4. Cruise Phase

Index to Section 4: Cruise Phase
4. Cruise Phase
    4.1. Interplanetary Trajectory
      4.1.1. Launch and Arrival Dates
      4.1.2. Trajectory Correction Maneuver Schedule
      4.1.3. TCM Implementation
      4.1.4. Propulsion System Operations for TCMs
    4.2. Initial Deployment and Acquisition
      4.2.1. Initial Deployment
      4.2.2. DSN Initial Acquisition
      4.2.3. Attitude Initialization
    4.3. Inner Cruise Activities
      4.3.1. Normal Spacecraft Cruise Operations
      4.3.2. Major Spacecraft Activities and TCMs
      4.3.3. Major Payload Activities
    4.4. Outer Cruise Activities
      4.4.1. Major Spacecraft Activities and TCMs
      4.4.2. Major Payload Activities


Cruise covers the time of ballistic flight between Earth and Mars. This phase will begin immediately after the spacecraft separates from the Delta third stage and will end at the beginning of the Mars orbit insertion burn over 300 days later. Primary activities during the cruise phase will include daily monitoring of the subsystems, navigation activities to determine and make minor corrections to the flight path to Mars, and limited science instrument checkout and calibration activities.

4.1. Interplanetary Trajectory

The main objective of the interplanetary trajectory design, including the selection of launch and arrival dates, involves maximizing the spacecraft dry mass delivered to the mapping orbit at Mars. Consequently, a Type 2 interplanetary trajectory was chosen over a Type 1 to minimize both the Earth departure and Mars arrival energy. This minimization translated into a lower Delta-V requirement and a higher injected dry mass.

4.1.1. Launch and Arrival Dates

MGS will take between 301 and 311 days to reach Mars on its Type 2 trajectory, depending on when the spacecraft leaves Earth during the 22 day launch period in November 1996. Under the current design, the launch period will open on 4 November 1996 and will close on 25 November 1996. These launch dates correspond to Mars arrival dates of 11 September 1997 and 22 September 1997, respectively. The following table lists the launch and arrival dates.

Launch Date   Arrival Date  Time of           Launch Date   Arrival Date  Time of       
                            Flight                                        Flight        
4-Nov-96      11-Sep-97     311 days          15-Nov-96     15-Sep-97     304 days      
5-Nov-96      11-Sep-97     310 days          16-Nov-96     16-Sep-97     304 days      
6-Nov-96      11-Sep-97     309 days          17-Nov-96     16-Sep-97     303 days      
7-Nov-96      12-Sep-97     309 days          18-Nov-96     17-Sep-97     303 days      
8-Nov-96      12-Sep-97     308 days          19-Nov-96     18-Sep-97     303 days      
9-Nov-96      13-Sep-97     308 days          20-Nov-96     18-Sep-97     302 days      
10-Nov-96     13-Sep-97     307 days          21-Nov-96     19-Sep-97     302 days      
11-Nov-96     14-Sep-97     307 days          22-Nov-96     20-Sep-97     302 days      
12-Nov-96     14-Sep-97     306 days          23-Nov-96     20-Sep-97     301 days      
13-Nov-96     14-Sep 97     305 days          24-Nov-96     21-Sep-97     301 days      
14-Nov-96     15-Sep-97     305 days          25-Nov-96     22-Sep-97     301 days      

Figure 4-1: MGS Cruise Timeline

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4.1.2. Trajectory Correction Maneuver Schedule

During cruise, a set of four trajectory correction maneuvers (TCMs) will adjust the interplanetary trajectory to ensure that the spacecraft reaches the proper velocity and position targets prior to the Mars orbit insertion burn. In general, the TCMs will predominantly be statistical maneuvers to correct for injection errors from the Delta third stage, orbit determination errors, unmodelled forces, and slight execution errors from previous TCMs. In addition, the TCMs will remove the Mars aim-point biasing introduced by the Delta for planetary protection purposes. This launch biasing intentionally aims the spacecraft away from Mars by about 50,000 km to guarantee a sufficiently low probability of the Delta third stage impacting the planet.

