2. Mission Background

Index to Section 2: Background
2. Mission Background
    2.1. Mission Synopsis
    2.2. Mission Phases
      2.2.1. Launch Phase
      2.2.2. Cruise Phase
      2.2.3. Orbit Insertion Phase
      2.2.4. Mapping Phase
      2.2.5. Relay Support Phase and Quarantine Period
      2.2.6. Velocity Change Budget
    2.3. Science Investigations and Descriptions
      2.3.1. Scientific Objectives and Payload Overview
      2.3.2. Magnetometer / Electron Reflectometer
      2.3.3. Mars Orbiter Camera
      2.3.4. Mars Orbiter Laser Altimeter
      2.3.5. Mars Relay
      2.3.6. Radio Science
      2.3.7. Thermal Emission Spectrometer
    2.4. Spacecraft Description
      2.4.1. Spacecraft Configuration
      2.4.2. Command and Data Handling
      2.4.3. Attitude Control
      2.4.4. Telecommunications
      2.4.5. Propulsion
      2.4.6. Power
    2.5. DSN Utilization


The Mars Global Surveyor (MGS) mission will deliver a single spacecraft to Mars for an extended orbital study of the planet's surface, atmosphere, and gravitational and magnetic fields. Achieving the scientific objectives of the mission will require delivering the spacecraft to a low-altitude, near-polar, sun-synchronous orbit and returning data over a complete Martian year. Data returned from six prime experiments on the spacecraft will provide for a better understanding of the geology, geophysics, and climatology of Mars. Five of those six will utilize a single scientific instrument. The sixth investigation will collect data about Mars by analyzing the spacecraft's radio signal as it reaches Earth.

2.1. Mission Synopsis

A Delta 2/7925A launch vehicle will boost the Mars Global Surveyor spacecraft from Cape Canaveral Air Force Station (CCAFS) during the November 1996 launch opportunity. The spacecraft will utilize a Type 2 transfer trajectory with a trans-Mars flight time of about ten months. After arriving at the red planet in September 1997, MGS will be propulsively inserted into an initial, highly elliptical capture orbit with a period of 48 hours. Over a the next five months, the spacecraft will be gradually lowered into the mapping orbit by the use of aerobraking methods. This technique works by dipping the spacecraft into Mars' upper atmosphere on every periapsis passage in order to slow down and lower apoapsis.

Mars Global Surveyor will utilize a sun-synchronous mapping orbit at a 378 km index altitude, and with a descending node orientation of 2:00 p.m. with respect to the fictitious mean Sun. In this 92.9&176 inclination orbit, the MGS spacecraft will circle the red planet once every 117.65 minutes. Once every seven Martian days (sols), the spacecraft will approximately retrace its ground track. After each seven sol cycle (88 orbits), the ground track pattern will be offset by 59 km eastward from the tracks on the previous cycle. If the orbit is perfectly maintained, then the entire network of ground track patterns will nearly repeat itself after a time span of about 26 sols (327 orbits), and almost exactly repeat after 550 sols (6917 orbits). Such a repeat cycle scheme will allow 99.9% global coverage to be built up from repeated instrument swaths across the planet.

Repetitive observations of the planet's surface and atmosphere from the mapping orbit will be conducted over a time span of one complete Martian year (687 Earth days). Throughout this entire period, the spacecraft remain in an orientation with the scientific instruments nadir pointed. Because MGS lacks a scan platform, any scanning capability will be provided by the instruments. The normal sequence of collecting science data will involve recording continuously for about 24 hours, and then playing it back through the Deep Space Network (DSN) during a 10-hour tracking pass once every day. Approximately every third day, an additional tracking pass will be scheduled to return high-rate, real-time data. Currently, primary mapping operations will occur between April 1998 and February 2000. From the end of mapping until the end of mission in January 2003, MGS is scheduled to support the Mars Surveyor Program by relaying data from various landers and atmospheric vehicles to the Earth.

Figure 2-1: General Timeline for MGS Mission

2.2. Mission Phases

Five mission phases have been defined to simplify the description of different periods of activity. These are the launch, cruise, orbit insertion, mapping, and relay. Several time epochs will be used throughout this document for defining activities and the boundaries of some of the mission phases and sub- phases. Launch (L) denotes the liftoff time, defined to be the instant of ignition of the Delta 2's first stage. Injection (TMI) represents the burn-out time of the Delta 2's third stage, used for the trans-Mars injection burn that place MGS onto a trajectory bound for the red planet. Mars Orbit Insertion (MOI) designates the time that the spacecraft begins the propulsive maneuver to slow down and enter Martian orbit after the completion of its interplanetary trajectory from Earth. End of Mission (EOM) symbolizes the end of ground operations to control spacecraft activities and collect data. The following table summarizes the dates of the mission phases and some key mission events. The dates in the table are specific to a mission that launches at the opening of the launch period on 4 November 1996. Under the current design, MGS can launch as late as 25 November 1996.

Event               Date                          Comments                             
Launch              4 Nov 1996                    Launch period opens on 4 Nov 1996    
                                                  and closes on 25 Nov 1996            
Inner Cruise Phase  4 Nov 1996 to 3 Jan 1997      Communications through LGA only      
                                                  because solar arrays must be         
                                                  pointed at a fixed angle to Sun      
TCM1                19 Nov 1996  (L+ 15 days)     Trajectory correct for injection     
Outer Cruise Phase  3 Jan 1997 to 10 Sep 1997     Communications through HGA, phase    
                                                  begins when Earth-MGS-Sun angle      
                                                  falls below 60 degrees               
TCM2                19 Mar 97 (TCM1+ 120 days)    Correct for execution errors from    
TCM3                18 Apr 97 (TCM2+ 30 days)     Correct for execution errors from    
TCM4                22 Aug 1997 (MOI- 20 days)    Final adjustment to MOI aim point    
Mars Orbit          11 Sep 1997                   MOI can vary from 11 Sep 97 to 22    
Insertion (MOI)                                   Sep 97 depending on exact launch     
Orbit Insertion     11 Sep 1997 to 5 Apr 1998     begins at MOI, and lasts about 5     
Phase                                             months to reach mapping orbit        
                                                  using aerobrake maneuvers            
Mapping Phase       9 Apr 1998 to 25 Feb 2000     Mars mapping operations for one      
                                                  Martian year, about 687 days.        
                                                  Launch at end of launch period       
                                                  will push start of mapping back to   
                                                  18 April 1998                        
Relay Phase         22 Feb 2000 to 1 Jan 2003     About 3 Earth years                  

2.2.1. Launch Phase

The launch phase extends from the start of the launch countdown to the separation of the spacecraft from the Delta third stage (PAM-D) after the injection maneuver. Separation always occurs 375 seconds after third stage-ignition and 287.86 seconds after burn-out. The time between lift-off and separation varies throughout the launch period and is dependent on the exact values of the departure conditions. On the average, about 45 to 50 minutes elapse between lift-off and spacecraft separation from the PAM. In between those two events, the MGS spacecraft and PAM-D will be launched into either a 28.7&176 or 36.5° low Earth parking orbit, and will coast in that orbit for less than one revolution to the appropriate point to begin the injection maneuver. Section 3 contains more detail.

