Mars Global Surveyor
Science Sampler Data Set Collection


June 26, 1998


The Mars Global Surveyor (MGS) Science Sampler Data Set Collection contains products derived from data collected during the Orbit Insertion Phase of the mission. Most of the data were acquired during the Assessment Subphase, orbits 19 through 36. This collection is intended to provide samples of data that will be available later in the MGS mission in significantly larger volumes. Included are products from the Mars Orbiter Camera (MOC), Mars Orbiter Laser Altimeter (MOLA), Thermal Emission Spectrometer (TES), and Magnetometer/Electron Reflectometer (MAG/ER). Mars gravity models based on Mariner 9 and Viking Orbiter data are also included as examples of future MGS Radio Science (RSS) products. Data products in this collection range from standard products in archival format, such as the MOLA PEDRs, to highly derived products in a form convenient for viewing but not necessarily intended for data analysis, such as the MAG/ER plots. It should be noted that most products in this collection were generated specifically for this publication and do not constitute an official MGS standard product delivery (see section 6, Archive Data Preparation).

This file is the primary documentation for the MGS Sampler Data Collection. It includes these sections:

1. Introduction
2. Mars Global Surveyor Mission
2.1 Mission History and Present Status
2.2 Mission Phases
2.3 Mission Objectives Summary
3. Mars Global Surveyor Spacecraft
4. Mars Global Surveyor Science Instruments
4.1 Mars Orbiter Camera (MOC)
4.2 Mars Orbiter Laser Altimeter (MOLA)
4.3 Thermal Emission Spectrometer (TES)
4.4 Magnetometer / Electron Reflectometer (MAG/ER)
4.5 Radio Science Subsystem (RSS)
5. Coordinate Systems and SPICE
5.1 Coordinate Systems
6. Archive Data Preparation
6.1 Mars Orbiter Camera (MOC)
6.2 Mars Orbiter Laser Altimeter (MOLA)
6.3 Thermal Emission Spectrometer (TES)
        6.3.1 Calibrated Spectral Radiance
        6.3.2 Global Temperature Images
6.4 Magnetometer / Electron Reflectometer (MAG/ER)
6.5 Radio Science Subsystem (RSS)
        6.5.1 The GMM-1 Gravity Solution
        6.5.2 The JPL-50C Gravity Solution
        6.5.3 The JPL-50C Gravity Map
7. Disk Directory Structure
8. File Names and Formats
9. References

For more information about the Mars Global Surveyor mission, instruments, and data sets on this volume, see the following articles in the March 13, 1998 issue of Science. The issue is available on the World-Wide Web at

The science instruments on Mars Global Surveyor were originally designed for the unsuccessful Mars Observer mission. Articles describing these instruments were published in the May 25, 1992, issue of the Journal of Geophysical Research (volume 97, number E5).

Current information about the status of the mission can be found on the MGS Web site at

This CD-ROM archive was produced by the Geosciences Discipline Node of NASA's Planetary Data System (PDS).

2. Mars Global Surveyor Mission

2.1 Mission History and Present Status

The MGS spacecraft was launched from Cape Canaveral, Florida, on November 7, 1996, aboard a Delta-2/7925 rocket. The 1062 Kg spacecraft, built by Lockheed Martin Astronautics, traveled nearly 750 million km over the course of a 300-day cruise to reach Mars on September 11, 1997.

Upon reaching Mars, MGS entered the Orbit Insertion Phase of the mission by firing its main rocket engine for a 25-minute Mars orbit insertion (MOI) burn. This maneuver slowed the spacecraft and allowed the planet's gravity to capture it into orbit. The initial MGS orbit was highly elliptical with a period of approximately 45 hours.

MGS performed a series of orbit changes to drop the low point of its orbit (i.e., periapsis) into the upper fringes of the Martian atmosphere at an altitude of about 110 km. During the periapsis atmospheric passes, the spacecraft velocity was slightly reduced because of drag. This effect also caused the spacecraft to decrease its apoapsis altitude. MGS was to use this aerobraking technique over a period of four months to lower the high point of its orbit from 56,000 km to near 400 km in altitude. However, at the low point of orbit 15, on October 8, 1997, the spacecraft experienced difficulties later diagnosed as due to excess vibrations of one of the solar panels. The problem was associated with a fracture of a panel damper arm [Albee et al., 1998]. While an evaluation of the solar array problem was underway, periapsis was raised to about 172 km on October 13, 1997 and remained near that altitude until November 7, 1997 (orbits 19 through 36). During this 26-day period the spacecraft instrument panel was pointed towards Mars during close approaches (i.e., near periapsis) and the first extensive set of science observations from MGS was collected. Orbits 19 through 36 are known as the assessment orbits, and the time period is known as the Assessment Subphase of the Orbit Insertion Phase. The science observations were acquired during the descending leg of each orbit; that is, as the spacecraft moved from north to south.

Aerobraking was restarted on November 8, 1997 (orbit 37), but with a periapsis approximately 10 km higher than that previously used. As of this writing, the mission plan is to conduct aerobraking at about 1/3 the rate originally planned and to place the spacecraft in a 2 AM Sun-synchronous mapping orbit in March 1999 rather than the planned 2 PM mapping orbit in March 1998. (The 2 PM orbit meant that the spacecraft would have crossed the equator in the descending leg of the orbit -- north to south -- at 2 PM, a desirable time for data collection for some instruments. This orbit could not be achieved given the new orbital characteristics. However, a 2 AM orbit is satisfactory because if the descending leg of the orbit crosses the equator at 2 AM, it means that the ascending leg, south to north, will cross the equator at the desired time of 2 PM.) Another aerobraking hiatus began on March 27, 1998, and will extend through September 11, 1998. This period, termed the Science Phasing Orbits of the mission, is necessary to ensure that the final two hour circular orbit has an equatorial crossing time of 2 AM. After a final period of aerobraking beginning in September, 1998, the Mapping Phase of the mission will begin in mid-March 1999. During mapping operations, the spacecraft will orbit Mars with a period of 118 minutes, at an average altitude of 400 km. For 687 Earth days (one Mars year), MGS will utilize this orbital vantage point to collect scientific data on a continuous basis.

2.2 Mission Phases

Six mission phases are defined for significant spacecraft activity periods. These included Pre-Launch, Launch, Cruise, Orbit Insertion, Mapping, and Relay Phases. The Cruise Phase includes both Inner and Outer Cruise components. The Orbit Insertion Phase includes the Aerobraking Phases, the Assessment Subphase (during which most of the data on this sampler volume were acquired), and the Science Phasing orbits. During the Mapping Phase, the spacecraft will approximately retrace its ground track once every seven Martian sols; these 88-orbit intervals are known as repeat cycles.

The following table lists the mission phase dates. The phases are further described below. Note that dates for events after 1998-03-27 are planned but have not yet occurred as of this writing.

Mission Phase   Start Date*   End Date*
Prelaunch   1994-10-12   1996-11-06
Launch   1996-11-06   1996-11-07
Cruise   1996-11-07   1997-09-02
    Inner Cruise   1996-11-07   1997-01-09
    Outer Cruise   1997-01-09   1997-09-02
Orbit Insertion   1997-09-02   1999-03-15
    Aerobraking Phase I   1997-09-17   1997-10-11
    Assessment Subphase   1997-10-13   1997-11-07
    Aerobraking Phase I continued   1997-11-08   1998-03-27
    Science Phasing Orbits   1998-03-27   1998-09-11
    Aerobraking Phase II   1998-09-12   1999-03-15
Mapping   1999-03-15   2001-01-31
Radio Relay   2001-01-31   2003-01-01


The Prelaunch Phase extended from beginning of the MGS mission until the start of the launch countdown at the Kennedy Space Center.

The Launch Phase extended from the start of launch countdown until completion of the injection into the Earth-Mars trajectory.

The Cruise Phase extended from injection into the Earth-Mars trajectory until 10 days before Mars orbit insertion. During the Inner Cruise, MGS aimed its solar panels toward the Sun and communicated through its low-gain antenna; during the Outer Cruise, the high-gain antenna could be used while the solar panels generated acceptable levels of power. The transition occurred on January 9, 1997.

The Mars Orbit Insertion Phase extends from 10 days before Mars orbit insertion until the spacecraft reaches the final mapping orbit and is declared ready for collection of science data. Orbit insertion occurred on September 11, 1997.

The Mapping Phase is the period of concentrated science data acquisition. It lasts for one Martian year.

The Radio Relay Phase begins at the end of the Mapping Phase and continues for the remainder of the spacecraft on-orbit lifetime. The radio system will be used to relay data from future landers on the surface of Mars. The nominal end of the MGS mission is scheduled for January 1, 2003.

