The spacecraft bus shall provide for the following subsystem functions:
The spacecraft shall provide command and data handling (C&DH) for control and reconfiguration of the spacecraft and to permit science and engineering data collection.
The spacecraft shall utilize the GFP payload data Subsystem (PDS), specified in JPL D-3419, Vol. 1, and JPL-D-6623. The spacecraft shall provide 5.25 W of power to the PDS at 28 +/-3 V, and up to 40 mA at 10 V. The spacecraft shall accommodate the PDS mechanical configuration given in Figure 5-1.
Figure 5-1. Payload Data Subsystem Mechanical Configuration
The spacecraft bus shall accept and process two classes of commands transmitted from the ground: real-time commands (RTCs) and stored-sequence command (SSC) files.
RTCs result in an immediate action by the spacecraft bus as determined by the operation code within the command. RTCs may command spacecraft hardware or software activities, or may command PDS actions or payload activities via the PDS. RTCs may consist of multiple words, but shall not be executed until the spacecraft has validated receipt of the entire command string.
An SSC file is a command file which is sent to the spacecraft during the course of the mission to effect the autonomous execution of a time- ordered series of spacecraft and/or instrument commands. An SSC file shall consist of a series of commands and/or calls to on-board stored scripts, each with an associated time tag. The time tags shall be either an absolute execution time or a time relative to the execution of the previous command or script call. Time tags shall be sized to encompass the duration of the mission.
A script is defined as a time-ordered series of spacecraft commands which is stored on- board in a library for use by other uplinked SSC files or by autonomous spacecraft housekeeping and fault protection functions.
The flight software shall provide the ability to specify a deterministic memory location for each SSC file or script at load time. There shall be an on onboard, automatic script load verification mechanism. It shall be possible to load a script or an SSC file and activate immediately upon satisfactory on-board verification, as well as by RTC or SSC file time-tagged activation commands.
Multiple scripts may execute simultaneously, and there shall be the capability to de- activate individual scripts.
The script execution software shall provide a capability to conditionally activate a script or terminate an already active script based on the following script flags:
1) RAM Safing Mode enabled flag 2) Late execution allowed flag 3) Active script restart allowed flag 4) Proceed on error flag 5) Load validation required flag
The flight software shall be capable of activating a previously uploaded script, specified by deterministic memory location, based on the following events:
1) Eclipse entry and exit detections 2) RAM Safing Mode Entry
The spacecraft bus shall be capable of accepting in- flight-switchable uplink command rates as defined in DM 514438, "Deep Space Command Detector Unit National Aeronautics and Space Administration Design Requirement For". A low rate of 7.8125 b/s shall be used in an emergency or backup mode.
The C&DH shall be capable of storing spacecraft bus and at least 1500 PDS and science instrument commands, plus timing information for each command. The transfer of a single RTC to the PDS shall be completed within 2 seconds. The C&DH shall ensure that the transmissions of RTC and SSC contents to the PDS do not interrupt each other. The C&DH shall provide for a minimum of 5k words of memory in cruise and a minimum of 13k words of memory in mapping for storing command sequences and a script library.
The flight hardware and software shall be designed to preclude the unintentional execution of critical or irreversible commands. The spacecraft shall reject all invalid commands, whether they are RTCs or contained within SSCs.
During mapping operation, the spacecraft shall provide a time reference to the PDS containing both the upcoming and subsequent ascending node equator crossing times approximately 7 minutes before the first of these two events.
The spacecraft bus shall provide for the data modes and streams shown functionally in Figure 5-2. A description of these modes and data streams follows.
. The spacecraft bus shall provide the following data streams: (1) S&E-1. S&E-1 is a combined science and engineering stream for recording which is intended to permit the continuous collection of science observations. The combined data stream shall be sent to the C&DH by the PDS as a complete transfer frame that will conform to applicable CCSDS standards. The spacecraft shall provide packetized engineering data to the PDS for insertion into the S&E-1 data stream at a rate or equivalent rate not to exceed 256 bits/second, and which shall meet the requirements of 3.4.3.
(2) S&E-2. S&E-2 will be similar to S&E-1 except that it will utilize higher data rates and is intended for real-time transmission.
(3) ENG. ENG shall be an all-spacecraft bus engineering data stream; it will be assembled by the C&DH and shall conform to the requirements of applicable CCSDS standards. The C&DH shall provide for variable rates and telemetry content as required for all spacecraft operations. An engineering dwell mode shall be provided, wherein normal telemetry is suspended while the downlinked telemetry dwells on selected points until commanded otherwise.
Science instrument data rates input to the PDS (or functional equivalent) are as defined in JPL D-4130, PDS- Instruments Interface Requirements Document . The PDS data rate outputs to the spacecraft are as defined in JPL D-3419, vol.1, Payload Data Subsystem Functional Requirements Document .