2.2.2. Cruise Phase

The cruise phase is the period of ballistic flight from Earth to Mars with a duration of about ten months. It is defined to extend from separation of the spacecraft from the PAM-D to the initiation of the orbit insertion burn at Mars. The cruise phase is divided into two sub-phases, defined below.

Inner Cruise Sub-phase

Inner cruise extends from PAM-D separation until sometime in January 1997, defined by the time period when communications with the Earth occurs though the low gain antenna (LGA). The reason is primarily due to the spacecraft configuration and solar panel geometry. Because the high gain antenna (HGA) sits on the spacecraft in a stowed, body-fixed orientation during cruise, communicating with the Earth through the HGA will require turning the spacecraft to point the antenna directly at Earth. However, such an orientation would push the incidence angle of sunlight on the panels past acceptable levels for minimum power generation. Therefore, communications through the LGA represents the only feasible option.

During inner cruise, initial deployment and checkout of the spacecraft and payload will be accomplished, and navigation tracking data will be taken to determine the flight path for the purpose of planning and executing the first of four planned trajectory correction maneuvers (TCMs). TCM1 always occurs 15 days after launch (L+ 15 days).

Outer Cruise Sub-phase

Technically, outer cruise will begin when the spacecraft switches from use of the low gain to the high gain antenna for communications with the Earth. The exact time when the switch becomes feasible depends on when the angle between the Sun and Earth as seen from the spacecraft (SPE) falls to a level low enough to allow good power while the spacecraft is oriented to point the HGA directly at Earth. This angle starts at about 120° at the time of launch and falls to less than 60° by February 1997.

Currently, the transition date to switch to the HGA from has not yet been determined to a greater fidelity than sometime during January 1997, and no later than 1 February 1997. Once the date has been identified, it will still be subject to change during flight as the spacecraft team evaluates the telemetry. In the interest of maintaining the highest possible communications link margin with Earth, switch over will occur as early as possible.

Most of this sub-phase consists of minimal activity as the spacecraft transits to Mars. The vast majority of the significant events will involve acquiring navigation and tracking data to support the remaining TCMs. During the last 30 days of approach to Mars, the focus will be on final targeting of the spacecraft to the proper aim point, and preparations for orbit insertion.

2.2.3. Orbit Insertion Phase

The orbit insertion phase starts with the Mars orbit insertion (MOI) burn and lasts until the spacecraft reaches the final mapping orbit and is declared ready for the collection of science data. Although orbit insertion consists of several sub-phases, aerobraking activities dominate most of the time that the spacecraft spends in this phase of the mission.

Mars Capture Sub-phase

The Mars capture sub-phase extends from MOI burn ignition and ends at the beginning of the first aerobrake walk-in maneuver (AB1). This burn will lower periapsis to an altitude of 138 kilometers and correct for slight inclination errors incurred during the orbit insertion burn. Currently, AB1 is scheduled to occur at the fourth apoapsis passage after MOI. Under normal circumstances, MOI will place the MGS spacecraft into a highly elliptical, 48 hour period capture orbit at a descending node orientation of approximately 5:45 p.m. Therefore, AB1 will nominally occur nine days after MOI.

Aerobrake Sub-phase

Aerobraking consists of three distinct sub-phases: walk-in, main phase, and walk-out. Walk-in represents the first of the three to occur and begins about nine days after MOI with the AB1 maneuver that lowers periapsis from the capture orbit periapsis altitude of 300 kilometers to 138 kilometers. Over the next month, three more walk in maneuvers (AB2, AB3, and AB4) will gradually lower the periapsis to 112 kilometers and reduce the period to about 40.6 hours.

After the completion of walk-in, the spacecraft will spend three months in the main phase of aerobraking. During this period of time, repeated periapsis passages through the upper fringes of the Martian atmosphere will gradually slow the spacecraft and lower the apoapsis. By the end of the aerobrake main phase, atmospheric drag will have lowered the apoapsis from the original altitude of 57,000 kilometers down to about 2,000 kilometers.

The three weeks of aerobraking following the main phase represent an extremely critical period as the spacecraft lowers its apoapsis down to 450 kilometers. During that time, the MGS spacecraft will slowly "walk-out" of the atmosphere by gradually raising its periapsis altitude to 143 kilometers. Daily aerobraking trim maneuvers (ABMs) will be performed as necessary to maintain a three day orbit lifetime. In other words, in the absence of ABMs due to unforeseen events that inhibit the ability of flight controllers to command the spacecraft, MGS will always be at least three days from a fiery crash.

Aerobraking will end with a termination burn (ABX) performed in late January 1998. This burn will raise the orbit periapsis out of the atmosphere, and to an altitude of approximately 400 kilometers. At this time, the spacecraft will be circling in a 400 x 450 kilometer orbit with a period slightly under two hours. In addition, the descending node location will have advanced from its original MOI position at 5:45 p.m. with respect to the fictitious mean Sun to nearly 2:00 p.m.

Transition-to-Mapping Sub-phase

Transition-to-mapping begins at the end of aerobraking and lasts until the final mapping orbit has been established and the spacecraft is declared ready to begin mapping operations. After the aerobrake termination burn (ABX), the spacecraft will circle Mars in a "trans-map" transition orbit for about 25 days. During this waiting period, the oblateness of Mars will alter the orbit and cause the location of periapsis to drift to a position almost immediately above the Martian south pole. At that time, the transition to mapping orbit (TMO) burn will be performed with the intent of "freezing" the periapsis location at the south pole, and establishing the proper altitude for mapping operations. Currently, TMO will occur sometime during mid-February 1998.