2.3 Mission Objectives Summary

The MGS Mission is part of the Mars Surveyor Program. This program focuses on understanding present and past climate conditions on Mars, determining whether Mars developed prebiotic compounds and life, and identifying resources of use during human expeditions to the surface. Determining the locations and states of water reservoirs today and in the past are key objectives. Missions in the Program are designed to make measurements from orbit, from the surface, and from returned samples. MGS represents a primary orbital component of the Program, collecting information on the characteristics and dynamics of the magnetosphere, atmosphere, surface and interior on a global basis [Albee et al., 1992; 1998].

In detail, the MGS science objectives are to:

3. Mars Global Surveyor Spacecraft

The MGS spacecraft was built by Lockheed Martin Astronautics (LMA). The spacecraft structure includes four subassemblies: the equipment module, the propulsion module, the solar array support structure, and the high-gain antenna (HGA) support structure.

The equipment module houses the avionics packages and science instruments. Its dimensions are 1.221 x 1.221 x 0.762 m in X, Y, and Z, respectively. With the exception of the Magnetometer, all of the science instruments are bolted to the nadir equipment deck, mounted above the equipment module on the +Z panel. The Mars Relay antenna is the tallest instrument rising 1.115 m above the nadir equipment deck.

Inside, two identical computers control almost all the spacecraft's flight activities. Although only one of the two units controls MGS at any one time, identical software runs concurrently in the backup unit in case of an emergency. Each computer consists of a Marconi 1750A microprocessor, 128 Kbytes of RAM for storage, and 20 Kbytes of ROM that contains code to run basic survival routines in the event that the computers experience a reset.

Additional storage for science and spacecraft health data is provided by two solid-state recorders with a combined capacity of 375 megabytes. MGS is NASA's first planetary spacecraft to use RAM exclusively (instead of a tape recorder) for mass data storage. This technological improvement reduces operational complexity and cost.

The equipment module also houses 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 can control the spacecraft attitude to high precision. A fourth reaction wheel, mounted in a direction skewed to the other three, provides redundancy and backup.

Sun sensors are placed at several locations about the spacecraft. They provide basic information on spacecraft attitude relative to the sun. Their primary use is during attitude reinitialization after a spacecraft anomaly.

The Inertial Measurement Unit (IMU) contains gyroscopes and accelerometers to measure angular rates and linear accelerations. Angular rate measurements are used to determine yaw attitude during the Mapping Phase. The IMU also provides inertial attitude control required during maneuvers. The linear accelerometers are useful for propulsive maneuvers, but are only marginally sensitive for aerobraking measurements.

The Mars Horizon Sensor Assembly (MHSA) determines the horizon as seen from the spacecraft; from this, an empirical nadir can be derived for pointing the science instruments. The MHSA is mounted to the +Z panel of the equipment module, next to the science instruments.

The Celestial Sensor Assembly (CSA) complements 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 will also be used when precise attitude knowledge is required during the Mapping Phase. The CSA is mounted to the +Z panel of the equipment module, next to the science instruments.

The propulsion module contains the propellant tanks, main engines, propulsion feed system and attitude control thrusters. It is a rectangular box 1.063 m on a side and is bolted to the equipment module on the latter's -Z panel. The propulsion module also serves as the adaptor to the launch vehicle.

Propulsion is provided by a dual mode bi-propellant system using nitrogen tetroxide (NTO) and hydrazine. This dual mode differs from conventional bi-propellant systems in that the hydrazine is used by both the main engine and the attitude control thrusters, rather than having separate hydrazine tanks for each. The main engine is the only one that used the bi-propellant system. The main engine maximum thrust is 659 N. It is 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, are provided. Each REM contains 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 will be used for Orbit Trim Maneuvers (OTM) during Mapping; they can also be used for attitude control during main engine burns. Two sets of four thrusters are on redundant strings so that one string can be isolated in the event of a failure. Four thrusters are provided for attitude control. In addition to their role during maneuvers, the 4.45 N thrusters are also used for momentum management.

At launch MGS carried about 385 kg of propellant; nearly 75 percent of that was used during MOI.

Two solar arrays, each 3.53 m long by 1.85 m wide, provide power. Each array is 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 adapter that holds the array to the propulsion module, the tip of each array stands 4.270 m from the side of the spacecraft. Rectangular, metal 'drag flaps' are mounted to the end of each array; these flaps increase the total surface area of the structure and added another 0.813 m to the overall dimensions. Magnetometers are mounted between each array and flap.

Each array consists 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 will vary from a high of 980 W at perihelion to a low of 660 W at aphelion.

While in orbit around Mars, the solar arrays provide power as MGS flies over the day side of the planet. When the spacecraft passes over the night side, energy flows from two nickel-hydrogen (NiH2) batteries, each with a capacity of about 20 Amp-hours. Eclipses last from 36 to 41 minutes per orbit; depth of battery discharge is limited to 27% except during emergencies.

The high-gain antenna structure is also bolted to the outside of the propulsion module. When fully deployed, the 1.5 m diameter antenna sits at the end of a 2 m boom which is mounted to the +X panel of the propulsion module. Two rotating joints (gimbals) hold the antenna to the boom and allow the antenna to track and point at Earth while the science instruments observe Mars.

One of the two main functions of the HGA is to receive command sequences sent by the flight operations team on Earth. During command periods, data flows 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 is 125 bits per second.

The other main function of the HGA is to send data back to Earth. All transmissions from MGS utilize an X-band radio link near 8.4 gigahertz. The transmitted power is about 25 W. Data rates as high as 85333 bits per second are used.

The spacecraft is also equipped with four low-gain antennas (LGA), two for transmit and two for receive. The LGAs are used in Inner Cruise, during special events such as maneuvers, during aerobraking, and for emergency communications following a spacecraft anomaly. The primary transmitting LGA is mounted on the HGA reflector, while the backup is mounted on the +X side of the propulsion module. The two receive LGAs are mounted on the -X panel of the equipment module and the +X side of the propulsion module.

The spacecraft is equipped with an experimental Ka-band downlink radio system. The transmitter converts the X-band signal to 32 GHz and amplifies it to about 0.5 W; the Ka-band output is radiated through the HGA.

4. Mars Global Surveyor Science Instruments

MGS carries four on-board science instruments. The Mars Orbiter Camera (MOC) has both a wide-angle mode for global coverage and a narrow-angle mode with resolution of 1.4 meters [Malin et al., 1992; 1998]. The Mars Orbiter Laser Altimeter (MOLA) gathers data that allow mapping the topography of the planet and the 1.064 micrometer reflectivity [Zuber et al., 1992; Smith et al., 1998]. The Thermal Emission Spectrometer (TES) measures infrared radiation, and is used to determine the general mineral composition of patches of ground as small as 9.0 square km. In addition, TES also scans the Martian atmosphere to provide data for the study of the clouds and weather [Christensen et al., 1992; 1998]. The Magnetometer/Electron Reflectometer (MAG/ER) is used to measure the global magnetic properties of Mars, which provide insight on internal structure [Acua et al., 1992; 1998].

An UltraStable Oscillator (USO) in conjunction with the on-board telecommunications equipment and ground equipment at stations of the NASA Deep Space Network (DSN) make up the Radio Science Subsystem (RSS). RSS measurements include radio tracking of the spacecraft to improve the gravity field model of Mars and radio occultation observations to study the structure of the atmosphere [Tyler et al., 1992].

A sixth "instrument" is the Mars Relay -- a cylindrically shaped antenna used to collect data transmitted to Surveyor from landers on the Martian surface. These landers will be carried to Mars by later Mars Surveyor Program spacecraft and operated after completion of the MGS mapping mission.

4.1 Mars Orbiter Camera (MOC)

MOC is a three-component imaging system (one narrow-angle and two wide-angle cameras) designed to take high spatial resolution pictures of the surface and to obtain lower spatial resolution, synoptic coverage of the surface and atmosphere [Malin et al., 1992; 1998]. The cameras are based on the 'push broom' technique, acquiring one line of data at a time as the spacecraft orbits the planet. Using the narrow-angle camera during the Mapping Phase of the mission, areas ranging from 2.8 x 2.8 km to 2.8 x 25.2 km (depending on available internal digital buffer memory) can be imaged at about 1.4 m/pixel. Additionally, lower-resolution pictures (to a lowest resolution of about 11 m/pixel) can be acquired by pixel averaging; these images can be much longer, ranging up to 2.8 x 500 km at 11 m/pixel. High-resolution data will be used to study sediments and sedimentary processes, polar processes and deposits, volcanism, and other geologic/geomorphic processes. The MOC wide-angle cameras are capable of viewing Mars from horizon to horizon and are designed for low-resolution global and intermediate resolution regional studies. Low-resolution observations can be made every orbit during the Mapping Phase, so that in a single 24-hour period a complete global picture of the planet can be assembled at a resolution of at least 7.5 km/pixel. Regional areas (covering hundreds of km on a side) may be imaged at a resolution of better than 250 m/pixel at the nadir. These images will be particularly useful in studying time-variable features such as lee clouds, the polar cap edge, and wind streaks, as well as acquiring stereoscopic coverage of areas of geological interest. The limb can be imaged at vertical and along-track resolutions of better than 1.5 km. Color filters within the two wide-angle cameras permit color images of the surface and atmosphere to be made to distinguish between clouds and the ground and between clouds of different composition.

