OZONE MAPPING PROFILING SUITE (OMPS )
Sensor Requirements Document (SRD)
for
NATIONAL POLAR-ORBITING OPERATIONAL ENVIRONMENTAL SATELLITE SYSTEM (NPOESS) SPACECRAFT AND SENSORS
Prepared by
Associate Directorate for Acquisition
NPOESS Integrated Program Office
17 March 1997
(Continuation of Document)
SRDX3.2.4.1-1
The mass properties of each sensor shall conform to performance, stability, and control requirements of the Launch Vehicle (LV) and the Space Segment
SRDX3.2.4.1.1-1
The sensor contractor shall provide the mass of the sensor to the spacecraft contractor for documentation in the ICD.
SRDX3.2.4.1.1-2
The mass of the sensor shall be measured to +/- 0.1 kg.
SRDX3.2.4.1.2-1
Sensor mass expulsion rates and substances, if any, shall be provided to the spacecraft contractor for documentation in the ICD.
SRDX3.2.4.1.3.1-1
Sensors shall be designed, where practical, to place the center of gravity location as near to the interface plane as possible unless excessive uncompensated momentum precludes this (sometimes the c.g. should be as close to a gimbal axis as possible to reduce uncompensated momentum).
SRDX3.2.4.1.3.1-2
The location of the sensor center of mass shall be provided using coordinates based on the spacecraft axes.
SRDX3.2.1.1.3.2-1
The launch and on-orbit center of mass of each sensor subsystem shall be measured and reported to +/- 5 mm, referenced to the sensor coordinate axes as documented in the ICD. Very heavy sensors may require a tighter tolerance.
SRDX3.2.4.1.4.1-1
The moments of inertia shall be defined using coordinates based on the spacecraft axes but passing through the sensor center of mass.
SRDX3.2.4.1.4.2-1
Moments of inertia values shall be accurate to within +/-10% (TBR)
SRDX3.2.4.1.4.3-1
The moments of inertia of each separately mounted subsystem of the sensor shall be provided to the spacecraft contractor for documentation in the ICD, referenced to the sensor coordinate axes.
SRDX3.2.4.1.4.4-1
If the sensor contains movable masses, expendable masses, or deployables, the respective moments of inertia variations shall be provided to the spacecraft contractor for documentation in the ICD.
All documents shall provide units in metric.
SRDX3.2.4.2-2
All interfaces shall be specified in the international system of units, System Internationale (SI), unless design heritage precludes this.
SRDX3.2.4.2-3
Dimensioning shall be in the as-designed units and identified when other than SI.
SRDX3.2.4.2-4
The design of the sensor shall meet the dimensional envelope constraints under a combination of static, dynamic, and thermal conditions encountered during factory assembly, system test, transportation and handling, launch, deployment, and on-orbit operations.
The spacecraft contractor is responsible for defining available sensor volume and making sure the spacecraft fits within the dynamic envelope of launch vehicle’s fairing. This is controlled with the spacecraft to launch vehicle Interface Control Document (ICD). Both the spacecraft contractor and the sensor contractor must work together to insure that the stowed, deploying, and final deployed positions of the sensor will clear all obstacles including obstacles on the spacecraft, other sensors, and the launch vehicle. If the sensor is to be deployed, all obstacles must be cleared in the stowed, deploying, and final deployed positions. If the sensor has moving assemblies, all obstacles must be cleared within the region of motion.
As a baseline, a 2.5 cm clearance between the sensor and surrounding structure will be maintained.
A critical clearance analysis will be conducted to identify areas where the 2.5 cm clearance rule may be violated, accounting for miscellaneous support hardware such as wire bundles and thermal blankets, deflections due to launch loads, launch vibrations, 1-g sag, thermal distortions, and misalignments, with all identified areas tracked in a critical clearance document.
SRDX3.2.4.1.1.1-1
The sensor contractor shall provide to the spacecraft contractor information on the sensor subsystem envelope (including thermal blankets) for documentation in the ICD. Documentation is to be in the form of engineering drawings with a set of "not to exceed" dimensions.
The mounting method is to accommodate manufacturing tolerance, structural, and thermal distortions.
SRDX3.2.4.2.1.2.1 -1
The sensor contractor shall comply with the mounting specification in the ICD.
3.2.4.2.1.2.2 Mounting Interface
The spacecraft mounting interface requirements for each sensor subsystem shall be delivered to the spacecraft contractor for documentation in the ICD.
SRDX3.2.4.2.1.2.2.2-1
The sensor contractor shall comply with the coordinates and dimensions of the holes for mounting hardware as specified at the mechanical interface and defined in the ICD.
SRDX3.2.4.2.1.2.3.1-1
If drill templates are used for simple planar interfaces, then sensor equipment, spacecraft, and test fixture interfaces shall be drilled using templates.
The drill template fabrication and functional requirements (e.g. material, use of inserts, etc.) are to be provided by the sensor contractor.
3.2.4.2.1.2.3.3 Drill Template Provider
SRDX3.2.4.2.1.2.3.3-1
The sensor contractor shall provide the spacecraft contractor with the drill template containing appropriate alignment and location reference information.
The spacecraft contractor is to provide all sensor mounting hardware including secondary structures.
Sensor mounting hardware is to be defined and documented in the ICD.
Finish and flatness requirements for the mounting surfaces are to be specified by the spacecraft contractor and documented in the ICD.
SRDX3.2.4.2.1.2.5-1
The sensor contractor shall support the spacecraft contractor in determining the location of the sensor on the spacecraft.
SRDX3.2.4.2.1.2.5-2
The sensor contractor shall provide necessary mounting information to the spacecraft contractor for documentation in the ICD.
SRDX3.2.4.2.1.3.1-1
The sensor contractor shall be responsible for measuring the alignment angles between the sensor boresight (line of sight), if applicable, and the sensor alignment reference.
The spacecraft contractor is responsible for aligning the sensor alignment reference to the spacecraft attitude reference.
SRDX3.2.4.1.3.2-1
The sensor contractor shall provide a sensor alignment reference.
SRDX3.2.4.1.3.2-2
The sensor alignment reference shall be viewable from two orthogonal directions.
The spacecraft contractor is to control the alignment of the sensor alignment reference with respect to the spacecraft attitude reference to within values specified by the sensor contractor.
The spacecraft contractor is to measure the alignment between the sensor alignment reference and the spacecraft attitude reference. The RMS uncertainty in the alignment knowledge is to be less than 25 arcsec per axis. This uncertainty is to include (if applicable), but not be limited to, measurement uncertainties, alignment shifts due to vibration environments in both ground processing and launch, uncompensated gravity effects, hygroscopic effects of composite materials, and component removal and replacement.
The spacecraft contractor is to limit the rms uncertainty in the alignment between the sensor alignment reference and spacecraft attitude reference caused by structural thermal distortion due to the on-orbit thermal environment to be less than 10 arcsec per axis.
The spacecraft is to supply a three-axis attitude of the spacecraft attitude reference for ground processing. The supplied attitude will be time-tagged and possess an angular rms accuracy per axis of 10 arcsec over a bandwidth of DC to 10 Hz.
The rms of all components of the attitude error of the spacecraft attitude reference with a frequency greater than 10 Hz will be less than 5 arcsec per axis.
The rms of the attitude reference control error over a bandwidth of DC to 10 Hz is to be less than 0.01 deg per axis.
The spacecraft will provide a spacecraft ephemeris estimate with an rms uncertainty of 25/25/25 meters for radial/in-track/cross-track components.
The spacecraft is to provide structural support for the sensor such that the loads transmitted across the interface into the sensor do not exceed interface limit loads to be determined by the spacecraft contractor.
SRDX3.2.4.2.1.4-1
The sensor and interface equipment shall be designed to meet the design load factors determined by launch vehicle acceleration levels.
For a deployable, the spacecraft contractor is to specify a deployed frequency such that the sensor will not saturate the spacecraft's control authority.
SRDX3.2.4.2.1.5-1
When the sensor is in its launch-locked configuration, the fundamental natural frequency of the sensor shall be 50 Hz or greater, axial and lateral.
SRDX3.2.4.2.1.5-2
The sensor contractor shall ensure that the sensor dynamic characteristics and control capability (e.g., a gimbaled sensor) will meet the requirements specified for the deployed frequency.
SRDX3.2.4.2.1.5-3
The lowest natural frequency for a deployed sensor shall be greater than 6 Hz (TBR).
SRDX3.2.4.3.1-1
Primary power distribution to power components shall be compatible with system and subsystem EMC performance requirements.
SRDX3.2.4.3.1-2
Secondary power distribution to power components shall be compatible with system and subsystem EMC performance requirements.
SRDX3.2.4.3.2-1
Sensor Suites shall be designed to operate from a 28 +/- 6volt dc (TBR ) power subsystem.
SRDX3.2.4.3.3.1-1
The electrical interfaces (Figure 3.2.4.3.3.1) shall include the following:
Figure 3.2.4.3.3.1. Spacecraft-Sensor Electrical Interfaces
SRDX3.2.4.3.3.1.2-1
The power source-generated and load-induced ripple, including repetitive spikes, shall not exceed 1.0 volts peak-to-peak as measured over the bandwidth of 30 Hz to 1.0 kHz, and 0.5 volts peak-to-peak from 1.0 kHz to 10 MHz when the power system is delivering the maximum rated current into loads.
SRDX3.2.4.3.3.1.3-1
Loads shall not produce reflected ripple greater than the limits of MIL-STD-461D, part 3, CEO1 and CEO3. CEO1 maximum levels apply to loads that are 15 amps/450 watts and greater.
SRDX3.2.4.3.3.1.3-2
CEO1 maximum emissions shall be reduced by 20 dB for each 20 amps reduction in current.
SRDX3.2.4.3.3.1.4-1
Positive and negative voltage surges shall decay to within steady state limits in less than 5 and 100 milliseconds, respectively.
SRDX3.2.4.3.3.1.4-2
All sensor components shall remain undamaged when subjected to step changes of the input voltage from 0% to 140% and from 120% to 0% of the nominal load voltage (28 volts). The step changes, exclusive of spikes, are the instantaneous surge amplitudes produced by load switching and the clearing of faults on the space-vehicle power bus.
SRDX3.2.4.3.3.1.4-3
With step changes from 0% to 100% of the nominal load voltage, the instantaneous inrush current shall not exceed 4-times the maximum average input current.
The spacecraft is to be able to remove bus power to all sensors if the bus voltage drops below 22 volts. Control heaters are to be turned off during these occurrences. This does not apply to survival heaters.
The spacecraft bus impedance at the interface looking back at the source is to be less than 100 milli-ohms resistive and 5 micro-henries inductive.
Three types of power are to be supplied to each sensor.
SRDX3.2.4.3.3.2.1-1
Direct bus connection shall be through a 5 ampere spacecraft fuse for 50 watt heater load maximum (TBR).
SRDX3.2.4.3.3.2.1-2
Thermal analysis shall be done to verify the adequacy of a 5 amp survival heater.
SRDX3.2.4.3.3.2.1-3
The sensor shall have two series thermostats to ensure fault-tolerant usage of this bus.
Bus connection is to be made through a 5 ampere fuse and relay switch in the spacecraft for a 50 watt heater load maximum (TBR).
Two supply circuits types are to be provided:
The primary power source (battery, power converter) is to be chassis grounded at only one point to avoid large structure current flow which might interfere with other spacecraft loads. The method used to reference the signal back to the secondary power return is dependent on the signal type. The goal is to minimize the voltage drop across the return.
SRDX3.2.4.3.3.3.1-1
The sensors shall have a single point ground.
SRDX3.2.4.3.3.3.1-2
Secondary power and signal returns shall be isolated from the primary power return by not less than 1 Meg-ohm when the sensor is disconnected from the interface/spacecraft and when measured at the sensor input. The secondary grounds may be grounded to structure if the local structure is conductive.
SRDX3.2.4.3.3.3.1-3
The impedance between the sensor and spacecraft single point ground shall be less than 10 Meg-ohms. This impedance may need to be smaller if the expected voltage drop and current flow exceed either the required error in the sensor reference or the radiated emission requirements.
The spacecraft contractor is to provide all spacecraft interface mating connectors.