The next table shows the TCM schedule along with the expected magnitude of the four burns. Current project policy calls for budgeting the TCMs at a 95% confidence level instead of a 99% level due to tight Delta-V budget situation. Specific dates listed represent the schedule for a launch at the open of the launch period on 4 November 1996.

Maneuver / (type)   Comments                      Time           Mean / (95%)           
TCM1 (main engine   Correct most of the           L+ 15 days     TBD (95% magnitude     
- biprop.)          injection errors, remove      (19-Nov-96)    TBD)                   
                    most of the launch biasing                                          
                    due to planetary quarantine                                         
TCM2 (main engine   Correct execution errors      TCM1+ 120      TBD (95% magnitude     
blow-down)          from TCM1, remove remaining   days           TBD)                   
                    launch injection errors and   (19-Mar-97)                           
                    planetary quarantine biasing                                        
TCM3 (ACS -         Correct execution errors      TCM2+ 30       TBD (95% magnitude     
monoprop)           from TCM2                     days           TBD)                   
TCM4 (ACS -         Final adjustment to MOI       MOI- 20 days   TBD (95% magnitude     
monoprop)           aimpoint                      (22-Aug-97)    TBD)                   
Grand Total for     Value for 95% magnitude                      64 m/s assumes worst   
TCMs                represents the                               case for launch on     
                    statistically combined                       24-Nov-96              
                    total for the entire cruise                                         
                    phase, not an algebraic sum                                         

The first and largest of the trajectory correction maneuvers, called TCM1, will always occur 15 days after launch. This maneuver will primarily correct for injection errors introduced by the Delta third stage and remove most of the launch biasing introduced into the trajectory for Mars planetary quarantine purposes. TCM2, scheduled for 45 days after TCM1, will correct for errors in TCM1 and remove the remaining planetary quarantine bias from the trajectory. Both of the first two TCMs will be performed with the main engine. However, the second TCM will be performed in "blow-down" mode with the residual pressure level remaining in the tanks after isolation of the high pressure helium line sometime after TCM1.

The last two correction maneuvers, called TCM3 and TCM4, will each consist of small burn designed to correct for slight execution errors in the previous maneuvers. In addition, TCM4 at 20 days prior to Mars orbit insertion (MOI) will precisely target the spacecraft to its MOI aim-point. Both of these burns will utilize the 4.4 N attitude control thrusters.

Contingency maneuver windows will exist at L+ 30 days for TCM1 and at MOI- 10 days for TCM4 if the primary opportunities are missed due to unforeseen circumstances.

4.1.3. TCM Implementation

Both TCM1 and TCM2 will utilize the main engine (596 N thrust) for primary thrust output with the smaller engines (4.4 N thrust) for attitude control. The small magnitude of the other two maneuvers, TCM3 and TCM4, will allow the attitude control thrusters to perform the burns.

Figure 4-2: MGS Trajectory to Mars

Before both TCM1 and TCM2, the spacecraft will turn from its nominal cruise attitude and point the -Z axis (engine thrust axis) opposite the direction of the desired Delta-V. In addition, the X and Y axes of the spacecraft in the maneuver attitude will point in a direction to optimize the Sun angle to the solar arrays. If this angle is too large, the spacecraft will need to operate on batteries during the maneuver.

Because TCMs are statistical maneuvers designed to correct errors in the trajectory, the spacecraft operations team will not know the exact direction and magnitude of the burn until several days before the scheduled maneuver date. Therefore, the turn may take up to 10 minutes if the magnitude of the turn angle measures as large as 180o. After reaching the maneuver attitude, the spacecraft will then hold a fixed inertial attitude while autonomous AACS checks occur. The spacecraft will then execute the burn autonomously with the capability to abort in the event of a malfunction.

