop Receiver (RIV) ----------------------------- The open loop receiver block diagram shown below is for the RIV system at 70-m and 34-m High-Efficiency (HEF) antenna sites. Input signals at both S- and X-band are mixed to approximately 300 MHz by fixed-frequency local oscillators near the antenna feed. Based on a tuning prediction file, the POCA controls the DANA synthesizer, the output of which (after multiplication) mixes the 300 MHz IF to 50 MHz for amplification. These signals in turn are down converted and passed through additional filters until they yield Output with bandwidths up to 45 kHz. The Output is digitally sampled and either written to magnetic tape or electronically transferred for further analysis. S-Band X-Band 2295 MHz 8415 MHz Input Input | | v v --- --- --- --- | X |<--|x20|<--100 MHz 100 MHz-->|x81|-->| X | --- --- --- --- | | 295| |315 MHz| |MHz v v --- -- 33.1818 --- --- | X |<--|x3|<------ MHz ------>|x11|-->| X | --- -- |115 | --- --- | |MHz | | | | | | 50| 71.8181 --- --- |50 MHz| MHz->| X | | X |<-10MHz |MHz v --- --- v --- ^ ^ --- | X |<--60 MHz | | 60 MHz-->| X | --- | approx | --- | 9.9 | 43.1818 MHz | 9.9 | | MHz ------------- MHz | | | ^ | | 10| v | v |10 MHz| --- ---------- --- |MHz |------>| X | | DANA | | X |<------| | --- |Synthesizr| --- | | | ---------- | | v v ^ v v ------- ------- | ------- ------- |Filters| |Filters| ---------- |Filters| |Filters| |3,4,5,6| | 1,2 | | POCA | | 1,2 | |3,4,5,6| ------- ------- |Controller| ------- ------- | | ---------- | | 10| |0.1 0.1| |10 MHz| |MHz MHz| |MHz v v v v --- --- --- --- 10 MHz -->| X | | X |<------ 0.1 MHz ------->| X | | X |<-- 10 MHz --- --- --- --- | | | | v v v v Output Output Output Output Reconstruction of the antenna frequency from the frequency of the signal in the recorded data can be achieved through use of one of the following formulas. Filters are defined below. FSant=3*SYN+1.95*10^9+3*(790/11)*10^6+Frec (Filter 4) =3*SYN+1.95*10^9+3*(790/11)*10^6-Fsamp+Frec (Filters 1-3,5,6) FXant=11*SYN + 7.940*10^9 + Fsamp - Frec (Filter 4) =11*SYN + 7.940*10^9 - 3*Fsamp + Frec (Filters 1,2,3,6) where FSant,FXant are the antenna frequencies of the incoming signals at S and X bands, respectively, SYN is the output frequency of the DANA synthesizer, commonly labeled the readback POCA frequency on data tapes, Fsamp is the effective sampling rate of the digital samples, and Frec is the apparent signal frequency in a spectrum reconstructed from the digital samples. NB: For many of the filter choices (see below) the Output is that of a bandpass filter. The sampling rates in the table below are sufficient for the bandwidth but not the absolute maximum frequency, and aliasing results. The reconstruction expressions above are appropriate ONLY when the sample rate shown in the tables below is used. DSCC Open-Loop Receiver (MMR at DSS 5 and 61) --------------------------------------------- The open loop receiver block diagram shown below is for MMR systems at the 34-m Standard (STD) DSS 61 antenna site and at the DSS 5 JPL DSN facility. Based on a tuning prediction file, the POCA controls the DANA synthesizer, the output of which (after multiplication) mixes input signals at both S- and X-band to fixed intermediate frequencies for amplification. These signals in turn are down converted and passed through additional filters until they yield Output with bandwidths up to 45 kHz. The Output is digitally sampled and either written to magnetic tape or electronically transferred for further analysis. S-Band X-Band 2295 MHz 8415 MHz Input Input | | v v --- --- | X |<------------- -------------->| X | --- 1995| |8115 --- | MHz| |MHz | | | | | | | --- | | | | X |<--800 MHz | | | --- | | | | | 300| | | |300 MHz| --- ---- |MHz | |x48| |x176| | v --- ---- v --- ^ ^ --- | X |<--290 MHz | | 290 MHz-->| X | --- | approx | --- | 9.9 | 41.56 MHz | 9.9 | | MHz ------------- MHz | | | ^ | | 10| v | v |10 MHz| --- ---------- --- |MHz |------>| X | | DANA | | X |<------| | --- |Synthesizr| --- | | | ---------- | | v v ^ v v ------- ------- | ------- ------- |Filters| |Filters| ---------- |Filters| |Filters| | 4-8 | | 1-3 | | POCA | | 1-3 | | 4-8 | ------- ------- |Controller| ------- ------- | | ---------- | | 10| |0.1 0.1| |10 MHz| |MHz MHz| |MHz v v v v --- --- --- --- 10 MHz -->| X | | X |<------ 0.1 MHz ------->| X | | X |<-- 10 MHz --- --- --- --- | | | | v v v v Output Output Output Output Reconstruction of the antenna frequency from the frequency of the signal in the recorded data can be achieved through use of one of the following formulas. Filters are defined below. FSant = 48*SYN + 300*10^6 - Fsamp + Frec (Filters 1,2,3,8) = 48*SYN + 300*10^6 + Frec (Filters 4,5,6,7) FXant = 176*SYN + 1100*10^6 - 3*Fsamp + Frec (Filters 1,2,3,8) = 176*SYN + 1100*10^6 + Frec (Filters 4,5,6,7) where the definition of terms and 'NB' are the same as for the RIV system (above). DSCC Open-Loop Receiver (MMR at DSS 7 and 42) --------------------------------------------- The open loop receiver block diagram shown below is for the MMR system at the 34-m Standard (STD) DSS 42 antenna site and the DSS 7 DSN facility at JPL. Based on a tuning prediction file, the POCA controls the DANA synthesizer, the output of which (after multiplication) mixes input signals at both S- and X-band to fixed intermediate frequencies for amplification. These signals in turn are down converted and passed through additional filters until they yield Output with bandwidths up to 45 kHz. The Output is digitally sampled and either written to magnetic tape or electronically transferred for further analysis. S-Band X-Band 2295 MHz 8415 MHz Input 800 MHz Input | | | v v 8115 v --- 1995 MHz --- MHz --- | X |<------------- | X |------------>| X | --- | --- --- | | | | | --- --- | | |x 3| |x11| | | --- approx --- | | | 665 MHz | | | ------------- | 300| | |300 MHz| --- |MHz | | X |<--600 MHz | v --- v --- ^ --- | X |<--290 MHz | 290 MHz-->| X | --- ----- --- | 9.9 |x 1.5| 9.9 | | MHz ----- MHz | | | ^ | | 10| v | v |10 MHz| --- ---------- --- |MHz |------>| X | | DANA | | X |<------| | --- |Synthesizr| --- | | | ---------- | | v v ^ v v ------- ------- | ------- ------- |Filters| |Filters| ---------- |Filters| |Filters| | 4-8 | | 1-3 | | POCA | | 1-3 | | 4-8 | ------- ------- |Controller| ------- ------- | | ---------- | | 10| |0.1 0.1| |10 MHz| |MHz MHz| |MHz v v v v --- --- --- --- 10 MHz -->| X | | X |<------ 0.1 MHz ------->| X | | X |<-- 10 MHz --- --- --- --- | | | | v v v v Output Output Output Output Reconstruction of the antenna frequency from the frequency of the signal in the recorded data can be achieved through use of one of the following formulas. Filters are defined below. FSant = (9/2)*SYN + 2100*10^6 - Fsamp + Frec (Filters 1,2,3,8) = (9/2)*SYN + 2100*10^6 + Frec (Filters 4,5,6,7) FXant = (33/2)*SYN + 7700*10^6 - 3*Fsamp + Frec (Filters 1,2,3,8) = (33/2)*SYN + 7700*10^6 + Frec (Filters 4,5,6,7) where the definition of terms and 'NB' are the same as for the RIV system (above). Filters - DSN ============= DSCC Open-Loop Receiver (RIV) ----------------------------- Nominal filter center frequencies and bandwidths for the RIV Receivers are shown in the table below. Recommended sampling rates are also given. S-Band X-Band ------------------------ ------------------------- Output 3 dB Sampling Output 3 dB Sampling Filter Center Band Rate Center Band Rate Freq Width Freq Width (Hz) (Hz) (sps) (Hz) (Hz) (sps) ------ ------ ------ -------- ------ ------ -------- 1 150 82 200 550 82 200 2 750 415 1000 2750 415 1000 3 3750 2000 5000 13750 2000 5000 4 1023 1700 5000 3750 6250 15000 5 75000 45000 100000 275000 45000 100000 6 37500 20000 50000 137500 20000 50000 DSCC Open-Loop Receiver (MMR) ----------------------------- MMR filters (DSS 5, 7, 42, and 61) and recommended sampling rates include the following: S-Band X-Band ------------------------ ------------------------- Output 3 dB Recommended Output 3 dB Recommended Filter Center Band Sampling Center Band Sampling Freq Width Rate* Freq Width Rate* (Hz) (Hz) (sps) (Hz) (Hz) (sps) ------ ------ ------ -------- ------ ------ -------- 1 150 100 200 550 100 200 2 750 500 1000 2750 500 1000 3 1500 1000 2000 5500 1000 2000 4 409 818 2000 1500 3000 6000 5 1023 2045 5000 3750 7500 15000 6 2045 4091 10000 7500 15000 30000 7 4091 8182 20000 15000 30000 60000 8 37500 20000 50000 137500 20000 50000 * Sampling rates depend on resolution of samples and number of analog-to-digital converters assigned to each channel -- see discussion of modes under 'DSCC Spectrum Processing Subsystem' below. The rates at which single A/D converters can operate with the MMR include: 8-bit samples: 12-bit samples: 16-bit samples: 200 200 1250 250 1000 400 1250 500 2000 1000 5000 1250 10000 2000 2500 3125 4000 5000 6250 10000 12500 15625 20000 25000 31250 50000 Detectors - DSN =============== DSCC Open-Loop Receivers ------------------------ Open-loop receiver output is detected in software by the radio science investigator. DSCC Closed-Loop Receivers -------------------------- Nominal carrier tracking loop threshold noise bandwidth at both S- and X-band is 10 Hz. Coherent (two-way) closed-loop system stability is shown in the table below: integration time Doppler uncertainty (secs) (one sigma, microns/sec) ------ ------------------------ 10 50 60 20 1000 4 Calibration - DSN ================= Calibrations of hardware systems are carried out periodically by DSN personnel; these ensure that systems operate at required performance