The following table summarizes MOC characteristics.

Camera Min wavelength
Max wavelength
Focal length
Aperture F number Resolution
(380 km, m/pixel)
Narrow angle 500 900 350.0 0.35 m 10 1.5
Wide angle red 600 630 1.1 1.7 mm 6.4 230
Wide angle blue 420 450 1.14 1.8 mm 6.3 230

During the Orbit Insertion Phase of the mission a number of MOC images have been acquired on a best efforts basis. Most of the data have been acquired in the high resolution mode. Nine of these images are included on this MGS Science Sampler volume and are stored as GIF files.

4.2 Mars Orbiter Laser Altimeter (MOLA)

The primary MOLA objective is to determine globally the topography of Mars at a level suitable for addressing problems in geology and geophysics [Zuber et al., 1992; Smith et al., 1998]. Secondary objectives include characterizing the 1.064 micrometer wavelength surface reflectivity of Mars to contribute to analyses of global surface mineralogy and seasonal albedo changes. Other objectives include addressing problems in atmospheric circulation and providing geodetic control and topographic context for the assessment of possible future Mars landing sites. Eighteen orbits of MOLA data were acquired during the Assessment Subphase of the mission. Coverage extends roughly from the north pole to the equator. The reason is that the spacecraft needed to be lower than approximately 750 km in altitude to sample the backscattered laser signal. Given that periapsis occurred at northern latitudes, data are necessarily focused in this hemisphere.

The principal components of MOLA are a diode-pumped, Nd:YAG laser transmitter that emits 1.064 micrometer wavelength laser pulses, a 0.5 m diameter telescope, a silicon avalanche photodiode detector, and a time interval unit with 10 nsec resolution. When in the Mapping Phase of the mission, MOLA provides measurements of the topography of Mars within approximately 160 m footprints and a center-to-center along-track footprint spacing of 300 m along the MGS nadir ground-track. The elevation measurements are quantized with 1.5 m vertical resolution before correction for orbit and pointing errors. MOLA profiles will be assembled into a global 0.25 x 0.25 degree grid referenced to Mars' center-of-mass with an absolute accuracy of approximately 30 m. Other standard data products will include a global grid of topographic gradients, corrected individual profiles, and a global 0.25 x 0.25 degree grid of 1.064 micrometer surface reflectivity.

The following table summarizes MOLA characteristics.

Parameter   Value   Unit
Physical Characteristics  
Volume   0.15   m3
Mass   26.18   kg
Power (TOTAL)   28.74   W
Heater Power   10.00   W
Laser Transmitter  
Laser type   Q-switched,
Wavelength   1.064   micrometer
Laser energy   40-30   mJ pulse-1
Laser power consumption   13.7   W
Pulse width   ~8.5   ns (FWHM**)
Pulse repetition rate   10   sec-1
Beam cross-section   25x25   mm2
Beam divergence   0.25   mrad
Altimeter Receiver  
Telescope type   Cassegrain  
Mirror composition   Gold-coated beryllium  
Telescope diameter   0.5   m
Focal length   0.74   m
Detector type   Silicon avalanche
(Si APD)
Sensitivity   1   nW
Optical filter   2.2   nm bandpass
Field of view   ~0.85   mrad
Receiver Electronics  
Receiver type   Match-filtered leading-edge trigger  
Time resolution   10   nsec
Range resolution   1.5   m
Pulse energy resolution   20%    
Footprint size (@ 400 km)   120   m
Footprint spacing (@ velocity = 3 km/sec)
(center-to-center, along-track)  
300 m
Type   80C86    
Data rate   617.14   bits sec-1

* Nd:YAG is neodymium-doped yttrium aluminum garnet.

** FWHM is full width at half maximum.

The MOLA instrument measures the round-trip time of flight of infrared laser pulses transmitted from the MGS spacecraft to the Martian surface. The instrument normally operates in a single autonomous mode, in which it will produce ranging measurements. Surface topography estimates can be derived from these data, given appropriate corrections for the position and attitude of the spacecraft.

All components of MOLA except for the laser and telescope have been designed, built and tested at NASA's Goddard Space Flight Center, Greenbelt, MD.

MOLA's transmitter is a Q-switched, Nd:YAG laser oscillator which is pumped by a 44 bar laser array. Each bar contains ~1000 AlGaAs (Aluminum, Gallium Arsenide) laser diodes. The Q-switch controls the emission of the laser, and Nd:YAG refers to the composition of the material that is optically excited to produce laser action: Neodymium-doped Yttrium Aluminum Garnet. The laser emits 8.5-ns-wide (full width at half the maximum pulse amplitude, FWHM) pulses at 1.064 micrometers. The pulse repetition rate is 10 Hz, and the expected pulse energy is 40 mJ at the beginning of the Mapping Phase and 30 mJ at end of mission. The laser consumes 13.7 W when operating, and its on-orbit lifetime is expected to be at least 6x108 laser pulses (~2 years).

The development of a space-qualified, long-lifetime laser represents one of the primary engineering challenges associated with MOLA. For comparison, the ruby flashlamp laser altimeters flown on the Apollo 15, 16 and 17 missions [Kaula et al., 1972, 1973, 1974] each operated for less than 104 laser pulses. High pulse-repetition-rate lasers with lifetimes of order 109 shots have been made possible due to breakthroughs in solid-state laser technology, resulting in improvements in the peak power, brightness, and availability of semiconductor diodes and arrays [Cross et al., 1987; Byer, 1988]. The key technological advance has been the replacement of the flashlamp, which is the device that has traditionally been used to pump optical energy into the laser rod, with a highly efficient array of laser diodes. While flashlamp lasers fail catastrophically, diode-pumped lasers such as MOLA's instead undergo a gradual degradation in energy output as individual pump diodes fail. Laser diodes also produce the required pump energy only in a narrow region near the laser rod's absorption band, which dramatically improves the laser's electrical to optical efficiency.

MOLA data derived from the Assessment Subphase of the mission are included on this MGS Science Sampler volume, including standard products consisting of along track estimates of heights and a suite of plots illustrating the data.

4.3 Thermal Emission Spectrometer (TES)

TES is designed to study the surface and atmosphere of Mars using thermal infrared (IR) spectroscopy, together with broadband thermal and solar reflectance radiometry. The specific objectives of the TES experiment are: (a) to determine and map the composition of surface minerals, rocks, and ices; (b) to study the composition, particle size, and spatial and temporal distribution of atmospheric dust; (c) to locate water-ice and CO2 condensate clouds and determine their temperature, height, and condensate abundance; (d) to study the growth, retreat, and total energy balance of the polar cap deposits; (e) to measure the thermophysical properties of the Martian surface materials; and (f) to characterize the thermal structure and dynamics of the atmosphere [Christensen et al., 1992; 1998].

The TES instrument consists of three sub-sections, the primary one being a Michelson interferometer that produces spectra from 1700 to 200 cm-1 (~6 to 50 micrometers), at either 5 or 10 cm-1 spectral resolution. The instrument cycle time, including collection of the interferogram, mirror flyback, and electronic reset, is 2 sec for 10 cm-1 operation, and 4 sec for 5 cm-1 operation. The interferometer includes a visible interferometer that is used to generate fringes which are used to control the linear drive servo and to determine position in the interferogram. This system uses two redundant neon lamps that produce an emission line at 703.2 nm for fringe generation and a continuum that is used for a quasi-white-light source for determination of zero path difference. The off-axis position of the six detectors results in self-apodization and a spectral shift that is a function of both distance from the axis and optical frequency. The resulting full-width half-maximum (FWHM) value is 12.14 cm-1 for 10 cm-1 operation. For the corner detectors and at the highest frequency (shortest wavelength) there is a significant departure from the ideal, with a worst-case degradation to a FWHM of 16 cm-1. Because all of the response functions have the same area there is no loss in signal when viewing a smooth continuum scene like Mars. However, there will be a slight loss in contrast of narrow spectral features due to broadening of the spectral width. Because the self-apodization is considerable, the data will generally be used without further apodization. Separate fast Fourier transform (FFT) algorithms are used for the center and edge detectors in order to partially correct for the different spectral shifts introduced into these detectors. As a result, the data generated by the two FFTs will have approximately the same frequency sampling.

