The spacecraft shall have a design lifetime of at least 5 years on-orbit and be capable of supporting science data collection in the mapping phase, supporting Mars relay operations during the relay operations phase, and achieving a quarantine orbit if necessary.
Appropriate block, functional, or alternative mode redundancy shall be employed to avoid single-point mission-critical failures. Specific exceptions to this requirement shall be identified and evaluated; they will be approved only if the failure mechanism is found to be acceptably improbable.
A mission-critical failure is defined to be a failure that results in the permanent loss of data from more than one scientific instrument during the mapping phase, loss of the relay capability during the relay phase, the failure to achieve and maintain the proper orbit or pointing control to within specified tolerances, the loss of science-critical engineering telemetry required for attitude determination, or the failure to achieve the quarantine orbit (if required) prior to the end of the mission.
The design shall also accommodate mission operation in degraded modes. A degraded mode of operation is defined to be one in which the primary scientific objectives of the mission can still be met, but at the expense of a loss of some scientific data and/or an increase in the complexity of the mission operations.
The spacecraft will be carried into Earth orbit by the Delta II 7925 (Delta II) and injected into the trans-Mars trajectory by the Delta third stage. The total injected spacecraft mass plus the despin hardware mass shall not exceed 1064 kg. Launch shall occur during the Mars opportunity of November 6-25, 1996, with launch possible on any day of the launch period.
During the period of November 6, 1996 through November 15, 1996, the spacecraft shall be capable of supporting two instantaneous daily launch windows which will have a minimum separation of 48 minutes and a maximum separation of 73 minutes.
Trajectory correction maneuvers (TCMs) will be executed between injection and arrival at Mars to correct injection errors and to adjust the arrival conditions.
The spacecraft shall be capable of executing up to four TCMs at any time during the cruise phase. The spacecraft shall be capable of executing the velocity increment for a TCM in any inertial direction. TCMs will not be planned earlier than injection plus 15 (I+15) nor later than 20 days prior to Mars Orbit Insertion (MOI-20.)
The spacecraft shall be delivered from a northern approach trajectory to Mars into capture orbit with a propulsive maneuver.
The spacecraft shall then transition to a 2:00 pm local mean solar time sun-synchronous mapping orbit with propulsive maneuvers and aerobraking.
The spacecraft shall be capable of performing periapsis altitude adjustment maneuvers as frequently as once a day during the aerobraking phase. The spacecraft shall be capable of establishing and operating within specification in any mapping orbit within the range of orbital elements shown in table 3-1.
The spacecraft shall be capable of providing a total mission delta-V of 1290 m/s, inclusive of finite burn losses from thrust vector misalignments, gravity losses, and all other maneuver inefficiencies.
After mapping orbit establishment but prior to commencement of mapping operations, a 7 day period will be allocated for data acquisition for Mars gravity calibrations. Spacecraft design shall be such that the spacecraft is as free from non-gravitational accelerations as possible during this gravity calibration period.
Table 3-1. Mapping Orbit Mean Elements
Minimum Semimajor Axis Maximum Semimajor Axis
Element (377.9 km index alt.) (400 km index altitude)
----------------------------------------------------------------
Semimajor Axis 3775.1 km 3797.2 km
Eccentricity 0.0072 +/- 0.007 0.0072 +/- 0.007
Inclination 92.87 deg 92.93 deg
-------------------------------------------------
Ascending Node -39.1664 + 0.524 T +/- 3 deg
Argument of Periapsis -90 +/- 10 deg
Coordinate System Mars mean equator and IAU vector of epoch
Epoch delta-T Earth days past 1/1/1998 at 0000 hours ET
NOTES:
a. Mean orbital elements consistent for truncated Balmino gravity
field (unnormalized.)
b. Index altitude is measured with respect to Mars equatorial
radius of 3397.2 km.
c. Ascending node located on the night side of Mars.
d. Sun-synchronous orbit requirement only applies to mapping phase.
The spacecraft shall be capable of performing orbit trim maneuvers (OTMs) as frequently as every 7 days. The total OTM delta-V requirements shall be taken from Figure 3-1. The spacecraft shall be capable of executing an OTM velocity increment in any inertial direction.
