urn:nasa:pds:context:instrument:vg2.eng 1.0 1.13.0.0 Product_Context 2021-05-13 1.0 The initial version of this context product, which describes the systems on the Voyager 2 spacecraft and (indirectly) the sensors that report status and performance. urn:nasa:pds:context:instrument_host:spacecraft.vg2 instrument_to_instrument_host Kohlhase, C., The Voyager Neptune Travel Guide, JPL Publication 89-24, 276 pp., Jet Propulsion Laboratory, Pasadena, CA, 1989. Summary of Voyager 2 plans at Neptune. Although written for Voyager 2, the chapter about the spacecraft applies equally well to Voyager 1 since the two spacecraft were built to be identical. Morrison, D., Voyages to Saturn, NASA SP-451, 227 pp., National Aeronautics and Space Administration, Washington, DC, 1982. Summary of Voyager 1 and 2 accomplishments at Saturn. Voyager 2 Engineering Spacecraft Sensor The aggregation of all spacecraft sensors on the Voyager 2 spacecraft. The aggregation of spacecraft sensors on the Voyager 2 spacecraft. The Voyager 2 spacecraft was built by the Jet Propulsion Laboratory (JPL). Its mass was 815 kilograms, including its own power, propulsion, and communications systems and its own science instruments. Spacecraft electrical power was supplied by Radioisotope Thermoelectric Generators (RTGs) that produced about 400 watts. The Attitude and Articulation Control Subsystem (AACS), Computer Command Subsystem (CCS), and Flight Data Subsystem (FDS) managed spacecraft operations. Thrusters and gyroscopes provided physical propulsion and attitude control. Communications between the spacecraft and Earth were carried out via a high-gain radio antenna using both S-band and X-band frequencies at data rates as high as 115.2 kilobits per second. A Digital Tape Recorder (DTR) could save up to 500 million bits when no Earth station was available for real-time data transmission. Voyager control systems could record sets of several thousand instructions, allowing autonomous operation for days or weeks at a time. Many of the systems were equipped with sensors, which fed engineering data to the ground where status and performance could be monitored. The spacecraft itself was built around its 'bus' -- a decagonal prism, which was about 2 meters in diameter and about 60 cm deep. Each of the ten sides of the bus was associated with a 'bay' containing engineering systems or science instrument electronics. Bay 1, for example, contained the radio transmitter. The High-Gain Antenna (HGA) was mounted to the end of the bus facing Earth. The bays were numbered 1 through 10 in a clockwise direction when viewed from Earth. Extending away from the bus were three booms: a science boom and scan platform to which most instruments were mounted, a magnetometer boom, and a boom to which the RTGs were mounted. Spacecraft Coordinate System ---------------------------- 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 science boom, supporting the scan platform, extended in the general direction of positive y; this boom was also defined as being at cone angle 90 degrees, clock angle 215 degrees and at azimuth 180 degrees, elevation 97 degrees. The boom supporting the RTGs was mounted on the bus in generally the negative y direction. The positive y-axis (yaw axis) of the spacecraft coordinate system passed through Bay 3; the negative y-axis passed through Bay 8. The x-axis (pitch axis) was in a direction which defined a right-handed rectangular coordinate system. The positive x-axis was at cone angle 90 degrees, clock angle 305 degrees (azimuth 270 degrees, elevation 90 degrees). Telecommunications Subsystem ---------------------------- The high-gain antenna was mounted to the spacecraft bus, pointing in the negative z-direction. It was a parabolic reflector 3.7 meters in diameter with a feed that permitted simultaneous operation at both S-band (13 cm wavelength) and X-band (3.6 cm). The half-power full-width of the antenna beam was 0.6 degrees at X-band and 2.3 degrees at S-band. 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. The Telecommunications Subsystem (TCS) electronics included a redundant pair of transponders, meaning that a failed functional unit in one transponder could be bypassed by swapping to the redundant unit. The TCS could transmit science data on the X-band link at rates between 4.8 and 115.2 kilobits per second and engineering data on the S-band link at 40 bits per second. It could receive instructions sent (uplinked) from ground stations at a rate of 16 bits per second. Commands were extracted from the uplink signal by the Command Detector Unit (CDU) and were then sent to the Computer Command Subsystem (CCS). 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. Voyager 2 differed from Voyager 1 in this locking mechanism, which failed after launch. Although lock could be achieved, the uplink signal had to be within a window only 100 Hz wide around the Voyager 2 receiver's rest frequency. Engineers discovered both the problem and the solution during cruise and were able to tune the uplink transmitted from the ground so that commands could be sent and data received for encounters with all four of the Voyager 2 target planets. Attitude and Articulation Control Subsystem ------------------------------------------- The Attitude and Articulation Control Subsystem (AACS) provided three-axis-stabilized control so that the spacecraft could maintain a fixed orientation in space. Attitude control was accomplished using gyroscopes or by celestial reference. The AACS also controlled motion of the scan platform, upon which the four 'remote sensing' instruments were mounted. Gyro control was used in special situations (e.g., trajectory corrections and solar conjunctions) for periods of up to several days. The