A pointing mirror capable of rotating 360 degrees provides views to space, both limbs, and to internal, full-aperture thermal and visible calibration targets, as well as image motion compensation. In addition to the spectrometer, the instrument has bore sighted bolometric thermal radiance (4.5 to ~100 micrometer) and solar reflectance (0.3 to 2.7 micrometer) channels. Each instrument sub-section has six instantaneous fields of view (IFOV) of ~8.5 mrad that provide a contiguous strip three elements wide with a spatial resolution designed to be 3 km from the final MGS mapping orbit altitude of 350 km. The outputs from all TES channels are digitized at 16 bits, processed, and formatted before being sent to the spacecraft Payload Data Subsystem (PDS). The outputs of the interferometer receive the following processing within the instrument before transfer to the PDS: 1) selectable apodization; 2) Fast Fourier Transformation (FFT) of data from all six interferometer channels; 3) correction for gain and offsets; 4) data editing and aggregation; 5) data compression; and 6) formatting for the PDS.

A separate 1.5 cm diameter reflecting telescope, collimated with the main telescope and using the same pointing mirror, is used for the thermal and visible bolometer channels. These channels have similar 3x2 arrays of detectors, that are bore sighted with the spectrometer array. The optical system consists of a single off-axis paraboloidal mirror operating at f/8. A reflecting resonant fork chopper operating at 30 Hz is used to separate the solar reflectance and thermal emission bands. Data from the visual bolometer channel have not undergone full calibration and are not included on this MGS Science Sampler volume.

The TES spectrometer has a noise equivalent spectral radiance near 1.2x10-8 W cm-2 str-1 cm-1. This corresponds to a signal-to-noise ratio (SNR) of 490 at 1000 cm-1 (10 micrometer) viewing a 270K scene. Calibration is achieved by periodic views of space and the internal reference surface using two-temperature calibration methods [Christensen and Harrison, 1993; Ruff et al., 1997]. Absolute radiometric accuracy was estimated from pre-launch data to be better than 4x10-8 W cm-2 str-1 cm-1. However, in-flight observations indicate that a small, systematic calibration offset with a magnitude of ~1x10-7 W cm-2 str-1 cm-1 is present in the TES data presented on this CD-ROM. This error is primarily due to slight variations in the instrument background energy between observations taken of space for calibration and those viewing the planet at an angle 90 degrees away. This error is not significant for surface observations at temperatures above ~240 K. However, for observations of the polar caps and the atmosphere above the limbs, where the radiance is low, this error can be significant.

Data from the Assessment Subphase orbits are included on this MGS Science Sampler Volume as ASCII files that contain calibrated spectral radiance. Note that the thermal bolometer data have undergone a preliminary calibration, although systematic evaluation of the calibration accuracy and precision has not been completed at this time. Also included on this volume are a set of 25-micron global temperature images derived from the calibrated radiance data.

4.4 Magnetometer / Electron Reflectometer (MAG/ER)

The MAG/ER instrumentation provides fast vector measurements (up to 16 samples/sec) of magnetic fields over a very wide dynamic range. The fundamental objectives of this investigation are: (a) establish the nature of the magnetic field of Mars; (b) develop appropriate models for its representation, and (c) to map the crustal remnant field to a resolution consistent with spacecraft orbit altitude and ground track separation. The instrument complement includes two redundant triaxial fluxgate magnetometers and an electron reflectometer. The vector magnetometers provide in-situ sensing of the ambient magnetic field in the vicinity of the MGS spacecraft over the range of 4 nT to 65,536 nT with a digital resolution of 12-bits. The electron reflectometer measures the local electron distribution function in the range of ~1 eV to 20 KeV and will remotely sense the strength of the magnetic field down to the top of the Martian atmosphere using directional information provided by the vector magnetometer. This synergistic combination was designed to increase significantly the sensitivity and spatial resolution achievable from Martian orbit with the vector magnetometer alone. Electron reflection magnetometry was first used on measurements from Apollo 15 and 16 Particles and Fields subsatellites [Anderson et al., 1975; Lin et al., 1976].

Unlike Mars Observer, MGS lacks a boom to separate sensors from the spacecraft body to reduce interference by spacecraft generated magnetic fields. Instead, each magnetometer sensor is placed at the outer edge of the articulated solar panels, approximately 5 meters from the center of the spacecraft bus. The electron reflectometer sensor is mounted directly on the spacecraft nadir panel. This twin magnetometer configuration does not allow the real time estimation of spacecraft fields [Behannon et al., 1977] but provides redundancy and the near real time detection and identification of spacecraft generated magnetic fields. The MAG/ER configuration was dictated by the reduced size of the MGS spacecraft compared to MO and the limited resources available as spares from the Mars Observer Project. It required the design and implementation of extremely magnetically "clean" solar array panels, a challenging task that was accomplished successfully through a close cooperation among engineers, scientists and project personnel. These panels are also used for aerobraking of the MGS spacecraft to achieve the final mapping orbit and are articulated about two orthogonal axes with respect to the spacecraft bus. Therefore, the orientation of the magnetic field sensors with respect to the spacecraft is variable and follows that of the solar panels which are controlled to satisfy a variety of engineering requirements. The magnetometer and electron reflectometer designs have extensive spaceflight heritage and similar versions have been flown in numerous missions: Voyager, AMPTE, Giotto, WIND, FAST, NEAR and recently ACE and Lunar Prospector. The instrument acquires a minimum of 2 to 16 vector samples per second of magnetic field data, depending on the telemetry rate supported. Data acquisition, transmission and command functions are microprocessor controlled and can be partially reprogrammed in flight. The total mass of the MGS MAG/ER instruments is 5.2 kg and the power consumption is 3.7 W.

As noted, the MGS spacecraft was inserted initially into a highly elliptical orbit with apoapsis greater than ten times the radius of Mars and periapses as low as 112 km above the surface. The desired circular mapping orbit will be achieved using atmospheric drag to reduce the spacecraft velocity ("aerobraking"). The significant advantages of aerobraking orbits to the MAG/ER investigation were recognized early in the planning of the observations. Under these conditions the spacecraft was assured to dip below the bottom of the Martian ionosphere and allow the MAG/ER experiment to achieve extraordinary sensitivity and spatial resolution for the detection of possible weak crustal fields. In addition, high plasma densities expected at these low altitudes required a different measurement technique so a Langmuir probe operational mode was added to the electron spectrometer, using the outer case of the electrostatic analyzer as a swept bias collector. MAG/ER results included on the MGS Science Sampler volume consist of plots of derived fields for data collected during the Assessment Subphase orbits. See Acua et al. [1998] for more details.

4.5 Radio Science Subsystem (RSS)

RSS investigations utilize instrumentation with elements on the spacecraft and at the NASA Deep Space Network (DSN). Much of this is shared equipment, used for routine telecommunications as well as for radio science. The performance and calibration of both the spacecraft and tracking stations directly affects the radio science data accuracy, and they play a major role in determining the quality of the results.

Science objectives of radio science investigations include measurement of small perturbations in spacecraft velocity from Doppler shifts on transmitted signals, followed by inference of detailed gravitational fields from solutions of systems of simultaneous equations based on such measurements. During occultations of the spacecraft by the atmosphere, small Doppler shifts can be interpreted as refractivity changes in the atmosphere; and from those, detailed temperature-pressure profiles of the atmosphere can be derived. Investigations of the gravity field and atmospheric structure have been conducted since the first spacecraft visited Mars; they will be continued and expanded with MGS.

The MGS spacecraft radio system was constructed around a redundant pair of X-band Mars Observer Transponders (MOTs). Other components include two redundant Low-Gain Receive antennas; two redundant Low-Gain Transmit antennas; two redundant Command Detector Units; two redundant Traveling Wave Tube Amplifiers; a single high-gain antenna; a single UltraStable Oscillator (USO); miscellaneous cables, connectors, waveguides, and switches; and a Ka-band Link Experiment.

The X-band capability reduced plasma effects on radio signals by a factor of 10 compared with previous S-band systems, but absence of a dual-frequency capability (both S- and X-band) means that plasma effects cannot be estimated and removed from radio data.

The spacecraft is capable of X-band uplink commanding and simultaneous X-band downlink telemetry. The MOT generates 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 behaves as a conventional transponder; its transmitted carrier frequency is derived coherently from the received uplink carrier frequency with a 'turn-around ratio' of 880/749. The nominal coherent downlink frequency is 8417716050 Hz. The coherent mode is used for routine spacecraft tracking; data collected in this mode are the primary inputs to modeling of gravitational fields.

In the non-coherent mode, the downlink carrier frequency is derived from one of the spacecraft's on-board crystal-controlled oscillators. After warm-up, the 'auxiliary' oscillator frequency is estimated to be 8417700000 Hz. A temperature-controlled USO is used as the frequency reference during one-way radio science observations. The measured USO frequency at 1997-08-12T11:30:00 was 8423152962.01 Hz with a drift of +0.182 Hz/day.

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.

For MGS, only X-band is used; earlier missions used lower frequencies or combinations of frequencies; Viking was equipped with both S- and X-band transmitters, but relied on S-band for its primary communications link.