levels -- for example, that antenna patterns, receiver gain, propagation delays, and Doppler uncertainties meet specifications. No information on specific calibration activities is available. Nominal performance specifications are shown in the tables above. Additional information may be available in [DSN810-5]. Prior to each tracking pass, station operators perform a series of calibrations to ensure that systems meet specifications for that operational period. Included in these calibrations is measurement of receiver system temperature in the configuration to be employed during the pass. Results of these calibrations are recorded in (hard copy) Controller's Logs for each pass. The nominal procedure for initializing open-loop receiver attenuator settings is described below. In cases where widely varying signal levels are expected, the procedure may be modified in advance or real-time adjustments may be made to attenuator settings. Open-Loop Receiver Attenuation Calibration ------------------------------------------ The open-loop receiver attenuator calibrations are performed to establish the output of the open-loop receivers at a level that will not saturate the analog-to-digital converters. To achieve this, the calibration is done using a test signal generated by the exciter/translator that is set to the peak predicted signal level for the upcoming pass. Then the output level of the receiver's video band spectrum envelope is adjusted to the level determined by equation (3) below (to five-sigma). Note that the SNR in the equation (2) is in dB while the SNR in equation (3) is linear. Pn = -198.6 + 10*log(SNT) + 10*log(1.2*Fbw) (1) SNR = Ps - Pn (SNR in dB) (2) Vrms = sqrt(SNR + 1)/[1 + 0.283*sqrt(SNR)] (SNR linear) (3) where Fbw = receiver filter bandwidth (Hz) Pn = receiver noise power (dBm) Ps = signal power (dBm) SNT = system noise temperature (K) SNR = predicted signal-to-noise ratio Operational Considerations - DSN ================================ The DSN is a complex and dynamic 'instrument.' Its performance for Radio Science depends on a number of factors from equipment configuration to meteorological conditions. No specific information on 'operational considerations' can be given here. Operational Modes - DSN ======================= DSCC Antenna Mechanical Subsystem --------------------------------- Pointing of DSCC antennas may be carried out in several ways. For details see the subsection 'DSCC Antenna Mechanical Subsystem' in the 'Subsystem' section. Binary pointing is the preferred mode for tracking spacecraft; pointing predicts are provided, and the antenna simply follows those. With CONSCAN, the antenna scans conically about the optimum pointing direction, using closed-loop receiver signal strength estimates as feedback. In planetary mode, the system interpolates from three (slowly changing) RA-DEC target coordinates; this is 'blind' pointing since there is no feedback from a detected signal. In sidereal mode, the antenna tracks a fixed point on the celestial sphere. In 'precision' mode, the antenna pointing is adjusted using an optical feedback system. It is possible on most antennas to freeze z-axis motion of the subreflector to minimize phase changes in the received signal. DSCC Receiver-Exciter Subsystem ------------------------------- The diplexer in the signal path between the transmitter and the feed horns on all three antennas may be configured so that it is out of the received signal path in order to improve the signal-to-noise ratio in the receiver system. This is known as the 'listen-only' or 'bypass' mode. Closed-Loop vs. Open-Loop Reception ----------------------------------- Radio Science data can be collected in two modes: closed- loop, in which a phase-locked loop receiver tracks the spacecraft signal, or open-loop, in which a receiver samples and records a band within which the desired signal presumably resides. Closed-loop data are collected using Closed-Loop Receivers, and open-loop data are collected using Open-Loop Receivers in conjunction with the DSCC Spectrum Processing Subsystem (DSP). See the Subsystems section for further information. Closed-Loop Receiver AGC Loop ----------------------------- The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. Ordinarily it is configured so that expected signal amplitude changes are accommodated with minimum distortion. The loop bandwidth is ordinarily configured so that expected phase changes can be accommodated while maintaining the best possible loop SNR. Coherent vs. Non-Coherent Operation ----------------------------------- The frequency of the signal transmitted from the spacecraft can generally be controlled in two ways -- by locking to a signal received from a ground station or by locking to an on-board oscillator. These are known as the coherent (or 'two-way') and non-coherent ('one-way') modes, respectively. Mode selection is made at the spacecraft, based on commands received from the ground. When operating in the coherent mode, the transponder carrier frequency is derived from the received uplink carrier frequency with a 'turn-around ratio' typically of 240/221. In the non-coherent mode, the downlink carrier frequency is derived from the spacecraft on-board crystal-controlled oscillator. Either closed-loop or open-loop receivers (or both) can be used with either spacecraft frequency reference mode. Closed-loop reception in two-way mode is usually preferred for routine tracking. Occasionally the spacecraft operates coherently while two ground stations receive the 'downlink' signal; this is sometimes known as the 'three-way' mode. DSCC Spectrum Processing Subsystem (DSP) ---------------------------------------- The DSP can operate in four sampling modes with from 1 to 4 input signals. Input channels are assigned to ADC inputs during DSP configuration. Modes and sampling rates are summarized in the tables below: Mode Analog-to-Digital Operation ---- ---------------------------- 1 4 signals, each sampled by a single ADC 2 1 signal, sampled sequentially by 4 ADCs 3 2 signals, each sampled sequentially by 2 ADCs 4 2 signals, the first sampled by ADC #1 and the second sampled sequentially at 3 times the rate by ADCs #2-4 8-bit Samples 12-bit Samples Sampling Rates Sampling Rates (samples/sec per ADC) (samples/sec per ADC) --------------------- --------------------- 50000 31250 25000 15625 12500 10000 10000 6250 5000 5000 4000 3125 2500 2000 1250 1000 1000 500 400 250 200 200 Input to each ADC is identified in header records by a Signal Channel Number (J1 - J4). Nominal channel assignments are shown below. Signal Channel Number Receiver (70-m or HEF) (34-m STD) --------------------- ------------- ---------- J1 X-RCP not used J2 S-RCP not used J3 X-LCP X-RCP J4 S-LCP S-RCP Location - DSN ============== Station locations are documented in [GEO-10REVD]. Geocentric coordinates are summarized here. Geocentric Geocentric Geocentric Station Radius (km) Latitude (N) Longitude (E) --------- ----------- ------------ ------------- Goldstone DSS 12 (34-m STD) 6371.997815 35.1186672 243.1945048 DSS 13 (develop) 6372.117062 35.0665485 243.2051077 DSS 14 (70-m) 6371.992867 35.2443514 243.1104584 DSS 15 (34-m HEF) 6371.9463 35.2402863 243.1128186 DSS 16 (26-m) 6371.9608 35.1601436 243.1264200 DSS 18 (34-m STD) UNK UNK UNK Canberra DSS 42 (34-m STD) 6371.675607 -35.2191850 148.9812546 DSS 43 (70-m) 6371.688953 -35.2209308 148.9812540 DSS 45 (34-m HEF) 6371.692 -35.21709 148.97757 DSS 46 (26-m) 6371.675 -35.22360 148.98297 DSS 48 (34-m STD) UNK UNK UNK Madrid DSS 61 (34-m STD) 6370.027734 40.2388805 355.7509634 DSS 63 (70-m) 6370.051015 40.2413495 355.7519776 DSS 65 (34-m HEF) 6370.021370 40.2372843 355.7485968 DSS 66 (26-m) 6370.036 40.2400714 355.7485976 Measurement Parameters - DSN ============================ Open-Loop System ---------------- Output from the Open-Loop Receivers (OLRs), as sampled and recorded by the DSCC Spectrum Processing Subsystem (DSP), is a stream of 8- or 12-bit quantized voltage samples. The nominal input to the Analog-to-Digital Converters (ADCs) is +/-10 volts, but the precise scaling between input voltages and output digitized samples is usually irrelevant for analysis; the digital data are generally referenced to a known noise or signal level within the data stream itself -- for example, the thermal noise output of the radio receivers which has a known system noise temperature (SNT). Raw samples comprise the data block in each DSP record; a header record (presently 83 16-bit words) contains ancillary information such as: time tag for the first sample in the data block RMS values of receiver signal levels and ADC outputs POCA frequency and drift rate Closed-Loop System ------------------ Closed-loop data are recorded in Archival Tracking Data Files (ATDFs), as well as certain secondary products such as the Orbit Data File (ODF). The ATDF Tracking Logical Record contains 117 entries including status information and measurements of ranging, Doppler, and signal strength. ACRONYMS AND ABBREVIATIONS - DSN ================================ ACS Antenna Control System ADC Analog-to-Digital Converter AMS Antenna Microwave System APA Antenna Pointing Assembly ARA Area Routing Assembly ATDF Archival Tracking Data File AZ Azimuth CMC Complex Monitor and Control CONSCAN Conical Scanning (antenna pointing mode) CRG Coherent Reference Generator CUL Clean-up Loop DANA a type of frequency synthesizer dB deciBel dBi dB relative to isotropic dBm dB relative to one milliwatt