inertial reference unit operated with tuned rotor gyros having an uncalibrated drift rate of less than 0.5 degrees per hour and a calibrated drift rate of less than 0.05 degrees per hour. Celestial control was based on viewing the Sun (through a sensor mounted on the high-gain antenna) and a single bright star (through a second sensor named the Canopus Star Tracker, after the star used most frequently as the reference). When the spacecraft attitude drifted by more than a small amount from the reference objects, the AACS fired small thrusters which returned the spacecraft to the proper orientation. The Sun sensor was an optical potentiometer with a cadmium sulfide detector; its error was less than 0.01 degrees and its limit cycle was +/-0.05 degrees. The Canopus Star Tracker was an image dissector tube with a cesium detector, an error of less than 0.01 degree, and a limit cycle of +/-0.05 degrees. Redundant (backup) sun sensors, star trackers, and computers were also part of the AACS. The non-redundant portions of the AACS were those controlling the pointing of the instrument scan platform, which had two degrees of freedom -- elevation and azimuth (see below). Propulsion Subsystem -------------------- The propulsion system was part of the AACS and consisted of 16 hydrazine thrusters. These thrusters were also used to control the three-axis stabilization of the spacecraft. Two thrusters on opposite sides of the spacecraft were used to perform positive roll turns around the +Z axis. Two oppositely pointed thrusters were used to perform negative roll turns. One thruster was used to perform positive yaw turns (around the +Y axis) and one was used to perform negative yaw turns. One thruster was used to perform positive pitch turns (around the +X axis) and one was used to perform negative pitch turns. A backup hydrazine system was connected to a redundant set of eight thrusters. Power Subsystem --------------- Spacecraft power was provided by three Radioisotope Thermoelectric Generators (RTGs) mounted on a boom in the negative y-direction. At Launch the three RTGs converted 7000 watts of heat into 475 watts of electrical power. RTG electrical output decreased by about 7 watts per year because of decay of the plutonium dioxide fissionable material and degradation of the silicon-germanium thermocouples. The difference between available electrical power and the power required to operate spacecraft subsystems was called the 'power margin.' Voyager Project guidelines required a power margin of at least 12 watts to guard against electrical transients and miscalculations; excess electrical power was dissipated as heat in a shunt radiator. Data Storage Subsystem ---------------------- The Digital Tape Recorder (DTR) was used to store data when real-time communications with Earth were either not possible or not scheduled. The DTR recorded data on eight tracks; rates were 115.2 kilobits per second (record only), 21.6 kilobits per second (playback only), and 7.2 kilobits per second (record and playback). Capacity of each track was 12 images or equivalent. Computer Command Subsystem -------------------------- The Computer Command Subsystem (CCS) consisted of two identical computer processors, their software algorithms, and associated electronic hardware. The CCS was the central controller of the spacecraft. During most of the Voyager mission the two CCS computers on each spacecraft were used non-redundantly to increase the command and processing capability of the spacecraft. Flight Data Subsystem --------------------- The Flight Data Subsystem (FDS) consisted of two reprogrammable digital computers and associated encoding hardware. The FDS collected and formatted science and engineering telemetry data for transmission to Earth. Convolutional coding was imposed on all data transmitted from the spacecraft. Additionally, both Golay encoding and Reed- Solomon encoding were available for use on spacecraft data. Data compression was also performed within the FDS. Science Boom ------------ The Voyager science instrument boom carried the plasma detector, cosmic ray detector and the low energy charged particle detector. The scan platform was mounted on the science boom. Scan Platform ------------- Four instruments (Imaging, PhotoPolarimeter, Infra-Red Interferometric Spectrometer, and Ultra Violet Spectrometer) were mounted on the scan platform, which could be slewed by motors and gears (called actuators). Elevation of the scan platform was measured with respect to a plane slightly offset (by approximately 7 degrees) from the spacecraft x-z plane; the spacecraft positive y-axis was at 97 degrees elevation (see Spacecraft Coordinate System above). The scan platform azimuth reference was defined by the y-z plane, with zero azimuth being in the negative z-direction. Drive actuators were controlled by fine feedback potentiometers; the error of each was less than 0.03 degrees, and the final pointing error of the scan platform was nominally +/-0.1 degrees (2-sigma per axis). Subsequent analysis by the Navigation and Ancillary Information Facility (NAIF) at JPL showed larger errors during at least the Jupiter and Saturn encounters. High rate slews of 1 deg/sec were discontinued after the azimuth drive mechanism on Voyager 2 temporarily froze a short time after Saturn closest approach. The medium slew rate was 0.33 deg/sec, and the low slew rate was 0.08 deg/sec. Magnetometer Boom ----------------- Two low-field magnetometers were mounted on a 13-meter-long boom that was unfurled and extended automatically after Launch. One low-field magnetometer was mounted at the end of the boom and a second was mounted about 3 meters from the end. Two high-field magnetometers were mounted at the base of the boom.