During all TCMs, the spacecraft will control the burn duration by using accelerometer measurements as inputs. At the end of the burn, the spacecraft will turn back to its original cruise attitude. For all maneuvers, at least one of the recorders will continuously record engineering data for subsequent playback. However, in the case of TCM1, the spacecraft will be close enough to the Earth to return data at the 2 ksps engineering rate, given a favorable maneuver direction in terms of sunlight on the solar array and low gain antenna (LGA) orientation with respect to the Earth.

Knowledge of the location of the Sun will factor heavily into the design of the TCMs because of a significant payload constraint on maneuver attitudes, required by the MOC and the MOLA instruments. These instruments do not have covers. Consequently, a flight rule exists that states that spacecraft's +Z axis (science instrument panel) cannot be pointed closer than 30o from the Sun. Due to the fact that the Delta-V direction for a TCM can be in any inertial direction, a performance penalty may exist for pointing the thrust in a non-optimal direction to satisfy the Sun pointing constraint.

4.1.4. Propulsion System Operations for TCMs

At launch the pressure in the entire propulsion system will sit at a blanket value of 100 psig. The mono-propellant lines will be launched wet down to the engine valves, allowing the attitude control thrusters to operate in blow-down mode to support spacecraft separation from the Delta and de-spin.

The main engine valve and latch valves below the tanks will remain closed at launch. Prior to TCM1, a dry-firing of the main engine for several minutes will vent the lines (between the latch and main engine valves) to space and bleed them dry. After venting, the ground team will command these latch valves to open. This action will allow the fuel and oxidizer to fill the lines up to the main engine valve at approximately the same blanket pressure used to operate the mono-propellant system in blow-down mode after launch.

Pressurization of the propulsion system will occur prior to TCM1. In order to pressurize the system, the high pressure latch valve located between the helium pressurant tank and the regulators will open, followed by the opening of one of the normally closed pyro valves located immediately below the pressurant tank. Tank pressurization will begin when the pressure downstream of the regulators reaches 175 psig and ruptures the two burst discs located upstream of the fuel and oxidizer tanks. Then, the regulator will come on line when the pressure in the tanks reaches the nominal value of 250 psig.

Due to concerns about regulator leakage with the tanks nearly full, one of the normally open pyro valves in the pyro ladder immediately downstream from the helium tank will be closed after TCM1, effectively isolating the high pressure helium from the rest of the system. Consequently, the remaining TCMs must occur in "blow-down" mode using the residual pressure level remaining in the tanks.

4.2. Initial Deployment and Acquisition

Initial deployment of the solar arrays and initial acquisition of the spacecraft's signal by the Deep Space Network (DSN) represents one of the most critical mission periods for two reasons. First, spacecraft survival critically depends on proper deployment of the solar arrays because the spacecraft can only operate on battery power for a limited time after launch. Second, the initial DSN acquisition must occur within about 100 minutes after separation because the X-band acquisition-aid antenna can only receive the spacecraft signal to a range of about 40,000 km.

The acquisition-aid antenna is a small wide-beam dish mounted on the 26 meter antenna at the Canberra tracking site. This small antenna will initially direct pointing of the larger, narrow-beam 34 meter high efficiency tracking antenna (34m HEF) normally used to track the spacecraft. Not acquiring the spacecraft's signal will prevent the navigation team from gathering the initial two-way Doppler data necessary to determine the spacecraft's trajectory for the purpose of producing an accurate ephemeris. In turn, lack of an accurate ephemeris will dramatically hamper the ability of the 34m HEF to find the spacecraft.

4.2.1. Initial Deployment

Before the spacecraft separates from the Delta third stage, a yo-yo cable device attached to the stage will deploy to reduce the spin rate from 60 rpm down to zero with a plus or minus two r.p.m. uncertainty. The device works by transferring the angular momentum of the third stage and spacecraft to the cable similar to how a spinning figure skater slows her spin by extending her arms. Although it is unlikely, in the event rotational rates exceeding 9° per second (1.5 rpm) occur resulting in saturation of one or more of the gyros, redundancy management is disabled to prevent the computer logic from incorrectly identifying the saturated gyros as failed.