SRDX3.2.4.3.3.4.1-1
For the standard electrical connector, separate D connectors, such as Cannon NM-K52 (rated at 5 amps, derated below 5 amps), military D subminiature non-magnetic/no-outgas connectors as described in MIL-C-24308, or Positronic SAD Series connector (rated at 7.5 amp, derated below 7.5 amps for use when load requires 5 amps) and Kem connector accessories shall be used for primary power, survival heater power, and analog thermistor returns (TBR).
SRDX3.2.4.3.3.4.1-2
The contractor shall derate electrical connectors using MIL-STD-1547 as a guide.
SRDX3.2.4.3.3.4.1-3
All interface circuits shall be categorized by signal type using DOD-W-83575 as a guide..
SRDX3.2.4.3.3.4.1-4
Primary and redundant connectors shall be differentiated by clearly marking all boxes and cables. Interface requirements for sensor electrical connectors are as follows: (TBR).
SRDX3.2.4.3.3.4.2-1
All power harnesses shall be #20 AWG, with 150 °C insulation. (TBR).
SRDX3.2.4.3.3.4.2-2
Twisted pairs shall be used to reduce magnetic contribution.
SRDX3.2.4.3.3.4.2-3
The wire current-handling capability shall be calculated at the ambient temperature. (TBR).
SRDX3.2.4.3.3.4.2-4
The sensor contractor shall determine the proper wire insulation requirements for any wire directly exposed to the space environment.
Fault isolation is to be included on the spacecraft side of the interface. The spacecraft is to be capable of removing any load in excess of 30 watts. The fault isolation is to either open the circuit to remove the load and short circuit from the spacecraft, or limit the current to the maximum specified load current. Fuses and circuit breakers are to be sized to protect wire between the bus and the sensor. The wire is to be sized to the maximum load. The fuses are to be derated by a factor of three.
SRDX3.2.4.3.3.5-1
Data and telemetry signals shall be segregated and routed from any power circuitry via a separate connector.
SRDX3.2.4.3.3.5-2
The voltage drop across any secondary return shall be less than the maximum allowable noise on the signal circuit reference.
SRDX3.2.4.3.3.5-3
Digital or analog cross talk between any two signal lines in data connectors shall be no greater than -20 dB at the maximum data rate.
The NPOESS System Survivability requirements are contained in Appendix B.
SRDX3.2.4.5-1
The on-orbit design life of the sensor, shall be no less than 7 years.
SRDX3.2.4.5-2
The design of the sensor shall be such that sensor storage, under controlled conditions, may be planned for as long as 8 years, including up to 3 years for intermittent testing.
SRDX3.2.4.5-3
The design service life of the sensor shall be at least 15 years. This includes the time allowed for test, storage, prelaunch checkout, launch and injection, on-orbit, recovery, and contingency time.
SRDX3.2.4.6-1
The finishes used shall ensure that completed devices are resistant to degradation caused by environmental conditions and galvanic action.
SRDX3.2.4.6-2
The sensors shall have special coatings for protection of surfaces against deterioration in space environments.
SRDX3.2.4.6-3
The sensors shall have special coatings for electrostatic discharge suppression in all environments.
SRDX3.2.4.6-4
The sensors shall not use cadmium or zinc platings.
SRD 3.2.4.6-5
Pure tin or tin alloy (>98% Sn) plating shall not be used on electrical devices and hardware for launch and space vehicles. The guiding document for this prohibition is MIL-STD-1547B, "Electronic Parts, Materials, and Processes for Space and Launch Vehicles."
SRDX3.2.4.6-6
Both metallic and insulating surfaces in electronic boxes, such as printed wiring assemblies, where contamination could cause electrical malfunction shall be conformally coated unless otherwise insulated or hermetically sealed.
SRDX3.2.4.6-7
This shall apply to electrical components in sensors and their associated ground equipment. MIL-I-46058, or equivalent can be used in selection of conformal coatings and their thicknesses.
SRDX3.2.4.6-8
Unjacketed flexible shielded cable and ground straps shall be specifically excluded from this conformal coating requirement.
SRDX3.2.4.6-9
Certain components will suffer significant performance degradation if conformally coated. In these situations, non-use of conformal coatings on electrical components and hardware shall be supported by a thorough analysis and be specifically approved by the government on a case by case basis.
SRDX3.2.4.6-10
There shall be no destructive corrosion of the completed devices if exposed to moderately humid or mildly corrosive environments during manufacture or handling.
SRDX3.2.4.7.1-1
All interface requirements specified in Section 3.2.4.7 shall be met at the mechanical interface.
The sensor thermal design shall provide for:
SRDX3.2.4.7.1-2
a. Maintaining the sensor within operating and survival temperature limits,
SRDX3.2.4.7.1-3
b. Maintaining the sensor at the minimum turn-on temperature via survival power,
SRDX3.2.4.7.1-4
c. Minimizing thermal gradients within the sensor,
SRDX3.2.4.7.1-5
d. Thermal decoupling of the sensor from the spacecraft.
The spacecraft is not be used as a heat source or sink.
SRDX3.2.4.7.2-1
The sensor shall be designed to maximize thermal isolation.
SRDX3.2.4.7.2-2
Sensor components shall be designed to maintain the sensor within its allowable temperature limits.
SRDX3.2.4.7.2-3
The thermal control unit shall be mounted on the sensor, where possible, or insulated in order to minimize thermal load to the spacecraft.
SRDX3.2.4.7.3.1-1
The heat transfer between the sensor (or interface plate) and the spacecraft shall not exceed 10.0 (TBR)watts maximum. Sensors with high power dissipation near the interface, or configuration requirements that do not lend themselves to thermal-isolation methods, require the contractor to develop mission sensor-specific heat-transfer rates.
SRDX3.2.4.7.3.1-2
For design purposes, the 10.0 watts heat transfer shall be applied in a worst-case scenario.
The spacecraft contractor is to provide the radiative loads to the sensor.
SRDX3.2.4.7.3.2-1
Incident radiation between the spacecraft and a sensor on any given surface shall be minimized.
SRDX3.2.4.7.3.2-2
The environmental fluxes, as shown in Table 3.2.4.7.3.2 below, shall add solar, albedo and earth IR hot fluxes for the hot case analysis and cold fluxes for the cold case analysis.
Table 3.2.4.7.3.2 Worse-Case Hot and Cold Environments
Hot Case Cold Case
BTU/hr-ft2 W/m2 BTU/hr-ft2 W/m2
Solar Radiation 444 1400 415 1308
Albedo 172 542 86 271
Earth IR 83 262 60 189
Radiation
SRDX3.2.4.7.4.1-1
For planning and preliminary design purposes, the interface temperature of the spacecraft is to be initially assumed to range from:
a. +5 C (TBR) to +40 °C during normal operations
b. -20 C to +50 C during survival modes
SRDX3.2.4.7.4.2-1
Thermal uncertainty margins used during the design and validation shall be applied to determine acceptance ranges in accordance with MIL-STD-1540C.
If heaters are employed, 25% heater control authority can be used in place of 11 °C uncertainty margin. Protoqualification ranges can be calculated by adding additional margins ±5 °C .
Temperature limits for sensor components during ground test and orbital operations shall be provided to the spacecraft contractor for documentation in the ICD.
SRDX3.2.4.7.4.3-2
Operating, non-operating, survival and turn-on temperature requirements shall be provided.
The spacecraft is to monitor and report in the spacecraft telemetry the temperature of the spacecraft at the sensor mechanical mounting interfaces.
All critical sensor temperatures are to be measured and reported in the health and status telemetry data.
SRDX3.2.4.7.5.3-1
The location of all sensor and mounting interface temperature sensors shall be provided to the spacecraft contractor for documentation in the ICD.
SRDX3.2.4.7.6-1
The sensor thermal-control subsystem shall be designed to minimize heater power requirements.
SRDX3.2.4.7.6-2
Use of passive thermal control techniques shall be maximized.
The sensor contractor shall provide to the spacecraft contractor information on sensor provided thermal control hardware for documentation in the ICD. The responsibility for providing the thermal control hardware is defined in Table 3.2.4.7.6.1.
Table 3.2.4.7.6.1. Thermal Control Hardware Responsibility
Hardware Responsibility Survival Heaters Sensor Provider Sensor, Thermal Control Sensor Provider Hardware, including blankets, louvers, and heat pipes Thermal insulation Blankets Spacecraft Provider to Interface between the Sensor Thermal Blankets and the Spacecraft Thermal Blankets
Electrical power for survival heaters is to be provided by the host spacecraft to accommodate at least two strings, a primary and secondary string, of sensor survival heaters.
SRDX3.2.4.7.6.2-1
Sensors shall use survival heaters to maintain temperature at the safe turn-on level.
SRDX3.2.4.7.6.2-2
Operational heaters shall be controlled by the sensor.
SRDX3.2.4.7.6.2-3
Survival heater circuits shall not exceed 0.5 (TBR) amperes per string.
SRDX3.2.4.7.6.2-4
Survival heater circuits shall be provided directly to thermostatically controlled heaters on the sensor side.
SRDX3.2.4.7.6.2-5
Survival heaters shall be capable of operation when the sensor power is off.
SRDX3.2.4.7.6.2-6
The interface shall also have the capability to accommodate up to five analog thermistors (3 kohms at 25 °C) per interface. These analog lines are separate and in addition to any state of health (SOH) input being transmitted over the serial data bus interface and are intended to provide insight during periods when the sensor power is off; therefore, excitation of thermistors is to be provided by the spacecraft.
When the sensor is unpowered, the survival heaters is to be controlled by the spacecraft through the sensor thermistor inputs.
SRDX3.2.4.7.6.2-7
Redundant thermostats shall be used.
The spacecraft contractor is to approve the Multilayer Insulation (MLI) selection in the Parts, Materials and Processes Control Board (PMPCB) review process.
SRDX3.2.4.7.6.3-1
MLI used in thermal control design shall have the following provisions: venting, interfacing with spacecraft thermal control surfaces, and electrical grounding to prevent Electro-Static Discharge (ESD).
SRDX3.2.4.7.6.4-1
Thermal control surfaces shall be cleanable to visibly clean or better.
SRDX3.2.4.7.6.4-2
Any sealed or closed system such as heat pipes, thermal control enclosures or fluid loops shall be analyzed to demonstrate that no safety hazard exists.
SRDX3.2.4.8.1.1-1
All signal interfaces shall use shielded conductors. Conductors may include, but are not limited to, twisted pair, coaxial, twinaxial, dual coaxial types, and fiber optics.
SRDX3.2.4.8.1.2-1
The sensor shall maintain electrical isolation of greater than 100 kohm between the primary and redundant interface circuitry within the sensor front end.
SRDX3.2.4.8.1.3-1
The sensor and spacecraft bus shall be tolerant of a single fault occurring in a signal interface circuit on either side of the interface.
The characteristics of the power bus with respect to power return will be as specified in the Electromagnetic Compatibility (EMC) Control Plan.
SRDX3.2.4.8.2.1-1
The Command and Telemetry Data bus shall be utilized as shown below and in Figure 3.2.4.8.2 .
- real time commands
- stored commands
- memory loads
- frame sync and time code data
- sensor health and status telemetry
- sensor diagnostic data
- low rate science data
Note: MIL-STD-1773 is being considered as an option.
Figure 3.2.4.8.2. Data Transfer Interface
SRDX3.2.4.8.2.2.-1
The Command and Telemetry (C&T) Data bus shall be a dual standby redundant data bus that complies in accordance with the requirements of MIL-STD-1553B, Notice 2, all sections.
SRDX3.2.4.8.2.2-2
Additional requirements shall be specified wherever necessary to select MIL-STD-1553 options and to eliminate ambiguities. MIL-STD-1773 is being considered as an option for the C&T and Low-Rate Data Bus.
The spacecraft C&DH is to perform the Bus Controller (BC) function for the 1553 data bus to send data to and collect data from the sensors/subsystems.
SRDX3.2.4.8.2.3-1
Sensors shall interface with the 1553 data bus via a Remote Terminal (RT) as shown in Figure 3.2.4.8.2.3.
SRDX3.2.4.8.2.3-2
Those sensors without an internal 1553 interface shall interface to the data bus via a Remote Terminal (RT).