DCO Digitally Controlled Oscillator DEC Declination deg degree DMC DSCC Monitor and Control Subsystem DSCC Deep Space Communications Complex DSN Deep Space Network DSP DSCC Spectrum Processing Subsystem DSS Deep Space Station DTK DSCC Tracking Subsystem E east EL Elevation FTS Frequency and Timing Subsystem GCF Ground Communications Facility GPS Global Positioning System HA Hour Angle HEF High-Efficiency (as in 34-m HEF antennas) IF Intermediate Frequency IVC IF Selection Switch JPL Jet Propulsion Laboratory K Kelvin km kilometer kW kilowatt L-band approximately 1668 MHz LAN Local Area Network LCP Left-Circularly Polarized LMC Link Monitor and Control LNA Low-Noise Amplifier LO Local Oscillator m meters MCA Master Clock Assembly MCCC Mission Control and Computing Center MDA Metric Data Assembly MHz Megahertz MMR Multi-Mission Receiver MON Monitor and Control System MSA Mission Support Area N north NAR Noise Adding Radiometer NBOC Narrow-Band Occultation Converter NIST SPC 10 time relative to UTC NIU Network Interface Unit NOCC Network Operations and Control System NSS NOCC Support System OCI Operator Control Input ODF Orbit Data File ODR Original Data Record ODS Original Data Stream OLR Open Loop Receiver POCA Programmable Oscillator Control Assembly PPM Precision Power Monitor RA Right Ascension REC Receiver-Exciter Controller RCP Right-Circularly Polarized RF Radio Frequency RIC RIV Controller RIV Radio Science IF-VF Converter Assembly RMDCT Radio Metric Data Conditioning Team RTLT Round-Trip Light Time S-band approximately 2100-2300 MHz sec second SEC System Error Correction SIM Simulation SLE Signal Level Estimator SNR Signal-to-Noise Ratio SNT System Noise Temperature SOE Sequence of Events SPA Spectrum Processing Assembly SPC Signal Processing Center SRA Sequential Ranging Assembly SRC Sub-Reflector Controller SSI Spectral Signal Indicator STD Standard (as in 34-m STD antennas) TID Time Insertion and Distribution Assembly TSF Tracking Synthesizer Frequency TWM Traveling Wave Maser UNK unknown UTC Universal Coordinated Time VF Video Frequency X-band approximately 7800-8500 MHz
  • instrument : Radio Science Subsystem for Cassini Orbiter
    The Cassini Radio Science instrument included elements on both the spacecraft and the ground. The spacecraft element (covered by this context product) was further distributed among several subsystems on the Orbiter, while the ground element included NASA Deep Space Network (DSN) complexes in California, Australia, and Spain and occasional support from other ground stations such as antennas of the European Space Agency tracking network. The Radio Science 'instrument' operated in two fundamental modes. For 'two-way' measurements, the 'uplink' signal from the ground could be a single carrier at either X-band (7.2 GHz) or Ka-band (34 GHz); or both carriers could be transmitted at the same time. The spacecraft radio equipment then acted as a repeater, collecting the carrier signal with the spacecraft High Gain Antenna (HGA) and transforming it to one or more 'downlink' frequencies (2.3 GHz, 8.4 GHz, or 32 GHz) by elements in the Radio Frequency Subsystem (RFS) and Radio Frequency Instrument Subsystem (RFIS). For 'one-way' measurements, the signal source was on board the Cassini Orbiter; the output from an Ultrastable Oscillator (USO) was transformed to downlinks at 2.3, 8.4, or 32 GHz. The downlink signals were amplified and radiated through the HGA toward Earth. After passing through the medium of interest (e.g., plasma, rings, a neutral atmosphere or gravitationally curved space), the perturbed signal was collected by a ground antenna, amplified, down-converted, and recorded for later analysis.
  • instrument : RADIO SCIENCE SUBSYSTEM for ODY
    Instrument Overview =================== There were no recognized radio science investigations on the 2001 Mars Odyssey (ODY) mission. But investigators on Mars Global Surveyor (MGS) requested access to ODY radio tracking data. To support them and future proposers to Mars data analysis programs (MDAPs), the Planetary Data System (PDS) accepted responsibility for archiving the ODY data with initial activities funded jointly by MGS. Radio science investigations utilize instrumentation with elements both on a spacecraft and at ground stations -- in this case, at the NASA Deep Space Network (DSN). For ODY much of this was equipment used for routine telecommunications. The performance and calibration of both the spacecraft and tracking stations directly affected the radio science data accuracy, and they played a major role in determining the quality of the results. The spacecraft part of the radio science instrument is described immediately below; that is followed by a description of the DSN (ground) part of the instrument. For more information, see [MAKOVSKY2001]. Instrument Specifications - Spacecraft ====================================== The 2001 Mars Odyssey spacecraft telecommunications subsystem served as part of a radio science subsystem for investigations of Mars. Many details of the subsystem are unknown; its 'build date' is taken to be 2001-04-01, which was near the end of the Prelaunch Phase of the ODY mission. Instrument Id : RSS Instrument Host Id : ODY Pi Pds User Id : UNK Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : 2001-04-01 Instrument Mass : UNK Instrument Length : UNK Instrument Width : UNK Instrument Height : UNK Instrument Manufacturer Name : UNK Instrument Overview - Spacecraft ================================ The spacecraft radio system was constructed around a redundant pair of X-band Small Deep Space Transponders (SDSTs). Other components included one low-gain receive antenna (LGA); one medium-gain transmit antenna (MGA); one steerable high-gain antenna (HGA) for both transmitting (Tx) and receiving (Rx); two redundant solid state power amplifiers (SSPAs); a diplexer; several switches; and cabling. The SDSTs were connected to redundant Command and Data Handling (C&DH) units in such a way that any pairing could be chosen. A functional block diagram is shown below. . . . . . . . . DIPLEXER . . --- . / . ---------------------| |------------/ . | --- . \ HGA . | _ BPF1 | . \ . ----------|_| | . . | | / \ | . / . | |_____| S1 |________________________/ . | \ / . \ MGA . | \_/ | . \ . | |_| | . . | | BPF2 | . . |\| |\| | . HGA ASSEMBLY. | \ | \ | . . . . . . . . | / SSPA_1 | / SSPA_2 | / |/| |/| | / | | | HGA GIMBAL | | | ASSEMBLY --------------- | | 3 dB HYBRID | | | COUPLER | | --------------- | | | | | | | ------ ------ | |SDST_1| |SDST_2| | ------ ------ | | \ / | | | X | | | / \ | - ------ ------ _ / \ _ NF1 ---|C&DH_A| |C&DH_B|-------|_| S2 |_|-- | ------ ------ NF2 \ / | | - | | | | / | ----------/ | | \ LGA ---------------------------------------- \ S1 was a waveguide transfer switch with positions:1. SSPA_1 to HGA and SSPA_2 to MGA2. SSPA_1 to MGA and SSPA_2 to HGA (Insertion loss, <0.05 dB) S2 was a coaxial transfer switch with positions:1. SDST_1 to LGA and SDST_2 to HGA2. SDST_1 to HGA and SDST_2 to LGA (insertion loss, <0.3 dB) BPF_1 and BPF_2 were bandpass filters (<0.2 dB insertion loss over 8400-8450 MHz) NF1 and NF2 were notch filters (>70 dB rejection over 8400-8450 MHz) The X-Band Diplexer insertion loss was 0.1 DB (Tx), 0.2 dB (Rx) End-to-end circuit losses are given in the following table: ======================================================== | Link/Direction | Elements | Value | +------------------+---------------+-------------------+ | X-Band Transmit | SSPA to HGA | -0.25 +/- 0.11 dB | +------------------+---------------+-------------------+ | X-Band Transmit | SSPA to MGA | -0.52 +/- 0.35 dB | +------------------+---------------+-------------------+ | X-Band Receive | HGA to SDST | -8.13 +/- 0.03 dB | +------------------+---------------+-------------------+ | X-Band Receive | LGA to SDST | -2.43 +/- 0.02 dB | ======================================================== SSPA output power design was for 15 W (41.8 dBm) at end of life. The ODY telecommunications system was designed to perform the following functions: 1) Receive an X-band uplink carrier from a DSN station and demodulate the command data and ranging signal if either were present; 2) Generate an X-band downlink carrier either by coherently multiplying the frequency of the uplink carrier by the turn-around ratio of 880/749 or by utilizing an auxiliary crystal oscillator (AUX OSC); 3) Phase modulate the downlink carrier with either (or both) of the following: a composite telemetry signal, consisting of a square wave subcarrier (25 kHz or 375 kHz) that was BPSK (binary phase shift keying) modulated by telemetry data provided by the C&DH subsystem; the ranging signal that was demodulated from the uplink (this is referred to as two-way, or turn-around, ranging); 4) Permit control of the telecom subsystem through commands to select signal routing and the operational mode of the subsystem either from the ground or from command sequences previously loaded on the spacecraft; 5) Provide telecom status for monitoring operating conditions of the subsystem; 6) Provide ON/OFF power control for all RF transmitters; 7) Assume a single well-defined operating mode (a known baseline state) after a Power-On-Reset (POR). The X-band capability reduced plasma effects on radio signals by a factor of 10 compared with older S-band systems, but absence of a dual-frequency capability (both S- and X-band) meant that plasma effects could not be estimated and removed from radio data. The