After the TMO burn, about one month will elapse before mapping operations begin. In this month, the navigation team will conduct gravity calibrations for seven days to ensure that the MGS spacecraft has reached the proper mapping orbit, and to update the Martian gravity field model using in-flight navigation data returned from the red planet. If necessary, an orbit trim maneuver (OTM) burn will be executed to refine the frozen orbit based on the results of the gravity calibration. Next, a ten day period spacecraft deployment and checkout period will follow the OTM to allow the operations team to configurethe spacecraft and its instruments for mapping operations

2.2.4. Mapping Phase

Mapping phase represents the period of concentrated return of science data from the mapping orbit. This phase will start on 9 April 1998 and last until 25 February 2000, a time period of one Martian year (687 Earth days), given a launch on 4 November 1996. Launching at the end of the launch period will push back the start of mapping until 19 April 1998.

Throughout the course of mapping, the spacecraft will circle Mars every 117.65 minutes with an aerocentric periapsis altitude 354 kilometers over the southern polar region, and an apoapsis altitude of 409 kilometers over the northern region. In addition, the mapping orbit will be "frozen" to keep the apsidal locations from drifting away from the poles, and to keep the altitudes constant at an index value of 378 kilometers (eccentricity value about 0.007).

During normal mapping operations, the scientific instruments will always remain nadir pointed, and will collect data on a continuous basis. Solid state recording devices on the MGS spacecraft will record this data for later transmission to Earth. Current plans call for 24 hours of recorded science data to be returned once per Earth day during a single 10 hour tracking pass at on the three DSN station locations.

2.2.5. Relay Support Phase and Quarantine Period

The relay operations phase begins at the end of mapping and continues for the remainder of the five Earth year on-orbit design life. During this phase, the spacecraft will function as a relay satellite for various Mars landers and orbiters in support of the Mars Surveyor program. End of mission will occur on 1 January 2003.

Before relay operations begin, a quarantine orbit raise maneuver will elevate the spacecraft to a near-circular orbit with an average altitude of 405 kilometers. This quarantine orbit is needed to reduce the probabilities of spacecraft impact with Mars before L+ 20 years and L+ 50 years to the values required for planetary protection.

2.2.6. Velocity Change Budget

Maneuvers will play a crucial role in transitioning the spacecraft between different phases of the mission. Due to the limited amount of propellant carried on the spacecraft, all of the maneuvers executed throughout the course of the MGS mission must fit within the 1290 m/s Delta-V budget. The general project policy is that statistically sized maneuvers will be budgeted at the 95% confidence level.

In general, maneuvers for this mission fall into two different categories. The first, called bi-propellant maneuvers, represents the larger of the two, will utilize the spacecraft's main engine, and will burn a combination of hydrazine and nitrogen tetroxide. The second type, called mono-propellant maneuvers, will employ the much smaller attitude control thrusters and will use only hydrazine. During bi-propellant maneuvers, the main engine will provide translational Delta-V and the smaller thrusters will provide attitude control (rotational Delta-V) in mono-propellant mode. For the smaller sized mono-propellant maneuvers, the attitude control thrusters will provide the translational Delta-V.

The next table shows the DV allocation for each maneuver in the MGS mission. "Trans." stands for translational, and "Rotat." stands for rotational. All velocities are expressed in meters per second.

Mission Phase       Maneuver Name /     Trans.    Rotat.      Comments                 
                    (Type)              Delta-V   Delta-V
Cruise              TCM 1 and 2 (bi)    54.7 3.0    0.3       Values based on 95%      
                    TCM 3 and 4 (mono)                        statistical confidence   
Orbit Insertion     MOI (bi)            977.1       5.9       Pitch-over burn,         
                                                              Delta-V includes      
                                                              finite burn loses, Rp    
                                                              = 3700 km                
Aerobrake Walk-in   AB1 (bi) AB2        10.0        0.1       1.5 m/s reserved for     
                    (mono) AB3 (mono)   1.5 0.5               Delta-i Periapsis =   
                    AB4 (mono)          0.5                   118 km Periapsis = 113   
                                                              km Periapsis = 112 km    
Aerobrake Main      ABM (mono)          4.0         5.0       Aerobrake corridor       
Phase                                                         control burns            
Aerobrake Walk-out  ABM (mono)          21.0        25.0      Used to slowly raise     
Transition to       ABX (bi) TMO (bi)   57.0        0.3 0.1   Raise periapsis to       
Mapping             Contingency (bi /   18.0                  about 400 km Establish   
                    mono)               14.7 /                mapping orbit            
                                        8.5                   Aerobrake emergencies    
Mapping             OTM drag (bi /      7.0 /        46.5     OTMs used primarily      
                    mono)  ACS          3.9                   for drag make-up, ACS    
                    Rotation (mono)                           for momentum wheel       
Quarantine Raise    PQ (mono)           12.0                  Performed at end of      
Relay               OTM drag (bi /      4.8          8.5      OTMs used primarily      
                    mono) ACS                                 for drag make-up, ACS    
                    Rotation (mono)                           for momentum wheel       
                    Sub-Total (bi /     1138 /       91.8     Grand total Delta-V   
                    mono) Sub-Total     60                    for entire MGS mission   
                    (mono)                                    is 1290 m/s              

2.3. Science Investigations and Descriptions

2.3.1. Scientific Objectives and Payload Overview

The general scientific objectives for the Mars Global Surveyor mission are to complete as fully as possible the original science objectives of the Mars Observer mission. The Mars Observer science objectives were derived from the recommendations of the Solar System Exploration Committee in their report Planetary Exploration Through Year 2000: A Core Program. These objectives were validated by the work of the Mars Observer Science Working Group. Specific scientific objectives for Mars Global Surveyor are listed as follows:

    a) Characterize surface morphology at high spatial resolution to quantify surface characteristics and geological processes.

    b) Determine the composition, map the distribution, and measure the surface thermo-physical properties of surface minerals, rocks, and ices.

    c) Determine globally the topography, geodetic figure, and gravitational field.

    d) Establish the nature of the magnetic field and map the crustal remnant field.

    e) Monitor global weather and thermal structure of the atmosphere.

    f) Study surface-atmosphere interaction by monitoring surface features, polar caps, polar thermal balance, atmospheric dust, and clouds over a seasonal cycle.

    g) Provide multiple years of on-orbit relay communications capability for Mars landers and atmospheric vehicles from any nation interested in participating in the International Mars Surveyor Program.

    h) Support planning for future Mars missions through data acquisitions with special emphasis on those measurements which could impact landing site selection.