MGS models of the Mars gravity field will begin with models derived from Mariner 9 and Viking Orbiter data. As the newer measurements are folded into the analysis, their higher precision will eventually dominate the solutions and the models will become, in practical terms, MGS models. The data included in this science sampler volume are only from Mariner 9 and Viking observations; they do not contain MGS measurements, and can be considered the baseline gravity models at the beginning of the MGS mission.

More information on the spacecraft and ground radio systems can be found in the MGS Telecommunications System Operations Reference Handbook [Mars Global Surveyor Project, 1996]. More information on the MGS radio science investigations can be found in Tyler et al. [1992].

5. Coordinate Systems and SPICE

5.1 Coordinate Systems

The diverse processing and display requirements for various observational quantities necessitates flexibility in the choice of coordinate system. Two systems are used to describe data products on this science sampler volume:

1. The areocentric coordinate system [Davies et al., 1994], more generally described as planetocentric, is body-centered, using the center-of-mass as the origin. Areocentric latitude is defined by the angle between the equatorial plane and a vector extending from the origin of the coordinate system to the relevant point on the surface. Latitude is measured from -90 degrees at the south pole to +90 degrees at the north pole. Longitude extends from 0 to 360 degrees, with values increasing eastward (i.e., it is a right-handed coordinate system) from the prime meridian [Davies et al., 1994]. This coordinate system is preferred for use in geophysical studies in which, for example, estimates of elevation or gravitational potential are generated mathematically.

2. The areographic system (more generally, the planetographic system) uses the same center-of-mass origin and coordinate axes as the areocentric coordinate system. Areographic latitudes are defined by a vector normal to a reference spheroid surface. Longitudes are measured from the prime meridian and increase toward the west since Mars is a prograde rotator [Davies et al., 1994]. This system is standard for cartography of Mars and most existing maps portray locations of surface features in areographic coordinates. For MGS, the following data have been adopted as standard for defining the reference spheroid for computing the areographic latitudes [Davies et al., 1994]:

Equatorial radius = 3397 km

Polar radius = 3375 km

Flattening = 0.0064763

Note that the flattening is computed as one minus the ratio of the polar radius to the equatorial radius. The relationship between areographic and areocentric latitudes is approximated as:

tan(lc) = (1-f) * (1-f) * tan(lg)


f = flattening

lg = areographic latitude

lc = areocentric latitude


NASA's SPICE system provides engineering data sets and allied software useful in analyzing space science data. (SPICE is also used for mission design, observation planning and data visualization.) SPICE data files--often called SPICE kernels--provide access to information such as spacecraft and planet/satellite (target body) ephemerides, target body size/shape/orientation, spacecraft orientation and instrument pointing, and sequence of events. SPICE kernels are included on the Science Sampler in the GEOMETRY directory.

The SPICE Toolkit, a subroutine library available in FORTRAN and C, provides access to the data in SPICE files plus the ability to use these data to compute derived (higher level) quantities such as lighting angles, altitude, sub-spacecraft latitude/longitude, etc. A scientist integrates the needed Toolkit subroutines into his/her own application program. The SPICE Toolkit is not included on the Science Sampler. It may be obtained from the Navigation and Ancillary Information Facility (NAIF) organization at NASA's Jet Propulsion Laboratory by visiting the Web site or by sending electronic mail to

PDS data sets generally contain labels that include a number of geometric parameters computed from early versions of SPICE files; sometimes these are sufficient for a scientist to both find and utilize the science data needed for some investigation. But someone interested in computing additional observation geometry parameters, or in using the very latest engineering data, can acquire the appropriate flight project SPICE files and the SPICE Toolkit software and use these to compute or update the parameters to be used.

This MGS Sampler CD contains SPICE kernel files for the Assessment Subphase of the MGS mission, in formats for both Unix and PC platforms. A brief summary of these files is shown in the tables below.

Kernel Type Kernel Name Size (Mb) Comments
SPK mgs_aero.bsp 10.8 Ephemeris for the MGS spacecraft (ID=-94) relative to the Mars barycenter (ID=4)
  de403s.bsp 3.4 Ephemeris for Mars barycenter (ID=4), Mars (ID=499), Earth barycenter (ID=3), Earth (ID=399) and Sun (ID=10) relative to solar system barycenter (ID=0). (Also contains data for all other planet barycenters and the Earth's moon.)
PCK pck00006.tpc 0.09 Size, shape and orientation for Mars (and all other planets and satellites)
CK m013.bc 2.1 Spacecraft orientation for a period of time centered at periapsis number 13. SPICE ID code for the spacecraft-fixed frame is -94000.
  m019-036.bc 25.0 Spacecraft orientation for a period of time centered at periapsis numbers 19 - 36
  m073.bc 2.7 Spacecraft orientation for a period of time centered at periapsis number 73
  m077-079.bc 3.4 Spacecraft orientation for a period of time centered at periapsis numbers 77-79
  m084.bc 1.9 Spacecraft orientation for a period of time centered at periapsis number 84
  m100.bc 1.7 Spacecraft orientation for a period of time centered at periapsis number 100
LSK naif0006.tls 0.004 Current Leapseconds kernel, used in covering between UTC and ephemeris time (ET) time systems
SCLK mgs_sclk.tsc 0.003 Current SPICE Spacecraft Clock file, used in converting between spacecraft clock (SCLK) and ET time systems

The CK files start prior to the apoapsis time preceding the (first) periapsis number shown and extend to past the apoapsis following the (last) periapsis.

The approximate time coverage of these SPICE kernels is shown below. As is typical for most flight projects and other endeavors where SPICE data are used, the files do not have the same coverage, and, in some cases, multiple files are needed to cover the time span or set of objects of interest. This is not a problem since SPICE Toolkit software is able to handle at run time multiple files of a given type (as well as all of the types of kernels).

Kernel Name UTC Start Time UTC Stop Time
mgs_aero.bsp 1997 SEP 12 01:42:00 1998 MAR 26 05:23:00
de403s.bsp 1979 DEC 02 2011 JAN 06
pck00006.tpc N/A N/A
m013.bc 1997 OCT 03 00:00:02 1997 OCT 04 23:59:58
m019-036.bc 1997 OCT 12 00:00:02 1997 NOV 07 23:59:58
m073.bc 1997 DEC 24 00:00:02 1997 DEC 26 23:59:57
m077-079.bc 1997 DEC 29 03:09:07 1998 JAN 01 23:59:57
m084.bc 1998 JAN 05 00:00:02 1998 JAN 06 21:44:17
m100.bc 1998 JAN 20 00:00:02 1998 JAN 21 22:43:01
naif0006.tls N/A N/A
mgs_sclk.tsc 1996 NOV 05 18:14:45 1998 MAY 22 13:45:00

The coverage for each C-kernel is not continuous--gaps of a few seconds to a few minutes are found in some files. A summary of the gaps is contained in the internal documentation area (the so-called "comment area") of each CK file. These comments can be viewed using either the "commnt" or the "spacit" utility program available in the SPICE Toolkit.

Instructions and examples for reading these SPICE kernel files are found in documents included with the SPICE Toolkit. Particularly noted are the following:

Document or Code File Name Document or Code Title and Description Most Useful SPICELIB Subroutines: Contains brief descriptions and some examples of how to use the most commonly used SPICELIB subroutines.
states A "cookbook" program giving a simple example of reading a SPICE ephemeris (SPK) file. See also the User's Guide document (
subpt A "cookbook" program giving an example of how to compute the sub-observer point (planetocentric latitude and longitude) on a target body. See also the User's Guide document (
spk.req Detailed reference document for the SPICE ephemeris system.
pck.req Detailed reference document for the SPICE planetary constants system.
ck.req Detailed reference document for the SPICE spacecraft orientation system.
sclk.req Detailed reference document for the SPICE spacecraft clock system.
kernels.req Overview document for the SPICE text kernels, including LSK, SCLK, PCK.
spkez.f, tipbod.f, sce2s.f, ckgp.f, str2et.f, et2utc.f These are the principal SPICELIB subroutines (FORTRAN source code modules) used in accessing the ephemeris, target orientation, spacecraft orientation and time conversion kernel files. Each has an extensive set of specifications for its use in a header located at the top of the file. (Each of these is also briefly mentioned in the "Most Useful SPICELIB Subroutines" document.)

A SPICE user must be prepared to write one's own application program, incorporating those SPICELIB subroutines needed to accomplish whatever task is at hand. Having some knowledge of mathematics and terminology commonly used in space science work would be very helpful.

All SPICE Toolkit code is COPYRIGHT California Institute of Technology, U.S. Government sponsorship acknowledged. All SPICE kernel files meet distribution and export criteria covered under GTDA and ITAR regulations.

Development of the SPICE system is carried out by the Jet Propulsion Laboratory, California Institute of Technology, under contract with the National Aeronautics and Space Administration.