spacecraft also carried redundant ultra-high frequency (UHF) transceivers for communication and relay with future missions. Since the UHF equipment was not used for radio science, it is not described here. Science Objectives ================== There were no radio science objectives for the 2001 Mars Odyssey mission. The radio tracking data could be used by others to improve knowledge of the Mars gravity field . Operational Considerations - Spacecraft ======================================= Descriptions given here are for nominal performance. The spacecraft transponder system comprised redundant units, each with slightly different characteristics. As transponder units age, their performance changes slightly. More importantly, the performance for radio science depended on operational factors such as the modulation state for the transmitters, which cannot be predicted in advance. The performance also depended on factors which were not always under the control of the 2001 Mars Odyssey Project. The telecom subsystem relied on C&DH to control its operating mode; that control could be done via real-time commands from the ground or via a stored sequence onboard the spacecraft. The only exception was the POR state, which would be entered directly after a Power-On-Reset. C&DH provided the data to be downlinked, it carried out the frame and packet formatting and the Reed-Solomon encoding, and it provided the clock to drive the encoding. The clock was either data clock X 2 for (7,1/2) encoding or data clock X 6 for (15,1/6) encoding C&DH also handled error control for the uplink data stream. Calibration Description - Spacecraft ==================================== All measurements below were made during the Prelaunch Phase of The mission. Antenna characteristics are listed below. Masses of MGA and HGA are combined. Gain and axial ratio are given for boresight. Beamwidth is between the 3 dB points. ========================================================= | Antenna Characteristics - 2001 Mars Odyssey | +-----------------+---------+--------+--------+---------+ | | MGA | HGA | LGA | | Parameter +---------+--------+--------+---------| | | Tx Only | Tx | Rx | Rx Only | +-----------------+---------+--------+--------+---------+ |Frequency (MHz) | 8406.851852 | 7155.377315 | +-----------------+---------+--------+--------+---------+ |Diameter (m) | N/A | 1.3 | N/A | +-----------------+---------+--------+--------+---------+ |Mass (kg) | 3.150 | 0.040 | +-----------------+---------+--------+--------+---------+ |Gain (dBi) | 16.5 | 38.3 | 36.6 | 7+/-4 | +-----------------+---------+--------+--------+---------+ |Axial Ratio (dB) | N/A | 1.35 | 1.24 | 3 | +-----------------+---------+--------+--------+---------+ |Beamwidth (deg) | 28 | 1.9 | 2.3 | 82 | ========================================================= Receiver Carrier Loop characteristics were as follows: ========================================================= | Parameter | Value | +------------------+------------------------------------+ |Noise Figure | 2.70 +0.60/-0.73 dB averaged over | | | lifetime aging, temperature,| | | and radiation | +------------------+------------------------------------| |Tracking | -155 to -156 dBm | | Threshold | | +------------------+------------------------------------| |Tracking Rate | 200 Hz/s for uplink Pt <= -120 dBm | +------------------+------------------------------------| |Capture Range | +/-1.3 kHz | +------------------+------------------------------------| |Tracking Range | +100 kHz/-200 kHz relative to best | | | lock frequency | +------------------+------------------------------------| |Carrier Loop | 20 Hz | | Threshold | | | Bandwidth | | +------------------+------------------------------------| |Strong Signal Open| 2.0e+07 | | Loop Gain | | +------------------+------------------------------------| |Predetection Noise| 12500 Hz | | Bandwidth | | +------------------+------------------------------------| |Loop Pole Time | 2258.6 s | | Constant | | +------------------+------------------------------------| |Loop Zero Time | 0.050 s | | Constant | | +------------------+------------------------------------| |Strong Signal Loop| 231.306 Hz two-sided at Pc/No = | | Noise Bandwidth | 100 dB-Hz | ========================================================= The SDST ranging performance is listed in the table below. One range unit was 0.947 nanoseconds for 2001 Mars Odyssey. ========================================================== | Parameter | Value (average over 3 devices) | +---------------------+----------------------------------+ |Range Delay | 1417.2 range units | +---------------------+----------------------------------+ |Temperature Variation| +/-4.0 ru (-25C to +30C) | +---------------------+----------------------------------+ |Carrier Suppression | 0.5 dB (17.5 deg range mod index)| | | 1.9 dB (35.0 deg range mod index)| +---------------------+----------------------------------+ |3 dB Bandwidth | 1.4 MHz | +---------------------+----------------------------------+ |Noise Equivalent | 2.0 MHz | | Bandwidth | | ========================================================== Platform Mounting Descriptions - Spacecraft =========================================== During the Launch, Cruise, Orbit Insertion, and Aerobraking phases of the mission, the HGA was stowed so that its boresight and the MGA boresight were along the +X axis. After aerobraking, the HGA was deployed and tracked the Earth using a pair of gimbals (azimuth and elevation) at the end of a boom. The MGA was mounted on the HGA dish so that the MGA and HGA boresights were equal. The SSPAs were mounted behind the HGA reflector to minimize circuit losses. Investigators ============= None. Instrument Section / Operating Mode Descriptions - Spacecraft ============================================================= Redundant components could be configured as desired. Each configuration had slightly different performance, but the quantitative differences are unknown. Instrument Overview - DSN ========================= Three Deep Space Communications Complexes (DSCCs) (near Barstow, CA; Canberra, Australia; and Madrid, Spain) comprise the DSN tracking network. Each complex is equipped with several antennas [including at least one each 70-m, 34-m High Efficiency (HEF), and 34-m Beam WaveGuide (BWG)], associated electronics, and operational systems. Primary activity at each complex is radiation of commands to and reception of telemetry data from active spacecraft. Transmission and reception is possible in several radio-frequency bands, the most common being S-band (nominally a frequency of 2100-2300 MHz or a wavelength of 14.2-13.0 cm) and X-band (7100-8500 MHz or 4.2-3.5 cm). Transmitter output powers of up to 400 kW are available. Ground stations have the ability to transmit coded and uncoded waveforms which can be echoed by distant spacecraft. Analysis of the received coding allows navigators to determine the distance to the spacecraft; analysis of Doppler shift on the carrier signal allows estimation of the line-of-sight spacecraft velocity. Range and Doppler measurements are used to calculate the spacecraft trajectory and to infer gravity fields of objects near the spacecraft. Ground stations can record spacecraft signals that have propagated through or been scattered from target media. Measurements of signal parameters after wave interactions with surfaces, atmospheres, rings, and plasmas are used to infer physical and electrical properties of the target. Principal investigators vary from experiment to experiment. See the corresponding section of the spacecraft instrument description or the data set description for specifics. The Deep Space Network is managed by the Jet Propulsion Laboratory of the California Institute of Technology for the U.S. National Aeronautics and Space Administration. Specifications include: Instrument Id : RSS Instrument Host Id : DSN Pi Pds User Id : N/A Instrument Name : RADIO SCIENCE SUBSYSTEM Instrument Type : RADIO SCIENCE Build Date : N/A Instrument Mass : N/A Instrument Length : N/A Instrument Width : N/A Instrument Height : N/A Instrument Manufacturer Name : N/A For more information on the Deep Space Network and its use in radio science see reports by [ASMAR&RENZETTI1993], [ASMAR&HERRERA1993], and [ASMARETAL1995]. For design specifications on DSN subsystems see [DSN810-5]. For DSN use with MGS Radio Science see [TYLERETAL1992], [TYLERETAL2001], and [JPLD-14027]. Subsystems - DSN ================ The Deep Space Communications Complexes (DSCCs) are an integral part of Radio Science instrumentation, along with the spacecraft Radio Frequency Subsystem. Their system performance directly determines the degree of success of Radio Science investigations, and their system calibration determines the degree of accuracy in the results of the experiments. The following paragraphs describe the functions performed by the individual subsystems of a DSCC. This material has been adapted from [ASMAR&HERRERA1993] and [JPLD-14027]; for additional information, consult [DSN810-5]. Each DSCC includes a set of antennas, a Signal Processing Center (SPC), and communication links to the Jet Propulsion Laboratory (JPL). The general configuration is illustrated below; antennas (Deep Space Stations, or DSS -- a term carried over from earlier times when antennas were individually instrumented) are listed in the table. -------- -------- -------- -------- -------- | DSS 25 | | DSS 27 | | DSS 14 | | DSS 15 | | DSS 16 | |34-m