One second after yo-yo release, pyrotechnic devices will fire to sever the connection between the spacecraft and third stage. A set of four springs will then uncoil to impart a relative separation velocity of between 0.6 and 2.4 m/s between the third stage and spacecraft. The actual mission elapse time of separation will vary as a function of launch day between the values of 45 and 56 minutes, but will always occur 370 seconds after third stage ignition. For the launch date of 4 November 1996, separation will occur about 50 minutes after launch.

On the spacecraft, detection of separation breakwires will activate the post-separation command sequence. The thrusters will be armed and enabled in the first two seconds after separation detection, allowing the launch despin control mode to be activated as soon as possible. The despin control software will "snapshot" the current attitude and fires the appropriate thrusters to despin the spacecraft and hold the attitude. Two minutes are allocated in the launch timeline for the attitude control system to damp the residual spin left over after yo-yo deploy, after which the thrusters will be disarmed for the solar array deployment.

Approximately two minutes after spacecraft separation from the Delta, the two folded solar arrays will deploy one after the other by first releasing the outer panel, and then the inner panel via solar array retention and release devices (SARR) two seconds later. Once released, the solar arrays will unfold to their fully deployed configuration by four pairs of spring-driven hinges. Once each hinge rotates approximately 180o, a latch will engage and lock the hinge and solar panel in place. Five minutes are allocated for the deployment of the solar arrays. After deployment, the attitude control thrusters will fire for up to 20 seconds to remove any rotation introduced by the process of releasing the solar panels.

At this time the spacecraft is commanded to begin acquiring the attitude for DSN initial acquisition of the spacecraft. Because the spacecraft is spinning when it separates from the upper stage, it does not have a three axis attitude reference. Additionally the initial spin vector at separation is moving before the thrusters can be activated to hold the spacecraft attitude. The only attitude knowledge the spacecraft possesses is the location of the sun as determined from the sun sensors (SSA). For DSN initial acquisition, the spacecraft is commanded to the sun acquisition and coning mode, "sun comm power", in which the +X axis is aligned to the measured sun vector. Five minutes are allocated for the spacecraft to slew about until the SSA detects the sun and another 3 minutes for the +X axis slew to the sun. Once the +X axis has been aligned with the sun vector, a 0.06 o/sec rate is commanded about this axis. Under normal conditions, the spacecraft will have acquired the DSN initial acquisition attitude within 18 minutes of spacecraft separation.

While the spacecraft is acquiring the required attitude, the solar panels are commanded to their cruise orientation by rotating the outboard (azimuth) gimbals 120° and the inboard (elevation) gimbals 90o. In this configuration, once the +X axis has been pointed directly at the Sun, the solar arrays will be swept forward towards the Sun, 30° above the Y axis in the +X direction. The spacecraft will hold this attitude to allow for initial acquisition from the tracking antennas of the Deep Space Network (DSN).

Four minutes prior to the expected attitude acquisition, the filament of the TWTA configured to the primary +X transmit low gain antenna (LGA) is turned on and warmed up for initial acquisition of the signal by the DSN.

Figure 4-3: Post-Separation Timeline

4.2.2. DSN Initial Acquisition Initial acquisition will begin about 18 minutes after the spacecraft separates from the Delta's third stage. At this time, the post-separation sequence script will command the spacecraft to begin transmitting realtime engineering data over the LGA at a rate of 2,000 bps. This transmission rate will allow the ground control team to instantly determine whether the spacecraft entered safe mode prior to initial acquisition. If safe mode entry occurs, then the first transmission seen by the ground team will appear at the slower, safe mode utilized rate of 10 bps.