The names of the scientific instruments for Mars Global Surveyor, and the measurement objectives for each instrument appear in the next table. More detailed descriptions of each instrument follow in subsequent paragraphs in this section of the mission plan.

Acronym       Full Name           Lead Center                Measurement Objective      

MAG / ER      Magnetometer /      Goddard Space Flight       Intrinsic magnetic field   
              Electron            Center (GSFC)              and solar wind             
              Reflectometer                                  interaction with Mars      

MOC           Mars Orbiter        Malin Space Science        Surface and atmospheric    
              Camera              Systems (MSSS)             imaging                    

MOLA          Mars Orbiter        Goddard Space Flight       Surface topography         
              Laser Altimeter     Center  (GSFC)                                        

MR            Mars Relay          France (CNES)              Support for future Mars    
                                                             missions, American and     

TES           Thermal Emissions   Arizona State University   Gravity field and          
              Spectrometer        (ASU)                      atmospheric refractivity   

USO           Ultra Stable        Stanford University        Mineralogy, condensates,   
              Oscillator for      (team leader)              dust, thermal              
              Radio Science                                  properties, and            
                                                             atmospheric measurements   

2.3.2. Magnetometer/Electron Reflectometer

The measurement objectives of this experiment are to determine the existence and characteristics of the global magnetic field, to characterize surface magnetic features, and to determine the nature of the solar wind interaction with Mars. The Magnetometer (MAG) will measure the three orthogonal components of the magnetic field, providing the direction and magnitude of the ambient field at the sensor. In contrast, the Electron Reflectometer (ER) will measures the energy spectrum and angular distribution of electrons arriving at the sensor. The two sensors share a common data processing unit, and the data will be combined and reduced before downlink. This experiment generates data at rates of 324, 648, or 1296 bps.

Depending on the ground track spacing and the amount of collected data, the Magnetometer can identify surface magnetic features of sufficient strength with a spatial resolution roughly equal to the orbit altitude. With careful calibration, the MGS spacecraft's contribution to the magnetic field can be subtracted from this data.

Figure 2-2: Footprints of MGS Instruments

The Electron Reflectometer will provide a gain of 10 to 100 in sensitivity and a gain in spatial resolution of about three times as compared to direct magnetic mapping for near-surface magnetic features. In order to determine the location of the magnetic field being sampled and the field strength of the absorbing layer in the atmosphere, the ER will require the independent measurement of the magnetic field.

2.3.3. Mars Orbiter Camera

The Mars Orbiter Camera (MOC) will acquire visual images of the surface and atmosphere of Mars for qualitative and quantitative photographic interpretation. The MOC actually consists of three optical assemblies that consist of one narrow-angle (high resolution) and two wide-angle (low resolution) cameras. Each optical assembly employs a CCD line array that will generate images in a "push-broom" manner by using the downtrack motion of the spacecraft to "sweep" the line array across the surface. All these cameras share common electronics for storing and processing photographic data.

The MOC's two wide-angle optical assemblies use 140° field-of-view "fisheye" lenses to see from horizon to horizon. These two cameras employ red (575-625 nanometers) and blue (400-450 nanometers) filters for color imaging, and their resolution varies from about 300 meters per pixel at the nadir to roughly two kilometers at the limb. Together, they function in two of the MOC's three data collection modes. The first mode, called global monitoring mode, will provide daily, full-planet observations of the atmosphere and surface to document changes over time. In contrast, the regional monitoring mode will observe selected atmospheric phenomena and surface features at resolutions as high as 300 meters per pixel.

In the third mode, called high resolution mode, the narrow-angle optics will sample important areas of the planet in great detail. The narrow-angle optical assembly will use a Ritchey-Chretien telescope to acquire images about 2.8 km across at a resolution of about 1.4 meters per pixel. Simultaneous wide-angle images will provide positional and meteorological context for the narrow-angle frames.

Figure 2-3: Mars Orbiter Camera (MOC)

MOC's electronics include a 12 megabyte buffer for data and sequence storage, and a 32-bit microprocessor to accommodate sophisticated software for acquiring, processing, and compressing the images. Average data rates for the instrument will vary from 700 bps to over 40 kbps, depending on the available downlink rate.

2.3.4. Mars Orbiter Laser Altimeter

The Mars Orbiter Laser Altimeter (MOLA) will generate high-resolution topographic profiles of Mars for studies of geophysical and geological structures and processes. This instrument determines the local altitude by firing a laser pulse at the Martian surface and counting the time that the beam takes to reach the surface and bounce back to the spacecraft. A secondary result of this experiment will be to measure surface reflectance profiles at 1.06 µ m.

MOLA consists primarily of a pulsed laser ranging system firing 10 pulses per second. The laser transmitter will utilize a diode-pumped Nd-YAG (Neodymium-Yttrium Aluminum Garnet) laser with a 1.06 µ m wavelength operating at an output power of 45 mJ per pulse. The receiving optics, a Cassegrain telescope with a 50-cm primary mirror, will focuses the return signal on a silicon avalanche photodiode detector.

Each laser spot will measure about 160 meters on the surface with a spacing of about 300 meters (0.1 second) per spot. The accuracy in measuring relative topography will vary from one to ten meters, with an absolute accuracy of about 30 meters. However, the absolute accuracy will ultimately depend on precise reconstruction of the spacecraft orbital position from the radio science gravity field results. MOLA will operate at a single data rate of 618 bps.

Figure 2-4: Mars Orbiter Laser Altimeter (MOLA)

2.3.5. Mars Relay

The Mars Relay (MR) consists of a CNES-provided radio system designed to return measurements and imaging data from instrumented balloons or landed packages. Data from the balloons or landed packages will be routed from the MR to the MOC data buffer for storage and subsequent return in the normal science data stream.

MR consists of a 1-meter helix antenna mounted on the nadir panel and all of the associated electronics. Unlike the main X-band telecommunications sub-system, this device operates at UHF frequencies near 400 MHz. The antenna pattern (-3 dB) takes the form of a 65° cone emanating from the tip of the antenna, and provides coverage to the horizon with a 4,700 km effective range for a 8 kbps data rate. However, the range drops to about 1,200 km for the 128 kbps data rate.

As the spacecraft orbits the red planet in relay mode, the MR will transmit a simple electronic beacon to the surface indicating to lander stations that the MGS spacecraft is approximately "overhead." This beacon will serve as an indicator for the ground stations to begin transmitting their data. At this time, detailed plans have not yet been formulated for the return of data from the various lander and atmospheric vehicles. Currently, the project is working out scenarios for possibly using the MR to facilitate the return of data from the landers of the Russian Mars 96 mission.