Data rates for the ENG data stream shall accommodate the emergency telemetry requirements specified in Paragraph 3.4.3.7, at 10 bps, and shall provide for 256 b/s and 2k b/s data rates for normal engineering-only telemetry.
The spacecraft bus shall provide the capability both to record and to playback the S&E-1 data stream. The capacity shall be adequate to record S&E-1 data at up to 16 ks/s data rate for a minimum of 24 hours. Simultaneous record and playback operations shall be as defined in JPL D-3419, vol.1, PDS Functional Requirements Document .
Figure 5-2. Command/Data Flow Diagram
The command and data handling subsystem shall produce one composite digital data stream to be routed to the telecommunications subsystem for transmission to Earth. S&E data shall be Reed-Solomon encoded. In addition, all data streams shall be convolutionally encoded and conform to the standards specified in the applicable CCSDS documentation.
The spacecraft bus shall provide the timing reference for the spacecraft. The spacecraft bus shall provide an unambiguous binary count to the PDS every 1 s and timing pulse every 125 ms. The time code supplied to the PDS and that which is embedded in the engineering source packets shall be coherent and derived from the same frequency reference. The time supplied shall be capable of being correlated with a known epoch to within 20 ms. The stability of the clock frequency source shall be such that the total drift over a 21-day period is predictable to an accuracy of 20 ms. The spacecraft shall provide an unambiguous count with each data frame in accordance with JPL CCSDS from launch through the end of the mission.
The spacecraft shall be capable of providing for on-board management of data storage and playback to prevent data losses resulting from Earth occultations and associated DSN lockup time requirements (approx. 5 min. required for lockup).
The flight software shall provide for memory readout of any portion of memory of any specified size.
In the event of a failure in a subsystem protected by fault Protection software, the software shall provide sufficient data to reconstruct the autonomous fault protection response. All autonomous fault protection actions taken by flight software shall be reversible by ground command. Ground command override of any autonomous fault protection routine shall be possible.
The spacecraft shall provide for DSN-compatible X-band communications to and from the Earth for radiometric tracking, telemetry, commanding, and radio science as described in JPL document 810-5, "Deep Space Network/Flight Project Interface Design Handbook." The spacecraft shall also provide for a Ka-band link experiment. The telecommunications design shall be capable of simultaneous radiometric tracking, telemetry, and commanding.
At all times during the mission, at least one receiver/CDU shall be powered; it shall not be possible to turn both receivers OFF at the same time. If two receivers/CDUs are powered, the spacecraft shall provide the capability to determine the strongest received signal and use the data from that receiver/CDU.
The spacecraft shall receive and phase lock to a modulated or unmodulated X-band uplink carrier signal. Channel frequency shall be in the range of 7145 to 7190 MHz. The spacecraft shall provide an automatic gain control (AGC) to adapt the receiver gain to the received X-band signal strength. The spacecraft receiver AGC and SPE shall be available in telemetry.
The spacecraft shall provide receiver phase lock operation to the uplink carrier, and the downlink carrier shall be generated from an internal auxiliary oscillator (AUX OSC) or from an external ultrastable oscillator (USO) during the two-way noncoherent tracking mode. The two- way noncoherent mode shall be selected by discrete spacecraft command.
The spacecraft shall provide the capability for turnaround ranging. In this mode, the uplink ranging data shall be demodulated from the X-band uplink carrier and remodulated on the X-band downlink carrier. The spacecraft shall provide a turnaround ranging mode ON/OFF capability.
Receiver tracking threshold shall be 154.7 dBm. Acquisition and tracking rate with an unmodulated signal level greater than -90 dBm shall be 550 Hz/s minimum.
The spacecraft shall be capable of receiving an X-band uplink carrier phase- modulated with a command subcarrier and ranging data from the DSN. The spacecraft shall be capable of providing a primary uplink path with a minimum G/T of -17.5 dB/K referenced to the antenna output terminals. This minimum value is sufficient for simultaneous high-rate commanding and two-way Doppler tracking via a 34-m Deep Space Station (DSS) at maximum range.
The spacecraft shall provide a reduced- performance uplink with a minimum G/T (mean-3 sigma) of -28.2 dB/K referenced to the low-gain antenna input terminals. This minimum value is sufficient for emergency low-rate commanding via the 34-m DSS at maximum range.
The spacecraft shall utilize the received uplink to coherently generate the X-band downlink carrier signal at a transmit/receive frequency ratio of 880/749 (two-way tracking). Transfer to AUX OSC or USO operation shall be automatic upon loss of receiver phase lock. Channel frequency shall be in the range of 8400 to 8450 MHz.