Use of the short coast launch trajectory will place the MGS spacecraft over the Canberra tracking site for initial acquisition. The DSN estimates that using the listen only, wide-beam, X-band acquisition aid (ACQ-AID) antenna, they will "lock-up" on the carrier portion of the signal within a few minutes after the spacecraft begins transmitting. After detection of the carrier, the DSN will then attempt to establish a coherent, two-way link with the spacecraft. In other to accomplish this task, they will track the spacecraft downlink with the ACQ-AID and use its pointing data to point the narrow-beam, 34m HEF antenna. Once the 34m HEF locks onto the spacecraft's signal, the ACQ-AID will no longer be needed. The DSN estimates that under normal circumstances, establishing a coherent, two-way lock with the spacecraft will require 30 minutes, but will probably occur sooner if historical performance is a valid predictor.

Under all circumstances, signal lock-up with the 34m HEF must occur within 100 minutes of the time that the spacecraft begins to transmit telemetry to the ground. After 100 minutes elapse, the spacecraft's range to the Canberra tracking site will exceed 40,000 km, a distance greater than the ACQ-AID's specified "listen range" given the MGS spacecraft's transmission link margin. If 100 minutes elapse and lock-up has not yet occurred, the 34m HEF can perform initial acquisition. However, such a task will be extremely difficult because of the HEF's narrow beam-width. Essentially, using the HEF to search without accurate knowledge of the spacecraft's position is analogous to finding a fly in the sky by looking through a straw.

During the initial acquisition period, the spacecraft will maintain its orientation of +X axis pointed directly at the Sun, solar panels swept forward 30° above the Y axis in the direction of +X, roll rate of one revolution every 100 minutes about the +X axis. The spacecraft will continue to hold this attitude for a total of two hours starting from the time that LGA begins transmitting. This time period exists primarily to allow the navigation team to collect coherent, two-way Doppler data for orbit prediction purposes. In addition, the ground operations team will examine the realtime engineering telemetry to assess the "health" and status of the spacecraft.

4.2.3. Attitude Initialization

Two hours after the start of the DSN initial acquisition period (138 minutes after separation), the spacecraft will begin its attitude initialization sequence. In order to establish a 3-axis attitude reference, the spacecraft must scan the celestial star sensor around the sky to identify known stars. This process will is accomplished by pointing the spacecraft +X axis 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." The coning motion will last for several 100 minutes revolutions to allow the star sensors (CSA) enough time to acquire a three axis attitude reference.

While the spacecraft cones around the Sun to perform attitude initialization, communications with the Earth will periodically fade in and out at 100 minute intervals. The reason is that during this time, the angle between the Sun and Earth, as seen from the spacecraft, will measure between 65° to 105° (depending on the launch day), and the extremes of the spacecraft's +X axis during the coning will place the axis plus or minus 60 degrees from the Sun. Because the LGA sits on the rim of the high gain antenna (HGA) and will point in the +X direction while the HGA sits in its stowed position, the LGA will cycle through positions of pointing almost directly at the Earth to pointing 125° or more away from the Earth. The impact of this periodic loss of telemetry on the navigation team's orbit prediction capability has not been evaluated at this time. However, under normal conditions, the navigation team will already have received up to two hours of coherent, two-way Doppler data.

The command sequence stored on the spacecraft will automatically transition the spacecraft to inner cruise mode after completion of the attitude initialization activities. Current estimates show that initialization will take at least two 100 minute coning revolutions to complete. However, a decision may be made at a future date to keep the spacecraft coning for a total of four revolutions to fully characterize the spacecraft's behavior in this mode.

4.3. Inner Cruise Activities

After completion of attitude initialization activities, the spacecraft will automatically transition the spacecraft's attitude to inner cruise orientation. In this mode, 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 because the undeployed HGA on the +X axis must point directly at the Earth for use.