2.3.6. Radio Science

Radio science experiments carried out with Mars Global Surveyor will advance two fields fundamental to the study of Mars. First, radio occultation observations of the atmosphere will provide consistent and accurate long term monitoring of the total gas content and the vertical structure of the neutral atmosphere. This data will complement and extend other types of MGS atmospheric observations in that they offer the potential for superior accuracy and vertical resolution (200 meters, with the possibility of 10 meters), and they can be obtained reliably in dusty atmosphere conditions. If detectable, signal scintillations will be studied to extend the understanding of small-scale dynamic processes. In addition, it may also be possible to characterize the main ionospheric layer.

Second, radio Doppler tracking of the spacecraft will provide improved information on the structure of the Martian gravitational field through measurement of its effect on the spacecraft's motion. The low altitude polar orbit, plus improvements in the spacecraft telecommunication subsystem (X-band uplink and downlink), will yield unprecedented spatial resolution, global coverage, and accuracy of the derived gravitational field model. This data will permit parameter estimation of Mars' internal structure and inferences of the red planet's geological evolution.

Unlike any other Mars Global Surveyor experiment team, the Radio Science Team does not rely on a specific scientific instrument. Instead, the spacecraft telecommunications subsystem, augmented by an ultra-stable oscillator, will operate together with one of several ground tracking stations of the DSN. Essentially, raw data will be acquired at the DSN station and then delivered to investigators via the Project Data Base.

Figure 2-5: Thermal Emission Spectrometer (TES)

2.3.7. Thermal Emission Spectrometer

The Thermal Emission Spectrometer (TES) will perform four main types of observations. First, TES will determine and map the composition of surface minerals, rocks, and ice. Second, the instrument will study the composition, particle size, and distribution of atmospheric dust. Third, it will locate and determine properties of water-ice and condensate clouds. Finally, it will also study the condensate properties, processes, and energy balance of the polar ice caps.

TES will function as a combined infrared spectrometer and radiometer designed to measure energy from the surface and atmosphere of Mars in the spectral range of 0.3 to 100 µ m. This instrument primarily consists of two nadir-pointing telescopes that share a common pointing mirror system. The larger of the two telescopes is a 15.24 cm diameter Cassegrain afocal design that feeds a two-port Michelson interferometer spectrometer with a spectral range from 6.25 to 50 µ m. The smaller of the two serves two bolometric channels (0.3 to 3.9 µ m and 0.3 to 100 µ m), measures only 1.5 cm in diameter, and takes the from of an off-axis paraboloidal telescope boresighted with the larger telescope.

Each telescope utilizes six detectors, each with an 8.3 by 8.3 mrad field of view. The six form a rectangular grid three frames wide in the cross-track direction, and two frames wide down-track. By employing a rotating scan mirror, TES will be able to view Mars from horizon to horizon (forward and backward along the ground track) despite the fact that the spacecraft will always remain nadir pointed. On the average, TES will operate at either 688, 1,664, or 4,992 bps.

Figure 2-6: View of MGS Spacecraft

2.4. Spacecraft Description

Mars Global Surveyor will provide a three-axis stabilized platform for observations of Mars by the science payload. The spacecraft design will be derived primarily from the Mars Observer spacecraft, with necessary modifications made to support aerobraking, and to incorporate the action plans from the Mars Observer failure reports. During the summer of 1994, Martin Marietta (MMTI) won the contract to build the MGS spacecraft in their Denver facility. Since then, they have merged with Lockheed and are now called Lockheed-Martin (LMA).

When fully loaded with hydrazine at the time of launch, the spacecraft will weigh no more than 1,052 kilograms under the current design and Delta-V budget. In order to meet this target mass, the spacecraft structure will consist of lightweight composite material divided into four sub-assemblies known as the equipment module, the propulsion module, the solar array support structure, and the high gain antenna support structure.

For the most part, the equipment module's main function involves housing the avionics packages and science instruments. The dimensions of this rectangular shaped module measure 1.221 x 1.221 x 0.762 meters in the X, Y, and Z directions, respectively. With the exception of the Magnetometer, all of the science instruments will bolted to the nadir equipment deck, mounted above the equipment module on the +Z panel. The Mars Relay (MR) is the tallest instrument and sticks up 1.115 meters above the nadir equipment deck.

The propulsion module serves as the adapter between the launch vehicle and contains the propellant tanks, main engines, and attitude control thrusters. This module bolts beneath the equipment module on the -Z panel. Essentially, the module consists of rectangular shaped box 1.063 meters on each side, with a .310 meter tall cylindrical shaped launch vehicle adapter sticking out from the bottom of the box. Each corner of the box portion of the module contains a small metal protrusion that houses attitude control thrusters. Including the length of these protrusions, the diagonal widths of the propulsion module measure 2.464 and 2.394 meters long.

Two solar arrays, each 3.531 meters long by 1.854 meters wide will provide power for the MGS spacecraft. Each array mounts close to the top of the propulsion module on the +Y and -Y panels, near the interface between the propulsion and equipment modules. Including the adapter that holds the array to the propulsion module, the tips of the arrays stick out 4.270 meters from the sides of the spacecraft. Rectangular shaped, metal drag "flaps" mounted onto the ends of both arrays add another .813 meters to the overall array structure. These "flaps" serve no purpose other than to increase the total surface area of the array structure to increase the spacecraft's ballistic coefficient during aerobraking. The two Magnetometer sensors mount on the end of each solar array, in between the array and the flap.

In addition to the solar arrays, the high gain antenna (HGA) structure also bolts to the outside of the propulsion module. When fully deployed, the 1.5 meter diameter HGA will sit at the end of a 2.0 meter long boom, mounted to the +X panel of the propulsion module.

2.4.1. Spacecraft Configuration

Throughout the mission, the Mars Global Surveyor spacecraft will utilize several different configurations and orientations depending on the mission phase and mode of operations. The main configurations are launch, DSN acquisition, inner cruise sun coning, inner cruise array normal spin (ANS), outer cruise array normal spin (ANS), maneuver, aerobrake nominal, and mapping.

Launch Configuration

Although the solar panels and HGA attach to the propulsion module, they will be initially stowed and folded upward against the rectangular equipment module at the time of launch. When attached to the Delta rocket, the XY plane of the spacecraft will lie perpendicular to the longitudinal axis of the Delta. Since the instruments will be pointed in the spacecraft's +Z direction, they will point upward along the Delta's longitudinal axis, toward the top of the booster's payload fairing during flight.