SRDX3.2.4.8.2.3-3
The sensors shall interface to the dual standby redundant data bus via dual redundant RT(s) to receive data from and send data to the spacecraft upon request.
Figure 3.2.4.8.2.3. Command and Data Handling Interface Topology
SRDX3.2.4.8.3.1-1
The electrical interface of the C&T Data bus shall comply with the requirements of MIL-STD-1553B, Notice 2, all sections.
SRDX3.2.4.8.3.1-2
Each electrical interface between the sensor/RT and the data bus shall be dual redundant.
SRDX3.2.4.8.3.1-3
Each RT shall be individually transformer coupled to both the primary and the redundant data buses.
SRDX3.2.4.8.3.1-4
No single failure in the data bus electrical interface circuit on either the sensor/RT side of the interface or the spacecraft data bus side of the interface shall cause the sensor to lose the capability to communicate with both the primary and the redundant data buses via each functionally distinct RT.
The Bus Controller (BC) is to monitor the data bus status so that no sensor/remote terminal or data bus failure prevents the Bus Controller from maintaining data flow over the Data Bus.
The spacecraft is to deliver the following data to the specified sensor RT-receive subaddresses by conducting single BC to RT Transfers or single RT to RT Transfers (from a spacecraft RT to an sensor RT):
SRDX3.2.4.8.4.1-1
The sensor shall be capable of accepting pulse and serial commands with the characteristics specified:
Pulse Command
a. Logic 0 TBR
b. Logic 1 TBR
c. Load Capacitance TBR
d. Pulse Width TBR
e. Voltage Rise Time TBR
f. Voltage Fall Time TBR
g. Noise Immunity TBR
h. Inductive Spike Suppression TBR
The serial command input will consist of Not Return to Zero (NRZ) data, clock and envelope signals.
SRDX3.2.4.8.4.1.2-1
The RT to Sensor serial command transfer shall consist of a three wire interface. Characteristics of the interface are TBD.
SRDX3.2.4.8.4.2-1
Unless otherwise specified, all commands and memory loads delivered to the sensor/remote terminal shall be formatted in accordance with the CCSDS Telecommand packet defined in CCSDS 203.0-B-1.
SRDX3.2.4.8.4.3-1
All sensor commands and memory load packet descriptions shall be delivered to the spacecraft contractor for documentation in the ICD.
SRDX3.2.4.8.4.4-1
Initiation of critical or hazardous functions shall use, as a minimum, separate enable and execute commands to prevent inadvertent execution of critical commands.
The spacecraft is to provide frame sync and time code data signals to the sensors.
The format of the vehicle time code words is to be based on the GPS UTC time representation. On-board absolute correlation of time is to be 1 millisecond or better with a correlation to 1 microsecond as a goal. Time representation is to be transmitted over the 1553 data bus or 1773 data bus once per second and is to correspond to the time of the rising edge of the time-of-day pulse.
SRDX3.2.4.8.5-1
Sensor health and status telemetry data shall include housekeeping data required for sensor status and health monitoring. Sensor health and status telemetry includes:
During sensor anomaly resolution, the spacecraft C&DH is to have the capability to dwell on particular telemetry measurands within the selected telemetry format in support of ground diagnostic investigation of the sensor anomaly. Dwell capability is to be a ground initiated process.
Low rate science data is defined as the user mission data from the sensors identified to produce output data rates less than 100 kbps. The spacecraft C&DH is to collect the low rate science data from the respective sensors through a sequence of data transfers over the 1553 data bus.
SRDX3.2.4.8.6.1-1
All telemetry and low rate data shall be packetized using the CCSDS Path Protocol Data Unit format in accordance with CCSDS 701.0-B-1.
The combined rate at which the spacecraft transmits commands, samples telemetry and collects low rate mission data to/from the sensors/subsystems, the maximum duration of a data transfer cycle and the minimum time gap between transfer cycles is to comply with the MIL-STD-1553, Notice 2 specification.
The bus sampling rates for each sensor are to meet the sensor service requirements.
SRDX3.2.4.9.1-1
A redundant High Rate (Mission Data) Bus shall be utilized to transfer the High Rate Science (Mission) Data from a high rate sensor(s) to the spacecraft C&DH.
A high-rate data bus is to be used for a sensor with data rates of > 100 Kbps or to a sensor suite with combined data rates of > 100 Kbps.
SRDX3.2.4.9.3-1
The high rate data bus shall be in compliance with MIL-STD-1773B. (TBR)
SRDX3.2.4.9.4-1
All data to be transferred to the spacecraft C&DH via the high rate data bus shall be packetized using the CCSDS Path Protocol Data Unit format defined in CCSDS 701.0-B-1.
Reliability is defined as the probability that an item can perform its intended function for a specified interval under stated conditions.
Each sensor suite's reliability shall be no less than 0.86 at the end of 7 years on orbit life. (TBR).
SRDX3.2.5.1-2
The sensor suite shall be operational 24 hours per day with no on-orbit repair capability.
The on-orbit design life of the sensor, shall be no less than 7 years.
SRDX3.2.5.1.1-2
The design of the sensor shall be such that sensor storage, under controlled conditions, may be planned for as long as 8 years, including up to 3 years for intermittent testing.
SRDX3.2.5.1.1-3
The design service life of the sensor shall be at least 15 years. This includes the time allowed for test, storage, prelaunch checkout, launch and injection, on-orbit, recovery, and contingency time.
The sensor suite design should include maintainability features to ensure timely replacement or test of sensor subsystems or modules prior to launch.
SRDX3.2.5.1.2-1
Only remove and replace maintenance actions shall be performed on the satellite and sensors after acceptance for shipment or storage by the procuring agency.
SRDX3.2.5.1.2-2
Except for software updates, the sensor suite shall not require maintenance or repair on-orbit.
SRDX3.2.5.1.2-3
Single-point failures of the sensors shall be eliminated where practical if they cause loss or serious degradation of the sensor's on-orbit mission.
SRDX3.2.5.1.2-4
Redundancy shall be provided where practical to eliminate critical single-point failures in the sensors and to ensure that the reliability requirements are satisfied.
SRDX3.2.5.1.2-5
For instances of on orbit failure of a sensor suite component, the sensor suite shall place itself into safe mode to await further commands by the spacecraft.
SRDX3.2.5.1.2-6
The contractor shall perform failure analysis to determine failures for which automatic switchover to redundant components is appropriate.
SRDX3.2.5.1.2-7
The sensor suite shall remain in a readiness condition following integration and system performance verification so that it will be available for launch within 60 days (objective is 45 days). The satellite and integrated sensors will support a launch event within 60 days of notification (objective is 45 days).
Specification of natural environment characteristics in the presence of which the sensor must meet all other requirements
SRDX3.2.6.1-1
The sensor shall be compatible with the natural environments for their operational orbits. The following references contain specifications of the natural environment: MIL-STD-1809 (USAF): Space Environments for USAF Space Vehicles; NASA SP-8031: NASA Space Vehicle Design Criteria / Structures; NASA Tech Memorandum 100471: Orbital Debris Environments for Spacecraft Designed to Operate in Low Earth Orbit; and the Handbook of Geophysics and Space Environments.
SRDX3.2.6.1.1-1
The sensor shall be capable of meeting the proton and electron total dose levels for a 7-year mission given in Table 3.2.6.1.1 below.
SRDX3.2.6.1.1-2
Two times the total dose shall be used to provide a design margin factor of two (Note: aE+N=a x 10N, e.g. 3.264E+06 = 3.264 x 106; one mil is 10-3 inch).
Table 3.2.6.1.1 Total Ionizing Dose Environment
SHIELDING Trapped Protons/7 Yr Trapped Electrons/7 Yr
Mils (AI)
100 6.50 E03 1.81 E04
200 4.79 E03 2.06 E03
400 3.67 E03 6.76 E01
600 3.05 E03 4.35 E01
1000 2.25 E03 3.04 E01
SRDX3.2.6.1.2.1-1
The sensor shall be capable of meeting all performance requirements in the Cosmic Ray and
High Energy Proton Radiation Environment specified in 3.2.6.1.2.1.1 and 3.2.6.1.2.1.2.
SRDX3.2.6.1.2.1-2
Predictions of single events (i.e. single event latch-up, single event upset and single event burn-out) induced by galactic cosmic ray ions and high energy protons shall be performed separately and the results combined.
The integral galactic cosmic ray linear energy transfer spectrum in (TBS) shall be used for prediction of ion-induced single events.
SRDX3.2.6.1.2.1.2-1
The differential proton fluence in (TBS), which consists of trapped protons and galactic cosmic ray protons shall be used for prediction of proton-induced single events in the absence of solar flares.
SRDX3.2.6.1.2.1.2-2
The differential proton fluence in (TBS), which consists of trapped protons, galactic cosmic ray protons and solar flare protons, shall be used for prediction of proton-induced single events with solar flares.
SRDX3.2.6.12.2.1.3-1
The sensor shall be capable of meeting all performance requirements when exposed to trapped proton (E 5 MeV) flux of (TBS) particles/cm2sec, trapped proton (E 0.5 MeV) flux of (TBS) particles/cm2sec with the following estimated solar flare proton peak fluxes and associated total event integral fluences for each extremely large solar flare:
Energy Flux Total Event
(MeV) (Particles/cm2sec) Integral Fluence
(Particles/cm)
>10 TBS TBS
>30 TBS TBS
>60 TBS TBS
> 100 - TBS
The total event integral fluence is accumulated within a time interval of a few hours to two days.
SRDX3.2.6.1.2.2-1
Prediction of proton-induced displacement damage (also known as the bulk damage) to Charge Coupled Device (CCD) detectors shall be based on the differential proton fluence in (TBS).
SRDX3.2.6.1.2.2-2
Where CCD detectors are used, the design shall incorporate features that minimize the effects of displacement damage.
The sensor shall be designed to withstand a payload fairing internal pressure decay rate of 20 mb/sec.
SRDX3.2.6.2-2
The sensor shall be designed to meet the launch environment.
SRDX3.2.6.2.1.1-1
The worst case effective internal environment within the fairing shall not exceed that caused by the internal fairing wall temperature profiles shown in Figure 3.2.6.2.1.1, with a surface emissivity of 0.1.
SRDX3.2.6.2.1.1-2
The combination of actual temperature and emissivity values may vary but in no case shall it expose the sensor to a thermal environment greater than that specified herein more than 10% of the time nor greater than 5% (in degrees F) above the maximum stated temperature.
Figure 3.2.6.2.1.1 Maximum PLF Inner Temperatures
SRDXX 3.2.6.2.1.3-1
The maximum instantaneous 3-sigma Free Molecular Heating on sensor surfaces perpendicular to the velocity vector at the time of fairing separation shall not exceed 0.1814 watts/cm2 (320 Btu/hr-ft2). Lower values may be achieved at the expense of LV performance and will be addressed on a case-by-case basis.
The quasi-static load factors for the MLV are shown in FigureFigure 3.2.6.2.3.
Figure 3.2.6.2.3 MLV Quasi-Static Load Factors
The maximum in-flight vibration levels will be provided in the LV to Satellite ICD, but are not defined in this SRD. Sensor design should be performed using the Expendable Evolved Launch Vehicle (EELV) acoustic data (provided in the next section).
SRDX3.2.6.2.5-1
The free-field maximum expected sound pressure levels (value at 95% probability within 50% confidence), from liftoff through payload separation shall not exceed those shown in Table 3.2.6.2.5 Maximum Acoustic Levels. These levels are shown graphically in Figure 3.2.6.2.5 for the MLV. The values shown are for fill factors of less than 60%. Higher fill factors may produce higher acoustic levels. Some sensors may choose to design to lower levels provided with the optional attenuation shown for MLV.
