Figures 5-1 and 5-2 show timelines for the baseline orbit insertion strategy, for a 3 November 1996 launch date and a 24 November 1996 launch date. These dates represent the old open and close dates for the launch period, since the aerobraking trajectory design has not been updated for the new launch period, which has shifted one day later. The last launch date drives the aerobraking design, due to the least amount of time available to reach a 2:00 p.m. orbit.
The spacecraft will utilize a "pitch-over" steering strategy for MOI in order to maximize the efficiency of the burn. Essentially, "pitch-over" works by using the attitude control thrusters to slew the spacecraft at a constant rate during the 20 to 25 minute burn in an attempt to keep the thrust tangent to the trajectory arc. Although this constant rate pitch will not allow the thrust vector to exactly follow an optimal steering profile, the strategy provides a more optimal solution than an inertially fixed burn.
Depending on the exact launch date, the spacecraft will arrive at Mars between 11 September 1997 and 25 September 1997. The incoming hyperbolic trajectory will be targeted for a Mars closest approach (periapsis) at roughly 1:00 a.m. UTC. MOI burn ignition will occur roughly 10 minutes before the targeted periapsis time and cut-off will occur about 10 minutes after periapsis. Exact burn times have not yet been calculated.
Figure 5-1: MGS Orbit Insertion Timeline (Open of Launch Period Scenario)
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Figure 5-2: MGS Orbit Insertion Timeline (Close of Launch Period Scenario)
For all arrival dates, Goldstone will serve as the primary DSN site for pre-MOI coverage, while Canberra will cover post-MOI tracking. The exact time of MOI will allow approximately one hour of overlapping coverage both before and after the burn. This dual station overlap strategy will provide tracking redundancy for last minute commanding in the case of an extreme spacecraft emergency, pre-burn monitoring of spacecraft "health" and status, and post-burn acquisition of telemetry for "health" checks and navigation data. The following table lists the Mars periapsis times corresponding to specific launch and arrival dates. All times are in Ephemeris Time (ET).
Launch Date Arrival Date Mars Launch Date Arrival Date Mars Periapsis Periapsis 4-Nov-96 11-Sep-97 01:27:53 15-Nov-96 15-Sep-97 01:21:07 5-Nov-96 11-Sep-97 01:27:53 16-Nov-96 16-Sep-97 01:19:29 6-Nov-96 11-Sep-97 01:27:53 17-Nov-96 16-Sep-97 01:19:29 7-Nov-96 12-Sep-97 01:26:09 18-Nov-96 17-Sep-97 01:17:52 8-Nov-96 12-Sep-97 01:26:09 19-Nov-96 18-Sep-97 01:16:14 9-Nov-96 13-Sep-97 01:24:28 20-Nov-96 18-Sep-97 01:16:14 10-Nov-96 13-Sep-97 01:24:28 21-Nov-96 19-Sep-97 01:14:40 11-Nov-96 14-Sep-97 01:22:47 22-Nov-96 20-Sep-97 01:13:05 12-Nov-96 14-Sep-97 01:22:47 23-Nov-96 20-Sep-97 01:13:05 13-Nov-96 14-Sep-97 01:22:47 24-Nov-96 21-Sep-97 01:11:32 14-Nov-96 15-Sep-97 01:21:07 25-Nov-96 22-Sep-97 01:10:01
The first periapsis after MOI is used as an aerobraking drag pass Operations Readiness Test (ORT). The command scripts required to configure and orient the spacecraft for aerobraking, as well as various flight software parameter updates are generated and uploaded. The predicted time of periapsis is determined and the "trigger" command generated and uploaded for the on-board drag pass command script which configures and orients the spacecraft for entry into the atmosphere (described in detail in Section 5.2).
For MOI the 590 N main engine is used to provide the required Delta-V while the 4.4 N thrusters will be used to control the spacecraft attitude during the burn. The burn sequence begins with the start of 2 kbps engineering telemetry recording on two of the recorders. The spacecraft will then turn to the pre-loaded MOI burn attitude using reaction wheel control. The TWTA beam is turned off to conserve power during the long burn and additionally because the spacecraft is occulted by Mars as viewed by Earth throughout much of the burn.
