System_Requirements_Document.docx

National Aeronautical and Space Administration

George C. Marshall Space Flight Center Marshall Space Flight Center, Alabama 35812

Systems Requirements Document for the NEA Scout Reaction Control System

August 22, 2016

Prepared By:

Andrew Heaton Date EV42/NEA Scout Guidance & Control (G&C) Subsystem Lead

Approved By:

Leslie McNutt Date FP30/NEA Scout Project Manager

Concurrence By:

Donald McQueen Date ER33/NEA Scout Chief Engineer

William B. Stein Date EV42/NEA Scout G&C RCS SME

Brandon Stiltner Date EV42/NEA Scout G&C

Dan Cavender Date ER23/NEA Scout RCS

Heather Koehler Date EV42/ Guidance, Navigation and Mission Analysis Branch Chief

Table of Contents

SECTION 1.0 SCOPE

This specification establishes the design, development, performance and verification requirements for the Near Earth Asteroid (NEA) Scout Reaction Control System (RCS). The RCS will be integrated to the NEA Scout spacecraft after acceptance of the RCS. Figure 1 shows the preliminary spacecraft configuration. Prior to launch, the RCS propellant will be loaded and the RCS will be subjected to an integrated spacecraft vibe/shock and thermal vacuum test. After integrated testing, the spacecraft will be mounted into the deployer mechanism and there will be no access to the RCS.

Figure 1: Preliminary Spacecraft Configuration The RCS will be used to provide the following guidance and control functions:

De-tumble the spacecraft from an initial spin.

Provide a propulsive maneuver.

De-saturate the NEA Scout reaction wheels before and after deployment of the NEA solar sail and throughout the duration of the mission.

1.1 CONVENTION AND NOTATION

The hardware shall meet specific technical requirements below. The convention used in this document, which indicates requirements, goals, and statements of facts, is as follows:

Every requirement containing a "shall" is binding and shall be verified. Goals and statements of fact are non-binding. Goals are included to guide trade studies and will be addressed at design reviews and technical interchange meetings (TIMs).

SECTION 2.0 DOCUMENTS

The following documents are applicable to the extent specified herein. In the event of a conflict between the referenced document and the contents of this specification, the contents of this specification will take precedence. If no revision or date is specified for the document, the latest issue at the time of contract award shall apply. NASA reserves the right to review and approve use of all equivalent specifications and standards.

2.1 APPLICABLE DOCUMENTS

SECTION 3.0. TECHNICAL REQUIREMENTS

3.1 ITEM DEFINITION

The following definitions provide detailed descriptions about the various systems and elements for the NEA Scout. Definitions are factual statements and do not require verification.

3.1.1 NEAR EARTH ASTEROID SCOUT

Near Earth Asteroid Scout is a Class D CubeSat mission per NPR 7120.5 (p. 26) that will launch as a secondary payload on the Space Launch System Exploration Mission 1 scheduled for launch in 2018. Its mission is to survey a near Earth asteroid. The NEA Scout spacecraft utilizes a Planetary Sciences Corporation 6U Canisterized Satellite Dispenser (CSD) for deployment from the SLS vehicle. The mass of the NEA Scout will be less than 14kg.

3.1.2 NEA REACTION CONTROL SYSTEM

The NEA Scout RCS uses thrusters to execute a propulsive maneuver and attitude control for the spacecraft. The RCS unit will fit within a volume envelope of approximately 2U and the remainder of the spacecraft will fit within the remaining volume envelope of approximately 4U. A notional schematic of the RCS is shown in Figure 2Figure 2.

Figure 2: Notional RCS Schematic

3.1.3 NEA RCS MISSION CONCEPT OF OPERATIONS

Below is a general outline of the concept of operations (ConOps) for the NEA Scout mission:

1. Deployment – Spacecraft powerup. G&C is quiescent until software is initialized.

2. G&C Initialization – Default G&C parameters loaded from onboard memory. Attitude control begins in default mode.

3. De-tumble – Spacecraft initial rates are damped to normal operating levels. Control will use reaction wheels and RCS to perform damping operations.

4. Solar Pointing – After rates are damped, ACS will acquire the sun to ensure that spacecraft power systems can charge. Reaction wheels will perform any slews and maintain solar pointing. RCS will be used to perform momentum dumping if wheels become saturated.

5. Communication Pointing – Spacecraft is reoriented to allow for communication with ground (inertial pointing). Primarily performed with RWs. RCS used for momentum dumping if necessary.