Figure 3.2.6.2.5 MLV Acoustic Levels
Table 3.2.6.2.5 Maximum Acoustic Levels
1/3 Octave Band MLV Internal PLF Sound MLV Internal PLF Sound Pressure
Center Frequency Pressure Level Level
(Hz) (dB re 20 micropascal) (dB re 20 micropascal)
-with optional blanketing-
32 118.0 -
40 123.4 123.0
50 123.0 122.5
63 124.5 124.0
80 126.0 124.5
100 128.2 126.5
125 129.1 126.0
160 130.0 127.0
200 131.1 127.0
250 130.5 126.5
315 130.0 126.0
400 130.0 125.0
500 129.8 123.5
630 128.3 122.0
800 126.9 119.5
1000 123.9 116.5
1250 122.0 114.0
1600 120.4 112.5
2000 120.9 114.0
2500 117.9 111.5
3150 117.2 110.0
4000 115.5 109.0
5000 114.5 108.5
6300 113.7 108.0
8000 113.9 109.5
10000 114.8 110.5
OASPL 140.5 136.6
SRDX3.2.7-1
The sensor suite and the support equipment that is to be transported with the sensor suite shall be designed for ground and air transportation in accordance with best commercial or military practices.
SRDX3.2.8.1-1
Addition and modification of computer resources in sensors of later flights shall be accommodated by the sensor designs.
For the purposes of this specification, the data processing subsystems of the operational sensor suite are defined to comprise all computer hardware and software.
SRDX3.2.8.1.1-1
The data processing subsystems of the sensor suite shall have 100 percent growth margin while meeting the original functional and performance computational requirements, including timing. This requirement allows the growth margin to be used if the government adds additional requirements.
SRDX3.2.8.1.1.1-1
Within the processing environment of the data processing subsystems of the space elements, each processor shall have an instruction execution rate sufficient to process a workload that is 100 percent greater than the worst case processor utilization workload that could load that processor.
SRDX3.2.8.1.1.2-1
Within the environment of the data processing subsystems of the space elements, the primary memory for each processor shall have 100 percent greater memory capacity than the worst case memory size requirement for that primary memory component, if operating under a non-virtual operating system.
SRDX3.2.8.1.1.2-2
Within the environment of the data processing subsystems of the space elements, the primary memory for each processor shall have, or be capable of having, memory added (through modification, addition, or replacement) to attain, a 200 percent greater memory capacity than the worst case memory size requirement for that primary memory component.
SRDX3.2.8.1.1.3-1
Within the environment of the data processing subsystems of the space elements, each peripheral data storage (secondary memory) component shall have 100 percent greater storage capacity than the worst case storage requirement for that peripheral data storage component.
SRDX3.2.8.1.1.3-2
Within the environment of the data processing subsystems of the space elements, each peripheral data storage (secondary memory) component shall have, or be capable of having, storage added (through modification, addition, or replacement) to attain, a 200 percent greater storage capacity than the worst case storage requirement for that peripheral data storage component.
SRDX3.2.8.1.1.4-1
Within the environment of the data processing subsystems of the space elements, each data transmission medium (e.g., local or global bus or channel) shall have sufficient capacity to support data throughput that is 50 percent greater than the worst case data throughput that could load that data transmission medium.
SRDX3.2.8.1.1.4-2
Within the environment of the data processing subsystems of the space elements, each data transmission medium (e.g., local or global bus or channel) shall have, or be capable of being augmented (through modification, addition, or replacement) to have, sufficient capacity to support data throughput that is 200 percent greater than the worst case data throughput that could load that data transmission medium.
SRDX3.2.8.1.1.5-1
Any hardware augmentations necessary to meet the expansion requirements shall, where practical, be designed so that the software and firmware in the data processing subsystems of the space elements are upward compatible with the implementation of those augmentations.
MIL-STD-1522A should be used as a guide for design and test of all pressurized systems.
SRDX3.3.1-1
Unless otherwise specified, the parts, materials, and processes shall be selected and controlled in accordance with contractor documented procedures to satisfy the specified requirements (reference MIL-STD-1543B).
SRDX3.3.1.1-1
The use of combustible materials or materials that can generate toxic outgassing or toxic products of combustion shall be compliant with applicable federal, state, and local laws and regulations.
Care should be exercised in the selection of materials and processes for the sensor to avoid stress corrosion cracking in highly stressed parts and to preclude failures induced by hydrogen embrittlement.
Parts, materials, and processes should be selected to ensure that any damage or deterioration from storage or the space environment or the outgassing effects in the space environment would not reduce the performance of the sensor beyond the specified limits.
SRDX3.3.1.2-1
Parts for space usage shall be chosen to meet the reliability and operational service life requirements. (Use MIL-STD-1547B and the Preferred Parts List PPL-21, Goddard Space Flight Center, as guides.)
SRDX3.3.1.2-2
Parts shall be selected in accordance with the contractor's Parts Management Plan.
SRDX3.3.1.2-3
The contractor shall be able to demonstrate via design and test or analysis that all parts meet the reliability and operational service life requirements.
SRDX3.3.1.3-1
Materials for the space equipment shall be selected for low outgassing using NASA SP-R-0 022A (NASA JSC) as a guide and resistance to the effects of incident radiation. All support facilities, including test facilities and launch base facilities, should comply with the grounding requirements of MIL-STD-1542B and NOAA S24.809.
SRDX3.3.1.3-2
Materials shall be selected that have demonstrated their suitability for the intended application.
SRDX3.3.1.3-3
Materials shall be corrosion resistant or be suitably treated to resist corrosion when subjected to the specified environments.
SRDX3.3.1.3-4
Where practicable, fungus inert materials shall be used.
SRDX3.3.1.3-5
A-basis material allowables shall be used for design. An A-basis allowable is defined as a value where 99 percent of a population of values is expected to equal or exceed the allowable, with a confidence of 95 percent.
SRDX3.3.1.3-6
Class I Ozone Depleting Substances (ODS) shall not be used in the design, test, manufacture, integration and assembly, handling, transportation, operations, maintenance, or disposal of the sensor..
SRDX3.3.1.3-7
Use of Class II ODS and Emergency Planning and Community Right to Know Act (EPCRA) Section 313 chemicals shall be identified and either eliminated or minimized, justified, and controlled.
SRDX3.3.1.3-8
A Hazardous Materials Management Program shall be developed in accordance with NAS 411.
SRDX3.3.2.1-1
The sensor shall have EMI input filters installed on the sensor side of the power interface. This does not apply to the survival heater circuits which are controlled on the spacecraft side of the interface.
SRDX3.3.2.1-2
The filters shall provide both common-mode and differential-mode filtering capable of meeting EMC requirements in accordance with MIL-STD-1541A.
SRDX3.3.2.1-3
The filters shall be designed to withstand and suppress electrical transients.
There are five macro-level interfaces to consider for EMC:
1) Interface between sensors and spacecraft bus
2) Interface between spacecraft and external environment
3) Interface between spacecraft and launch vehicle
4) Interface between spacecraft and ground support equipment
5) Interface between spacecraft bus/sensors and test equipment
There are four EMC interfaces:
1) Conducted Emissions/Susceptibility
2) Radiated Emissions/Susceptibility
3) Grounding
4) Wiring
SRDX3.3.2.2.2.1-1
The Electromagnetic Compatibility (EMC) requirements shall be in accordance with MIL-STD-461D and MIL-STD-1541A.
SRDX3.3.2.2.2.1-2
The sensor shall not impact the search and rescue mission during any mode of operation.
SRDX3.3.2.2.2.1-3
The spacecraft bus, sensors, ground support equipment, and test equipment shall be operated without performance degradation with each other, and the external environment.
SRDX3.3.2.2.2.2-1
Each interface margin shall be the larger of the following:
1) At least 12 dB.
2) At least large enough to cover manufacturing variations from spacecraft to spacecraft and end-of-life variations.
SRDX3.3.2.2.2.2-2
Electro-explosive devices circuits shall have at least a 20 dB margin.
SRDX3.3.2.2.3-1
The system shall operate without performance degradation in the following external environment.
Frequency V/m
10 k-100M TBD
100 M - 1 G TBD
1 - 10 G TBD
10 G - 40 G TBD
SRDX3.3.2.2.3-2
The intended receivers shall operate without performance degradation for the external environment outside their pass band
SRDX3.3.2.2.3-3
The intended receivers shall survive and automatically recover for the external environment inside their pass band.
SRDX3.3.2.2.2.3-1
The sensor shall operate without performance degradation due to surface charging, bulk charging, and deep charging in accordance with MIL-STD-1541A except paragraph 6.5.2.4.1.
SRDX3.3.2.3.3-1
Power cables, supply and return wires shall be twisted to reduce electromagnetic contribution.
SRDX3.3.2.3.3-2
The power wiring shall be shield twisted pairs with EMI backshells.
SRDX3.3.2.3.3-3
The shield shall be terminated on the backshell.
SRDX3.3.2.3.3-4
Each power wire shall have a dedicated return.
SRDX3.3.2.3.4-1
The interfaces shall meet the requirements (including CE101, CE102, CE106, CS101, CS103, CS104, CS105, CS114, CS116, RE101, RE102, RS101, and RS103) of MIL-STD-461D as tailored by MIL-STD-1541A as tailored within. CS114 and CS116 only apply to power cables. CS103, CS104, CS105, and CE106 apply only to subsystems with transmit/receive antennas. The upper frequency range of RE102 and RS103 shall be extended to envelope all sensor, transmitter, and receiver frequencies.
SRDX3.3.2.3.4.1-1
The requirements shall be tailored by the magnetometer requirements (TBD).
SRDX3.3.2.3.4.2-1
The requirement shall be less than 100 dBµV/m except as tailored for the search and rescue receiver, the UHF receiver, the Satellite Ground Link System (SGLS) receiver, sensor receivers, and launch vehicle receivers (TBD).
SRDX3.3.2.3.4.3-1
The requirements shall be tailored by the magnetometer requirements (TBD).
SRDX3.3.2.3.4.4.-1
The requirements shall be tailored by the external RF environment and the UHF transmitter, the SGLS transmitter, and the launch vehicle transmitter (TBD).
SRDX3.3.4-1
Critical steps of fabrication which are item-peculiar shall be detailed in drawing notes and figures which include appropriate workmanship criteria.
SRDX3.3.4-2
Workmanship relating to all other aspects of fabrication shall be in accordance with the Quality Control Plan approved for each manufacturing facility.
SRDX3.3.5-1
All sensor subsystems shall be configured for modular replacement of components to expedite maintenance and repair.
SRDX3.3.5-2
All components, assemblies, subassemblies, and modules that are identical with respect to fit, form, and function shall be interchangeable.
SRDX3.3.5-3
Parts that are not functionally, electrically and dimensionally interchangeable shall have different part numbers.
SRDX3.3.6-1
Hazards to personnel, hardware, or the environment shall be identified, controlled, or eliminated during design, test, manufacture, integration and assembly, handling, transportation, and operations of the sensor.
SRDX3.3.6-2
Design and operational safety requirements shall be developed and implemented to eliminate or control personnel, hardware, or environmental hazards.
SRDX3.3.6-3
Sensor suites developed for this program shall comply in accordance with EWR 127-1 in the areas of design safety, flight termination, launch integration, and ground operations.
SRDX3.3.6-4
Software controlling hazardous systems or operations (e.g., propulsion systems, electro-explosive devices, electromechanical release devices, etc.) shall be assessed for hazard severity and probability in accordance with AFM 91-201 and using MIL-STD-882C as a guide.
SRDX3.3.6-5
A sensor safety program shall be established and maintained using MIL-STD-822C as a guide.
SRDX3.3.6.1-1
Interfaces shall be designed in accordance with the safety requirements in EWR 127-1.
SRDX3.3.6.1-2
The use of electro-explosive devices (EEDs) shall be avoided. EEDs may be used where the use of such devices can be shown to reduce risk.
SRDX3.3.6.1-3
Paraffin and other non-explosive actuators (NEA) shall be activated through the standard command and data interface, and within the sensor envelope.
SRDX3.3.6.1-4
Dedicated EED circuits shall not be included in the baseline standard interface.
SRDX3.3.6.1-5
Space debris shall not be generated.
SRDX3.3.6.1-6
Actuating circuitry shall be two-fault tolerant to unanticipated deployment or release.
SRDX3.3.6.1-7
All applicable safety-related explosive ordnance design requirements shall be met.
SRDX3.3.6.1-8
Explosive ordnance to be installed on a sensor shall be in accordance with DOD-E-83578A.
SRDX3.3.7-1
The operator-hardware and operator-software interfaces shall be designed to maximize safety, efficiency, and usability, and minimize number of personnel, resources, skills, and training.