The spacecraft shall generate a downlink carrier from either an AUX OSC or from a USO, when the receiver is not phase-locked to an X-band uplink signal (one-way tracking). Operation in the USO mode shall be by discrete spacecraft command. The spacecraft shall also provide an exciter ON/OFF function.
. During the period from Delta II third stage separation and attitude initialization until DSN initial acquisition has been verified, the spacecraft shall provide a minimum effective isotropic radiated power (EIRP) of 37.5 dBm. DSN acquisition will be within 30 minutes after spacecraft comes in view of DSN station with spacecraft transmitter on.
During cruise and orbit insertion phases the spacecraft shall support a 4k bits/sec downlink, and shall provide an EIRP at least as great as indicated in figure 5-3.
Figure 5.3. Cruise Minimum EIRP Requirements
From the beginning of the mapping phase through the end of the mission, the spacecraft shall be capable of providing an EIRP that is sufficient to meet navigation requirements defined in section 3.3 and engineering and science telemetry data return requirements defined in section 3.4. The required EIRP at maximum range shall be derived from the data in Figure 5-4 on the basis of minimum S&E data rates described in JPL D-3419, vol.1, Payload Data Subsystem Functional Requirements Document .
Figure 5-4. Minimum EIRP vs. Data Rate
For emergency engineering telemetry, the spacecraft shall provide a minimum EIRP of 46.0 dBm at the emergency rate.
The spacecraft shall provide the capability to phase modulate the RF drive for X-band downlink carrier with composite telemetry supplied by the spacecraft telemetry data system.
The spacecraft shall provide two square-wave subcarriers; one for symbol rates greater than 500 symbols per second and one for symbol rates of 500 symbols per second or less. The subcarriers shall be biphase modulated with convolutionally encoded data. For each subcarrier, there shall be a range of modulation indices which are in-flight selectable independent of the data rate. Channel coding shall conform to the standards specified in the CCSDS standards. The subcarrier frequencies shall be greater than 20 kHz and shall be at least 1.5 times the maximum downlink symbol rate in symbols per second. For symbol rates of 500 symbols per second or less, the subcarrier frequency shall not exceed 40 kHz.
The spacecraft shall accommodate a Ka-band link experiment (KABLE) that will generate a coherent downlink carrier at four times the X-band downlink frequency with minimum impact on the performance and reliability of the X-band communication link. As a goal, the KABLE shall have an output of 50 dBm, including beam- pointing loss. The KABLE spaceborne hardware shall not be considered mission critical.
The spacecraft bus shall have sufficient control authority to automatically maintain attitude orientation and stability of the spacecraft during all phases of the mission, including and following separation from the Delta II, and to support the required spacecraft functions associated with communications, power, thermal control, and propulsive maneuvers.
During the mapping phase, the nadir orientation of the spacecraft shall be maintained as specified below. The AACS shall provide a spacecraft attitude that supports continuous and simultaneous data taking by all science instruments during mapping.
The AACS shall also be capable of pointing the body mounted science instruments at arbitrary targets in the celestial frame. The pointing profile shall be able to incorporate an open loop slew.
The orbital reference coordinate system in Mars orbit is shown in Figure 5-5.
Figure 5-5: Orbital Reference Coordinate System
The fundamental planet reference direction is the nadir direction. The +Z axis (nadir direction) is defined by a line passing through the spacecraft perpendicular to the Mars mapping reference spheroid (polar radius = 3375.7 km, equatorial radius = 3393.4 km). The +X axis is defined to lie in a plane perpendicular to the nadir direction and along the projection of the velocity vector on this plane. The +Y axis also lies in this plane and is orthogonal to both the +Z and +X axes, forming a right- handed coordinate system. As defined, the +X axis will be close to, but not always coincident with, the direction of the spacecraft velocity vector; and the +Y axis will be close to, but not always coincident with, the orbit normal.
"Yaw" shall be taken to refer to a rotation about the Z axis, "roll" to a rotation about the X axis, and "pitch" to a rotation about the Y axis. Pointing control (in terms of roll, pitch, and yaw) shall refer to the +X, +Y, and +Z axes as defined above.
During mapping, the spacecraft shall be able to control the pointing of the baseplate of the body-mounted science instruments to within +/-10 mrad (per axis, 3- sigma) with respect to the orbital reference coordinate system.
When pointing at arbitrary celestial targets, the spacecraft bus should, as a goal, be able to control the pointing of the baseplate of body mounted science instruments to within +/-6 mrad (per axis, 3-sigma) with respect to the celestial frame. The AACS should be capable of holding an arbitrary spacecraft attitude for at least 1 hour, barring orientations that violate sun avoidance constraints.
