Instrument Host Overview    ========================      For most Mars Global Surveyor experiments, data were collected      by instruments on the spacecraft.  Those data were then relayed      via the telemetry system to stations of the NASA Deep Space      Network (DSN) on the ground.  Radio Science experiments (such      as radio occultations) required that DSN hardware also      participate in data acquisition.  The following sections      provide an overview first of the spacecraft and then of the      DSN ground system as both supported Mars Global Surveyor      science activities.      Instrument Host Overview - Spacecraft    =====================================      The Mars Global Surveyor (MGS) spacecraft was built by      Lockheed Martin Astronautics (LMA).  The spacecraft structure      included four subassemblies: the equipment module, the      propulsion module, the solar array support structure, and the      high-gain antenna (HGA) support structure.       The equipment module housed the avionics packages and science      instruments.  Its dimensions were 1.221 x 1.221 x 0.762 meters      in X, Y, and Z, respectively.  With the exception of the      Magnetometer, all of the science instruments were bolted to      the nadir equipment deck, mounted above the equipment module on      the +Z panel.  The Mars Relay antenna was the tallest      instrument rising 1.115 meters above the nadir equipment deck.       Inside, two identical computers orchestrated almost all of      the spacecraft's flight activities.  Although only one of the      two units controlled Surveyor at any one time, identical      software ran concurrently in the backup unit in case of an      emergency.  Each computer consisted of a Marconi 1750A      microprocessor, 128 Kbytes of RAM for storage, and 20 Kbytes      of ROM that contained code to run basic survival routines in      the event that the computers experienced a reset.       Additional storage for science and spacecraft health data      was provided by two solid-state recorders with a combined      capacity of 375 megabytes.  Mars Global Surveyor was NASA's      first planetary spacecraft to use RAM exclusively (instead of      a tape recorder) for mass data storage.  This technological      improvement reduced operational complexity and cost.       The equipment module also housed three 'reaction wheels'      mounted at right angles to each other.  By transferring angular      momentum to and from the rapidly spinning reaction wheels, MGS      flight computers could control the spacecraft attitude to high      precision.  A fourth reaction wheel, mounted in a direction      skewed to the other three, provided redundancy and backup.       Sun sensors were placed at several locations about the      spacecraft.  They provided basic information on spacecraft      attitude -- namely, a rough vector toward the Sun.  Their      primary use was during attitude reinitialization after a      spacecraft anomaly.       The Inertial Measurement Unit (IMU) contained gyroscopes      and accelerometers to measure angular rates and linear      accelerations.  Angular rate measurements were used to      determine yaw attitude during the Mapping Phase.  The IMU      also provided inertial attitude control, as might be      required during maneuvers.       The Mars Horizon Sensor Assembly (MHSA) determined the horizon      as seen from the spacecraft; from this, an empirical nadir      could be derived for pointing the science instruments.  The      MHSA was mounted to the +Z panel of the equipment module, next      to the science instruments.       The Celestial Sensor Assembly (CSA) complemented the IMU by      providing attitude data based on determination of positions      of known stars.  It was used during the Cruise Phase and Orbit      Insertion Phase for both attitude determination and control.      It was also used when precise attitude knowledge was required      during the Mapping Phase.  The CSA was mounted to the +Z panel      of the equipment module, next to the science instruments.       The propulsion module contained the propellant tanks, main      engines, propulsion feed system and attitude control      thrusters.  It was a rectangular box 1.063 meters on a side      and was bolted to the equipment module on the latter's -Z      panel.  The propulsion module also served as the adaptor to      the launch vehicle.       Propulsion was provided by a dual mode bi-propellant system      using nitrogen tetroxide (NTO) and hydrazine.  This dual mode      differed from conventional bi-propellant systems in that the      hydrazine was used by both the main engine and the attitude      control thrusters, rather than having separate hydrazine      tanks for each.  The main engine was the only one that used      the bi-propellant system.  The main engine maximum thrust      was 659 N.  It was used for major maneuvers including large      trajectory corrections during Cruise, Mars orbit injection      (MOI), and transfer to the Mapping orbit (TMO).       Four rocket engine modules (REM), each containing three 4.45 N      thrusters, were provided.  Each REM contained two aft-facing      thrusters and one roll control thruster.  Four of the eight      aft-facing thrusters were used for the smaller trajectory      corrections during Cruise and for Orbit Trim Maneuvers (OTM)      during Mapping; they could also be used for attitude control      during main engine burns.  Two sets of four thrusters were on      redundant strings so that one string could be isolated in the      event of a failure.  Four thrusters were provided for      attitude control.  In addition to their role during maneuvers,      the 4.45 N thrusters were also used for momentum management.       MGS carried about 385 kg of propellant; nearly 75 percent of      that was used during MOI.       Two solar arrays, each 3.53 meters long by 1.85 meters wide      provided power.  Each array was mounted close to the top of      the propulsion module on the +Y and -Y panels and near the      interface between the propulsion and equipment modules.      Including the adaptor that held the array to the propulsion      module, the tip of each array was designed to stand 4.270      meters from the side of the spacecraft.  During initial      deployment, the -Y solar array yoke was damaged leaving its      exact position and orientation in some doubt (and leading to      several changes in mission design).  Rectangular, metal      'drag flaps' were mounted to the end of each array; these      flaps increased the total surface area of the structure and      added another 0.813 meters to the overall dimensions.      Between each array and flap was mounted a magnetometer sensor.       