SRDX3.3.7-2
The operator-software interface shall be developed using open systems technology.
SRDX3.3.7-3
The operator-software interface shall be designed using an iterative design methodology consisting of rapid prototyping, usability evaluation, and feedback to design.
SRDX3.3.8-1
Provisions shall be made for the control of all nuclear materials, such as radioactive sources, if used in manufacturing, calibration, and checkout of certain mission sensors.
Communications security (COMSEC) measures provide protection for the transmission of sensitive information.
SRDX3.3.9.1-1
The probability that the sensor will accept the following as a valid command shall be no more than once over the life of each sensor :
a. An invalid command
b. Noise
SRDX3.3.9.1-2
The sensor shall check for invalid commands and execute only valid commands.
SRDX3.3.9.1-3
Invalid commands shall be reported in telemetry.
Any sensor element that processes multiple security levels of data should comply with DOD 5200.28-STD, paragraph 3.1.1.3.
The government will provide standard scenes for sensor/EDR evaluations, where applicable.
Computer resources include all computer software and the associated computational equipment included within the sensor.
The computational equipment includes processing units; special-purpose computational devices; main storage; peripheral data storage; input and output units.
SRDX3.3.11.1.3.1.1-1
All software and firmware shall be implemented with an internal identifier (embedded in the executable program) that can be included in the sensor engineering data.
SRDX3.3.11.1.3.1.1-2
This identifier shall be keyed to the configuration management process so that the exact version of software and firmware residing in the sensor can be determined at any time.
SRDX3.3.11.1.3.1.2-1
Loading of the sensor microprocessor via hardline shall take no longer than 10 minutes following a reset or power-up.
SRDX3.3.11.1.3.1.3-1
Flight software shall be designed so that complete or partial revisions can be installed and verified on-orbit.
The government intends to determine the standard software language for operational application software after preliminary designs are complete. The government intends to do the same for flight software.
SRDX3.3.11.1.3.2-1
Sensor software shall be written in a higher order language except where assembly language is necessary for the satisfaction of sensor performance requirements or where its use is cost effective over the life of the sensor. (TBR)
SRDX3.3.11.1.4-1
Code shall be written such that no code is modified during execution.
SRDX3.3.11.1.5-1
All information technology items shall be Year 2000 compliant, or non-compliant items shall be upgraded at no additional cost to be Year 2000 compliant by (TBS, but NLT December 31, 1999). Year 2000 compliant means information technology that accurately processes date/time data (including but not limited to, calculating, comparing and sequencing) from, into and between the twentieth and twenty-first centuries, and the years 1999 and 2000 and leap year calculations.
SRDX3.3.11.5.1-2
Year 2000 compliant information technology, when used in combination with other information technology shall accurately process date/time data if the other information technology properly exchanges date/time data with it.
SRDX3.3.12.1-1
The primary support structure for the sensor equipment shall possess sufficient strength, rigidity, and other characteristics required to survive the critical loading conditions that exist within the envelope of handling and mission requirements.
SRDX3.3.12.2.1-1
The structure shall be designed to have sufficient strength to withstand simultaneously the yield loads, applied temperature, and other accompanying environmental phenomena for each design condition without experiencing yielding or detrimental deformation.
SRDX3.3.12.2.2-1
The structure shall be designed to withstand simultaneously the ultimate loads, applied temperature, and other accompanying environmental phenomena without failure.
SRDX3.3.12.3.1-1
The structural dynamic properties of the equipment shall be such that its interaction with the sensor control subsystem does not result in unacceptable degradation of performance.
SRDX3.3.12.3.2-1
Stiffness of the structure and its attachments shall be controlled by the equipment performance requirements and by consideration of the handling and launch environments.
SRDX3.3.12.3.2-2
Special stowage provisions shall be used, if required, to prevent excessive dynamic amplification during transient flight events.
SRDX3.3.12.3.3-1
The fundamental resonant frequency of a component weighing 23 kilograms or less shall be 50 Hertz or greater when mounted on its immediate support structure.
SRDX3.3.12.3.3-2
Detailed analyses of the potential responses of the component to inputs from the adjoining structure(s) shall be required for components weighing 23 kilograms or less and having fundamental resonant frequencies of less than 50 Hertz (TBR).
A test-verified model is preferred when available, and shall be used if the sensor lowest frequency is less than 50 Hz as shown by analysis.
The factor of safety of the structure is the ratio of the limit load to the allowable load.
SRDX3.3.12.4.1-1
The design factors of Table 3.3.12.4.1 below shall be applied to all loading conditions.
Table 3.3.12.4.1 Structural Design Factors of Safety
Design Factor of Safety of Limit
Loads
Design and Test Options Yield Ultimate
(FSy) (FSu)
Unmanned
Events
1. Dedicated Test 1.10 1.25
Article
2. Test One Flight 1.25 1.40
Article
3. Proof Test Each 1.10 1.25
Flight Article
4. No Static Test 1.60 2.00
SRDX3.3.12.4.1-2
The level of required analysis increases significantly with increased option number. For the no-static-test option, a detailed and comprehensive structural analysis shall be conducted. (TBR)
SRDX3.3.12.4.1-3
The structural analysis documentation shall be available for review by the spacecraft contractor. (TBR)
SRDX3.3.12.4.1-4
The flight hardware shall be capable of withstanding all worst-case load conditions to which it may be exposed during ground (handling and transportation), test, pre-launch, launch, and on-orbit operations.
SRDX3.3.12.4.1-5
Positive structural margins of safety shall be maintained so that the sensor can meet all design requirements after being subjected to the worst case load combinations.
SRDX3.3.12.4.1-6
In those cases involving maintenance of sensor critical components for on-orbit operations, the precision elastic limits shall be used for structural materials.
SRDX3.3.12.4.1-7
All composite structures and structural bonded joints shall be proof tested, regardless of safety factor, however, a metallic structure is usually qualified such that each unit will not have to be tested, or it is protoqualed. The no-static-test option allows the capability of the structure to be determined via purely analytical methods. The analytical models are not being verified by test, but verified by the integrator/government for accuracy.
SRDX3.3.12.4.2-1
Sensors with pressurized systems shall follow the requirements in accordance with EWR 127-1 for the design of pressurized systems and using MIL-STD-1522A for guidance.
SRDX3.3.12.4.2-2
Factors of safety for pressure loads shall be determined individually for each pressure vessel, based on tests to establish material characteristics and an analysis of life requirements and other environmental exposure.
SRDX3.3.12.4.2-3
Proof and burst pressure factors shall be established at levels that ensure structural integrity, structural life, and safety throughout all phases. The values listed in Table 3.3.12.4.2 below are to be considered as limiting lower bounds.
Table 3.3.12.4.2 Factors of Safety for Pressurized Components
Design Acceptance Qualificatio
n
Component Ultimate (Proof) (Burst)
Pneumatic Vessels (SVE)a 2.00 1.50b 2.00b
Pneumatic Vessels (GSE)a 4.00 2.00b 4.00
Lines, Fittings, and Hoses
Less than 3.81 cm 4.00 2.00b 4.00b
diameter
3.81 cm diameter and 1.50 1.10b 1.50b
larger
Other Pressurized Components 2.50 2.00b 2.50b
Notes:
a. Factors of safety shown are minimum values applicable to
metallic pressure vessels for which ductile fracture mode
is predicted via a combination of stress and fracture
mechanics analyses. Design of metallic pressure vessels
for which brittle fracture mode is predicted by these
analyses should be in accordance with fracture mechanics
methodology wherein the proof factor as well as the design
ultimate factor of safety shall be established to provide a
minimum of four times the specified service life against
mission requirements. In addition, a fracture control
program should be established to prevent structural failure
due to the initiation or propagation of flaws or crack-like
defects during fabrication, testing, and service life.
b. No measurable (TBR) yielding is permitted at acceptance
(proof) test pressure and no rupture at qualification
pressure.
SRDX3.3.12.5-1
The sensor equipment shall be capable of withstanding all design load conditions to which it is exposed in all mission phases, as applicable: ground, prelaunch, erection, post-launch, boost and orbit.
SRDX3.3.12.5-2
During the orbit phase, all of the following shall be considered: maneuvering loads, vehicle spin, meteoroid environment, radiation environment, and other environmental factors, such as thermal effects due to internal heating, solar heating, eclipses, and extreme cold due to ambient space environment.
SRDX3.3.12.6.1-1
Tubing design shall incorporate provisions for cleaning and to allow proof testing.
SRDX3.3.12.6.2-1
Separable fittings shall have redundant sealing surfaces, such as double "O" rings, and be of the "parallel loaded" type. "Parallel loaded" means that the fitting contains a compressed element which exerts outward pressure on the other elements of the fitting such that both seals are maintained even if relaxation occurs.
SRDX3.3.12.6.2-2
Separable fittings shall have provisions for locking.
SRDX3.3.12.6.2-3
Separable fittings shall be accessible for leak tests and for torque checks.
SRDX3.3.12.6.2-4
Separable fittings shall not be designed or assembled with lubricants or fluids that could cause contamination or could mask leakage of a poor assembly.
SRDX3.3.12.6.2-5
Separable fluid fittings shall not use "B" nuts.
The only fitting assemblies permitted shall meet all of the following constraints:
SRDX3.3.12.6.2-6
(1) The fittings shall be comprised of one or more compressed or internally pressure-energized members which maintain a seal even if stress relaxation occurs in any of the other components.
SRDX3.3.12.6.2-7
(2) The fittings shall have redundant seals in series.
SRDX3.3.12.6.2-8
(3) The fittings shall be lockwired to prevent any rotation between the fitting and the nut.
SRDX3.3.12.7-1
Deployment mechanisms, sensor mechanisms, pointing mechanisms, drive mechanisms, de-spin mechanisms, separation mechanisms, and other moving mechanical assemblies on sensors shall be in accordance with MIL-A-83577B in order to increase reliability of Moving Mechanical Assembly (MMA's) and facilitate integration and test activities.
SRDX3.3.12.7-2
System level test impacts shall be considered in the sensor design; for example, gimballing in a 1 ‘g’ environment, momentum compensation, and the use of GSE.
The spacecraft contractor is to provide estimates of allowable disturbance torque, vibration, and end-of-travel or latch-up loads to the sensor contractor. The design will be an iterative process to accommodate the requirements of both the spacecraft contractor and sensor contractor.
SRDX3.3.12.7.2-1
The sensor contractor shall provide data that will allow the spacecraft contractor to ensure that the "swept" or deployed volume is verified to ease integration and operation, accounting for all distortions and misalignments.
SRDX3.3.12.7.2-2
Early estimate of gimbaled masses, inertia’s and cg’s shall be provided to size the control components to meet pointing and stability requirements.
All sensor mechanisms which require restraint during launch shall be caged during launch without requiring power to maintain the caged condition.
SRDX3.3.12.7.4-1
Each sensor having movable components shall not exceed an uncompensated momentum contribution to be defined and agreed to in an ICD between the sensor contractor and the spacecraft contractor.
SRDX3.3.12.7.5.1-1
The magnitude of the disturbance torque that the sensor imparts to the spacecraft shall be in the acceptable region of Figure 3.3.12.7.5.1.
Figure 3.3.12.7.5.1 Allowable Transmitted Torque
SRDX3.3.12.7.5.2-1
The sensor contractor shall provide the actual sensor torque versus time profile to the spacecraft contractor for documentation in the ICD.
SRDX3.3.12.7.5.3-1
The sensor contractor shall provide to the spacecraft contractor the magnitude and direction of net thrust resulting from the expulsion of expendables by the sensor for documentation in the ICD. (TBR)
The use of large quantities of magnetic materials should be avoided where possible.
SRDX3.3.12.8-1
If magnets are inherent to the sensor design, early estimates of magnetic fields and residual magnetic dipole moments shall be provided to the spacecraft contractor.
SRDX3.3.12.8-2
A full magnetic survey shall be conducted by the sensor contractor if the sensor has a total residual uncompensated magnetic moment greater than (TBD) ampere-turn-meter-square.
SRDX3.3.12.9.1-1
The sensor contractor shall provide to the spacecraft contractor access requirements for documentation in the ICD.
SRDX3.3.12.9.2-1
All items to be installed, removed, or replaced at the spacecraft level shall be accessible without disassembly of the unit.