Figure 5-3: Trajectory Plot of MOI
The solar arrays are then gimballed to their MOI positions, 30deg. from the +/-Y axis towards the -Z axis. This is accomplished by rotating the outboard gimbals -90deg. from their cruise configuration. Then, the catbed heaters for the 4.4 N ACS thrusters and the main engine heater are turned on at the appropriate times to ensure the optimal operating temperatures for the burn. At the required burn time, the flight software maneuver task begins firing the main engine. The burn duration varies between 20 to 25 minutes. The spacecraft executes a "pitch-over" steering strategy for MOI in order to maximize the efficiency of the burn. The attitude control thrusters are used to slew the spacecraft at a desired rate during the 20 to 25 minute burn in an attempt to keep the thrust tangent to the trajectory arc. The maneuver continues until the required DV is sensed from accelerometer data. After the burn, the spacecraft will slew back to the Earth-pointed attitude and resume the array-normal spin attitude mode. Downlink telemetry will be turned on after the burn with realtime 2 kbps engineering data. A playback of the recorded maneuver data will then follow.
The spacecraft has several levels of fault protection that can impact the success of the MOI maneuver. If position or rate thresholds are exceeded during the burn, the flight software will suspend the maneuver, swap to the alternate gyro configuration and restart the maneuver. Redundancy management software is enabled for all critical hardware components and upon detection of a fault will switch to the redundant hardware. The spacecraft also has special system level fault protection modes, designated as safe mode and contingency mode, which if entered would abort the on-board MOI sequence. These modes will be disabled at appropriate times prior to the maneuver. These times are determined by the predicted recovery times needed to reestablish the spacecraft to the required configuration for MOI.
Post-MOI Capture Orbits
With the exception of the aerobraking ORT on the first periapsis after MOI, the spacecraft remains in its normal cruise configuration throughout this period. For the aerobraking ORT, the spacecraft is commanded to its drag pass configuration and attitude as described in Section 5.2.
It is expected that two Russian Mars 96 landers will reach the Martian surface during this time. There is a Project level requirement to support the return of the lander telemetry to the Earth. However, working requirements have not currently been provided by the Russians. For now, lander support windows will be defined within which the MR and MOC will be powered on and the spacecraft slewed to point the +Z axis at nadir when within range of the lander sites. The lander data is received by the MR and inserted into the MOC data buffer. The MOC packets are then collected by the PDS and recorded onto a recorder for playback later in the orbit, when the MAG and ER data is played back.
Aerobraking is the utilization of atmospheric drag on the spacecraft to reduce the energy of the orbit. The friction caused by the passage of the spacecraft through the atmosphere provides a velocity change at periapsis, which results in the lowering of the apoapsis altitude. The rate at which the apoapsis altitude decreases is determined by how much drag is generated and the resulting velocity change at periapsis. Going deeper into the atmosphere will provide greater drag and reduce the orbit faster, but will consequently generate higher S/C temperatures and dynamic pressures.
One of the primary constraints driving the aerobraking timeline is illustrated in Figure 5-4 which shows the direction of approach to Mars for the interplanetary trajectory (latest launch and arrival dates) and the desired 2:00 p.m. orientation for the near-polar mapping orbit. The approach directions for the Type 2 interplanetary trajectories result in the initial orbit's dayside node being located near 5:30 p.m. if the most efficient, in-plane orbit insertion strategy were used. Rapid transition to the 2:00 p.m. orientation is beyond the spacecraft propulsive capability. This transition can be achieved much more slowly, but for no Delta-V penalty, by using the intermediate elliptical non sun-synchronous orbits during the aerobraking phase. For large elliptical drift orbits, the Sun will appear to move eastward with respect to the orbit line of nodes at a rate of about 0.5 degrees per day due to the orbital motion of Mars. The latest arrival date results in the shortest amount of time available to achieve the 2:00 p.m. orbit. Correspondingly, it represents the worst case aerobraking timeline with respect to the amount of drag and resulting dynamic pressures and temperatures on the spacecraft.