Pointing knowledge error is defined as the difference between the actual pointing vector and the estimated pointing vector. The spacecraft bus shall provide sufficient engineering telemetry to obtain a non-real-time, reconstructed pointing knowledge error of the body-mounted instruments during mapping of less than 3 mrad (per axis, 3-sigma). Pointing knowledge shall be defined relative to the orbital reference coordinate system.
During the mapping phase of the mission, there are two spacecraft bus Areodetic pointing stability requirements: one for short instrument integration times (0.5 s), and one for longer integration times (12 s). Over a 0.5 second period of time, the attitude excursions shall be less than 0.5 mrad in pitch and roll (orthogonal to the instrument boresight), respectively, and 1.0 mrad in yaw (about the instrument boresight). Over a 12 second period of time, the attitude excursions shall be less than 3.0 mrad in pitch, roll, and yaw respectively.
Attitude angular rates outside these values shall be minimized in number and duration. Additionally, the attitude angular rate information shall be derivable from the engineering telemetry.
During periods when antenna pointing is required, the spacecraft bus shall be capable of automatically controlling the pointing of its antenna such that the transmitting and receiving requirements of sections 3.3 and 3.4 are met.
The spacecraft shall be capable of generating, storing, supplying, controlling, converting, regulating, and distributing all primary electrical power required for spacecraft functions. Primary electrical power consists of all power from any power source(s) to the primary side of any DC-DC converters. This capability is required continuously from spacecraft internal power- on prior to launch through all subsequent mission phases. The power subsystem shall support normal spacecraft operations in the worst-case orbit conditions.
The spacecraft shall provide the necessary power for spacecraft operation, maneuvers, and instrument activities.
The spacecraft shall be able to provide continuous power to the payload, on an orbital average basis, as specified in Section 4. Power distributed to the payload shall meet the payload power regulation requirements in Section 4.
The spacecraft shall be capable of accepting externally provided power during ground and launch operations as necessary. Primary power shall be provided to the spacecraft by GSE during ground testing. This power shall mimic each of the power subsystem sources' output characteristics overtheir full operating ranges. The transfer from GSE- supplied external power to spacecraft internal power shall be accomplished without power interruption and shall be reversible and repeatable.
The spacecraft shall incorporate circuitry to protect it from short- circuit-type faults in the externally provided power source.
Appropriate overload protection shall be provided by the spacecraft for all electrical loads external to the power subsystem.
No single fault from high-side or low-side of the primary power distribution system to the chassis shall cause a mission-critical failure. The primary power distribution system consists of all hardware and cabling required to distribute electrical power to the spacecraft loads, from the power source(s) up to and including the primary side of and DC- DC converters.
The capability to maintain spacecraft components within their flight-allowable operating and/or nonoperating temperature ranges shall be provided for all mission phases.
Thermal control surfaces shall not have a direct line of sight path to the exit aperture of propulsion thrusters.
The spacecraft shall provide for instrument thermal control; the thermal interface requirements between the instruments and the spacecraft are defined in the unique interface control documents for each instrument. The spacecraft shall maintain the operating and nonoperating temperature of each instrument within allowable temperature levels as defined in section 4.6.
Attitude control pointing error budgets for the spacecraft shall include misalignments resulting from thermal distortions.
The spacecraft structure shall perform without compromising the mission or the functional integrity of the spacecraft during all phases of the mission.
The minimum factor of safety for ultimate stress (FSu) on structure shall be 1.25. Unstressed structural members may require higher factor of safety. Analytical flight system safety factor requirements are contained in table 5-1 of JPL D-11510 , Spacecraft Performance Assurance Provisions .
The spacecraft bus shall be designed to accommodate 75.6 kg of payload mass, plus an initial contingency of 2%.
Attitude control pointing error budgets for the spacecraft shall include any structural misalignments.
The spacecraft shall provide the propulsive capability to (1) maintain spacecraft attitude control throughout all phases of the mission, (2) execute TCM's, (3) place the spacecraft into the capture orbit at Mars, (4) support insertion into mapping orbit, (5) maintain mapping orbit for the mission duration, and (5) place the spacecraft in the planetary quarantine orbit at the end of the mission if necessary.
Liquid propulsion subsystems shall include protection against overpressurization during the long periods between propulsive maneuvers. Propellant vapor and liquid shall be isolated from the pressurization system prior to initial system pressurization. Propulsion system shall prevent fuel or oxidizer migration or condensation in pressurization lines for all mission phases.
Pyrotechnics systems, excluding initiators, shall meet the design requirements portion of MIL-STD-1576, "Electroexplosive Subsystem Safety Requirements and Test Methods for Space Subsystems."
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