The spacecraft design shall accommodate the payload as described in Table 4-1:
Table 4-1. Mars Observer Payload
Instrument (Abbreviation)
The spacecraft shall accommodate the physical, orientation, and field of view requirements of the payload as defined in the payload Interface Control Documents (ICD) referenced in section 2.
The payload ICDs include a physical representation of each payload assembly. Three-view layouts with body-fixed orthogonal axes (X', Y', Z') are defined for each instrument assembly that has a specific orientation requirement. Physical dimensions are supplied for those assemblies that have no specific orientation requirements.
All instruments for which body-fixed axes (X', Y', Z') are defined in the payload ICDs shall be oriented such that the +Z' axis is in the nadir direction, parallel to the orbital reference +Z axis. The +X' axis shall be in the downtrack direction, parallel to the orbital reference +X axis and velocity vector. The +Y' axis shall be normal to the +X' and +Z' axes, forming a right-handed coordinate system. Instruments shall be aligned relative to these axes to within 1 mrad with knowledge of the alignment to better than 0.5 mrad. Note: In the payload ICDs, each instrument orientation is based on an orbit with the ascending orbit node on the dark side of the planet.
Instrument locations on the spacecraft shall accommodate physical characteristics and fields-of-view as described in the payload ICDs.
The allocated mass and volume requirements for the instruments are specified in the payload ICDs. The spacecraft shall provide thermal blanketing and associated thermal hardware and equipment required for maintaining thermal control.
The view direction (VDIR), instrument field of view (IFOV), and stray radiation field of view (SRFOV) requirements are listed in the payload ICDs with respect to the instrument coordinates axes (X', Y', Z'). The spacecraft shall provide fields of view for the science instruments as specified.
To protect the instruments from damage, spacecraft operations shall insure that the +Z-axis is never pointed within 30 degrees of the sun, unless a slew rate greater than or equal to 0.4 degrees per second is maintained while the +Z-axis is within 30 degrees from the sun.
The spacecraft telecommunications subsystem shall utilize the USO described in the USO ICD. The USO shall be located in a stable thermal environment (delta-T less than or equal to 3°/hr.) The USO has two identical outputs (one for each transponder). When in the USO mode, the transponder will use the USO output as a frequency reference for the X-band downlink.
The spacecraft shall provide the capability for X-band transmission in the USO transponder mode from approximately 4 minutes prior to Earth occultation ingress to approximately 1 minute after, and from approximately 1 minute prior to Earth occultation egress to approximately 4 minutes after Earth occultation. This capability shall be provided for each Earth occultation event during each tracking period of the mapping phase, within the limitations of the HGA gimbal travel. During this period the spacecraft shall provide the capability for X-band transmission in either the USO or two-way coherent transponder mode with telemetry modulation switched OFF and the transmitter ON.
The composite digital data stream and subcarrier shall be capable of being turned off during radio science atmospheric observations.
When the downlink carrier is referenced to the USO for occultation measurements, the spacecraft shall not degrade the specified frequency stability and phase noise performance of the USO by more than 20% (1 dB). This includes telecommunications subsystem, motion of antenna relative to center of spacecraft mass and thermal/magnetic environment of USO.
Modulation sidebands or spurious signals in the downlink signal transmitted by the spacecraft during occultation measurements shall be less than -60 dBc within 2000 to 500 Hz of the carrier and less than -70 dBc within 500 Hz of the carrier.
The unmodeled contribution of spacecraft effects on received power of the downlink carrier shall be less than 0.1 dB over any 50 second interval during occultation measurement.
When the spacecraft is seen at the DSN above 25 degrees elevation (or approximately 50 dB/Hz), the spacecraft contribution shall be less than 0.1 mm/sec (3-sigma) unmodeled error in the Doppler tracking measurements with a 10 second integration time. This includes telecommunication subsystem, instability of onboard oscillator if coherent tracking is not available, and motion of antenna relative to spacecraft center of mass.
Spacecraft telemetry data for maneuvers or momentum unloading thruster firings, etc. shall be of sufficient accuracy to reconstruct spacecraft velocity changes to 1.0 mm/sec (3-sigma ).
The spacecraft shall provide for a payload engineering checkout to assess instrument health early in the cruise phase, after TCM-1. All instruments shall be exercised during the checkout period. With the exception of the magnetometer and MOC calibrations, no specific maneuvers or spacecraft attitude are required.
The spacecraft shall provide for instrument health checks, calibrations, and a bake-out as defined below, while maintaining adequate attitude control. Achieving the mapping orbit and maintenance of center- of-gravity control shall take precedence over these operations.
The MAG sensor calibration will require a minimum of 30 days of data collection (as continuously as possible) in the cruise phase and 30 days in the transition orbit. Additionally a calibration maneuver will be performed during cruise consisting of 2 consecutive revolutions around one spacecraft principal axis, followed by two revolutions around an axis nearly orthogonal to the first.
The MOC requires a period for bake-out and two calibrations during the cruise phase. The bake-out is required for a period of 60 days early in the mission. Prior to and subsequent to the bake-out period, two periods of up to seven days each of MOC calibration shall be provided to determine the effectiveness of the bake-out procedure. The two calibration periods each require rotation of the spacecraft and instrument data recorded and returned.