SRDX3.3.12.10.1-1
The sensor contractor shall provide proof tested handling fixtures for each subsystem.
SRDX3.3.12.10.1-2
Handling fixtures shall be designed to 5 times limit load for ultimate and 3 times limit load for yield.
SRDX3.3.12.10.1-3
Handling fixtures shall be tested to 2 times working load.
SRDX3.3.12.10.2-1
Sensors shall be capable of being mounted to the spacecraft with the spacecraft in the horizontal or vertical position.
SRDX3.3.12.10.3-1
Sensors shall be capable of being mounted or removed without removal of other sensors or subsystem.
SRDX3.3.12.10.4-1
The sensor contractor shall provide to the spacecraft contractor information on all items to be installed or removed prior to flight, for identification in the ICD.
SRDX3.3.12.11-1
Sensor contractors shall provide to the spacecraft contractor the location, size, path and operation time of vents in the sensors for inclusion in the ICD.
SRDX3.3.12.11-2
Sensor purge requirements, including type of purge gas, flow rate, gas purity specifications, filtration and desiccant requirements, and the limit for purge interruption, shall be provided to the spacecraft contractor for documentation in the ICD.
Acceptance and flight certification of sensor equipment is based primarily on an evaluation of data from the manufacturing process.
SRDX3.3.15.1-1
The manufacturing process for sensor equipment shall be accomplished in accordance with documented procedures and process controls which assure the reliability and quality required for the mission.
SRDX3.3.15.1-2
These manufacturing procedures and process controls shall be documented to give visibility to the procedures and specifications by which all processes, operations, inspections, and tests are to be accomplished by the contractor.
SRDX3.3.15.1-3
Each lot shall be manufactured, tested and stored with sequential lot numbers that indicate the date of manufacture assigned to each lot. (Typically, use three digits for the day of the year and two digits for the year.
SRDX3.3.15.1-4
Contractor documentation shall include each material required.
SRDX3.3.15.1-5
Contractor documentation shall include the point it enters the manufacturing flow.
SRDX3.3.15.1-6
Contractor documentation shall include the controlling specification or drawing.
SRDX3.3.15.1-7
The documentation shall indicate required tooling corresponding to each applicable process or material listed.
SRDX3.3.15.1-8
The documentation shall indicate facilities and test equipment corresponding to each applicable process or material listed.
SRDX3.3.15.1-9
The documentation shall indicate the manufacturing check points corresponding to each applicable process or material listed.
SRDX3.3.15.1-10
The documentation shall indicate the quality assurance verification points corresponding to each applicable process or material listed.
SRDX3.3.15.1-11
The documentation shall indicate the verification procedures corresponding to each applicable process or material listed.
SRDX3.3.15.1-12
The specifications, procedures, drawings, and supporting documentation shall reflect the specific revisions in effect at the time the items were produced.
SRDX3.3.15.1-13
Flow charts and the referenced specifications, procedures, drawings, and supporting documentation become the manufacturing process control baseline and shall be retained by the contractor for reference.
SRDX3.3.15.1-14
All changes to the baseline processes used, or the baseline documents used shall be recorded by the contractor following the establishment of the manufacturing baseline. It is recognized that many factors may warrant making changes to this documented baseline; however, these changes provide the basis for flight accreditation of manufactured or of subsequent flight items.
SRDX3.3.15.1-15
The manufacturing process and control documents for sensor equipment shall provide a contractor-controlled baseline that assures that any subsequent failure or discrepancy analysis that may be required can identify the specific manufacturing materials and processes that were used for each item. In that way, changes can be incorporated to a known baseline to correct the problems.
SRDX3.3.15.1.1-1
To the extent practicable, parts for use in sensor equipment shall be grouped together in individual assembly lots during the various stages of their manufacture to assure that all devices assembled during the same time period use the same materials, tools, methods, and controls.
SRDX3.3.15.1.1-2
Parts and devices for sensor equipment that cannot be tested adequately after assembly without destruction of the item, such as explosive ordnance devices, some propulsion components, and complex electronics, shall have lot controls implemented during their manufacture to assure a uniform quality and reliability level of the entire lot.
A system level contamination plan is to be developed by the spacecraft contractor. Sensor and other sensor equipment and facility contamination and cleanliness requirements are to be considered along with the spacecraft requirements and the resulting contamination budget provided to each sensor contractor.
SRDX3.3.15.1.2-1
A contamination control program shall be developed and implemented.
SRDX3.3.15.1.2-2
The contamination requirements shall be provided to the spacecraft contractor for inclusion in the ICD.
SRDX3.3.15.1.2.1-1
The sensor contractor shall perform independent contamination analyses to identify, locate and size components sensitive to contamination and assess, calculate or measure the maximum allowable particulate and molecular film (nonvolatile residue, or NVR) contamination consistent with top level mission performance and lifetime specifications.
SRDX3.3.15.1.2.1-2
The sensor contractor shall provide contamination limits to the spacecraft contractor for documentation in the ICD.
SRDX3.3.15.1.2.2-1
The sensor contractor shall describe the required integration and test environments in accordance with the definitions of FED-STD-209E.
The sensors is to be integrated with the spacecraft in a Class 10,000 (TBR) cleanroom environment and maintained in that environment as much as possible during the integration and test flow.
SRDX3.3.15.1.2.2-2
The facility requirements shall be documented and should include air cleanliness, air flow and recirculation rates, temperature and humidity, and tolerance for out-of-spec conditions (i.e. intermittent spikes) as a minimum.
SRDX3.3.15.1.2.2-3
Requirements shall include verification by standard testing methods to be performed at regular, specified intervals.
3.3.15.1.2.2.1 Ground Support Equipment (GSE) Cleanliness Requirements
SRDX3.3.15.1.2.2.1-1
The sensor contractor shall document the need for contamination control of all GSE entering cleanrooms.
SRDX3.3.15.1.2.2.1-2
All GSE used inside thermal/vacuum facilities shall be cleaned and vacuum compatible.
SRDX3.3.15.1.2.3-1
Inspections and cleaning by the sensor contractor during Integration and Test (I&T) shall be coordinated and defined in the ICD. Sensor contractors should consider having defined "pass/fail" criteria consistent with top level mission performance and lifetime specifications and the derived contamination budget for that I&T milestone. Sensor contractors should pay particular attention to any I&T activity utilizing thermal/vacuum test facilities.
SRDX3.3.15.1.2.4-1
Sensor purge requirements, including type of purge gas, flow rate, gas purity specifications, filtration and desiccant requirements, and the tolerance for purge interruption, shall be provided to the spacecraft contractor for documentation in the ICD.
SRDX3.3.15.1.2.5-1
Fabrication and handling of sensor equipment shall be accomplished in a clean environment.
SRDX3.3.15.1.2.5-2
Attention shall be given to avoiding non-particulate (chemical) as well as particulate air contamination.
SRDX3.3.15.1.2.5-3
To avoid safety and contamination problems, the use of liquids shall be minimized in areas where initiators, explosive bolts, or any loaded explosive devices are exposed.
The particulate cleanliness of internal moving subassemblies should be maintained to at least level 500 as defined in MIL-STD-1246C.
SRDX3.3.15.1.2.6-1
External surfaces shall be visibly clean.
The spacecraft contractor is to coordinate handling of the issue of material outgassing
SRDX3.3.15.1.2.7-1
Sensor contractors shall identify and characterize all sources of contamination that can be emitted from the sensor.
SRDX3.3.15.1.2.7-2
At a minimum, the characterization shall include the material name, the amount, the emission rate, and its location.
SRDX3.3.15.1.2.7-3
The extent to which outgassing products have access to exterior surfaces shall also be considered.
SRDX3.3.15.1.2.7-4
Data from the American Society for Testing and Materials (ASTM) E-595 test for percent total mass loss (%TML) and percent collected volatile condensable material (%CVCM) shall be used. Materials with the following properties are recommended: Total Mass Loss (TML) less than 1.0%; production of Collected Volatile Condensable Material (CVCM) less than 0.1% when tested under conditions of ASTM E595-93 or equivalent. Composite materials are an exception.
SRDX3.3.15.1.2.7-5
The outgassing chemical species and any tendency to photodeposit in an Ultraviolet (UV) or energetic particle radiation environment shall be identified and quantified.
SRDX3.3.15.12.7-6
The sensor contractors shall consider effects on materials due to the planned space environment, including radiation, vacuum, thermal cycles, and atomic oxygen.
SRDX3.3.15.1.2.7-7
If voltage over 60 V (TBR) are present, the design shall be protected (e.g., potting, pressure vessel) and tested for arcing.
SRDX3.3.15.1.2.7-8
Items that might otherwise produce deleterious outgassing while on orbit shall be baked for a sufficient time to drive out all but an acceptable level of outgassing products prior to installation in the experiment or sensor.
Sensor materials selection should minimize the generation of particulate and molecular film contamination via interaction with atomic oxygen (AO). Atomic oxygen fluence is shown in (TBS).
SRDX3.3.15.1.2.8-1
Sensor contractors shall consider the effects of atomic oxygen in the space environment.
SRDX3.3.15.1.2.8-2
The sensor shall meet performance requirements during exposure to atomic oxygen experienced during a 833 km polar orbit for seven years.
SRDX3.3.15.1.3-1
Appropriate provisions shall be made to avoid and to protect against the effects of static electricity generation and discharge in areas containing electrostatic sensitive devices such as microcircuits, initiators, explosive bolts, or any loaded explosive device. DOD-HDBK-263 provides examples of appropriate provisions.
SRDX3.3.15.1.3-2
There shall be a capability to ground both equipment and personnel working on and around the sensor, subsystems, and components.
The sensor contractor is responsible for documenting the functional and physical requirements for the sensors in a hierarchical set of specifications, comprising sensor, subsystem, and component levels. MIL-STD-961D, Notice 1 provides guidance on hierarchical specifications. Lower level specifications may be used to define requirements for software or individual units.
The spacecraft contractor is responsible for developing the Sensor to Spacecraft ICD. The sensor contractor is responsible for providing interface design description data to the spacecraft contractor for inclusion in the Sensor/Spacecraft ICD.
SRDX3.4.3-1
Equipment designed for the sensor shall be documented in drawings and associated lists.
Sensor software and databases should be developed and managed in accordance with MIL-STD-498 and NOAA S24.806 software standards. MIL-STD-498 will take precedence.
Technical manuals should meet the requirements of NOAA Standards S24.801,S24.806, and S24.809 (TM 86-01).
Integrated Logistics Support (ILS) will minimize the impact on NPOESS of the existing support infrastructure while ensuring the lowest NPOESS life cycle cost and while providing full and timely logistics response.
All maintenance procedures will be approved by the government.
The ground support equipment should be addressed by the sensor contractor.
The need for unique support equipment should be minimized by the careful selection of Commercial-Off-The-Shelf (COTS) hardware and software.
The contractor should provide a PHS&T plan which will cover all PHS&T issues for the sensor.
Existing government facilities may be available for operations, maintenance, or storage of sensors.
The high reliability required of the sensor suite is achieved by the designs, design margins, and by the manufacuring process controls imposed at each and every level of assembly. The design and design margins should ensure that the sensor suite is capable of performing its mission in the space environment. The manufacturing process controls are intended to ensure that a known quality product is manufactured to meet the design requirements and that any changes required can be made based on a known baseline. To ensure successful operation of the sensor suite, attention to every detail is required at every level of assembly throughout development, manufacture, qualification, and testing, starting with the parts, materials, and processes used.
SRDX4.1-1
The sensor suite contractor shall ensure that quality assurance requirements flow down to all subcontractors.
Parts, materials, and process controls are to be applied during production of all items to ensure that a reliable sensor is fabricated.
SRDX4.1.1.1.1-1
The tools, processes, parts and materials used in the fabrication of the sensor shall be controlled and inspected to ensure compliance with approved manufacturing processes and controls.
SRDX4.1.1.1.2-1
The contractor shall maintain records documenting the status of the sensor following assignment of serial numbers.
SRDX4.1.1.1.2-2
Each sensor item shall have inspection records and test records maintained by serial number to provide traceability from sensor usage to production lot data for the devices.
SRDX4.1.1.1.2-3
The contractor shall maintain complete records for the sensor items and have them available for review during the service life of the sensor.