Included on the Science Sampler in the GEOMETRY directory are files containing sub-spacecraft locations for each Assessment Subphase orbit, given in both areocentric and areographic coordinates. Also in the GEOMETRY directory is a summary file of the time and sub-spacecraft location at periapsis for each orbit. These files were generated by the Navigation and Ancillary Information Facility using SPICE.

6. Archive Data Preparation

Most of the data sets on this volume have been generated specifically for inclusion in the Science Sampler Data Set Collection. They are designated by the word "Sampler" as part of the Data Set ID field in their PDS labels. The products in these data sets will be re-released at a later date, as part of each instrument team's required standard product delivery to the Planetary Data System. Formats of standard products may differ from the formats of products in the Sampler collection.

The Radio Science gravity models are included on this volume as examples of future MGS products. They are based on Mariner 9 and Viking data, and do not contain any MGS data (see section 4.5, Radio Science Subsystem).

The following sections describe the data sets in the Sampler collection, including discussion of the processing applied to the data. For TES products, a formula is given for converting calibrated radiance to effective surface kinetic temperature.

All images on this volume are in GIF format, except the Radio Science gravity map. While GIF will not be used as an archive format for standard products on future volumes, it is used on the Science Sampler to provide a convenient means for a quick look at the data, as GIF images can easily be viewed with a Web browser.

None of the data products on this volume are compressed.

6.1 Mars Orbiter Camera (MOC)

The nine MOC narrow angle images contained on the MGS Science Sampler volume have been cosmetically corrected to reduce artifacts found in the raw image data. The digital image processing procedures applied to these data include corrections for the varying response of each detector of the push-broom imaging system, an aspect ratio correction, and a high-pass boxcar filter to balance the scene brightness of the image. Although the processing has greatly improved the appearance and interpretability of an image, the procedures do not provide for a rigorous radiometric, geometric, or photometric rectification that would allow quantitative interpretation of the data.

Each detector in the push-broom imaging system has a varying response in sensitivity and bias. The visual result seen in an unprocessed image is dark and light vertical (top to bottom) "streaks". A first order correction has been applied to the images to minimize the "streaks". The corrective procedure first determined the average brightness value for each column in the image array. Next, each pixel of the column was divided by the column average and then multiplied by the average brightness of the entire scene. This method worked well in most circumstances; however, scene dependent brightness variations can cause residual "streaking" in the processed image.

The downtrack and crosstrack pixel resolutions of a MOC image are different, resulting in pixels that are rectangular on the surface rather than square. The crosstrack to downtrack ratio, or aspect ratio, typically ranges from 5.0 to 1.0 but is generally about 1.5 for the MOC images on the MGS Science Sampler volume. The crosstrack resolution is dependent on spacecraft altitude and the camera focal length and pointing, whereas the downtrack resolution is dependent on the integration time of the push-broom array and the relative velocity of the spacecraft as it sweeps over the Martian surface. Other factors that affect the resolution are the camera operating modes that allow for pixel summing in either the crosstrack or downtrack direction. A simple aspect ratio correction has been applied by resampling the image to adjust the crosstrack resolution to match the downtrack resolution. The correction decreases the resolution in the crosstrack direction (resulting in fewer pixels) using a simple bilinear interpolation method. An alternate aspect ratio correction was considered that would expand the pixels in the downtrack direction but this would have provided for much larger images that are already very large. Many images on the MGS Science Sampler volume are more than 4,000 rows (lines) and it was thought that increasing the size of the images would potentially make it more difficult for investigators to display and use the images.

Many of the MOC images contain brightness variations due to solar illumination differences over the image array. A high-pass boxcar filter was applied to adjust the scene brightness in order to reduce these solar illumination variations. A 301 x 301 high-pass boxcar filter was applied to the data and an average of the original scene and the high-pass filtered scene was used to create the final images. This technique additionally adjusted large scene brightness differences due to surface albedo variations.

The nine narrow angle MOC images included in the volume are GIF files with detached PDS labels. The labels include areographic coordinates for centers of the frames, camera gain and offset information, and timing of data acquisition. The latitude and longitude value in the labels are for the most part based on locating the MOC images in the Viking MDIM data set. Analysis of MOLA data and the location of the Mars Pathfinder landing site indicate that locations in the MDIM data set have an uncertainty of several tenths of a degree or about 10 to 20 km. Thus, the MOC image locations listed in the PDS labels on this sampler CD-ROM should be considered as approximate locations.

Nine "finder" frames, generated from Mars Digital Image Mosaics, are included in the BROWSE directory of this volume. These are also GIF files. They show the approximate locations of the MOC footprints. PDS labels for the finder frames have image coordinates and map projection information. (Note that image coordinates are given by the PDS keywords MINIMUM_LATITUDE, MAXIMUM_LATITUDE, EASTERNMOST_LONGITUDE, and WESTERNMOST_LONGITUDE. The keywords CENTER_LATITUDE and CENTER_LONGITUDE refer to the center of the map projection, which is not necessarily the center of the image.) The finder frames can be viewed using a web browser See section 7, Disk Directory Structure, for directions for exploring the volume with a web browser using files in the BROWSE directory.

6.2 Mars Orbiter Laser Altimeter (MOLA)

The topography of Mars relative to the gravitational potential was determined by the MOLA instrument during 18 assessment orbits over the northern hemisphere, i.e., near periapsis. Overall more than 221,000 valid laser ranges were collected. While the instrument was designed to collect data in a mapping orbit of approximately 400 km altitude, MOLA performed well out to its 785 km maximum range. This science sampler volume contains the complete archival delivery of MOLA data for the Assessment Subphase of the MGS Mission. The primary standard products are the Aggregated Experiment Data Record (AEDRs) and the Precision Experiment Data Record (PEDR). The Software Interface Specification (SIS) documents for both of these standard products are included on this volume. Further, a simplified version of the PEDR is included as a set of ASCII tables.

AEDRs contain raw MOLA engineering and science data, in a time-ordered series. The PEDR tables contain instrument science data, sub-spacecraft location (in both areocentric and areographic coordinates), estimates of the planetary radii, and elevations relative to the areoid. Areoid heights were calculated using the GMM-1 gravity model for a mean equatorial radius of 3396 km.

The PEDRs incorporate the best multi-arc orbital solutions derived from the Goddard Mars potential model GMM-1, and the available tracking. The latest (January 20, 1998) spacecraft SCLK timing corrections have been applied. The ranges account for instrument delays and the leading edge timing biases, estimated by the receiver model of Abshire and Sun [1997]. This model assumes a Gaussian shape for the transmitted and surface-scattered pulse waveforms, using the detector threshold settings and the observed pulse width and energy measurements between the threshold crossings to infer the true pulse centroid, width, and amplitude. The eccentric orbit brought MOLA much closer to the surface of Mars than the design called for, thus the pulse width and energy measurements were saturated for much of each pass. Caution must be exercised when interpreting these measurements. Laser energies are calculated according to the transmitter model of Afzal et al. [1997]. A post-launch calibration to the MOLA oscillator frequency has been applied, based on the difference between the spacecraft high-resolution timer and the MOLA clock, resulting in an estimated frequency of 99,996,232 5 Hz. Barring new information about the receiver characteristics, the ranges are final.

Time tags are given in ET seconds of MOLA pulse receipt time, a few milliseconds after the laser fire time. Timing of the shots is resolved to ~100 microseconds, an essential element due to the highly elliptical contingency-orbit geometry. While the absolute spacecraft time is not guaranteed to this accuracy, we have no way as yet to assess timing errors independently. The resolution of the data is about 40 cm vertically, and about 330 m along-track, limited by the 10 Hz firing rate of the laser. The absolute, long-wavelength radial orbit error is estimated to be about 30 m. The uncertainty in absolute ground spot location is limited by the attitude knowledge of the spacecraft, and is estimated to be about 400 m at a nominal range of 400 km.

Due to the inverse-square-law energy return in the link equation [Zuber et al., 1992], the instrument detector was saturated during a part of the periapsis approach. Received pulse energy and pulse width are resolved during the portion of the pass when the detector is not saturated. The absolute accuracy of these quantities is about 5%.

There is a table entry for each non-zero shot range detection for all in-range packets in the contingency pass data stream. Pass 35 contains 13 corrupted range values in packet 739, otherwise, shots are either valid ranges or false detections due to the intrinsic noise characteristics of the receiver.

Plots of the elevation derived from MOLA data, relative to an areoid model, are given in the BROWSE/MOLA directory of this volume. The plots show elevation in kilometers versus latitude, with colors black, red, green and blue corresponding to MOLA first trigger channels 1, 2, 3, and 4 respectively.