BWG| |34-m HSB| | 70-m | |34-m HEF| | 26-m | -------- -------- -------- -------- -------- | | | | | | v v | v | --------- | --------- --------->|GOLDSTONE|<---------- |EARTH/ORB| | SPC 10 |<-------------->| LINK | --------- --------- | SPC |<-------------->| 26-M | | COMM | ------>| COMM | --------- | --------- | | | v | v ------ --------- | --------- | NOCC |<--->| JPL |<------- | | ------ | CENTRAL | | GSFC | ------ | COMM | | NASCOMM | | MCCC |<--->| TERMINAL|<-------------->| | ------ --------- --------- ^ ^ | | CANBERRA (SPC 40) <---------------- | | MADRID (SPC 60) <---------------------- GOLDSTONE CANBERRA MADRID Antenna SPC 10 SPC 40 SPC 60 -------- --------- -------- -------- 26-m DSS 16 DSS 46 DSS 66 34-m HEF DSS 15 DSS 45 DSS 65 34-m BWG DSS 24 DSS 34 DSS 54 DSS 25 DSS 26 34-m HSB DSS 27 DSS 28 70-m DSS 14 DSS 43 DSS 63 Developmental DSS 13 Subsystem interconnections at each DSCC are shown in the diagram below, and they are described in the sections that follow. The Monitor and Control Subsystem is connected to all other subsystems; the Test Support Subsystem can be. ----------- ------------------ --------- --------- |TRANSMITTER| | | | TRACKING| | COMMAND | | SUBSYSTEM |-| RECEIVER/EXCITER |-|SUBSYSTEM|-|SUBSYSTEM|- ----------- | | --------- --------- | | | SUBSYSTEM | | | | ----------- | | --------------------- | | MICROWAVE | | | | TELEMETRY | | | SUBSYSTEM |-| |-| SUBSYSTEM |- ----------- ------------------ --------------------- | | | ----------- ----------- --------- -------------- | | ANTENNA | | MONITOR | | TEST | | DIGITAL | | | SUBSYSTEM | |AND CONTROL| | SUPPORT | |COMMUNICATIONS|- ----------- | SUBSYSTEM | |SUBSYSTEM| | SUBSYSTEM | ----------- --------- -------------- DSCC Monitor and Control Subsystem ---------------------------------- The DSCC Monitor and Control Subsystem (DMC) is part of the Monitor and Control System (MON) which also includes the ground communications Central Communications Terminal and the Network Operations Control Center (NOCC) Monitor and Control Subsystem. The DMC is the center of activity at a DSCC. The DMC receives and archives most of the information from the NOCC needed by the various DSCC subsystems during their operation. Control of most of the DSCC subsystems, as well as the handling and displaying of any responses to control directives and configuration and status information received from each of the subsystems, is done through the DMC. The effect of this is to centralize the control, display, and archiving functions necessary to operate a DSCC. Communication among the various subsystems is done using a Local Area Network (LAN) hooked up to each subsystem via a network interface unit (NIU). DSCC Antenna Mechanical Subsystem --------------------------------- Multi-mission Radio Science activities require support from the 70-m, 34-m HEF, and 34-m BWG antenna subnets. The antennas at each DSCC function as large-aperture collectors which, by double reflection, cause the incoming radio frequency (RF) energy to enter the feed horns. The large collecting surface of the antenna focuses the incoming energy onto a subreflector, which is adjustable in both axial and angular position. These adjustments are made to correct for gravitational deformation of the antenna as it moves between zenith and the horizon; the deformation can be as large as 5 cm. The subreflector adjustments optimize the channeling of energy from the primary reflector to the subreflector and then to the feed horns. The 70-m and 34-m HEF antennas have 'shaped' primary and secondary reflectors, with forms that are modified paraboloids. This customization allows more uniform illumination of one reflector by another. The BWG reflector shape is ellipsoidal. On the 70-m antennas, the subreflector directs received energy from the antenna onto a dichroic plate, a device which reflects S-band energy to the S-band feed horn and passes X-band energy through to the X-band feed horn. In the 34-m HEF, there is one 'common aperture feed,' which accepts both frequencies without requiring a dichroic plate. In the 34-m BWG, a series of small mirrors (approximately 2.5 meters in diameter) directs microwave energy from the subreflector region to a collection area at the base of the antenna -- typically in a pedestal room. A retractable dichroic reflector separates S- and X-band on some BWG antennas or X- and Ka-band on others. RF energy to be transmitted into space by the horns is focused by the reflectors into narrow cylindrical beams, pointed with high precision (either to the dichroic plate or directly to the subreflector) by a series of drive motors and gear trains that can rotate the movable components and their support structures. The different antennas can be pointed by several means. Two pointing modes commonly used during tracking passes are CONSCAN and 'blind pointing.' With CONSCAN enabled and a closed loop receiver locked to a spacecraft signal, the system tracks the radio source by conically scanning around its position in the sky. Pointing angle adjustments are computed from signal strength information (feedback) supplied by the receiver. In this mode the Antenna Pointing Assembly (APA) generates a circular scan pattern which is sent to the Antenna Control System (ACS). The ACS adds the scan pattern to the corrected pointing angle predicts. Software in the receiver-exciter controller computes the received signal level and sends it to the APA. The correlation of scan position with the received signal level variations allows the APA to compute offset changes which are sent to the ACS. Thus, within the capability of the closed-loop control system, the scan center is pointed precisely at the apparent direction of the spacecraft signal source. An additional function of the APA is to provide antenna position angles and residuals, antenna control mode/status information, and predict-correction parameters to the Area Routing Assembly (ARA) via the LAN, which then sends this information to JPL via the Ground Communications Facility (GCF) for antenna status monitoring. During periods when excessive signal level dynamics or low received signal levels are expected (e.g., during an occultation experiment), CONSCAN should not be used. Under these conditions, blind pointing (CONSCAN OFF) is used, and pointing angle adjustments are based on a predetermined Systematic Error Correction (SEC) model. Independent of CONSCAN state, subreflector motion in at least the z-axis may introduce phase variations into the received Radio Science data. For that reason, during certain experiments, the subreflector in the 70-m and 34-m HEFs may be frozen in the z-axis at a position (often based on elevation angle) selected to minimize phase change and signal degradation. This can be done via Operator Control Inputs (OCIs) from the LMC to the Subreflector Controller (SRC) which resides in the alidade room of the antennas. The SRC passes the commands to motors that drive the subreflector to the desired position. Pointing angles for all antenna types are computed by the NOCC Support System (NSS) from an ephemeris provided by the flight project. These predicts are received and archived by the CMC. Before each track, they are transferred to the APA, which transforms the direction cosines of the predicts into AZ-EL coordinates. The LMC operator then downloads the antenna predict points to the antenna-mounted ACS computer along with a selected SEC model. The pointing predicts consist of time-tagged AZ-EL points at selected time intervals along with polynomial coefficients for interpolation between points. The ACS automatically interpolates the predict points, corrects the pointing predicts for refraction and subreflector position, and adds the proper systematic error correction and any manually entered antenna offsets. The ACS then sends angular position commands for each axis at the rate of one per second. In the 70-m and 34-m HEF, rate commands are generated from the position commands at the servo controller and are subsequently used to steer the antenna. When not using binary predicts (the routine mode for spacecraft tracking), the antennas can be pointed using 'planetary mode' -- a simpler mode which uses right ascension (RA) and declination (DEC) values. These change very slowly with respect to the celestial frame. Values are provided to the station in text form for manual entry. The ACS quadratically interpolates among three RA and DEC points which are on one-day centers. A third pointing mode -- sidereal -- is available for tracking radio sources fixed with respect to the celestial frame. Regardless of the pointing mode being used, a 70-m antenna has a special high-accuracy pointing capability called 'precision' mode. A pointing control loop derives the main AZ-EL pointing servo drive error signals from a two- axis autocollimator mounted on the Intermediate Reference Structure. The autocollimator projects a light beam to a precision mirror mounted on the Master Equatorial drive system, a much smaller structure, independent of the main antenna, which is exactly positioned in HA and DEC with shaft encoders. The autocollimator detects elevation/cross- elevation errors between the two reference surfaces by measuring the angular displacement of the reflected light beam. This error is compensated for in the antenna servo by moving the antenna in the appropriate AZ-EL direction. Pointing accuracies of 0.004 degrees (15 arc seconds) are possible in 'precision' mode. The 'precision' mode