Parameter Open of Launch Period Close of Launch Period Case Case Launch Date 4 November 1996 25 November 1997 Arrival Date 11 September 1996 22 September 1997 Number of Aerobrake 482 385 Orbits Aerobrake Completion Jan 23, 1998 Jan 19, 1998 Date Walk-In Main 22 days in duration 81 22 days in duration 75 Phase days in duration 23 days in duration Walk-out Total days in duration 126 14 days in duration days in duration 111 days in duration Mean Local Solar Time 14.06 13.97 (hrs) Periapsis Latitude 9deg. south 33deg. north Apoapsis Altitude 450 km 450 km Periapsis Altitude 143 km 144 km Propulsive Delta-V 7.58 Walk-In Main Phase m/s 3.83 7.56 3.14 15.10 Walk-out Total m/s 22.60 25.80 m/s 34.01
The walk-in phase is 22 days in duration and represents the initial phase of aerobraking where the periapsis altitude is lowered in four steps (via apoapsis propulsive maneuvers) to the desired main phase altitude. The need for a gradual walk-in is due to the large uncertainty in the atmospheric density model of Mars. Additionally, the start of the aerobraking phase coincides with the start of the dust storm season. The effect of the dust storms on the atmosphere is increased atmospheric heating, which propagates to the top of the atmosphere and causes the atmospheric density to increase. How rapidly the atmospheric density changes at the main phase periapsis altitude as a result of these dust storms is not well known. Thus the time between the walk-in maneuvers allows for multiple orbits in which navigation can determine not only a measure of the atmospheric density but additionally how much it changes on an orbit to orbit basis.
Figure 5-4: Orbit Insertion Geometry
The table in sub-section 2.2.6 lists the DV allocation and resulting orbit for each of the four walk-in maneuvers. AB1 is the first walk-in maneuver and occurs 9 days after MOI, to lower periapsis altitude down to the upper fringe of the atmosphere. Additionally, AB1 is used to correct MOI delivery errors. The magnitude of this maneuver requires the use of the main engine. The actual targets for the three remaining walk-in maneuvers (AB2, AB3, and AB4) will be based on what is observed at the previous level. The step sizes for these maneuvers are much smaller and will be performed with the RCS thrusters. The number of days from the previous maneuver are eight, six, and eight, respectively for these three maneuvers.
The main-phase is where the majority of the orbital energy is removed and is 81 days in duration for the opening of the launch period and 75 days in duration for the latest arrival date. The main-phase is defined by the use of small propulsive maneuvers (ABMs), executed at apoapsis, to maintain periapsis within a well-defined periapsis altitude corridor, which is low enough to produce enough drag to reduce the orbit within the time constraints imposed by the 2:00 p.m. orbit, yet high enough to avoid spacecraft heating limits and maximum allowable dynamic pressure to maintain control authority over aerodynamic torques. Due to the oblateness of the Mars, the altitude of periapsis tends to rise during the main phase, so that the ABMs will be in the "down" direction to lower the periapsis altitude.
Figure 5-5: Phases of Aerobraking
The walk-out phase is 22 and 15 days, respectively, for the open and close of the launch period, and represents the final stages of aerobraking when the apoapsis altitude rapidly approaches the desired 450 km. The periapsis altitude is raised with daily ABMs, to decrease the dynamic pressure in order to maintain a 3 day orbit decay period. The 3 day orbit decay period is critical for recovery from an anomaly (e.g. contingency mode entry) which would prevent the execution of the next periapsis raise maneuver. When the desired apoapsis altitude is reached, the ABX maneuver is performed to raise periapsis to 350.8 km, ending the aerobraking phase.
Aerobraking Trajectory Design Summary and Plots
Figures 5-6 through 5-11 are plots for both the open and close of launch period cases, illustrating the key trajectory characteristics described above. A complete set of trajectory design plots for both the open and close of launch period cases are provided in the Trajectory Characteristics Document (Ref. TBD).