6. Initial Calibration – Based on data from detumble and other early ACS operations, gains and other parameters may be adjusted to ensure proper functionality and control response.

7. Perform Trajectory Correction Maneuver (TCM) 1 – RCS will perform correction maneuver sequence. Reaction wheels, attitude control thrusters, and differential axial thruster firings will be used to control and maintain desired burn attitude profile. Maneuver will likely be broken into several burns due to power constraints. In between burns, attitude will be changed to provide proper battery charging. Momentum dumping performed with RCS if necessary.

8. Translunar Coast – Desired attitude profile maintained during cruise periods to ensure proper power state of charge and for communication opportunities. Momentum dumping with RCS if necessary.

9. G&C Calibration – Perform tests to calibrate G&C sensor and actuator performance. Ground will upload any parameter changes required.

10. Resume cruise – Continue translunar coast.

11. Perform Further TCMs – Repeat TCM profile if further corrections are required.

12. Lunar Flyby 1 – Spacecraft maintains desired attitude for flyby.

13. Sail Deployment – Attitude maintained during sail deployment sequence. Once calibrated, AMT will assume role to perform momentum management or RWs. RCS will now be used only as a backup.

14. Sail Calibration – Sail thrust characterization and calibration will be performed to adjust/determine thrust model. AMT movements will be performed as needed to compensate for CM/CP offsets.

15. Earth/Lunar Cruise – Spacecraft pointing will be dictated by solar sail thrust profile to shape trajectory for second lunar flyby. Attitude slews will also be performed as needed for power system charging and ground comm passes.

16. Lunar Flyby 2 – Spacecraft attitude/sail thrusting will be maintained for desired flyby conditions.

17. Heliocentric Cruise – Pointing will be maintained to achieve desired solar sail thrust except for power, safing, or communication with ground events.

18. Science Observations – Attitude slews will be performed to acquire and track NEA.

19. NEA Flyby – Desired observations slews will be performed for NEA observations during flyby. Attitude profile table will be provided by ground for flyby.

20. Extended Mission – Perform extended mission operations (optional).

21. Disposal – RCS propellant will be depleted. Spacecraft G&C will cease.

Note: Prior to launch, the NEAS will be containerized and loaded on SLS for up to a year prior to activation with a contingency of up to three years stored in a fueled state. There will be no contact (mechanical/electrical/com) with the system once integrated on SLS prior to ejection.

3.1.4 COORDINATE SYSTEM

The coordinate system is described in Figure 3. The origin of the reference frame is defined in Planetary Systems Payload Spec 2002367C.

Figure 3: Spacecraft Body Coordinate Frame with RCS unit outlined in orange.

3.1.5 NEA SCOUT RCS INTERFACE DEFINITION

3.1.5.1 Mechanical Interface with NEA Scout Spacecraft

Figure 4Figure 4 shows a notional design concept for the RCS unit’s mechanical interface with the NEA Scout spacecraft which depicts the negative Z-face of the RCS unit. The following requirements apply to this interface:

3.1.5.1.1 Circular Cutout

The interface shall be a circular cutout on the –Z-axis face of the RCS unit, 3.83mm in depth (Z-axis). This cutout defines the XY plane for all thruster orientations

3.1.5.1.2 Bolts

Seven bolts shall secure the RCS unit to the NEA Scout spacecraft. Details will be provided for the bolt pattern and critical dimensions for the mechanical interface.

3.1.5.2 Mechanical Interface with Planetary Systems 6U CSD

The RCS unit shall interface with the Canisterized Satellite Dispenser per Planetary Systems document 2002367C. (Document 2002367C specifies AL 6061 or AL 7075).

The RCS volume envelope is provided in Section 3.2.2. In particular, the mechanical interface between the CSD and RCS unit are visualized in Figure 4 as flanges on the positive and negative X-faces of the unit. Detailed dimensions of the flanges will be provided.

Figure 4: Notional RCS design to illustrate the mechanical interfaces.

3.1.5.3 Mechanical Interface with Solar Arrays

3.1.5.3.1 Mechanical Interface Conformance

The RCS unit shall conform to the mechanical interface with the NEA Scout solar arrays. (There will be four solar arrays on the NEA Scout spacecraft which will notionally attach to the RCS unit as shown in Figure 3Figure 3. A detailed interface drawing will be provided separately.)

3.1.5.3.2 Mechanical Interface Threaded Inserts

The RCS unit shall provide an interface to the solar arrays in the form of threaded inserts, as are notionally shown in Figure 5Figure 5. The interface details will be provided in a separate interface drawing.