Shortly after the Delta jettisons its first stage, it will also jettison the payload fairing and expose the MGS spacecraft to the ambient environment. Because one of the mission flight rules dictates that the +Z axis (science instruments) can never be pointed within 30° of the Sun under nominal conditions, the Delta must fly an ascent trajectory that keeps its longitudinal axis at least 30° from the Sun.

DSN Initial Acquisition

After completion of the trans-Mars injection burn by the Delta's third stage, the spacecraft will be de-spun from the upper stage's spin stabilized rate of 60 r.p.m. before separating from the stage. At this time, the spacecraft will not know its inertial attitude because the third stage's spin rate is sufficiently high to saturate the gyroscopes in the spacecraft's inertial measurement units (IMUs). The only course of action will involve using the sun sensors to point the spacecraft's +X axis directly at the Sun.

In this attitude, the undeployed HGA will pointed directly at the Sun, and the solar arrays will be swept forward toward the Sun, 30o above the Y axis in the +X direction. The spacecraft will then hold this attitude for two hours to allow for initial acquisition from the tracking antennas of the Deep Space Network (DSN). Since the HGA will point at the Sun during in this mode, initial acquisition must occur on the low gain antenna (LGA). Throughout this time period, the spacecraft will spin at the rate of one revolution every 100 minutes about its +X axis.

Inner Cruise Sun Coning

This next configuration mode will allow the spacecraft to initialize its star catalog after completing DSN initial acquisition. This initialization must be performed to give the spacecraft a fixed "reference" point from which to determine attitude. Without it, the spacecraft will not be able to assume any arbitrary attitude other than those derived from the vector to the Sun.

During inner cruise sun coning, the solar arrays will be swept forward 30° above the Y axis in the +X direction, exactly the same as in the DSN initial acquisition configuration. However, the undeployed HGA on the +X axis will always point somewhere along the edge of an imaginary cone with a half-angle of 60° and longitudinal axis along the vector to the Sun. In this orientation, the spacecraft will spin so that its +X axis traces a path around this 120° Sun exclusion cone once every 100 minutes. Therefore, the spacecraft +X axis is said to be "coning 60° off the Sun." This coning motion will last for at least 200 minutes to allow the star sensors (CSA) enough time to initialize the catalog.

Inner Cruise Array Normal Spin (ANS)

After initializing the star catalog, the spacecraft will transition to inner cruise array normal spin (ANS). In this orientation, the solar arrays will lie in the same position relative to the spacecraft body (30° swept forward toward +X axis) as during inner cruise sun coning. However, the difference is that the +X axis will point to a position halfway between the Earth and Sun, and the spacecraft will roll at a rate of one revolution every 100 minutes around the +X axis.

Spinning halfway between the Earth and Sun represents a compromise between needing to point the +X axis directly at the Earth for maximum communications link margin, and needing to point the solar arrays at the Sun for adequate power generation. Communications with the Earth will always occur through the low gain antenna during inner cruise ANS because the undeployed HGA on the +X axis must point directly at the Earth for use.

Figure 2-7: MGS Spacecraft Configurations (flaps not shown)

Outer Cruise Array Normal Spin (ANS)

About two months after launch and sometime in January 1997, the angle between the Earth and Sun as seen from the spacecraft will allow for pointing the undeployed HGA directly at Earth and still provide for adequate illumination of the solar arrays. From this time until Mars orbit insertion, excluding trajectory correction maneuvers, the nominal spacecraft configuration will be outer cruise array normal spin in order to take advantage of the decreasing angles between the Earth and Sun. When configured in this mode, the spacecraft solar panels will lie swept forward 30° above the Y axis toward the +X direction, +X axis of the spacecraft will point directly at the Earth, and the spacecraft will roll about the +X axis one revolution every 100 minutes.

In the inner cruise Sun coning, inner cruise array normal spin, and outer cruise array normal spin, the solar arrays will lie in the same orientation, and the spacecraft will turn at the same rate of 0.01 revolutions per minute. The only difference is whether the spacecraft spins or cones, and the location of the primary axis of revolution.

Maneuver Configuration

Executing maneuver burns will require a different configuration from the previous three modes discussed. All of the large maneuvers will use the main engine, located on the bottom (-Z panel) of the propulsion module. During the burns, the solar panels will be swept back 30° below the Y axis, toward the - Z direction. The reason for this choice is that the thrust direction of the main engine lies along the -Z axis. By sweeping the panels in that direction , as opposed to the +X direction, the spacecraft center of mass will be better aligned with the main engine's thrust axis. In all cases, the active side of the solar panels will also point in the -Z direction.

The spacecraft will use the maneuver configuration for all four trajectory correction maneuvers (TCMs) and Mars orbit insertion (MOI). In addition, this configuration will also be used for all major orbit change maneuvers at Mars such as OTMs, TMO, and ABx burns whether or not the burn is performed with the main engine or attitude control system.

Aerobrake Nominal Configuration

The nominal spacecraft configuration during aerobrake drag passes will resemble the maneuver configuration, except the solar panels will point in the opposite direction. In this case, the panels will be swept 30° above the Y axis, toward the +Z direction, and the active side of the solar array will point in the +Z instead of the -Z direction. Because the spacecraft will fly through the drag pass -Z axis (main engine end) forward during the drag pass, this orientation will maximize protection for the arrays by keeping their active side away from the oncoming airflow. In addition, the +X axis will remain nadir pointed throughout the duration of an aerobrake drag pass.

For all propulsive maneuvers and for the aerobraking drag passes, the spacecraft will turn under 3- axis control to the maneuver attitude. Then, upon completion of the maneuver, the spacecraft will return to the attitude and configuration required for the current mission phase.

Mapping Configuration

After insertion into the mapping orbit after the aerobrake phase of the mission, the MGS spacecraft will be configured for mapping operations. In this mode, the spacecraft will be 3-axis controlled, using input from the horizon sensors to maintain the science instruments on the +Z panel pointed in the nadir direction. Also, the +X side of the spacecraft will point forward in the direction of orbital motion, and the spacecraft will complete one revolution about the Y axis once per orbit.