Figure 5-6 shows the fictitious mean Sun local solar time at the descending node. The requirement is to reach a 2:00 p.m. (14:00 hours) mapping orbit. The close trajectory follows essentially the same path as for the open, only it begins later because the arrival date is 11 days later but the node is about the same. The initial slope which continues to day 100 is due to the orbital motion of Mars around the Sun. The orbit plane moves by only a few degrees relative to inertial space during this period. The curves flatten out at 14:00 hours because the orbit becomes nearly Sun synchronous at the end of the aerobraking phase, where the apoapsis is low enough for the gravitational perturbation to precess the orbit node at the same rate as the average motion of Mars around the Sun.
Figure 5-6 (Top): Fictitious Mean Local Solar Time at the Descending Node
Figure 5-7 (Bottom): Periapsis Altitude vs. Time During Aerobraking
Figure 5-7 shows the periapsis altitude versus the number of days since 10 September 1997, the arrival date for the open of the launch period. The vertical axis has been limited to the period during aerobraking, following the AB1 maneuver to begin the walk-in phase. The sharp slope in the plot at the beginning of aerobraking represent the four walk-in maneuvers, as the periapsis altitude is lowered down to desired main phase altitude. The main phase begins following AB4, and the plots shows a series of ABMs are required to keep periapsis at the desired altitude. The desired altitude is determined by the aerodynamic heating rate, which depends on atmospheric density and spacecraft velocity. Although the heating rates for open and close are different, the periapsis altitudes overlap considerably during the main phase. The walkout phase begins where periodic propulsive maneuvers begin to drive the periapsis altitude higher to maintain a 3 day margin relative to the "non-recoverable" orbit. The periapsis altitude drops rapidly at the end because the last periapsis raise maneuver is one day prior to ABX (not shown).
Figure 5-8: Aerodynamic Heating Rate at Periapsis During Aerobraking
Figure 5-8 shows the aerodynamic heating rate at periapsis for the open and close trajectories. The aerodynamic heating is equal to the aerodynamic pressure times the velocity and represents the kinetic energy flux of the atmospheric molecules relative to the spacecraft. Because the planet is oblate and the equator crossing is a few minutes after periapsis during the main phase, the peak aerodynamic heating rate is as much as 1.5% greater than the value shown for periapsis at a time soon after periapsis. The aerodynamic heating rate at periapsis increases in stages during the walk-in phase. The aerodynamic heating during the main phase gradually decreases in order to counteract the growing drag duration and changing geometry relative to the Sun. The aerodynamic heating rate drops off rapidly at the start of the walkout phase, where orbit lifetime, rather than temperature drives the design.
The aerodynamic heating rate is larger for the close of the launch period because both cases follow the same local solar time plot (Figure 5.6) to reach the 2:00 p.m. orbit, but fewer days are available for aerobraking for the close case because it arrives at the planet 11 days later than the open case, but must finish aerobraking at about the same date. (Actually, the close of the launch period trajectory finishes aerobraking four days earlier than the open.) Thus, more energy must be removed each orbit in order for the close trajectory to remove the same amount of energy from the orbit in fewer days.
Figure 5-9 shows the shrinking period. The slope is slightly steeper for the close of the launch period case because greater heating rates were required to reach the 2:00 p.m. goal on essentially the same date as for the open of the launch period.
Figure 5-9: Period vs. Time During Aerobraking
Figure 5-10 shows the argument of periapsis, measured from the ascending node in the IAU Mars equatorial system. The approximate latitude of periapsis is shown on the scale on the right. Periapsis moves slowly when the eccentricity is large, but moves more rapidly as the orbit period shrinks below five hours. On about day 120, periapsis moves by the North pole from the day side of the planet onto the night side. The arrival geometry for the close of launch period trajectory puts periapsis closer to the equator, which means that the date where periapsis is closest to the North pole is about five days later than for the open of the launch period. Since the need for the periapsis raise maneuvers to maintain a three day lifetime are loosely correlated to the point where periapsis is near the North pole, this difference effectively adds five days to the main phase for the close of the launch period, partially offsetting the later arrival which removes 11 days from the main phase. Thus the main phase heating rates associated with the close of the launch period are only slightly higher than for the open of the launch period.