Each array consisted of two panels, an inner and outer panel,      comprised of gallium arsenide and silicon solar cells,      respectively.  During mapping operations at Mars, the amount      of power produced by the arrays varied from a high of 980      Watts at perihelion to a low of 660 Watts at aphelion.       While in orbit around Mars, the solar arrays provided      power as MGS flew over the day side of the planet.  When      the spacecraft passed over the night side, energy flowed      from two nickel-hydrogen (NiH2) batteries, each with a      capacity of about 20 Amp-hours.  Eclipses lasted from 36 to      41 minutes per orbit; depth of battery discharge was limited      to 27% except during emergencies.       The high-gain antenna structure was also bolted to the      outside of the propulsion module.  When fully deployed, the      1.5-meter diameter antenna sat at the end of a 2-meter boom      which was mounted to the +X panel of the propulsion module.      Two rotating joints (gimbals) held the antenna to the boom      and allowed the antenna to track and point at Earth while      the science instruments observed Mars.       One of the two main functions of the HGA was to receive      command sequences sent by the flight operations      team on Earth.  During command periods, data flowed to MGS      at rates in multiples of two from 7.8125 bits per second      (emergency rate) to 500 bits per second (750 commands per      minute); the nominal rate was 125 bits per second.       The other main function of the HGA was to send data back      to Earth.  All transmissions from MGS utilized an X-band      radio link near 8.4 gigahertz.  The transmitted power was      about 25 watts.  Data rates as high as 85333 bits per second      were used.       The spacecraft was also equipped with four low-gain antennas      (LGA), two for transmit and two for receive.  The LGAs were      used in Inner Cruise, during special events such as maneuvers,      during aerobraking, and for emergency communications following      a spacecraft anomaly.       The primary transmitting low-gain antenna (LGT1) was mounted      on the traveling wave tube amplifier (TWTA) enclosure, which      was mounted on the rim of the HGA reflector; its boresight      was aligned with the HGA boresight, which was in the +X      direction until HGA deployment.  The backup (LGT2) was also      mounted on the TWTA enclosure.  LGT2 boresight was aligned      at a cant angle approximately 160 degrees away from the      shared boresights of the HGA and LGT1.  This angle was chosen      to minimize the consequences of a gimbal failure once      articulation commenced after deployment of the HGA boom in      mapping orbit.  LGT2 was not used prior to HGA deployment      because its orientation and proximity to the nadir payload      deck would lead to irradiation of the payload instruments      while the HGA was in its stowed position.  One receiving LGA      (LGR) was mounted on the -X panel of the equipment module;      the other was on the +X side of the propulsion module.       The spacecraft was equipped with an experimental Ka-band      downlink radio system.  The transmitter converted the X-band      signal to 32 Ghz and amplified it to about 0.5 watts; the      Ka-band output was radiated through the HGA.       The spacecraft +Z axis vector was normal to the nadir equipment      deck; the main engine was aimed in the -Z direction.  The -X      axis vector was in the direction of the velocity vector during      nominal Mapping (e.g., May 1999).  +X was in the direction of      the HGA boresight during Cruise, and the HGA boom was mounted to      the +X panel of the propulsion module.  The +Y axis completed an      orthogonal rectangular coordinate system.  The +/-Y axes defined      generally the deployment directions of the solar panels.  The      solar cells themselves were on the -Z sides of the panels.       There were three levels of anomaly response in the spacecraft      flight software.  The first, emergency mode, was entered in      response to a command-loss timeout.  Entry into emergency mode      reconfigured the telecom subsystem to its lowest data rate      settings to enhance the chances of successful contact from      Earth.  After a programmable period of time in emergency mode,      the spacecraft transitioneds to contingency mode.       Contingency mode was entered by four paths: failure to regain      contact with Earth while in emergency mode, power-related faults      such as gimbal faults and low battery state of charge, loss of      inertial reference, and explicit ground command.  Contingency      mode sets telecom rates to their minimum values, turneds off      non-essential power loads (including the payload), disableds      stored sequences not explicitly specified as enabled for this      mode, and changeds the spacecraft attitude to sun-coning to      optimize power and communications.       Safe mode was the deepest level of anomaly response.  It couldan be      be entered by three paths: failures of key spacecraft components      that could cannot be corrected by normal fault protection, power-on      reset of both Spacecraft Control Processors (SCPs), or explicit      ground command.  The response to safe mode entry was similar to      that of contingency mode.  Safe mode program code for the SCP was      executed from Programmable Read-Only-Memory (PROM).       For more information on the spacecraft and mission see      [JPLD-12088].      Instrument Host Overview - DSN    ==============================      The Deep Space Network is a telecommunications facility managed      by the Jet Propulsion Laboratory of the California Institute of      Technology for the U.S.  National Aeronautics and Space      Administration (NASA).       The primary function of the DSN is to provide two-way      communications between the Earth and spacecraft exploring the      solar system.  To carry out this function it is equipped with      high-power transmitters, low-noise amplifiers and receivers,      and appropriate monitoring and control systems.       The DSN consists of three complexes situated at approximately      equally spaced longitudinal intervals around the globe at      Goldstone (near Barstow, California), Robledo (near Madrid,      Spain), and Tidbinbilla (near Canberra, Australia).  Two of      the complexes are located in the northern hemisphere while the      third is in the southern hemisphere.       Each complex includes several antennas, defined by their      diameters, construction, or operational characteristics:      70-m diameter, standard 34-m diameter, high-efficiency 34-m      diameter (HEF), and 34-m beam waveguide (BWG).       For more information see [ASMAR&RENZETTI1993].