Because the mapping orbit is Sun-synchronous with respect to 2:00 p.m. of the fictitious mean Sun, the Sun will always shine (except during eclipse periods) on the +Y spacecraft side at an angle that varies between 50° and 74° from the +Y axis (or 16° to 40° from +X), depending on the Martian time of year. The solar arrays gimbal drive control will be enabled to automatically track the Sun as the spacecraft progresses around the orbit. In addition, the HGA boom will be deployed and its gimbal drive control will allow the antenna to track the Earth around each orbit. This configuration will allow the HGA to point directly at Earth without rotating the entire spacecraft.

2.4.2. Command and Data Handling

The Command and Data Handling (C&DH) sub-system is built around the following six major sub- components:

    a) Two redundant flight computers called standard controls processors (SCP), each with 128K words RAM and 20K words PROM.

    b) Controls interface unit (CIU) that connects the computers to other spacecraft components.

    c) Engineering data formatter (EDF) that formats engineering telemetry.

    d) Payload data subsystem (PDS) that provides the command and data interface to the science instruments.

    e) Cross-strap unit (XSU) that routes telemetry to the solid-state flight recorders and to the telecommunications sub-system.

    f) Two solid state recorders (SSR) units that record science and engineering telemetry, each with a 1,500 Mbit storage capacity.

Almost all of the spacecraft activities will be orchestrated by the SCP flight software. While one SCP is in active control of the spacecraft, the identical software will run concurrently in the backup unit. The SCP flight software functions include attitude and articulation control, command processing, some telemetry functions, power management and battery charge control, thermal monitoring and heater control, and fault protection initiation and execution.

General Description of Anomaly Modes

Fault protection will automatically transition the spacecraft into a safe, neutral attitude and configuration in the event of an in-flight anomaly. The purpose is to keep the spacecraft from further executing functions that could lead to hardware damage or loss of mission, and to provide ground control an opportunity to recover from the anomaly.

Three basic fault protection modes exist. Emergency mode is the least serious level of emergency response and will assist the spacecraft in times of communications difficulties, especially in the event of a command-loss timeout. Contingency mode represents a more extreme response than emergency mode and is designed to combat anomalies in telecommunications, power, or attitude control. Safe mode is the most serious of the three and places the spacecraft into what can qualitatively be described as a "minimal effort survival mode." Essentially, safe mode will provide a final fallback mode of reduced operations in the event of major sub-system problems, or loss or corruption of the RAM code. As such, safe mode software is located in the SCP PROM. Entry into safe mode will occur if a key spacecraft component fails or if the flight computers encounter a power-on-reset condition.

Data Rates

The C&DH sub-system will handle three different types data streams of science data, engineering data, or a combination of both.

    (a) Science and Engineering 1 (S&E-1), a combined science and engineering data stream that can be recorded for later playback, returned in real-time, or both.

    (b) Science and Engineering 2 (S&E-2), a combined science and engineering data stream for returning high-rate science in real-time only.

    (c) Engineering (ENG), an engineering only data stream that can be recorded for later playback, returned in real-time, or both.

Data rates that correspond to the three data streams are defined in the table below. Here, "sps" stands for symbols per second. A symbol is essentially a Reed-Solomon encoded (250:218 ratio) data bit. Therefore, it takes approximately 1.147 bits worth of storage space to encode one bit raw data.

Data Stream      Record Rate      Playback Rate      Real-time Rate     
S&E-1            4000 sps         21333.333 sps      4000 sps           
                 8000 sps         42666.667 sps      8000 sps           
                 16000 sps        85333.333 sps      16000 sps          
S&E-2            n/a              n/a                40000 sps          
                 n/a              n/a                80000 sps          
ENG              2000 bps         8000 bps           2000 bps           
                 n/a              n/a                250 bps            
                 n/a              n/a                10 bps             

Notice that in S&E1, playback of recorded data always occurs at a rate 5.333 times faster than the record rate. This ratio will allow the spacecraft to return 24 hours of recorded data during a single 10 hour DSN tracking pass. In addition, the 80,000 sps real-time S&E2 is a capability, but not an official project requirement. It will be implemented on a "best-effort" basis only.

The spacecraft will carry two solid state recorders (SSR) for redundancy. Each SSR contains two 0.75 Gbit recorders that support simultaneous record and playback. Therefore, each SSR can store up to 104 hours of data at the 4,000 sps record rate, and over 26 hours at the 16,000 sps record rate.

Bit allocations for each science instrument will vary depending on the data stream mode, and the specific data rate. Detailed information regarding these allocations can be found in Appendix ??.

2.4.3. Attitude Control

The Attitude and Articulation Control Sub-system (AACS) will provide attitude determination and control, and will control pointing of the HGA and solar array. Software to run this sub-system resides in the SCPs. Nominally, spacecraft pointing control authority, except during maneuvers, will be provided by three reaction wheels (RWAs), all mounted in orthogonal directions. A fourth RWA, mounted skewed to the other three, will serve as a redundant unit for backup.

Four different types of sensors will provide attitude information. The Sun sensors located in multiple locations on the spacecraft are the simplest and will provide extremely basic attitude knowledge, namely the location of the Sun. They will come into use for attitude acquisition after separation from the upper stage of the Delta, and to begin attitude re-initialization in the event of an anomaly.

Sensor type number two, the Inertial Measurement Unit (IMU), contains gyroscopes and accelerometers that will measure angular rates and linear accelerations. These measurements will be used to determine yaw attitude in the mapping phase. In addition, the IMU measurements will also aid the spacecraft in pointing inertially, for both fixed pointing and attitude slews during maneuvers.

Like the IMUs, the Mars Horizon Sensor Assembly (MHSA) will also provide key attitude data during the mapping phase. This device will primarily look at the atmospheric horizon to define the nadir direction by sensing roll and pitch errors.

Finally, the Celestial Sensor Assembly (CSA) will complement the IMUs by also providing inertial attitude data. The CSA works by scanning the position of distant stars, and will provide attitude control and knowledge during cruise, orbit insertion, and when precise knowledge is required during mapping. Both the MHSA and CSA mount to the +Z panel of the equipment module, next to the science instruments.

2.4.4. Telecommunications

All spacecraft communications with Earth will utilize X-band with the Mars Observer Transponders (MOT), the GFP Command Detector Units (CDU), the 25-Watt RF power amplifiers (RPA), the HGA and four LGAs (two for receive and two for transmit). The LGAs are used early in cruise and for emergency communications in anomaly modes. The primary LGA is mounted on the HGA antenna, while the backup transmitter 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 1.5- meter HGA provides high-rate communications during the outer cruise, orbit insertion, and mapping phases. In the mapping configuration, the HGA is deployed on a 2.0-meter boom to provide clearance over the solar arrays to point to the Earth. The spacecraft can receive uplink commands at data rates in multiples of two between 7.8125 bps (emergency) and 500 bps. The 125 bps commanding rate will normally be used.