Users of MOLA data must be aware of two important differences between the MOLA coordinate system and the Viking-era coordinates. These differences are significant when comparing MOLA groundtracks to MDIMs, USGS DTMs, or maps. MOLA uses the areocentric coordinate frame (section 5). MOLA areocentric latitudes should be converted to areographic latitudes using the equation provided above. (Note that Viking data was processed assuming different radii than given in this document: equatorial radius = 3393.40 km and polar radius = 3375.73 km for a flattening of 1/192.) The ASCII tables on this CD-ROM provide both areocentric and areographic latitudes for each MOLA shot; areographic latitudes are calculated for the same flattening as the MDIMs. There appears to be a residual discrepancy in latitude of less than 0.1 degree (6 km) magnitude, and variable sign, between MGS and Viking coordinates. MOLA longitudes are also areocentric, with positive degrees east. However, there is an additional eastward offset of the Viking-era coordinate system relative to the present MGS inertial frame. The magnitude of this offset ranges from about 0.1 to 0.3 degrees (<20 km). More than one factor may contribute to this discrepancy; the primary reason is a change in the IAU coordinate system. Other possible effects are a drift of the prime meridian due to uncertainties in the martian rotation period or errors in the Viking spacecraft orbital position that propagated through the image processing [Smith et al., 1998]. Adding 0.2 degrees east longitude to the MOLA coordinates is a first-order correction for comparison to Viking data.

6.3 Thermal Emission Spectrometer (TES)

6.3.1 Calibrated Spectral Radiance

The TES calibrated radiance data ("single scans") for the assessment orbits cover a spectral range of 148.1323 to 1650.6176 cm-1. The data set includes 143 spectral values over this range resulting in a spectral resolution of 10.5809 cm-1. The wavenumber values corresponding to the radiance estimates are found in the file TESWAVEN.TAB in the TES data directory on this volume.

Note that in the calibrated radiance data, latitude and longitude values are given in areocentric coordinates, with longitude increasing toward the west. This is in contrast to the standard areocentric coordinate system described in section 5, in which longitude increases toward the east.

A simple algorithm was performed on each TES spectrum in order to estimate the effective surface kinetic temperature using the TES spectrometer data. (This temperature is included in the TES calibrated radiance data products, in the field labeled "surface_temperature".) The primary use of this temperature is for emissivity determination in which only a first-order estimate of the surface temperature is required. Therefore, no attempt is made to model mixtures of surface materials at different kinetic temperatures, nor to remove atmospheric effects. The best spectral region to determine surface kinetic temperature is in the 6-8 micrometer region where most geologic materials have emissivities close to unity. Unfortunately, for ground temperatures less than approximately 225 K, the signal-to-noise of the TES instrument is low in this spectral region, and accurate temperatures cannot be determined. Therefore, an algorithm has been developed that uses the 6-8 micrometer region for warm surfaces, and uses the 20-30 micrometer region for surface temperatures below 215 K. The following steps outline the details of this algorithm.

1) Convert the calibrated radiance to brightness temperature at each wavenumber assuming that:

a) the emissivity is unity (temp. = TB); and

b) the emissivity is 0.95 and dividing the calibrated radiance by this value before determining the brightness temperature (temp. = TB'). Filter the brightness temperatures using a boxcar filter 7 samples wide to reduce noise effects.

2) Find the maximum brightness temperature over the sample ranges from:

a) 300 to 500 cm-1 and 800 to 1350 cm-1. This range was selected to include both the long and short wavelength portions of the spectrum, to include the wavenumber typically with the highest brightness temperature (~1300 cm-1) as determined by both the Mariner 9 IRIS and the preliminary TES data, and to eliminate the CO2 15 micrometer feature.

b) 300 to 500 cm-1. This range covers only the long wavelength portion of the spectrum.

3) If TB is > T2 (225 K), set Teffective to TB; If TB' is < T1 (215 K) set Teffective to TB'. Else set Teffective to weighted average of TB and TB'. Weighting is determined by:

Weight1 = 1 - ( (T2-TB) / (T2-T1) )

Weight2 = 1 - ( (TB'-T1) / (T2-T1) )

If Weight1 or Weight2 < 0, then they are set to 0.

4) Finally:

Teffective = ( (TB*Weight1) + (TB'*Weight2) ) / (Weight1 + Weight2)

6.3.2 Global Temperature Images

During the majority of the elliptical orbit around Mars, the spacecraft is in array normal spin (ANS) mode. In this mode the spacecraft executes a rotation about the +x axis once every 100 minutes. This rolling of the spacecraft rotates the payload platform (the +z axis) through a full 360 degrees. Depending on the distance between the spacecraft and Mars, it is possible for the TES mirror to scan Mars for approximately 7 minutes during each ANS roll. Using multiple mirror angles to scan across the disk of the planet and the rotation of the spacecraft to offset the cross-planet scans in the "down-planet" direction it is possible to build up a full disk data set during each ANS roll. The full disk collection of data is termed 'sample' mode. During the two rolls close to periapsis (first outbound and last inbound) the instrument collects higher spatial resolution data by constraining the step size of the mirror scan, producing a narrow strip image instead of the full disk. The higher resolution collection of data is termed 'map' mode. One roll for each orbit is collected in the 4 sec. cycle time ("double", i.e. twice the spectral resolution), and all the rest are collected using the 2 sec. cycle time ("single").

Full spectral coverage is acquired during each ANS scan. The 25 micron band was used to produce the global temperature images included on this volume for two reasons: (1) for cold temperatures (mainly viewing the south pole during the majority of the rolls) the 25 micron band has the best signal to noise ratio, and (2) it is approximately similar to the Viking IRTM 20 micron band.

A color/temperature scale bar is located at the bottom of each image.

The table below summarizes the contents of the temperature images.

of Rolls
Rolls with
Map Mode Data
21   20   1, 20   6    
22   21   1, 21   7    
23   21   1, 21   7    
24   20   1, 20   7    
25   21   1, 20   7    
26   21   1, 20   7    
27   21   1, 21   7   roll 14 missing
28   21   1, 21   7    
29   21   1, 21   7    
30   14   14     roll 1 sequence error
31   20   1, 20   6    
32   20   1, 20   6   roll 17 missing
33   18   1, 18    
34   20   1, 20   6  
35   19   1, 19   6  
36   19   1, 19     roll 11 and most of roll 13 missing

6.4 Magnetometer / Electron Reflectometer (MAG/ER)

MAG/ER data included on this volume consist of a suite of GIF files that are plots of derived electron and magnetic field observations acquired during the assessment orbits [see: Acua et al., 1998]. The plots include electron fluxes at 10, 50, 130, 300, and 1000 eV, a color spectrogram of the same data, and magnetic field amplitude and rms, and spacecraft altitude. The numerical data will be released later on MAG/ER archive volumes, after further validation and calibration procedures have been applied.

The start and stop times in the PDS labels for the MAG/ER files are approximate and are based on inspection of the plots.

6.5 Radio Science Subsystem (RSS)

There are two spherical harmonic gravity solutions in this archive. One model is GGM50A01.SHA (or GMM-1), which was produced at Goddard Space Flight Center under the direction of David E. Smith. It is a 50th degree and order gravity model. The second, JGM50C01.SHA (or JPL-50C), is also a 50th degree and order model produced at the Jet Propulsion Laboratory under the direction of William L. Sjogren. The two models are essentially identical (i.e., the coefficient values) out to the 20th degree and order (i.e., lower frequencies), but were processed slightly differently by the two groups (i.e., arclength, adjustable parameters). The big difference though was in terms of how the solutions were obtained, specifically what type of a priori constraint was used to obtain the two solutions. A global constraint was used in producing the GGM50A01.SHA solution, whereas, a spatial constraint was used in the JGM50C01.SHA solution. Specifically, in the JGM50C01 solution, the Kaula constraint was removed over certain regions of the planet.

6.5.1 The GMM-1 Gravity Solution

The gravitational signature of Mars was determined from velocity perturbations of the Mariner 9 and Viking 1 and 2 spacecraft as measured from the Doppler shift of the S-band radio tracking signal. The Doppler data from the DSN stations were acquired at the rate of one point per 60 seconds near periapsis and one point per 600 seconds near apoapsis.

The GMM-1 gravity solution (GGM50A01.SHA) consists of 232,323 observations, of which 49,878 were contributed by Mariner 9, while 95,370 and 87,075 were contributed by Vikings 1 and 2 respectively. The data were divided into 270 spans or independent arcs (spanning 1142 days) based on considerations of data coverage and timing of maneuvers. The table below summarizes the number of observations and arcs from each spacecraft:

Satellite   Periapsis
Altitude (km)
of Arcs
Avg. Arc
Length (days)
Mariner 9   1500 64.4 32 4.3 49,878
Viking 1   1500 38.2 29 4.2 31,393
Viking 1   300 39.1 95 4.8 63,977
Viking 2   1400 55.4 12 3.8 11,878
Viking 2   1500 75.1 11 4.7 10,467
Viking 2   778 80.1 54 3.8 35,375
Viking 2   300 80.2 37 4.2 29,355
Total       270   232,323

For each arc certain parameters were determined: the spacecraft state (position and velocity), a solar radiation pressure coefficient, Doppler biases for each station over the arc to account for frequency biases, and for the 300 km periapsis altitude arcs, a drag coefficient was adjusted on each day. The a priori force model that was used included a preliminary gravity solution obtained from the analysis of these data, a 50 x 50 solution designated mgm0340. In addition, third body perturbations were modeled, the solar radiation pressure perturbation was applied, and the appropriate relativistic perturbations in the measurement model, and the force model were included. The DE200 set of planetary and lunar ephemerides was used in these analyses. The planet orientation model was based on the 1988 IAU system of constants, and a reference radius of 3394.2 km.