is not available on 34-m antennas -- nor is it needed, since their beamwidths are twice as large as on the 70-m antennas. DSCC Antenna Microwave Subsystem -------------------------------- 70-m Antennas: Each 70-m antenna has three feed cones installed in a structure at the center of the main reflector. The feeds are positioned 120 degrees apart on a circle. Selection of the feed is made by rotation of the subreflector. A dichroic mirror assembly, half on the S-band cone and half on the X-band cone, permits simultaneous use of the S- and X-band frequencies. The third cone is devoted to R&D and more specialized work. The Antenna Microwave Subsystem (AMS) accepts the received S- and X-band signals at the feed horn and transmits them through polarizer plates to an orthomode transducer. The polarizer plates are adjusted so that the signals are directed to a pair of redundant amplifiers for each frequency, thus allowing simultaneous reception of signals in two orthogonal polarizations. For S-band these are two Block IVA S-band Traveling Wave Masers (TWMs); for X-band the amplifiers are Block IIA TWMs. 34-m HEF Antennas: The 34-m HEF uses a single feed for both S- and X-band. Simultaneous S- and X-band receive as well as X-band transmit is possible thanks to the presence of an S/X 'combiner' which acts as a diplexer. For S-band, RCP or LCP is user selected through a switch so neither a polarizer nor an orthomode transducer is needed. X-band amplification options include two Block II TWMs or an HEMT Low Noise Amplifier (LNA). S-band amplification is provided by an FET LNA. 34-m BWG Antennas: These antennas use feeds and low-noise amplifiers (LNA) in the pedestal room, which can be switched in and out as needed. Typically the following modes are available: 1. downlink non-diplexed path (RCP or LCP) to LNA-1, with uplink in the opposite circular polarization; 2. downlink non-diplexed path (RCP or LCP) to LNA-2, with uplink in the opposite circular polarization 3. downlink diplexed path (RCP or LCP) to LNA-1, with uplink in the same circular polarization 4. downlink diplexed path (RCP or LCP) to LNA-2, with uplink in the same circular polarization For BWG antennas with dual-band capabilities (e.g., DSS 25) and dual LNAs, each of the above four modes can be used in a single-frequency or dual-frequency configuration. Thus, for antennas with the most complete capabilities, there are sixteen possible ways to receive at a single frequency (2 polarizations, 2 waveguide path choices, 2 LNAs, and 2 bands). DSCC Receiver-Exciter Subsystem ------------------------------- The Receiver-Exciter Subsystem is composed of two groups of equipment: the closed-loop receiver group and the open-loop receiver group. This subsystem is controlled by the Receiver-Exciter Controller (REC) which communicates directly with the DMC for predicts and OCI reception and status reporting. The exciter generates the S-band signal (or X-band for the 34-m HEF only) which is provided to the Transmitter Subsystem for the spacecraft uplink signal. It is tunable under command of the Digitally Controlled Oscillator (DCO) which receives predicts from the Metric Data Assembly (MDA). The diplexer in the signal path between the transmitter and the feed horn for all three antennas (used for simultaneous transmission and reception) may be configured such that it is out of the received signal path (in listen-only or bypass mode) in order to improve the signal-to-noise ratio in the receiver system. Closed Loop Receivers: The Block V receiver-exciter at the 70-m stations allows for two receiver channels, each capable of L-Band (e.g., 1668 MHz frequency or 18 cm wavelength), S-band, or X-band reception, and an S-band exciter for generation of uplink signals through the low-power or high-power transmitter. The closed-loop receivers provide the capability for rapid acquisition of a spacecraft signal and telemetry lockup. In order to accomplish acquisition within a short time, the receivers are predict driven to search for, acquire, and track the downlink automatically. Rapid acquisition precludes manual tuning though that remains as a backup capability. The subsystem utilizes FFT analyzers for rapid acquisition. The predicts are NSS generated, transmitted to the CMC which sends them to the Receiver-Exciter Subsystem where two sets can be stored. The receiver starts acquisition at uplink time plus one round-trip-light-time or at operator specified times. The receivers may also be operated from the LMC without a local operator attending them. The receivers send performance and status data, displays, and event messages to the LMC. Either the exciter synthesizer signal or the simulation (SIM) synthesizer signal is used as the reference for the Doppler extractor in the closed-loop receiver systems, depending on the spacecraft being tracked (and Project guidelines). The SIM synthesizer is not ramped; instead it uses one constant frequency, the Track Synthesizer Frequency (TSF), which is an average frequency for the entire pass. The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. It will be configured such that the expected amplitude changes are accommodated with minimum distortion. The loop bandwidth (2BLo) will be configured such that the expected phase changes can be accommodated while maintaining the best possible loop SNR. Open-Loop Receivers (OLR): The OLR utilized a fixed first Local Oscillator (LO) frequency and a tunable second LO frequency to minimize phase noise and improve frequency stability. The OLR consisted of an RF-to-IF downconverter located at the feed , an IF selection switch (IFS), and a Radio Science Receiver (RSR). The RF-IF downconverters in the 70-m antennas were equipped for four IF channels: S-RCP, S-LCP, X-RCP, and X-LCP. The 34-m HEF stations were equipped with a two-channel RF-IF: S-band and X-band. The IFS switched the IF input among the antennas. DSCC Transmitter Subsystem -------------------------- The Transmitter Subsystem accepts the S-band frequency exciter signal from the Receiver-Exciter Subsystem exciter and amplifies it to the required transmit output level. The amplified signal is routed via the diplexer through the feed horn to the antenna and then focused and beamed to the spacecraft. The Transmitter Subsystem power capabilities range from 18 kW to 400 kW. Power levels above 18 kW are available only at 70-m stations. DSCC Tracking Subsystem ----------------------- The Tracking Subsystem primary functions are to acquire and maintain communications with the spacecraft and to generate and format radiometric data containing Doppler and range. The DSCC Tracking Subsystem (DTK) receives the carrier signals and ranging spectra from the Receiver-Exciter Subsystem. The Doppler cycle counts are counted, formatted, and transmitted to JPL in real time. Ranging data are also transmitted to JPL in real time. Also contained in these blocks is the AGC information from the Receiver-Exciter Subsystem. The Radio Metric Data Conditioning Team (RMDCT) at JPL produces an Archival Tracking Data File (ATDF) which contains Doppler and ranging data. In addition, the Tracking Subsystem receives from the CMC frequency predicts (used to compute frequency residuals and noise estimates), receiver tuning predicts (used to tune the closed-loop receivers), and uplink tuning predicts (used to tune the exciter). From the LMC, it receives configuration and control directives as well as configuration and status information on the transmitter, microwave, and frequency and timing subsystems. The Metric Data Assembly (MDA) controls all of the DTK functions supporting the uplink and downlink activities. The MDA receives uplink predicts and controls the uplink tuning by commanding the DCO. The MDA also controls the Sequential Ranging Assembly (SRA). It formats the Doppler and range measurements and provides them to the GCF for transmission to NOCC. The Sequential Ranging Assembly (SRA) measures the round trip light time (RTLT) of a radio signal traveling from a ground tracking station to a spacecraft and back. From the RTLT, phase, and Doppler data, the spacecraft range can be determined. A coded signal is modulated on an uplink carrier and transmitted to the spacecraft where it is detected and transponded back to the ground station. As a result, the signal received at the tracking station is delayed by its round trip through space and shifted in frequency by the Doppler effect due to the relative motion between the spacecraft and the tracking station on Earth. DSCC Frequency and Timing Subsystem ----------------------------------- The Frequency and Timing Subsystem (FTS) provides all frequency and timing references required by the other DSCC subsystems. It contains four frequency standards of which one is prime and the other three are backups. Selection of the prime standard is done via the CMC. Of these four standards, two are hydrogen masers followed by clean-up loops (CUL) and two are cesium standards. These four standards all feed the Coherent Reference Generator (CRG) which provides the frequency references used by the rest of the complex. It also provides the frequency reference to the Master Clock Assembly (MCA) which in turn provides time to the Time Insertion and Distribution Assembly (TID) which provides UTC and SIM-time to the complex. JPL's ability to monitor the FTS at each DSCC is limited to the MDA calculated Doppler pseudo-residuals, the Doppler noise, the SSI, and to a system which uses the Global