The relay orbit is not required to be sun-synchronous
Figure 3-1. Total OTM delta-V Requirement (Mapping and Relay Phases)
The spacecraft shall be capable of controlling maneuver magnitude and direction per the limits set forth in table 3-2. For each component of each maneuver, the allowable error shall be less than the RSS of the 3-sigma fixed and proportional error limits shown.
Table 3-2. Maneuver Execution Error Limits (3-sigma) (note a)
Error Source Error Limit
Proportional Magnitude 2.0 %
Fixed Magnitude 0.05 m/s
Proportional Side Velocity (Total)
(note b) 2.5 %
Fixed Side Velocity (Total) 0.01 m/s
NOTES:
a. Maneuver execution errors have two components: (1) magnitude
errors that are measured parallel to the desired impulse direction
and (2) side velocity errors that are measured perpendicular to
that direction. The error model assumes that the error in each
component is made up of a fixed error, which is independent of the
impulse magnitude, and a proportional error, which is proportional
to the impulse magnitude.
b. The proportional side velocity error corresponds to a pointing
error of 25 mrad.
During the mapping phase, if any spacecraft function depends on the ability of the MOS to predict the ephemeris of the spacecraft, the analysis of that function shall be based on the ephemeris prediction errors given in Table 3-3.
Table 3-3. Mapping Orbit Ephemeris Prediction Errors - Position
Position
Component Uncertainty in the Predicted Position of the Spacecraft
after 14 Days (km) (notes a,b)
----------------------------------------------------------------------
Downtrack 25
Crosstrack 9
Radial 8
NOTES: a. Excludes solar conjunction.
b. Ephemeris prediction errors during the period surrounding Mars
perihelion will increase due to atmospheric density errors. During these periods
more frequent emphemeris updates will be required to achieve the values in this
table.
During the mapping phase, if any spacecraft capability depends on the ability of the MOS to provide after-the-fact knowledge of the ephemeris of the spacecraft, the analysis of that capability shall be based on the ephemeris reconstruction errors given in Table 3-4.
Table 3-4. Mapping Orbit Ephemeris Reconstruction Errors
Position Component 3-sigma Uncertainty in the Reconstructed Position
of the Spacecraft (km)
----------------------------------------------------------------------
Downtrack 9
Crosstrack 5
Radial 2
Note: Excludes solar conjunction
The flight propellant load, including pressurant, shall be sufficient to accommodate propulsion system performance uncertainties and provide the specified delta-V with a 99% confidence.
The spacecraft shall provide engineering telemetry to allow reconstruction of maneuver, attitude control and momentum management events for navigational and radio science purposes.
The spacecraft shall be capable of supporting the navigational tracking schedule shown in Table 3-5, excluding periods of maneuvers and solar conjunction.
Table 3-5. Navigation Tracking Requirements
Mission Phase Daily Tracking
Coverage Data Type
-----------------------------------------------------------------------------
Cruise
Initial Acquisition
to L+30 days 24 hours 2-way coherent Doppler & ranging
L+30 to MOI-90 days 10 hours (note a) 2-way coherent Doppler & ranging
TCM-3 to TCM+3 days 24 hours 2-way coherent Doppler & ranging
MOI-90 to MOI-30 days 20 hours 2-way coherent Doppler & ranging
MOI-30 to MOI 24 hours 2-way coherent Doppler & ranging
Orbit Insertion (note b)
MOI to mapping orbit 24 hours 2-way coherent Doppler & ranging
Mapping
Daily One DSN pass 2-way coherent Doppler & ranging
Every 3rd day Two DSN passes
(note c) 2-way coherent Doppler & ranging
A period of 4 weeks Two DSN passes 2-way coherent Doppler,
duration, nominally & ranging.
centered on the date
of the edge-on orbital
configuration (note c)
Relay Phase
Every 2nd day One DSN pass 2-way coherent Doppler & ranging
NOTES:
a. A single daily DSN pass and 2 consecutive DSN passes per week.
b. Within limits of power available.
c. Every third day, the 2nd pass will be reserved for real-time science and engineering data transmission.
