INSTRUMENT_HOST_ID = MGS INSTRUMENT_HOST_NAME = Mars Global Surveyor INSTRUMENT_HOST_TYPE = Spacecraft 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). 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). References ========== Asmar, S.W., and N.A. Renzetti, The Deep Space Network as an Instrument for Radio Science Research, Jet Propulsion Laboratory Publication 80-93, Rev.1, 15 April 1993. Mars Global Surveyor Project, Mission Plan, Final Version (MGS 542-405), JPL Document D-12088, Jet Propulsion Laboratory, Pasadena, CA, 1995.