The data in GMM-1 were weighted at 0.8 to 4.1 cm/s, and the data weights were calibrated using the subset calibration method of Lerch [1991]. Although each data arc was typically fit to the level of a few mm/s, the data were downweighted in this fashion in order to attenuate the power of the high degree terms, and account for any systematic mismodeling that might still be present in the data. The solution was also derived using a power law rule (Kaula constraint of 13 x 105/L2) [Kaula, 1966], where L is the spherical harmonic degree. Without this constraint, the high degree terms develop excessive power.

More information on this Mars gravity solution can be found in Smith et al. [1993].

6.5.2 The JPL-50C Gravity Solution

The gravitational signature of Mars was determined from velocity perturbations of the Mariner 9 and Viking 1 and 2 spacecraft as measured from the Doppler shift of the S-band radio tracking signal. The Doppler data from the DSN stations were acquired at the rate of one point per 60 seconds.

The JPL-50C gravity solution (JGM50C01.SHA) consists of 325,005 observations, of which 55,984 were contributed by Mariner 9, while 157,125 and 111,896 were contributed by Vikings 1 and 2 respectively. The data were divided into 372 spans or independent arcs based on considerations of data coverage and timing of maneuvers. The table below summarizes the number of observations and arcs from each spacecraft:

Satellite Number
of Arcs
Avg. Arc
Length (hours)
Mariner 9   50 72 55,984
Viking 1   189 96 157,125
Viking 2   133 96 111,896
Total   372   325,005

For each arc certain parameters were determined: the spacecraft state (position and velocity), a solar radiation pressure coefficient, Doppler biases for each station over the arc to account for frequency biases, and the mismodeling of the effects of the troposphere and ionosphere on the Doppler signal. The a priori force model that was used included the GMM-1 (Goddard Mars gravity model), and included the third-body perturbations due to the Sun, all the planets (including Earth), the solar radiation pressure perturbation, and appropriate relativistic effects. The DE200 set of planetary and lunar ephemerides was used in the analyses.

The data in JPL-50C were weighted at 2 to 5 mm/s. Although each data arc was typically fit to the level of 1 mm/s, the data were downweighted in this fashion in order to attenuate the power of the high degree terms, and account for any systematic mismodeling that might still be present in the data. The solution was also derived using a power law rule (Kaula constraint of 13 x 105/L2) [Kaula, 1966], where L is the spherical harmonic degree. Without this constraint, the high degree terms develop excessive power. A degree strength a priori was also used in selected areas at the poles and Olympus Mons. This is a spatial constraint and is described in Konopliv and Sjogren [1995], in which a detailed report is presented for the entire determination of this gravity field.

In the process of deriving JPL-50C, extensive experiments were performed in order to select the a priori weights for the sets of data in the solution - and care was taken to downweight or delete data that produced spurious signals in the anomaly maps. The gravity anomalies in this model were evaluated at the Martian surface.

6.5.3 The JPL-50C Gravity Map

The JPL gravity map (JGM50C01.IMG) contains a digital map of Mars gravity accelerations derived from radio tracking of the Mariner and Viking spacecraft, based on the JPL-50C gravity model. The map shows vertical gravity in milligals at the surface referenced to a sphere of 3394.2 km radius.

7. Disk Directory Structure

The volume and directory structure of this CD-ROM conforms to the ISO-9660 standard, except that files do not contain any Extended Attribute Records (XAR). The directory content, file format, and supporting documentation conform to PDS standards version 3.2 [Planetary Data System, 1995]. The AAREADME.TXT file in the top-level directory has an outline of the archive directory structure. File names within this archive conform to the "8.3" convention.

In the top-level directory of this volume, a brief documentation file appears in both a plain text (AAREADME.TXT) and a hypertext version (AAREADME.HTM) as an introduction to the volume. Instrument data, browse data, ancillary files and documentation are located in separate directories, as follows.

The MGS science data are located in five subdirectories, one for each instrument: MAGER, MOC, MOLA, RSS, and TES. The MOLA directory is further divided into three subdirectories, one for each MOLA data set (AEDRs, PEDRs, and PEDR ASCII tables). The TES directory is also divided in two subdirectories, one for the calibrated radiance data products and another for the temperature images.

The BROWSE directory contains a suite of files useful for visualizing the science data using a Web browser. To display these files, begin by using a Web browser to open the file INDEX.HTM in the BROWSE directory. This file provides an introduction to the products and hypertext links to data and documentation files.

The CATALOG directory contains PDS catalog objects defining various map projections used for some images on this volume. PDS labels for images in the BROWSE/MOC directory and for the Radio Science gravity map in RSS/IMG/JGM50C01.IMG refer to these catalog objects.

The INDEX directory contains an index table listing the data set ID, instrument name, orbit number, and other information for all the data product files present on this volume, in a format suitable for loading into a data base management system.

The DOCUMENT directory includes this document in a plain text version (VOLINFO.TXT) and a hypertext version (VOLINFO.HTM) for use with a Web browser. It also includes the MOLA AEDR and PEDR Software Interface Specifications as hypertext files and as Adobe PDF files.

The GEOMETRY directory contains SPICE kernels relevant to the data products on this volume (see section 5.2, SPICE). It also contains a set of tables listing the sub-spacecraft latitude and longitude at one-minute intervals for orbits 19 through 36, in both areocentric and areographic coordinates (see section 5.1, Coordinate Systems). Also in this directory is a summary of the periapsis times and coordinates for orbits 2 through 100, in the file ORBITSUM.TAB.

The LABEL directory contains files that are referenced by the PDS labels for MOLA AEDR and PEDR data products.

8. File Names and Formats

File names for products included on this volume follow a standard naming convention. Specifically, the name begins with an acronym that identifies the instrument, followed by the orbit number (orbit number 1 is the first orbit after insertion) and then a subfield showing the type of data product or the observation number within the orbit, as appropriate. File name extensions identify the format and version number. The following table summarizes the file naming convention.

Instrument File Name Notes
MOC NAxxxxyy.GIF narrow angle, xxxx = orbit number, yy = image number
MOLA AAxxxxxF.B AEDR file names
  APxxxxxV.B PEDR file names
  APxxxxxV.TAB ASCII equivalent of PEDRs
TES TESxxxxR.TAB r = radiance
  TESxxxxT.GIF t = temperature
MAG/ER MAGxxxxP.GIF p = plot
RSS GGM50A01.SHA Goddard Model
  JGM50C01.SHA JPL Model
NAIF ORBITSUM.TAB summary of all orbits
  ORBxxxxC.TAB areocentric data for each orbit
  ORBxxxxG.TAB areographic data for each orbit

All documentation, detached PDS labels, and hypertext (*.HTM) files on this volume are stream format files, with a carriage return (ASCII 13) and a line feed character (ASCII 10) at the end of the record. All of these files can be viewed with any standard text editor or Web browser.

All images on this volume are in GIF format. They also can be viewed with any standard Web browser. Data in tabular form can be viewed in a Web browser but might be more conveniently used when loaded into a spreadsheet program. The accompanying PDS labels for these files contain descriptions of the table format. The only data products not directly viewable in a Web browser are the MOLA AEDR and PEDR binary files. However, ASCII text versions of the PEDRs in tabular form have been provided.

Note that to view PDS labels and table data in a Web browser, the browser may need to be configured to recognize files with extensions ".LBL" and ".TAB" as text files.

Each data file on this volume is described by a PDS label, either embedded at the beginning of the file or detached in a separate file of the same name but with the extension ".LBL".

PDS labels are object-oriented. The object to which the label refers (e.g., IMAGE, TABLE, etc.) is denoted by a statement of the form:

^object = location

in which the carat character (^, also called a pointer in this context) indicates that the object starts at the given location. For an object in the same file as the label, the location is an integer representing the starting record number of the object (the first record in the file is record 1). For an object located outside the label file, the location denotes the name of the file containing the object, along with the starting record or byte number. For example:

^IMAGE = ("C102.IMG",3)

indicates that the IMAGE object begins at record 3 of the file C102.IMG, in the same directory as the detached label file. Additional information about PDS labels is found in the Planetary Data System Standards Reference [1995].

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