Mechanically induced antenna phase center motion shall not be greater than 0.1 mm/sec (1sigma) with respect to the spacecraft center of gravity over a sixty second integration time, except during momentum management periods
If any part of the spacecraft is jettisoned after injection into the trans-Mars trajectory, it shall be shown by analysis that the probability of accidental impact of any part of the spacecraft on Mars shall be less than 10-2 up to 20 years after launch and 0.05 for an additional 30 years
At the end of the mission, the spacecraft shall be capable of being raised to the quarantine orbit.
The spacecraft shall be capable of performing its required functions through the use of on-board stored operating programs, stored sequence commands and ground-transmitted real-time commands.
During pre-launch and after the spacecraft and fairing are mated to the launch vehicle, the spacecraft shall be capable of receiving mission critical commands through the spacecraft umbilical. The spacecraft design shall provide for operation for at least 48 hours following separation without the need for ground commands.
In the event of a failure that disrupts normal ground-to- spacecraft communications, the spacecraft shall provide for autonomous on-board action that makes it continuously receptive to low-rate commands in any spacecraft orientation.
The spacecraft shall not require time-critical transmission of ground commands.
. Any spacecraft quantity or function capable of being updated by the ground during the launch phase shall be verifiable in telemetry. The spacecraft shall be capable of storing launch and ascent spacecraft telemetry and subsequently downlinking the stored telemetry stream.
During these phases, the spacecraft shall be capable of gathering and returning data to characterize the performance and health of the spacecraft. These data may be returned either in real-time or non-real time.
The spacecraft shall be capable of gathering and returning data to support in- cruise payload calibration and science data return. These data can be returned either in a real- time or a non-real-time mode.
During the mapping phase of the mission, the spacecraft shall be capable of continuously gathering, storing and transmitting science and engineering data.
The spacecraft shall be capable of transmitting these data to Earth during a daily DSN pass which will allow a total of 4.5 noncontinuous hours of downlink. In addition, the spacecraft shall be capable of returning real-time Radio Science data (non-coherent carrier using the USO without telemetry modulation) for up to ten minutes per orbit during scheduled DSN passes.
During the relay phase of the mission, the spacecraft shall be capable of continuously gathering, storing and transmitting relay data as described in section 4.
. The spacecraft shall provide for a minimum of two telemetry subcommutation maps which shall be selectable and modifiable by ground command.
The spacecraft shall provide engineering data for incorporation in the telemetry stream that enables correlation of the spacecraft's state, performance, attitude and environment with the science data.
The spacecraft shall provide data to perform an assessment of the current health of the spacecraft through the telemetry downlink during each DSN pass.
. The spacecraft engineering telemetry shall provide measurements to allow determination of the quantity of propellants remaining on the spacecraft.
The spacecraft shall be capable, via ground command, of reading out any or all of its memory through telemetry. The spacecraft shall also provide for validating its memory without having to perform a complete memory readout.
The spacecraft shall provide information in the telemetry stream to allow determination of command receipt and execution.
In the event of abnormal conditions onboard the spacecraft, the emergency telemetry data volume within a single DSN pass shall be sufficient to determine the spacecraft state.
The emergency telemetry data transfer frame length shall be of sufficient size to allow a minimum of 5 full frames to be received by the DSN during a single orbit.
Due to the adverse effect of interplanetary distances and solar conjunction on the commandability of the spacecraft, autonomous spacecraft operations shall be provided for as follows:
The spacecraft shall provide for autonomous management of the following:
(1) orbital position determination with respect to eclipse. While in orbit around Mars, the spacecraft shall have the capability to autonomously detect eclipse entry. Upon eclipse entry, the spacecraft shall initiate execution of a stored command sequence designed for eclipse ingress, and upon eclipse egress shall initiate execution of an independent stored command sequence designed for eclipse egress. It shall be possible to enable/disable this capability by real-time or stored sequence command.
(2) battery state-of-charge.
(3) nadir pointing during mapping and relay phases.
(4) spacecraft momentum.
(5) thermal control of subsystems and payload where small thermal time constant would require real-time corrective action in less than 72 hours.
(6) high-gain antenna pointing by stored program.
(7) data storage.
(8) payload sun-avoidance during spacecraft slews for propulsive maneuvers or slews to sun acquisition or coning orientations.