Positioning System (GPS). GPS receivers at each DSCC receive a one-pulse-per-second pulse from the station's (hydrogen maser referenced) FTS and a pulse from a GPS satellite at scheduled times. After compensating for the satellite signal delay, the timing offset is reported to JPL where a database is kept. The clock offsets stored in the JPL database are given in microseconds; each entry is a mean reading of measurements from several GPS satellites and a time tag associated with the mean reading. The clock offsets provided include those of SPC 10 relative to UTC (NIST), SPC 40 relative to SPC 10, etc. Radio Science Receiver (RSR) ---------------------------- A radio frequency (RF) spacecraft signal at S-band, X-band, or Ka-band is captured by a receiving antenna on Earth, down converted to an intermediate frequency (IF) near 300 MHz and then fed via a distribution network to one input of an IF Selector Switch (IFS). The IFS allows each RSR to select any of the available input signals for its RSR Digitizer (DIG). Within the RSR the digitized signal is then passed to the Digitial Down Converter (DDC), VME Data Processor (VDP), and Data Processor (DP) [JPLD-16765]. \ ----------- ------ ----- ----- ----- \ | RF TO IF | | |----| | | | | | |----| DOWN |----| |----| |----| DIG | | DP | / | CONVERTER | | |----| | | | | | / ----------- | IF |----| IFS | ----- ----- ANTENNA --| DIST |----| | | | 300 MHz IF --| | .. | | ----- ----- FROM OTHER --| |----| | | | | | ANTENNAS --| | ----- | DDC | | VDP | ------ | | | | ----- ----- | | ------- In the DIG the IF signal is passed through a programmable attenuator, adjusted to provide the proper level to the Analog to Digital Converter (ADC). The attenuated signal is then passed through a Band Pass Filter (BPF) which selects a frequency band in the range 265-375 MHz. The filtered output from the BPF is then mixed with a 256 MHz Local Oscillator (LO), low pass filtered (LPF), and sampled by the ADC. The output of the ADC is a stream of 8-bit real samples at 256 Msamples/second (Msps). DIG timing is derived from the station FTS 5 MHz clock and 1 pulse per second (1PPS) reference; the DIG generates a 256 MHz clock signal for later processing. The 1PPS signal marks the data sample taken at the start of each second. The DDC selects one 16 MHz subchannel from the possible 128 MHz bandwidth available from the DIG by using Finite Impulse Response (FIR) filters with revolving banks of filter coefficients. The sample stream from the DIG is separated into eight decimated streams, each of which is fed into two sets of FIR filters. One set of filters produces in-phase (I) 8-bit data while the other produces quadrature-phase (Q) 8-bit data. The center frequency of the desired 16 MHz channel is adjustable in 1 MHz steps and is usually chosen to be near the spacecraft carrier frequency. After combining the I and Q sample streams, the DDC feeds the samples to the VDP. The DDC also converts the 256 MHz data clock and 1PPS signals into a msec time code, which is also passed to the VDP. The VDP contains a quadruply-redundant set of custom boards which are controlled by a real-time control computer (RT). Each set of boards comprises a numerically controlled oscillator (NCO), a complex multiplier, a decimating FIR filter, and a data packer. The 16 Msps complex samples from the DDC are digitally mixed with the NCO signal in the complex multiplier. The NCO phase and frequency are updated every millisecond by the RT and are selected so that the center frequency of the desired portion of the 16 MHz channel is down-converted to 0 Hz. The RT uses polynomials derived from frequency predictions. The output of the complex multiplier is sent to the decimating FIR filter where its bandwidth and sample rate are reduced (see table below). The decimating FIR filter also allows adjustment of the sub-channel gain to take full advantage of the dynamic range available in the hardware. The data packer truncates samples to 1, 2, 4, 8, or 16 bits by dropping the least significant bits and packs them into 32-bit data words. Q-samples are packed into the first 16 bits of the word, and I-samples into the least significant 16 bits (see below). In 'narrow band' operation all four sets of sets of custom boards can be supported simultaneously. In 'medium band' operation no more than two channels can be supported simultaneously. In 'wide band' operation, only one sub-channel can be recorded. |============================================================| | RSR Sample Rates and Sample Sizes Supported | |================+=======+======+=================+==========| | Category | Rate | Size | Data Rate |Rec Length| | | (ksps)|(bits)|(bytes/s) (rec/s)| (bytes) | |================+=======+======+=========+=======+==========| |Narrow Band (NB)| 1 | 8 | 2000 | 1 | 2000 | | | 2 | 8 | 4000 | 1 | 4000 | | | 4 | 8 | 8000 | 1 | 8000 | | | 8 | 8 | 16000 | 1 | 16000 | | | 16 | 8 | 32000 | 2 | 16000 | | | 25 | 8 | 50000 | 2 | 25000 | | | 50 | 8 | 100000 | 4 | 25000 | | | 100 | 8 | 200000 | 10 | 20000 | | | 1 | 16 | 4000 | 1 | 4000 | | | 2 | 16 | 8000 | 1 | 8000 | | | 4 | 16 | 16000 | 1 | 16000 | | | 8 | 16 | 32000 | 2 | 16000 | | | 16 | 16 | 64000 | 4 | 16000 | | | 25 | 16 | 100000 | 4 | 25000 | | | 50 | 16 | 200000 | 10 | 20000 | | | 100 | 16 | 400000 | 20 | 20000 | |Medium Band (MB)| 250 | 1 | 62500 | 5 | 12500 | | | 500 | 1 | 125000 | 5 | 25000 | | | 1000 | 1 | 250000 | 10 | 25000 | | | 2000 | 1 | 500000 | 20 | 25000 | | | 4000 | 1 | 1000000 | 40 | 25000 | | | 250 | 2 | 125000 | 5 | 25000 | | | 500 | 2 | 250000 | 10 | 25000 | | | 1000 | 2 | 500000 | 20 | 25000 | | | 2000 | 2 | 1000000 | 40 | 25000 | | | 4000 | 2 | 2000000 | 100 | 20000 | | | 250 | 4 | 250000 | 10 | 25000 | | | 500 | 4 | 500000 | 20 | 25000 | | | 1000 | 4 | 1000000 | 40 | 25000 | | | 2000 | 4 | 2000000 | 100 | 20000 | | | 250 | 8 | 500000 | 20 | 25000 | | | 500 | 8 | 1000000 | 40 | 25000 | | | 1000 | 8 | 2000000 | 100 | 20000 | |Wide Band (WB) | 8000 | 1 | 2000000 | 100 | 20000 | | | 16000 | 1 | 4000000 | 200 | 20000 | | | 8000 | 2 | 4000000 | 200 | 20000 | |============================================================| |============================================================| | Sample Packing | |=================+==========================================| | Bits per Sample | Contents of 32-bit Packed Data Register | |=================+==========================================| | 16 | (Q1),(I1) | | 8 | (Q2,Q1),(I2,I1) | | 4 | (Q4,Q3,Q2,Q1),(I4,I3,I2,I1) | | 2 | (Q8,Q7,...Q1),(I8,I7,...I1) | | 1 | (Q16,Q15,...Q1),(I16,I15,...I1) | |============================================================| Once per second the RT sends the accumulated data records from each sub-channel to the Data Processor (DP) over a 100 Mbit/s ethernet connection. In addition to the samples, each data record includes header information such as time tags and NCO frequency and phase that are necessary for analysis. The DP processes the data records to provide monitor data, such as power spectra. If recording has been enabled, the records are stored by the DP. NCO Phase and Frequency ----------------------- At the start of each DSN pass, the RSR is provided with a file containing a list of predicted frequencies. Using these points, the RT computes expected sky frequencies at the beginning, middle, and end of each one second time interval. Based on the local oscillator frequencies selected and any offsets entered, the RT computes the coefficients of a frequency polynomial fitted to the DDC channel frequencies at these three times. The RT also computes a phase polynomial by integrating the frequency polynomial and matching phases at the one second boundaries. The phase and frequency of the VDP NCO's are computed every millisecond (000-999) from the polynomial coefficients as follows: nco_phase(msec) = phase_coef_1 + phase_coef_2 * (msec/1000) + phase_coef_3 * (msec/1000)**2 + phase_coef_4 * (msec/1000)**3 nco_freq(msec) = freq_coef_1 + freq_coef_2 * ((msec + 0.5)/1000) + freq_coef_3 * ((msec + 0.5)/1000)**2 The sky frequency may be reconstructed using sky_freq = RF_to_IF_LO + DDC_LO - nco_freq + reside_freq where RF_to_IF_LO is the down conversion from the microwave frequency to IF (bytes 42-43 in the data record header) DDC_LO is the down-conversion applied in the DIG and DDC (bytes 40-41 in the data record header) Resid_Freq is the frequency of the signal in the VDP output Detectors - DSN =============== Nominal carrier tracking loop threshold noise bandwidth at X-band is 10 Hz. Coherent (two-way) closed-loop system stability is shown in the table below: integration time Doppler uncertainty (secs) (one sigma, microns/sec) ------ ------------------------ 10 50 60 20 1000 4 For the open-loop subsystem, signal detection is done in software. Calibration - DSN ================= Calibrations of hardware systems are carried out periodically by DSN personnel; these ensure that systems operate at required performance levels -- for example, that antenna patterns, receiver gain, propagation delays, and Doppler uncertainties meet specifications. No information on specific calibration activities is available. Nominal performance specifications are shown in the tables above. Additional information may be available in [DSN810-5]. Prior to each tracking pass, station operators perform a series of calibrations to ensure that systems meet specifications for that operational period. Included in these calibrations is measurement of receiver system temperature in