Inner cruise will be devoted to characterizing the operation of the spacecraft, performing the first TCM, checking out the spacecraft, and calibrating the science instruments. Adequate link margins and continuous DSN coverage during this time period will support the return of data rates as high as the 40,000 sps S&E-2 realtime rate for payload operations and check-out.

4.3.1. Normal Spacecraft Cruise Operations

Normal operations of the spacecraft during cruise will consist of fairly uncomplicated tasks. Except for TCMs and limited science activities, minimal commanding will be sufficient to maintain the engineering sub-systems. The primary ground activity will involve analyzing trends in the performance of various sub-systems based on engineering telemetry and radio-metric data returned during the daily DSN tracking pass.

Although the spacecraft can operate for many days without commanding from the ground, a command loss timer will initiate fault protection responses if a command fails to arrive within a set amount of time programmable from the ground. The only routine uplinks required to maintain the spacecraft will consist of routine "no-op" command loss timer reset commands and a bi-weekly star catalogue and planetary ephemeris update. Data from this ephemeris will assist in the pointing of the spacecraft's +X axis in the array-normal-spin attitude mode, while the star catalog is used to update and maintain the spacecraft's inertial reference.

Figure 4-4: Timeline for Inner Cruise

GIF = 204 KBytes

4.3.2. Major Spacecraft Activities and TCMs

Activities during the first half-month of flight will focus only on spacecraft check-out and preparations for executing the first trajectory correction maneuver (TCM) at 15 days after launch. No "special" check-out procedures exist other than planned intensive monitoring of each subsystem with contingency recording of telemetry on one of the recorders. In addition, telemetry recorded during launch will be replayed during this time.

Propulsion system activities will begin at L+ 7 days with the "dry-firing" of the main engine. The "dry-firing" will vent the lines (downstream of the tanks) to space and bleed them dry. Pressurization of the tanks with the high pressure helium will follow the next day in preparation for the first TCM at L+ 15 days. This schedule implementation will provide seven days of margin before the maneuver in the event of an anomaly during pressurization.

Upon successful completion and verification of TCM1, one of the pyro valves immediately downstream of the helium tank will be closed to effectively isolate both the oxidizer and fuel tanks from the high pressure assembly upstream. Re-pressurization of the tanks will not occur until just prior to Mars orbit insertion (MOI).

4.3.3. Major Payload Activities

Approximately two weeks of payload check-out will follow after the successful execution and validation of TCM1. The first event, check-out of the payload data sub-system (PDS), will occur on days 17 and 18. This activity will involve uploading the PDS memory and operating the system at the different data output rates and modes. The PDS remains at the 8 ksps real-time and record mode (S&E-1 MRC) for the remainder of the payload checkout period.

Following the PDS checkout, the required checkout of the MAG/ER will be performed, scheduled from L+ 20 to L+ 29 days. The MAG/ER will be activated, the ER cover will be opened and the ER high voltage turned on. The MAG/ER will be secured at L+ 29 days, as continuous DSN coverage is no longer available and the data rate drops to 2,000 bps for contingency TCM1.

The MOC, MR, MOLA, and TES instruments all request short checkout periods to verify instrument operation after launch. The MOC instrument electronics will be powered on at L+ 21 days for a five day period in order to determine single event upset/single event latchup rates (SEU/SEL) and to perform a focus check of the MOC Narrow Angle (NA) optics prior to the instrument bakeout. The purpose of the focus check is to determine how much out of focus the NA camera is owing to water absorption by the composite tube. The focus check requires taking multiple pictures of a selected star at various settings of the focus heaters located around the rim and hub of the primary mirror. The heaters are used to induce thermal gradients sufficient to vary the focus. The focus check is spread out over the five days, so that each day four pictures of the star are taken at one of the five heater settings. In order for the MOC to take images, the spacecraft must be rotating about the +Y axis as opposed to the normal cruise mode in which the spacecraft rotates about the +X axis. Continuous LGA coverage may not be possible during the image maneuver, so the MOC images will be stored on a recorder and played back at the end of the focus check period at the 8,000 kbps playback rate. Toward the end of the MOC checkout period, after completion of the focus check, the MR will be powered on and MOC will be configured to receive MR data. The MR will cycle through different modes in order to verify launch survival and correct telemetry frequencies. After this checkout, the MR and MOC will be turned off.