(1) Spacecraft Health In the event of a failure, the spacecraft shall be capable of autonomously maintaining for at least 72 hours the minimum functions required for safe system operation, including payload protection from abnormal spacecraft states or attitudes.
(2) Telecommunications Spacecraft shall provide for autonomous initiation of continuous low-rate telemetry in the event of an on-board failure which would interfere with normal high-gain antenna communications. The spacecraft shall provide for autonomous initiation of emergency telemetry in the event that no commands are received within a ground-selectable time.
(3) Power-On Reset (POR) The spacecraft shall enter a known and verifiable state upon spacecraft bus POR.
The spacecraft shall be able to function without ground commands and maintain the minimum functions required for safe system operation during the period when the Earth-Spacecraft-Sun angle is less than 2 degrees.
The spacecraft shall be designed to meet the functional requirements as specified in other sections of this document when operating in the expected mission environment described in JPL D-11513 "Spacecraft Environmental Estimates" , with the design margins specified herein and when under test in accordance with the provisions of JPL D- 11510, "Performance Assurance Provisions." Except where specified otherwise, all environmental design requirements shall equal or exceed the corresponding protoflight test requirements in JPL D-11510. Table 3-6 summarizes the required design margins to be applied to the environmental estimates of JPL D-11513.
The spacecraft bus shall be assembled and maintained in a class 100,000 clean room per Federal Standard 209.
The spacecraft shall be designed to operate within specifications after exposure to the ambient thermal and humidity environment expected during ground operations, including assembly, handling, transportation, and storage. If provisions are made to maintain the temperature and humidity of assemblies within that defined for the controlled environment as specified in JPL D-11513, then values for the controlled environment should be used for design, and design verification tests for the controlled environment are not required. If the temperature and humidity of assemblies are uncontrolled, then the values for the uncontrolled environment as specified in JPL D-11513 should be used for design, and design verification tests at the upper and lower levels of the uncontrolled environment are required.
The spacecraft shall be designed to operate within specifications after exposure to the vibration, acceleration, and shock experienced during ground operations and handling. Special shipping and handling equipment shall be utilized so that transportation levels do not exceed flight levels.
. The spacecraft shall be designed and fabricated to be compatible with the external and self-induced electromagnetic environments that will exist and to function without degradation when operated in any combination of modes. Each assembly shall be designed and constructed so as not to cause electromagnetic interference (EMI) to other assemblies or to itself. The design shall provide a susceptibility margin of 9 dB above expected electromagnetic radiated and conducted environments and shall provide an emission margin at least 9 dB (20 dB for pyro devices) below the expected environment specified in JPL D-11513.
The spacecraft shall be compatible with and shall not produce levels in excess of the Delta II electromagnetic interference (EMI) and electromagnetic compatibility (EMC) requirements contained in MDCH3224 as specified in JPL D-11513.
Table 3-6 Summary of Environmental Design Requirements
ENVIRONMENT DESIGN REQUIREMENT
------------------------------------------------------------------
Dynamics
Sine Vibration (note a) MEF (note b) x 1.5, 5-100 Hz
Random Vibrationa MEF + 4dB (note c)
Acoustics MEF + 4dB
Pyro Shock (note a) MEF + 4dB
Thermal and Vacuum
Thermal/Vacuum Greater of -30/85 C or MEF +35 C
(note d)
Thermal Shock (note e) Greater of -20/75 C or MEF +25 C
Launch Pressure Profile MEF x 1.5
EMC/EMI
Conducted Susceptibility MEF + 9 dB
Conducted Emissions MEF
Radiated Susceptibility MEF + 9 dB (note f)
Conducted Emissions MEF
NOTES:
a. Any axis
b. MEF = Maximum expected flight levels from JPL D-11513
c. For random vibration, MEF shall be at at least minimum workmanship
levels
d. Worst case analysis at design temp, Parts Stress Analysis at
Qual/Protoflight test temperature
e. Only for assemblies with dT/dt>50 C/Hour
f. For pyro devices a design margin of 20 dB is required
The spacecraft shall operate safely in the explosive atmosphere that may exist during launch preparations and launch.