the configuration to be employed during the pass. Results of these calibrations are recorded in (hard copy) Controller's Logs for each pass. The nominal procedure for initializing open-loop receiver attenuator settings is described below. In cases where widely varying signal levels are expected, the procedure may be modified in advance or real-time adjustments may be made to attenuator settings. Operational Considerations - DSN ================================ The DSN is a complex and dynamic 'instrument.' Its performance for Radio Science depends on a number of factors from equipment configuration to meteorological conditions. No specific information on 'operational considerations' can be given here. Operational Modes - DSN ======================= DSCC Antenna Mechanical Subsystem --------------------------------- Pointing of DSCC antennas may be carried out in several ways. For details see the subsection 'DSCC Antenna Mechanical Subsystem' in the 'Subsystem' section. Binary pointing is the preferred mode for tracking spacecraft; pointing predicts are provided, and the antenna simply follows those. With CONSCAN, the antenna scans conically about the optimum pointing direction, using closed-loop receiver signal strength estimates as feedback. In planetary mode, the system interpolates from three (slowly changing) RA-DEC target coordinates; this is 'blind' pointing since there is no feedback from a detected signal. In sidereal mode, the antenna tracks a fixed point on the celestial sphere. In 'precision' mode, the antenna pointing is adjusted using an optical feedback system. It is possible on most antennas to freeze z-axis motion of the subreflector to minimize phase changes in the received signal. DSCC Receiver-Exciter Subsystem ------------------------------- The diplexer in the signal path between the transmitter and the feed horns on all antennas may be configured so that it is out of the received signal path in order to improve the signal-to-noise ratio in the receiver system. This is known as the 'listen-only' or 'bypass' mode. Closed-Loop Receiver AGC Loop ----------------------------- The closed-loop receiver AGC loop can be configured to one of three settings: narrow, medium, or wide. Ordinarily it is configured so that expected signal amplitude changes are accommodated with minimum distortion. The loop bandwidth is ordinarily configured so that expected phase changes can be accommodated while maintaining the best possible loop SNR. Coherent vs. Non-Coherent Operation ----------------------------------- The frequency of the signal transmitted from the spacecraft can generally be controlled in two ways -- by locking to a signal received from a ground station or by locking to an on-board oscillator. These are known as the coherent (or 'two-way') and non-coherent ('one-way') modes, respectively. Mode selection is made at the spacecraft, based on commands received from the ground. When operating in the coherent mode, the transponder carrier frequency is derived from the received uplink carrier frequency with a 'turn-around ratio' typically of 880/749. In the non-coherent mode, the downlink carrier frequency is derived from the spacecraft on-board crystal-controlled oscillator. Either closed-loop or open-loop receivers (or both) can be used with either spacecraft frequency reference mode. Closed-loop reception in two-way mode is usually preferred for routine tracking. Occasionally the spacecraft operates coherently while two ground stations receive the 'downlink' signal; this is sometimes known as the 'three-way' mode. Location - DSN ============== Station locations are documented in [GEO-10REVD]. Geocentric coordinates are summarized here. Geocentric Geocentric Geocentric Station Radius (km) Latitude (N) Longitude (E) --------- ----------- ------------ ------------- Goldstone DSS 13 (34-m R&D) 6372.125125 35.0660185 243.2055430 DSS 14 (70-m) 6371.993286 35.2443527 243.1104638 DSS 15 (34-m HEF) 6371.966540 35.2403133 243.1128069 DSS 24 (34-m BWG) 6371.973553 35.1585349 243.1252079 DSS 25 (34-m BWG) 6371.983060 35.1562594 243.1246384 DSS 26 (34-m BWG) 6371.993032 35.1543411 243.1269849 Canberra DSS 34 (34-m BWG) 6371.693561 -35.2169868 148.9819620 DSS 43 (70-m) 6371.689033 -35.2209234 148.9812650 DSS 45 (34-m HEF) 6371.675906 -35.2169652 148.9776833 Madrid DSS 45 (34-m BWG) 6370.025429 40.2357708 355.7459008 DSS 63 (70-m) 6370.051221 40.2413537 355.7519890 DSS 65 (34-m HEF) (see next paragraph) The coordinates for DSS 65 until 1 February 2005 were 6370.021697 40.2373325 355.7485795 In cartesian coordinates (x, y, z) this was (+4849336.6176, -0360488.6349, +4114748.9218) Between February and September 2005, the antenna was physically moved to (+4849339.6448, -0360427.6560, +4114750.7428) Measurement Parameters - DSN ============================ Closed-loop data are recorded in Archival Tracking Data Files (ATDFs), as well as certain secondary products such as the Orbit Data File (ODF). The ATDF Tracking Logical Record contains 150 entries including status information and measurements of ranging, Doppler, and signal strength. ACRONYMS AND ABBREVIATIONS - DSN ================================ ACS Antenna Control System ADC Analog-to-Digital Converter AGC Automatic Gain Control AMS Antenna Microwave System APA Antenna Pointing Assembly ARA Area Routing Assembly ATDF Archival Tracking Data File AUX Auxiliary AZ Azimuth BPF Band Pass Filter bps bits per second BWG Beam WaveGuide (antenna) CDU Command Detector Unit CMC Complex Monitor and Control CONSCAN Conical Scanning (antenna pointing mode) CRG Coherent Reference Generator CUL Clean-up Loop DANA a type of frequency synthesizer dB deciBel dBi dB relative to isotropic dBm dB relative to one milliwatt DCO Digitally Controlled Oscillator DDC Digital Down Converter DEC Declination deg degree DIG RSR Digitizer DMC DSCC Monitor and Control Subsystem DOR Differential One-way Ranging DP Data Processor DSCC Deep Space Communications Complex DSN Deep Space Network DSP DSCC Spectrum Processing Subsystem DSS Deep Space Station DTK DSCC Tracking Subsystem E east EIRP Effective Isotropic Radiated Power EL Elevation FET Field Effect Transistor FFT Fast Fourier Transform FIR Finite impulse Response FTS Frequency and Timing Subsystem GCF Ground Communications Facility GHz Gigahertz GPS Global Positioning System HA Hour Angle HEF High-Efficiency (as in 34-m HEF antennas) HEMT High Electron Mobility Transistor (amplifier) HGA High-Gain Antenna HSB High-Speed BWG IF Intermediate Frequency IFS IF Selector Switch IVC IF Selection Switch JPL Jet Propulsion Laboratory K Kelvin Ka-Band approximately 32 GHz KaBLE Ka-Band Link Experiment kbps kilobits per second kHz kilohertz km kilometer kW kilowatt LAN Local Area Network LCP Left-Circularly Polarized LGR Low-Gain Receive (antenna) LGT Low-Gain Transmit (antenna) LMA Lockheed Martin Astronautics LMC Link Monitor and Control LNA Low-Noise Amplifier LO Local Oscillator LPF Low Pass Filter m meters MCA Master Clock Assembly MCCC Mission Control and Computing Center MDA Metric Data Assembly MGS Mars Global Surveyor MHz Megahertz MOLA Mars Orbiting Laser Altimeter MON Monitor and Control System MOT Mars Observer Transponder MSA Mission Support Area N north NAR Noise Adding Radiometer NBOC Narrow-Band Occultation Converter NCO Numerically Controlled Oscillator NIST SPC 10 time relative to UTC NIU Network Interface Unit NOCC Network Operations and Control System NRV NOCC Radio Science/VLBI Display Subsystem NSS NOCC Support System OCI Operator Control Input ODF Orbit Data File ODR Original Data Record ODS Original Data Stream OLR Open Loop Receiver OSC Oscillator PDS Planetary Data System POCA Programmable Oscillator Control Assembly PPM Precision Power Monitor RA Right Ascension REC Receiver-Exciter Controller RCP Right-Circularly Polarized RF Radio Frequency RIC RIV Controller RIV Radio Science IF-VF Converter Assembly RMDCT Radio Metric Data Conditioning Team RMS Root Mean Square RSR Radio Science Receiver RSS Radio Science Subsystem RT Real-Time (control computer) RTLT Round-Trip Light Time S-band approximately 2100-2300 MHz sec second SEC System Error Correction SIM Simulation SLE Signal Level Estimator SNR Signal-to-Noise Ratio SNT System Noise Temperature SOE Sequence of Events SPA Spectrum Processing Assembly SPC Signal Processing Center sps samples per second SRA Sequential Ranging Assembly SRC Sub-Reflector Controller SSI Spectral Signal Indicator TID Time Insertion and Distribution Assembly TLM Telemetry TSF Tracking Synthesizer Frequency TWM Traveling Wave Maser TWNC Two-Way Non-Coherent TWTA Traveling Wave Tube Amplifier UNK unknown USO UltraStable Oscillator UTC Universal Coordinated Time VCO Voltage-Controlled Oscillator VDP VME Data Processor VF Video Frequency X-band approximately 7800-8500 MHz
  • data set : NEW HORIZONS REX JUPITER ENCOUNTER CALIBRATED V1.0
    Calibrated data taken by New Horizons Radio Science Experiment instrument during the JUPITER mission phase. This is VERSION 1.0 of this data set.
  • data set : NEW HORIZONS REX PLUTO CRUISE CALIBRATED V1.0
    Calibrated data taken by New Horizons Radio Science Experiment instrument during the PLUTOCRUISE mission phase. This is VERSION 1.0 of this data set.
  • data set : NEW HORIZONS REX JUPITER ENCOUNTER RAW V1.0
    Raw data taken by New Horizons Radio Science Experiment instrument during the JUPITER mission phase. This is VERSION 1.0 of this data set.

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