The TES checkout lasts for 24 hours and is scheduled at L+ 26 days. TES will be turned on, its cover opened, and orbit period commands will be uploaded to the instrument for testing. A 3-hour MOLA checkout will occur at L+ 27 days. Continuous DSN coverage permits all science data to be returned in the realtime 8,000 sps data rate.

From L+30 to the end of inner cruise, there are few planned science or spacecraft activities. The MOC bakeout heater will be turned on at L+ 31 days and left on for 60 days to drive out accumulated moisture in the composite structure. Also in this period, Radio Science requests twice monthly tests to monitor USO performance. For these tests, the spacecraft transponder is placed in the USO mode for two hours. The Doppler sample rate at the DSN needs to be one sample per second with open-loop recording preferred. DSN elevation angles above 30 degrees are preferred. The test at L+ 56 days will be done in the occultation configuration with telemetry modulation off. Radio Science also requests a tracking system calibration at L+ 89 days. Radiometric data will be used to assess the performance of the spacecraft and the DSN and to test data reduction software. The spacecraft transponder should be in the coherent tracking mode with ranging on. Again, the Doppler sample rate at the DSN needs to be one sample per second, and DSN elevation angles above 30° are preferred.

4.4. Outer Cruise Activities

About two months after launch and sometime in January 1997, the angle between the Earth and Sun as seen from the spacecraft (values of 60° or less) will allow for pointing the undeployed high gain antenna (HGA) directly at Earth and still provide for adequate illumination of the solar arrays. This geometry will allow the spacecraft to transition from using the LGA to the HGA.

From this time until Mars orbit insertion, excluding trajectory correction maneuvers, the normal spacecraft configuration will be outer cruise array normal spin (ANS) 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.

Primary activities during outer cruise will involve routine monitoring of the spacecraft, collection of navigation data, and execution of the remaining three trajectory correction maneuvers. Because the spacecraft HGA will point directly at Earth during outer cruise, substantial capability will exist for returning data from the science payload. However, only a limited number of science data collection activities and calibrations are currently planned.

4.4.1. Major Spacecraft Activities and TCMs

Until preparations for Mars orbit insertion (MOI) begin toward the end of outer cruise, no major spacecraft activities will occur other than the execution of TCM2, TCM3, and TCM4. Due to the isolation of the high pressure helium tanks after TCM1, the three remaining maneuvers will occur in "blow-down" mode using the residual pressure level remaining in the fuel and oxidizer tanks. Seven days prior to MOI, the propulsion system is repressurized to 250 psig by firing one of the normally closed pyro valves immediately upstream of the regulator system. Seven days allows margin in the event of an anomaly associated with the repressurization of the propulsion system.

4.4.2. Major Payload Activities

Payload activities in outer cruise phase are limited to calibrations necessary to characterize the operation of the instruments for mapping. Only two of the instruments currently have calibration requirements, the MAG and the MOC. The MAG calibration is required only if the spacecraft magnetic field cannot be completely characterized during pre-flight integration and testing. Currently it appears that the in-flight calibration will not be required.

Upon completion of the 60 day MOC bakeout event, another focus check is performed identical to the pre-bakeout check described in Section 4.3.2. The post-bakeout focus measurements are both absolute and relative to the baseline focus measurements of the pre-bakeout focus check. Approach images of Mars are also planned to be taken by the MOC at MOI- 120, 90, 60 and 30 days. Similar to the star pictures taken for the focus checks, the approach image opportunities establish focus heater control authority over a range of operating temperatures.

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