SRDX4.1.1.1.2-4
The records shall indicate all relevant test data; all rework or modifications; and all installations and removals for whatever reason.
SRDX4.1.1.1.2-5
Ground test equipment items shall have inspection records and test records maintained by serial number for the service life of the item.
SRDX4.1.1.1.3-1
The contractor shall subject each level of assembly to in-process manufacturing and assembly screens to ensure compliance with the specified requirements to the extent practicable.
SRDX4.1.1.1.3-2
At each level of assembly, the contractor shall subject each completed unit to visual inspection to ensure that it is free of obvious defects and is meets specified physical limits.
SRDX4.1.1.1.4-1
The contractor shall reject all material, components, or assemblies that do not meet the established tolerance limits set for the acceptance limits in the in-process screens. Any rejected material, component, or assembly may be reworked and re-screened in accordance with established procedures, if the rework is not so extensive as to jeopardize the lot identity of the material or assembled unit. If the reworked material or assembled unit subsequently passes the in-process screens, it can again be considered part of the lot.
Non-conforming material or assembled units that do not satisfy these rework criteria will be considered scrap.
SRDX4.1.1.1.4-2
Reassignment of assembled units to a different lot shall not be made.
SRDX4.1.1.1.5-1
The contractor shall perform design verification testing to demonstrate that new or modified designs comply with the specified performance margins.
NPOESS will employ the following test strategy: One flight unit will be subjected to protoqualification level testing, subsequent flight units will be acceptance level tested.
Tests and evaluations of the sensor elements may be conducted at in-plant test facilities, which may include subcontractor's facilities and/or at a government-approved test facility. Tests of parts, materials, software units, ground equipment and computer software may also take place at in-plant test facilities, (including subcontractor's facilities or at a government-approved test bed).
SRDX4.2.3-1
The EDU shall have the form, fit and function of the production unit.
SRDX4.2.3-2
A EDU , developed and fabricated by the sensor contractor, shall be provided for the sensor suite that will be integrated on the spacecraft.
SRDX4.2.3-3
The EDU shall accurately represent (within 0.1Kg) the mass and mass distribution of the actual flight sensor but not necessarily the correct moments of inertia.
The requirement for mass models will be determined prior to PDR. (TBR)
The interface requirements for the mathematical models are defined in this section.
Thermal and structural models are required for the spacecraft and each of the mission sensors to accurately define the thermal, structural and dynamic loads at each sensor/spacecraft interface.
These models are to be used to analyze the structural and dynamic characteristics of each sensor. The models will be used to determine the sensor structural adequacy to withstand transportation, launch and on-orbit loads. Also, they will be used to predict sensor structural resonant frequencies. The satellite/launch vehicle loads analysis will be used in the development and definition of the sensor/spacecraft interface loads.
SRDX4.2.4.1-1
A structural Finite Element Model using NASTRAN shall be provided for the sensor suite.
SRDX4.2.4.1-2
If the sensor has any structural frequencies less than 50 Hz, a test-verified sensor dynamic (modal) model shall be provided.
Two mathematical models, provided by the sensor contractor, will be used in defining each sensor/spacecraft thermal interface. The sensor model will be integrated into a satellite model to define the satellite/launch vehicle interface. All models will use software and formats that are acceptable to the spacecraft contractor. All models shall be fully documented to permit ease of use by other contractors in the system.
SRDX4.2.4.2-1
The sensor contractor shall develop a sensor surface geometric model and a reduced node thermal math model.
SRDX4.2.4.2-2
The contractor shall provide adequate documentation for both models for easy incorporation in system-level models and analyses.
SRDX4.2.4.2-3
The contractor shall develop a geometric math model (GMM), including 50 external surfaces or less, in TRASYS or compatible format.
SRDX4.2.4.2-4
The contractor shall develop a thermal math model (TMM) of 50 nodes or less in a SINDA compatible format to correlate pretest temperature predictions with the test data from the thermal balance test.
SRDX4.2.4.2-5
The TMM shall represent all external surfaces.
SRDX4.2.4.2-6
The thermal models shall include an adequate level of detail to predict, under worst case hot and cold conditions, all critical temperatures, including those that drive operational and survival temperature limits and heater power. Worst case conditions include variations in season, orbit selection, orbital time, and environmental flux parameters (seasonal and spatial) and a rational combination of the effects of design tolerances, fabrication uncertainties, material differences, and degradation due to aging. See Table 3.2.4.7.3.2 for cold and hot worst case environmental parameters.
SRDX4.2.4.2-7
Models shall use conservative values for conduction, absorption, emission, and MLI effective emittance, and consider contact resistance.
SRDX4.2.4.2-8
A thermal math model shall be used to correlate pretest temperature predictions with the test data from the thermal balance test. As a goal, correlation of test results to the thermal model predictions should be within ± 3C.
SRDX4.2.4.2-9
The sensor contractor shall document the nodal description and results of the detailed thermal analysis.
The spacecraft contractor is to provide mission-specific information for maximum equivalent loads to the sensor contractor for his static loads analyses.
SRDX4.2.5-1
A structural analysis using maximum equivalent loads shall be conducted on the sensor suite.
SRDX4.2.5-2
Those sensors with modes under 50 Hz shall have a full modal survey test completed in a base fixed configuration to obtain all mode shapes and frequencies to correlate the dynamics model.
SRDX4.2.5-3
Modal surveys shall be conducted in accordance with MIL-STD-1540C (TBR).
SRDX4.2.5-4
A static loads test shall be conducted in accordance with MIL-STD-1540C (TBR), on the first production sensor if the loads predicted by the model exceed the maximum equivalent flight values.
SRDX4.2.5-5
The sensor contractor shall coordinate on the spacecraft contractor's analyses. The two parties should reach an agreement about assumptions and model development either by test or by close coordination between the two structural analysis groups.
SRDX4.2.6-1
Developmental testing shall be conducted in accordance with MIL-STD-1540C (TBR).
A comprehensive sensor test program, conducted in conjunction with the spacecraft test program, will demonstrate that the sensor can meet its performance requirements and will ensure that all interface requirements are satisfied. These interface requirements will include interface structural and thermal loads, electrical power, electrical signals and other interface performance characteristics for ground handling, launch, deployment (where applicable), and on-orbit operations as well as for worst case systems tests conducted after delivery to the spacecraft contractor. ALL of the tests will be conducted by the sensor contractor before delivery of the instruments to the spacecraft contractor. Additional tests will be conducted at the satellite level after integration of the sensor onto the spacecraft. The types of testing to be performed by the spacecraft contractor include:
• Thermal vacuum and Thermal cycling
• EMI/EMC characterization to understand and measure radiative and conductive emissions and susceptibility
• Static and Dynamic structural testing (including pressure vessel and ordnance testing)
• Electrical and Mechanical functional testing to demonstrate performance
• Calibration: Radiometric and Geometric
SRDX4.2.7-1
Protoqualification level testing shall be required for one flight unit and its subsystems, with acceptance level testing for all remaining flight units and their subsystems.
SRDX4.2.7-2
All protoqualification and acceptance shall be done in accordance with MIL-STD-1540C (TBR) unless specified otherwise in this document
SRDX4.2.7-3
All protoqualification tests shall be conducted with hardware of the final design that has passed the in-process production screens.
SRDX4.2.7-4
The sensor contractor shall perform comprehensive radiometric/geometric calibration at multiple levels of assembly as approved by the government.
The random vibration test levels are dependent on the payload fairing internal acoustic environment and design of the spacecraft bus.
The test levels found in Table 4.2.7.1.1 and Figure 4.2.7.1.1 are considered a conservative estimate of the random vibration environment on a representative spacecraft bus. The test levels shown in Table 4.2.7.1.1 are the minimum test levels recommended to detect workmanship defects.
SRDX4.2.7.1.1-1
The test duration shall be 1 minute per axes for acceptance level testing.
SRDX4.2.7.1.1-2
In no case shall the acceptance test levels for the sensor or its subsystems be less than those shown in Table 4.2.7.1.1.
Table 4.2.7.1.1 Random Vibration - Acceptance Test Levels
Frequency Acceleration Spectral Density (G2/Hz)
20 0.01
20 to 160 +3 dB/oct
160 to 250 0.08
250 to 2000 -3 dB/oct
2000 0.01
Overall 7.4 Grms
The plateau acceleration spectral density (ASD) level may be reduced for components between 25 kg and 200 kg according to the component weight (W) up to a maximum of 9 dB as follows:
dB reduction = 10 LOG(W/25)
ASD(plateau)level = 0.08 x (25/W)
where W = component weight in kg
SRDX4.2.7.1.1-3
The sloped portions of the spectrum shall be maintained at ± 3 dB/oct. Therefore, the lower and upper break points, or frequencies at the ends of the plateau become:
FL = 160 (25/W) FL = frequency break point low end of plateau
FH = 250 (W/25) FH = frequency break point high end of plateau
SRDX4.2.7.1.1-4
The test spectrum shall not go below 0.01 G2/Hz. For components whose weight is greater than 200 kg, the workmanship test spectrum is 0.01 G2/Hz from 20 to 2000 Hz with an overall level of 4.4 Grms.
Figure 4.2.7.1.1 Random Vibration - Acceptance Levels
The test levels found in Table 4.2.7.1.2 and Figure 4.2.7.1.2 are considered a conservative estimate of the random vibration environment on a representative spacecraft bus. The test levels shown in Table 4.2.7.1.2 are the minimum test levels recommended for protoqualification.
SRDX4.2.7.1.2-1
The protoqualification test duration shall be 2 minutes per axis.
SRDX4.2.7.1.2-2
In no case shall the acceptance test levels for the sensor or its components be less than those shown in Table 4.2.7.1.2.
Table 4.2.7.1.2 Random Vibration - Protoqualification Levels
Frequency Acceleration Spectral Density (G2/Hz)
20 0.026
20 to 50 +6 dB/oct
50 to 800 0.16
800 to 2000 -6 dB/oct
2000 0.026
Overall 14.1 Grms
The acceleration spectral density (ASD) level may be reduced for components more than 25 kg according to:
dB reduction = 10 LOG(W/25)
ASD(25 to 400) = 0.16 x (25/W)
where W = component mass in kg
SRDX4.2.7.1.2-3
The slope shall be maintained at ± 6 dB/Oct for components up to 65 kg.
SRDX4.2.7.1.2-4
Above 65 kg, the slopes shall be adjusted to maintain an ASD level of 0.01 G2/Hz at 20 and 200 Hz.
SRDX4.2.7.1.2-5
For components over 200 kg, the test specification shall be maintained at the level for 200 kg
Figure 4.2.7.1.2 Random Vibration - Protoqualification Levels
SRDX4.2.7.2-1
This test shall be conducted with the sensor suite in the launch configuration.
SRDX4.2.7.2-2
There shall be one sweep from 5 Hz to 50 Hz for each axis.
SRDX4.2.7.2.1-1
The sensor shall be acceptance tested to the sine vibration test levels specified in Table 4.2.7.2.2 and in Figure 4.2.7.2.2 in each of three orthogonal axes.
SRDX4.2.7.2.1-2
The acceptance test sweep rate shall be 4 oct/min except in the frequency range of 25-35 Hz, where the sweep rate shall be 1.5 oct/min.
SRDX4.2.7.2.2-1
For protoqual testing, the sine vibration levels shall be the same as the acceptance test levels specified in Table 4.2.7.2.2 in each of three orthogonal axes, however the sweep rates shall be reduced by a factor of two to 2 oct/min and 0.75 oct/min respectively.
Table 4.2.7.2.2 Sinusoidal Test Levels
Frequency Amplitude/Acceleration
5 to 18 Hz Displacement = 12 mm (double
amplitude)
18 to 50 Hz 8G peak
Figure 4.2.7.2.2 Sinusoidal Protoqualification Test Levels
SRDX4.2.7.2.3-1
The sensor contractor shall test the sensor structure in accordance with MIL-STD-1540C (TBR).
SRDX4.2.7.3-1
Sensor flight hardware shall be designed to withstand a maximum acceleration of 0.015g on orbit without permanent degradation of performance.