In addition, images of Mars shall be secured within two months of MOI and during the orbit insertion phase by properly orienting the spin plane of the spacecraft while preserving other instrument FOV requirements.
The spacecraft shall be capable of providing for low-rate real-time and playback of recorded instrument calibration data throughout the cruise phase as necessary to support check-out and calibration activites.
The spacecraft shall provide independently switchable and fused primary power lines for each instrument and as required to control replacement heaters, decontamination heaters, and pyrotechnic functions.
The spacecraft shall provide replacement heater power to each instrument during nonoperating conditions. The individual power requirements shall be specified in the payload ICDs.
The spacecraft shall be capable of providing power to the payload for necessary instrument bakeouts and calibrations as specified in Section 4.4. Cruise power requirements for instrument calibrations will not exceed the mapping power shown in Table 4-2, except for MOC bakeout. MOC bakeout will require additional power as detailed in the unique instrument ICD.
The spacecraft shall provide orbital average and peak power to each instrument as shown in Table 4-2. The total payload peak power capability shall be equal to the sum of the peak powers of all of the instruments flown. The peak power requirements in Table 4-2 do not apply to load current ripple, startup transients, or other transient loads.
The spacecraft shall provide sufficient power to operate the MOC and MR as listed in Table 4-2.
The spacecraft shall accommodate the instrument transient characteristics as specified in the unique ICDs.
Table 4-2. Payload Power Requirements
Instrument Orbital Ave Power Peak Power Operating Voltage
(W) (W) (Vdc)
--------------------------------------------------------------------
MAG/ER 4.63 4.63 28 +/- 0.56
MOC and MR
(note a) 22.75 29.75 28 +/- 6 and 28 +/- 0.56
MOLA 34.20 35.94 28 +/- 2
TES 12.29 18.20 26 - 32
USO 3.00 3.00 (note b) 24 - 30 (note c)
NOTES:
a. The MOC average and peak allocations are 22.75 and 29.75 watts for the
entire mapping phase of the mission with the MR off. When the MR is
active,
the MOC average and peak allocations drop to 10.25 and 17.25 watts,
resulting
in no net change for the combination of MOC and MR.
b. The USO peak power is for a maximum of 2 hours at turn-on.
c. The USO input voltage drift rate shall be less than one volt per hour.
Spacecraft thermal control shall maintain the temperature application point of the PDS, USO, MR, TES, and MAG/ER within the ranges shown in Table 4-3. The temperature application point is defined as the location of the flight thermistor, as specified in the instrument unique ICD.
Table 4-3. Instrument Application Point Temperature Limits
Temperature Limits (deg C)
Instrument (note a) Operating Nonoperating
----------------------------------------------------------
USO -20 to +30 -30 to +40
MR Electronics -15 to +40 -30 to +50
MR Antenna Base -100 to +100 -100 to +100
TES Electronics -20 to +40 -20 to +50
MAG/ER Electronics -10 to +40 -20 to +40
MAG Sensor -20 to +45 -20 to +45
ER Sensor -15 to +35 -30 to +35
NOTE:
a. The MOLA and MOC instruments will maintain operating and nonoperating
temperatures as stated in the instrument unique interface control
documents.
The spacecraft shall monitor the temperature of each payload element. The location of these measurements shall be specified in the unique interface control document for each payload element. Such monitoring shall be incorporated into the spacecraft telemetry (ENG) so that the temperature information can be obtained, consistent with paragraph 3.4.3.
The MAG/ER experiment has three sensors: two magnetometer sensors and one electron reflectometer. Magnetic control requirements shall be established to ensure that the magnetic field due to any spacecraft-related cause is less than 3 nT (static) and less than 0.3 nT (peak-to- peak) for magnetic field variations at frequencies less than 10 Hz, as measured by the magnetometer most distant from the spacecraft center. The use of magnetic materials and devices shall be minimized.
Contamination is defined as molecular and particulate material that has the potential to degrade the performance of the science instruments, attitude sensors, solar arrays, and thermal control surfaces. The spacecraft, together with test, handling, storage, launch and flight procedures, shall be designed to limit contamination on the sensitive instrument surfaces through the end of the mapping mission. The spacecraft external surfaces shall be visibly clean at payload encapsulation. Visibly clean is defined as the absence of all particulate and non-particulate matter visible to the normal unaided (except corrected vision) eye when inspected at a distance between 6 and 18 inches under a surface illumination of 100 ft-candles minimum. Particulate matter is defined as matter of miniature size with observable length, width and thickness. Non-particulate matter is a film without definite dimension.
The spacecraft design shall provide for continuous nitrogen purge for the TES, MOC and MOLA from instrument-spacecraft integration until Delta liftoff. The spacecraft contractor shall provide for nitrogen purge for the ER, with replenishment of the nitrogen environment around the ER every three days at a minimum.
The TES requires that the microphonic environment at the nadir panel instrument interface not exceed 0.005 g in the frequency range from 10 Hz to 120 Hz.
The MOC requires that the total magnitude of the acceleration at the instrument feet be below the following to keep the variation of pixel position from nominal to less that 2%:
2 Hz: less than 0.26 g
2-20 Hz: -25 dB/decade
20-100 Hz: less than 0.014 g
100-400 Hz: -30 dB/decade
400-2000 Hz: +80 dB/decade
The relay payload interface requirements are identical to those stated in the applicable MBR and MOC Interface Control Documents.
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