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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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%) Magnitudes 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) (18-Apr-97) 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.
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.
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.
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.
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.
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.
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.
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
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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).
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.
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.
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.