SRDX4.2.7.4-1
Sensors shall be designed and tested to survive, without permanent performance degradation, the environment shown in Figure 4.2.7.4 to account for externally induced shocks.
SRDX4.2.7.4-2
Shock testing is required at the sensor level if there are any self induced shocks (i.e., launch lock releases, pin pullers, etc.).
SRDX4.2.7.4.2-1
Protoqualification testing is accomplished by actuating the device two times for each self-induced shock source to account for the scatter associated with the actuation of the same device.
Frequency (Hz)
Figure 4.2.7.4 Shock Spectrum (Q=10)
SRDX4.2.7.5-1
Acoustic testing shall be performed for sensors with large surfaces (units with surface to mass ratio greater than 150 cm2/kg (50 in2/lb)), which could be excited by the acoustic field directly. For sensors less than 180 kg, vibration testing may be substituted for acoustic testing.
SRDX4.2.7.5-2
Acoustic testing shall be performed in accordance with MIL-STD-1540C (TBR).
SRDX4.2.7.5.1-1
The acceptance test acoustics levels shall be as defined in Table 4.2.7.5.1.
SRDX4.2.7.5.1-2
The acceptance test duration shall be one minute.
Table 4.2.7.5.1 Acceptance Acoustics Levels
One-Third Octave Noise Level (dB)
Center Frequency (Hz) ref: 0 dB = 20Pa
25 118
32 123
40 127
50 130
63 132
80 133
100 133.5
125 134
160 134
200 134
250 134.5
315 135.5
400 135.5
500 133
630 128.5
800 127
1000 124
1250 122
1600 120
2000 119
2500 118
3150 116.5
4000 115.5
5000 114.5
6300 114
8000 113
10000 112
Overall 145dB
4.2.7.5.2 Protoqualification Level Acoustic Testing
SRDX4.2.7.5.2-1
The protoqualification test levels shall be the levels shown in Table 4.2.7.5.1 increased by 3 dB.
SRDX4.2.7.5.2-2
The protoqualification test duration shall be two minutes.
SRDX.2.7.6-1
Thermal vacuum testing shall be performed in accordance with MIL-STD-1540C (TBR) using the environmental temperature ranges specified in SRD Paragraph 3.2.4.
SRDX4.2.7.6-2
Thermal cycle testing shall be performed in accordance with MIL-STD-1540C (TBR) using the environmental temperature ranges specified in SRD Paragraph 3.2.4.
SRDX4.2.7.6-3
Thermal balance testing shall be performed on the fist production unit in accordance with MIL-STD-1540C (TBR) using the environmental temperature ranges specified in SRD Paragraph 3.2.4.
SRDX4.2.7.6-4
An assessment on the need for Solar Simulation testing shall be accomplished and the result presented to the government for concurrence. .
SRDX4.2.8-1
Electromagnetic verification testing shall be conducted in accordance with MIL-STD-462D (TBR).
SRDX4.2.8-2
The standard interface shall undergo the following EMI/EMC tests:
1. Conducted susceptibility using CS101 (30 Hz to 100 kHz), CS114 (10 kHz to 400 MHz).
2. Radiated susceptibility using RS103 (20 kHz to 10 MHz).
3. Conducted emissions using CE101 (30 Hz to 20 kHz) and CE102 (20 kHz to 1 MHz).
4. Radiated emissions using RE102 (20 Hz to 10 Mhz).
SRDX4.2.8-3
The contractor shall perform electromagnetic testing to verify that the interface will operate properly if subjected to conducted or radiated emissions from maximum expected external sources, and to verify that the design of the interface does not result in deleterious conducted or radiated signals that might affect other mission elements. Sensors that fail to meet all MIL-STD-461D requirements may be suitable for flight if detailed analysis and system-level EMI/EMC testing shows no impact on mission operations.
SRDX4.2.8-4
The contractor shall perform conduction and radiation tests on the sensor, operating with expected power levels, current, and data rates.
SRDX4.2.8-5
All requirements shall be verified by test.
SRDX4.2.8-6
Sensor charging verification shall be conducted in accordance with MIL-STD-1541A except paragraph 6.5.2.4.1.
SRDX4.2.8-7
The contractor shall verify that his sensor does not impact the Search and Rescue mission during any mode.
SRDX4.2.9-1
Electrical current margins on all electro-explosive ordnance circuits shall be demonstrated by a test. The test will verify that no less than the minimum recommended firing current (twice all-fire) will be delivered to the electro-explosive devices under worst conditions of minimum voltage and maximum circuit and electro-explosive device resistance.
SRDX4.2.9-2
The test shall verify that the maximum current delivered to the electro-explosive device does not exceed its maximum qualified firing current under worst conditions of maximum voltage and minimum circuit and electro-explosive device resistance.
SRDX4.2.10-1
The deployment and latching devices shall be tested to demonstrate adequate functioning under worst case environments in accordance with MIL-A-83577B, Assemblies, Moving Mechanical for Space/Launch Vehicles (TBR).
Outgassing evaluation tests shall be conducted for materials, components, and sub systems whose properties are not known.
SRDX4.2.12-1
Items that incorporate extensive changes in design, manufacturing processing, environmental levels, or performance requirements shall be requalified. However, methodology from MIL-HDBK-340 may be used to show that existing designs or items (previously qualified for other applications) have adequately demonstrated compliance to all qualification requirements for the new designs. Deficiencies in meeting some requirements may be fulfilled by supplementing existing data with new test data.
SRDX4.2.12-2
Qualification by similarity shall be permitted only with the concurrence of the contracting officer.
SRDX4.2.12-3
Waiver of qualification or requalification requirements shall be approved by the contracting officer.
SRDX4.2.13-1
Life time testing shall be conducted in accordance with MIL-STD-1540C (TBR).
SRDX4.2.13-2
Explosive ordnance devices and other components whose performance may degrade with time shall have life extensions based upon passing either an aging surveillance test or an accelerated aging test.
To the extent practicable, pre-launch sensor-level inspections and tests will be conducted to verify by end-to-end testing that each critical path in the launch system, spacecraft, and sensor is satisfactory.
Whether electrical, mechanical, or both, all critical paths or circuits will be verified from the application of the initiating signal through completion of each event. This testing verifies:
Once successfully accomplished, that particular critical path or circuit is considered validated. Not all end-to-end tests can be performed with only flight hardware, as in the case in which an explosive event is involved. In cases in which end-to-end testing cannot be performed with the flight hardware and software, appropriate simulation devices will exercise the flight hardware and software to the maximum extent possible. Simulation devices will be controlled carefully and will be permitted only when there is no feasible alternative for conducting the test. All of the events that occur during the mission profile should be tested in the flight sequence to the extent that is practical. Redundant components and subsystems also should be validated in the same manner.
Pre-launch validation tests shall be conducted on sensor equipment in accordance with the applicable requirements of MIL-STD-1540C (TBR).
The government will use standard datasets and sensor performance models to test the performance of contractor sensor designs and algorithms, as well as supply standard datasets to contractors for their use in developing designs which meet the TRD requirements. For all sensors, the government will develop a set of up to 10 (TBR) standard datasets in each of the climate categories relevant to the sensor and its allocated EDRs. Up to half of these will be available to the contractors for design development. The remainder will be retained by the government to evaluate designs which are delivered at PDR. The sensor-specific details of the types and characteristics of the standard datasets are provided in the SRD (see paragraph 3.2.1).
SRDX4.3.1-1
The contractor shall provide to the government sufficient design-specific information and science algorithms to allow the government to accurately model the performance of the design, including (TBR) parameters for each sensor.
SRDX4.3.2-1
Methods of verification shall be selected from the following:
a. Inspection. An observation or examination of the item against the applicable documentation to conform compliance with requirements.
b. Analysis. A process used in place of or in addition to testing to verify compliance with specifications. The techniques typically include an interpretation or interpolation or extrapolation of analytical or empirical data under defined conditions or reasoning to show theoretical compliance with stated requirements.
c. Demonstration. An exhibition of the operability or supportability of an item under intended service-use conditions. These verifications are usually non-repetitive and are oriented almost exclusively toward acquisition of qualitative data. Demonstrations may be accomplished by computer simulation.
d. Test. An action that verifies an item's operability, supportability, performance capability or other specified qualities when subjected to controlled conditions that are real or simulated. These verifications may require use of special test equipment and sensor to obtain quantitative data for analysis as well as qualitative data derived from displays and indicators inherent in the item(s) for monitor and control.
e. Similarity. Similarity is the process of comparing a current
item with a previous item, taking into consideration configuration,
test data, application and/or environment. The evaluation shall
be documented and shall include: the test procedures/reports of
the item to which similarity is claimed; a description of the
difference(s) between the items; and the rationale for verification
by similarity.
All in orbit experience shall be documented and available for
review.
f. Not Applicable. Use of the term "Not Applicable" shall be limited to those paragraph/paragraph headings for which there is no method of verification or where verification is accomplished in subparagraphs.
SRDX4.3.3-1
EDR requirements shall be validated by the contractor by analysis, modeling, and/or simulation based on the instrument design and performance characteristics and the contractor's scientific algorithms. The standard scenes provided by the government may be used for this purpose, but need not be relied upon exclusively by the contractor.
SRDX4.3.3-2
All relevant sources of error, including those associated with the scene radiance, instrument, spacecraft, data transmission, and algorithms, shall be taken into account. The spacecraft contractor is to provide errors associated with spacecraft and data transmission.
SRDX4.3.3-3
The contractor's analysis, modeling, and/or simulation shall be sufficiently extensive in scope to verify that EDR requirements are met under a broad range of conditions (TBR) that are representative of those occurring in nature, including both typical and extreme conditions. EDR requirements should be validated for any value of the geophysical parameter within the specified measurement range, any latitude, any time of day (subject to specified daytime or nighttime only restrictions), any season, any climate, any level of cloudiness (subject to specified clear/cloudy restrictions), any sensor viewing geometry, and any degree of horizontal or vertical spatial structure in the observed scene consistent with nature. The contractor is not relieved from this requirement by any limitations in the standard scenes provided by the government (See Section 3.2.1.34.) This requirement may be satisfied by a judicious choice of typical and stressing test cases for analysis, modeling, and/or simulation, and should not be construed to entail exhaustive validation throughout the space of parameters and conditions mentioned in this paragraph.
SRDX4.3.3-4
For simulations involving random variate generation, a sufficient number of iterations shall be performed for each test case or standard scene to ensure that statistical errors are negligible compared to the EDR attribute value being validated.
SRDX4.3.4-1
The contractor shall include in his requirements flowdown analysis uncertainties in data from any data bases that are relied upon in generating EDRs.
SRDX4.3.4-2
The contractor shall notify the government
The government will determine the appropriate remedial action.
SRDX4.3.4-3
The contractor shall generate a new data base or partial data base if a fixed data base, (e.g., one addressing terrain) is needed, and existing data bases are not adequate to allow meeting EDR thresholds. The IDPS (TSPR) contractor will be responsible for developing any new fixed data bases needed by the operational algorithms.
SRDX4.3.4-4
The contractor shall identify and quantify any EDR performance degradation resulting from the lack of availability of any data base or other ancillary data.
SRDX4.3.5-1
The sensor shall have the capability of being externally tested, while in storage and on the launch pad, to verity its performance and operational readiness.
SRDX4.3.5-2
Each sensor shall have the capability to perform self-testing using built-in test (BIT) functions to determine its functionality, performance, and operational readiness, both on the ground and on-orbit.
TBD
SRDX5.1-1
The contractor shall pack and handle deliverable items so as to protect them against vibrations, shocks, moisture, and contamination associated with ground or air transport.
SRDX5.1-2
The contractor shall provide protection against natural environmental conditions using containers, shrouds, or covers.
SRDX5.1-3
The contractor shall provide access provisions for inspection and handling for all deliverable items.
SRDX5.1-4
All deliverable items shall include positive means to verify compliance with shock, temperature, and moisture requirements.
The contractor should mark deliverable units in accordance with MIL-STD-129K as appropriate for the item being prepared.
SRDX5.2-1
Nameplates for hardware shall contain the item or configuration item number, serial number, lot number (or contract number), manufacturer, and nomenclature.
SRDX5.2-2
Software media shall be marked to display software configuration item number, serial number, contract number, manufacturer, and nomenclature.