The spacecraft shall function within specification after exposure to the acoustic, vibration, shock, and acceleration environments experienced during launch and separation of the flight system from the Delta II separation as specified in JPL D-11513 with a design margin of 1.5 for sine vibration and 4 dB for random vibration, acoustics and shock.
Those elements of the spacecraft that are required during the launch phase shall function within specification during exposure to the launch phase dynamic environment.
The spacecraft shall maintain temperatures within the environmental estimates of JPL D- 11513 when exposed to the environmental extremes encountered during these phases as specified in that document. The spacecraft assembly designs shall be based on the environmental extremes as specified in JPL D-11513 margined by +/- 35 degrees C.
. The spacecraft shall withstand the pressure environments and pressure changes experienced during launch as specified in JPL D-11513 with a design margin of 1.5.
Surfaces exposed to the atmosphere in low Earth orbit shall be designed to perform within specification during and after exposure to neutral atomic oxygen.
. Contamination due to either the Delta II, the spacecraft, or the environment shall not unacceptably degrade the spacecraft performance at any time during the mission. Acceptable levels of payload contamination are given in Section 4.
The spacecraft shall maintain temperatures within the environmental estimates of JPL D-11513 when exposed to the environmental extremes encountered during these phases as specified in that document. The spacecraft assembly designs shall be based the environmental extremes as specified in JPL D- 11513 margined by +/- 35 degrees C.
The spacecraft shall employ protective means or damage-tolerant design against failures due to solid particle (micrometeoroids) impact during the mission. For internal components, protection shall provide a 0.95 probability of no penetration based on a Poisson distribution and the expected meteoroid flux and fluence as specified in JPL D-11513. The expected meteoroid flux and fluence shall be margined by a factor of three for surface material effects.
The spacecraft shall be designed to operate within specification when exposed to the ambient magnetic fields encountered during the mission as specified in JPL D-11513 .
The spacecraft shall be designed to accommodate the natural particle radiation encountered during the mission as specified in JPL D-11513 . The levels for radiation design of assemblies to total ionizing doses shall be expected radiation values margined by a factor of two. For surface effects the design margin shall be 1. The spacecraft shall be designed so that single-event effects (SEEs), including single event upsets (SEUs) and single event latchups (SELs), do not cause mission-critical failures.
The spacecraft design shall minimize the possibility of internal electrostatic discharge, and be immune to any electrostatic discharges occurring on external surfaces. Metallized layers of thermal blankets shall be strapped together and grounded to the spacecraft structure. All electrical cables from the interior of the spacecraft shall be shielded with the shield conductively grounded to the spacecraft structure.
The spacecraft shall be able to accomodate the aerodynamic heating rates and dynamic pressures encountered during the passes through the atmosphere.
Critical functions of the spacecraft and interactions between the spacecraft and the MOS, Delta II and DSN shall be testable at a level of assembly adequate to verify the function or interface to be tested. Normal and fault protection/correction hardware and software states and sequences shall be testable at the system level. Spacecraft and test system shall provide for safe ground testing of all hazardous spacecraft commands.
The spacecraft design shall provide for functional verification of the payload operation at the system level. Hardware and software interactions among GFP subsystems and between GFP subsystems and the spacecraft shall be testable at the system level. Access to and electrical monitoring of payload interfaces shall be possible at the system level. Access for external stimulation of payload sensor elements shall be possible at the system level. Payload interfaces shall be designed so as not to limit installation/removal cycles or test duration at the system level.
The spacecraft, in its launch configuration, shall be fully compatible with the Delta launch vehicle.
The spacecraft and GSE shall permit installation of a shroud or other equipment to maintain a cleanliness level of class 100,000 around the spacecraft during launch site processing. The spacecraft and GSE shall be capable of routing and/or sending spacecraft telemetry and command data through a facility communication interface prior to encapsulation in the Delta payload fairing. After encapsulation in the payload fairing, all telemetry and command links shall be via a payload ground umbilical.
The spacecraft and GSE shall meet the requirements and be subject to the constraints set forth in ERR 127-1, "Range Safety", AFSPACECOM letter dated 23 November 1993, and GP-1098, KSC Ground Operations Safety Plan.
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