Figure 5-10 (Top): Argument of Periapsis vs. Time During Aerobraking
Figure 5-11 (Bottom): Propulsive Maneuvers During Aerobraking
Finally, Figure 5-11 shows the propulsive maneuvers required to keep periapsis at the appropriate altitude. All of the walk-in and main-phase maneuvers lower periapsis. (During actual operations it is not inconceivable that a few maneuvers may be required to raise periapsis in response to short term variations in the atmospheric density.) When the orbit period is large, small ABMs can make a big change in periapsis altitude. This effect contributes to the larger size of the walk-out maneuvers as compared to those for walk-in. During the walkout phase, daily maneuvers are required to add one day of orbit lifetime per maneuver, resulting in a maneuver magnitude slightly more than 1 m/sec.
The aerobrake drag pass sequence begins with the catbed heaters for the thrusters being warmed up for normally 20 minutes. Engineering 2 kbps telemetry is recorded on one of the recorders. The transmitter is normally turned off to conserve power during the drag pass. The spacecraft then turns to the "tail-first" drag attitude under reaction wheel control. The desired drag entry attitude has the -Z axis aligned with the velocity vector, which keeps the science instruments and solar cells out of the incoming flow. Once the turn has been achieved, the solar arrays are then gimballed to their drag pass positions, 30deg. from the +/-Y axis towards the +Z axis. This is accomplished by rotating the inboard or elevation gimbals 90deg.. The 30deg. cant on the arrays provides for maximum aerodynamic stability, with minimal drag reduction, by moving the center-of-pressure aft of the center-of-gravity. This configuration is aerodynamically stable in that the aerodynamic torque will push the attitude back toward the aerodynamic null where the velocity lines up with the -Z axis. Additionally, this orientation puts the arrays against the hard stops to prevent back driving the gimbal motors.
At the predicted start of the pass (plus timing margin), attitude control is switched from reaction wheel to thruster control, and the wheel speeds zeroed to avoid attitude drift caused by wheel spin down. Due to navigation timing errors, thruster control is required during the drag pass as a result of the aerodynamic torques on the spacecraft being larger than the reaction wheels can counteract. In order to minimize the number of thruster firings and the amount of fuel consumed controlling the spacecraft through the drag pass, the attitude control pointing deadbands are opened up to +/-15deg.. This allows the spacecraft to oscillate around the aerodynamic null without using any propellant.
During the drag pass, steering is done using a CSA/aerobraking attitude control mode. This mode utilizes the Mars ephemeris to provide position knowledge about Mars and controls the spacecraft using the CSA and the IMU. The spacecraft -Z axis is maintained along the spacecraft velocity vector, while the +X axis is pointed in the nadir direction. Star processing is disabled during the drag pass, since the buffeting of the spacecraft by the atmosphere makes star mis-identification likely, increasing the risk of contingency mode entry. The inertial attitude reference is propagated by the gyros. At the predicted end of drag the attitude control deadbands are tightened back up to reduce the spacecraft body rates below the limits for reaction wheel control authority. After allowing for an additional delay period for timing uncertainties when coming out of the drag pass, reaction wheel control is re-enabled and the spacecraft is commanded back to the outer cruise "array-normal-spin" mode, in which the +X axis is aligned along the Earth vector, while rotating about the +X axis at 0.01 rpm. For the very small orbit periods, less than 3 hours, there is a concern that Mars may be in the field of view of the CSA upon re-acquiring the cruise attitude, extending even further the amount of time the inertial reference is not updated by star crossings. Therefore an option exists to do an set up attitude in which a desired clock angle about the X axis is acquired that will ensure the CSA is not obscured by Mars prior to commanding back to "array-normal-spin".
After commanding the turn back to "array-normal-spin", the solar arrays are subsequently repositioned back to their normal cruise configuration. The transmitter is then turned back on and a period of real-time transmission allowed for spacecraft health assessment. Following the real-time period, the engineering telemetry stored on the recorder is played back for analysis.
As described earlier, the periapsis altitude is maintained within a prescribed corridor, balancing drag levels to achieve the final orbit period against thermal and dynamic pressure constraints. The maneuvers to adjust the periapsis altitude are performed at apoapsis. It is anticipated that several maneuver sequences will be pre-loaded on-board. The maneuvers will be initiated by real-time command as dictated by ground analyses. These "pre-canned" maneuvers will have fixed delta-V amounts, but the burn directions will require periodic updating due to the location of periapsis changing throughout the aerobraking phase.