2.4.5. Propulsion

The propulsion system is a dual mode bi-propellant system, using nitrogen tetroxide (NTO) and hydrazine. The dual mode differs from a conventional bi-propellant system in that the hydrazine is used by both the main engine and the attitude control thrusters, rather than having a separate hydrazine tank for each. The main engine is the only one that utilizes the bi-propellant system. The main engine delivers a nominal Isp of 318 seconds (minimum 317 seconds) and 596 N thrust. The main engine will be used for the larger maneuvers, including TCM1, TCM2, MOI and 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 will be used for the smaller TCMs and OTMs in a pulse-off mode, as well as providing attitude control during the main engine burns in a pulse-on mode. Two sets of four thrusters are on redundant strings in the event of a failure (e.g. stuck open/close engine valve or a failed closed latch valve) which requires one string to be isolated. Four thrusters are provided for roll control. In addition to maneuvers, the 4.45 N thrusters are also used for momentum management.

2.4.6. Power

Two solar arrays, each 6.0 m2 provide the power for the spacecraft. Each solar array has two panels, the inner panel comprised of gallium arsenide (GaAs) cells, and the outer panel comprised of Silicon (Si) cells. When the spacecraft is in eclipse or turned away from the Sun, energy is taken from two nickel-hydrogen (NiH2) batteries, each with a capacity of 20 Amp-hours. During launch the two solar arrays are fully deployed , and the available power varies from about 1100 Watts after launch to a minimum of about 660 Watts. The array output varies with the solar range and the angle to the Sun, since the spin axis is pointed to Earth. After insertion into the mapping orbit, autonomous gimbal drive control is enabled and each array will track the Sun around each orbit under AACS software control. The orbit-averaged array output will vary from about 1,660 Watts at perihelion (April 1994) to about 1,120 Watts at aphelion (March 1995). On each orbit the spacecraft will be in eclipse from 36 to 41 minutes, and power will be supplied from the batteries.

To achieve the required lifetime, the battery depth of discharge in the mapping orbit cannot normally exceed 27%. This limits the amount of time the spacecraft can transmit during eclipse. The spacecraft is required to support up to ten minutes of transmission in eclipse for the radio science occultation experiment.

2.5. DSN Utilization

The 34 meter high efficiency (34m HEF) antennas of the Deep Space Network (DSN) will provide almost all of the tracking coverage for the MGS mission. This type of antenna was selected for its capability to both transmit and receive X-band signals. Project use of other antenna types, such as the 34 meter beam waveguide (34m BWG), will only be accepted on a negotiated case by case basis.

During periods of "normal" operation in cruise, mapping, and relay phase, the project's requirement for DSN support are modest at one 10 hour track per day. However, critical operation such as launch, maneuvers (TCMs and MOI), and aerobraking operations will require continuous coverage. In addition, the project will also require continuous coverage for special, week-long science campaigns during mapping to take advantage of special orbit geometry conditions or to observe unique seasonal changes on Mars. The following table lists the MGS project's requirements for DSN tracking coverage.

The following table provides a profile of the tracking support required by the Mars Global Surveyor Project. Stations other than, but equivalent to a 34m HEF, such as a 34m BWG or 70m with comparable up and downlink performance, may only be substituted with negotiation from the project.

Time Period              Antenna     Tracking Coverage        Data Types               

Launch to L+ 30 days     34m HEF     Continuous               2-way coherent Doppler   
(includes TCM1 at L+                                          and ranging, angular     
15 days)                                                      data from launch to L+   
                                                              1 day, acquire           
                                                              tracking data as soon    
                                                              as possible after        

L+ 30 days to MOI- 90    34m HEF     10 hours/pass 1          2-way coherent           
days (except for TCM2,               pass/day                 Doppler, 3-way Doppler   
TCM3)                                                         and ranging              

TCM2 (TCM1+ 120 days)    34m HEF     Continuous for a         2-way coherent Doppler   
TCM3 (TCM2+ 30 days)                 period of 3 days         and ranging              
                                     before TCM to 3 days                              
                                     after TCM                                         

MOI- 90 days to MOI-     34m HEF     10 hours/pass 2          2-way coherent Doppler   
30 days                              pass/day                 and ranging              

MOI- 30 days to start    34m HEF     Continuous               2-way coherent           
of mapping (includes                                          Doppler, 3-way Doppler   
TCM4 at MOI- 20 days)                                         and ranging              

MOI- 24 hours to MOI+    70m         Continuous               2-way coherent Doppler   
24 hours                                                      and ranging              

Routine Mapping          34m HEF     10 hours/pass 1          2-way coherent Doppler   
Operations (about                    pass/day plus 1          and ranging, 3-way       
5-Apr-98 to 21-Feb-00)               additional pass every    Doppler and ranging,     
(see notes after this                3rd day for real time    open loop recording      
table)                               data                     during atmospheric       

Science Campaigns        34m HEF     Continuous for 88        Same as during routine   
A: 2-Mar-98 to 9-Mar-98              orbits (approximately    mapping                  
B: 29-Jun-98 to 6-Jul-98             7.2 days)                                         
C: 26-Oct-98 to 2-Nov-98
D1:5-Jan-99 to 12-Jan-99                                                                  
D2:20-Jan-99 to 27-Jan-99
D3:3-Feb-99 to 10-Feb-99
D4:18-Feb-99 to 25-Feb-99                                                                 
E: 3-May-99 to 10-May-99
F: 27-Sep-99 to 4-Oct-99
G:13-Dec-99 to 20-Dec-99                                                                 
(see notes after this                                                                  

Diametric Occultations   34m HEF     10 hour/pass 2           During the overlap       
Edge-on Orbital                      pass/day (w/ 2 hour      period, simultaneous     
Configuration (28 day                overlap) 28-Oct-98 and   2-way coherent Doppler   
duration centered on                 19-Feb-99 require        and 3-way Doppler.       
dates of 28-Oct-98,                  continuous coverage      Otherwise, same as       
19-Feb-99)                                                    during routine mapping   

Communications Relay     34m HEF     10 hours/pass 1          2-way coherent Doppler   
Phase (22-Feb-00 to                  pass/day                 and ranging              

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