Figure 5-12: Typical Aerobraking Orbit Profile With ABM
On the first day, the HGA will be deployed to its final mapping position as described in Section 5.3.4. The next day, an HGA calibration is performed to determine the exact position of the HGA boom and the HGA gimbal zero-reference point. On day three, the Mars Horizon Sensor Assembly (MHSA) is powered on for 48 hours prior to use in order to characterize its performance before initiating mapping control. The following day is devoted to flight software parameter uploads required to configure the spacecraft for mapping operations. Finally on day five the spacecraft is commanded to begin mapping nadir pointing. The spacecraft is monitored over the next several days to characterize its operations in the mapping configuration. Day 6 begins the payload checkout. All of the instruments and the PDS are powered on and their memory loads uplinked. Real-time data collection will take place for the next five days. During this time the routine data collection/return strategy of continuous recording at 4 ksps with daily playbacks at 21.33 ksps will be used. Upon completion of the spacecraft and payload checkout, the spacecraft will be declared ready for mapping and the mapping phase initiated.
After reaching the mapping orbit and completing the gravity calibration period, the high-gain antenna (HGA) will be fully deployed and verified. The HGA deploy sequence begins with a 20 minute warm up of the thruster catbed heaters. Next communications are switched from the HGA to the backup LGA transmit antenna (spacecraft body mounted) to ensure downlink throughout most of the deployment and in the event the boom does not deploy properly. There is insufficient link margin to transmit the telemetry over the LGA, so telemetry modulation is turned off and one of the recorders is commanded to begin recording 2,000 bps engineering telemetry for later playback. A couple of minutes prior to the deployment, the spacecraft is commanded to the "deploy/despin" attitude control mode and actuator control is switched to the thrusters. In this mode, the four reaction wheels are held in tach hold at or above 200 r.p.m. to protect them from possible shock damage when the retention and release devices are fired. The spacecraft controls to a desired attitude throughout the deployment. The HGA is deployed by simultaneously actuating three retention and release devices. The boom/antenna assembly rotates roughly 150° and latches into place within 10 minutes. Actuator control is switched back to reaction wheels and the normal cruise "array-normal-spin" mode is reacquired. The HGA is then rotated around to align the boresight back to the Earth. This is accomplished by rotating the outboard or azimuth gimbal roughly 180° followed by rotating the inboard or elevation gimbal about 30° . Upon completion of the gimbal rotation, the primary LGA (mounted to the HGA) is selected to determine if the boom deployed properly. Based on the strength of the signal a measure of any boom displacement can be determined. An HGA calibration is planned (details have yet to be determined) in order to determine the exact position of the boom. Upon determination of the actual boom position, the gimbal zero reference point is updated as required and communications reenabled over the HGA.
After verification of the HGA deployment, the IMU is commanded to the "low rate" mode to meet required mapping pointing accuracy. Additionally, the MHSA is turned on 48 hours prior to initiation of mapping nadir pointing control, in order to verify its health and characterize its operation. Various flight software parameters updates for mapping operations are subsequently uploaded to the spacecraft.
At this point the spacecraft is now ready to acquire the mapping nadir pointing attitude. In order to do this, the spacecraft is first commanded to the "inertial slew/hold" attitude control mode to slew the spacecraft such that the +Z axis is pointed at Mars. Using this mode provides autonomous payload sun avoidance protection during Mars acquisition, which is a capability not available in the mapping attitude control modes. Attitude control is then switched to "CSA/Mapping" mode to point and maintain the spacecraft +Z axis along the velocity vector. Once the MHSA has acquired "Mars Lock" in which all four quadrants are viewing Mars, attitude control is autonomously switched to "Primary" mode. In primary mode, roll and pitch control are maintained using the MHSA, while yaw is controlled using the IMU as a gyrocompass. Additionally, using the planetary ephemeris, autonomous HGA Earth tracking and solar array sun tracking are enabled. After a period of on-orbit characterization, the spacecraft is declared ready for mapping.