Figure 5: Notional RCS unit depicting the notional locations of the solar pannel attachment points.

3.1.5.4 Mechanical Interface for Loading into the CSD

3.1.5.4.1 Mechanical Interface Handles

The RCS unit shall provide a mechanical interface for attaching two handles to be used for loading the NEA Scout into the CSD. These handles will be temporarily attached to the positive Z-face of the RCS unit. After the NEA Scout has been loaded into the CSD, the handles will be removed, and therefore are not flight hardware.

3.1.5.4.2 Handles

The RCS unit shall provide an interface to the handles in the form of threaded inserts as notionally depicted in Figure 6Figure 6. The interface details will be provided in a separate interface drawing.

3.1.5.5 Mechanical Interface for NEA Scout Sun Sensor

3.1.5.5.1 Sun Sensor Interface

The RCS unit shall provide mechanical interfaces to accommodate mounting a sun sensor near the center of the unit’s positive Z-face. A notional design for the mounting points is depicted in Figure 6Figure 6, and the sun sensor is depicted in Figure 7Figure 7.

3.1.5.5.2 Sun Sensor Interface Threaded Inserts

The RCS unit shall provide an interface with sun sensor in the form of threaded inserts. The interface details will be provided in a separate drawing.

3.1.5.6 Mechanical Interface for NEA Scout Transmit/Recieve (Tx/Rx) Units

The RCS shall provide mechanical interfaces to accommodate mounting the Tx/Rx units. A notional design for the mounting points are shown in Figure 6Figure 6, and the Tx/Rx units are shown in Figure 7Figure 7. The RCS unit shall provide an interface with the Tx/Rx units in the form of raised mounting studs (standoffs) with threaded inserts. The r details will be provided.

Figure 6: Notional RCS unit illustrating potential locations for mounting the CSD installaiton handles, Tx/Rx units, and sun sensor.

3.1.5.7 Electrical Interfaces

The interface details will be provided in a separate drawing.

3.2 PROPULSION SYSTEM REQUIREMENTS

3.2.1 MASSES

3.2.1.1 Total Mass

The total wet mass of the RCS shall be no more than 2.50 kg.

3.2.1.2 Reserved

3.2.1.3 Propellant Mass

The propellant mass of the RCS shall be greater than 1.25 kg

3.2.1.4 Usable Propellant

The usable propellant of the RCS shall be greater than 99%.

3.2.2 VOLUMES

3.2.2.1 Volume Envelope

The RCS shall fit within an envelope of 226.4mm (X-axis) x 105mm (Y-axis) x 66.9mm (Z-axis). A graphical depiction of the notional volume can be found in Figure 7Figure 7.

Figure 7: Notional RCS concpet depicting the OML dimensions of the unit.

3.2.2.2 Wiring Harness Accomodation

The RCS volume shall accommodate the passage of a wiring harness oriented in the z-axis direction. A conceptual image of the wiring harness passage is provided in Figure 4Figure 4.

3.2.3 SYSTEM ELECTRONICS

3.2.3.1.1 Data Interface

The RCS data interface to the NEAS spacecraft shall be RS-422.

3.2.3.1.2 Data Rate

The RS-422 data interface shall function at 115,200 bps

3.2.3.1.3 Command Latency

Commands to the RCS shall be processed within 100 ms

3.2.3.1.4 Data Packages

The RCS shall be capable of sending customized telemetry which at a minimum consist of a “large” and “small” size package.

3.2.3.2 Input Voltage

The propulsion system heaters and valves shall meet all the performance requirements in this specification using input voltages between 9.0 V and 12.6 V.

3.2.3.3 Power

3.2.3.3.1 Idle

RCS power draw shall be no more than 1.1 W when in Standby mode (powered on but not heating propellant or firing thrusters).

3.2.3.3.2 Propellant Heating Cycle

RCS power draw shall be no more than 55 W when heating propellant but not firing.

3.2.3.3.3 Propellant Heating Rate

RCS propellant heater power draw shall be adjustable in flight between 13 W and 52W with a 1W minimum resolution.

3.2.3.3.4 Thruster Operation

RCS power draw shall be capable of operating fully while drawing no more than 9W (The system can function with the propellant heaters turned off).

3.2.3.4 System Sensors

3.2.3.4.1 RCS Temperature Sensors

The RCS shall provide a means of determining temperature in the RCS unit which includes temperature measurements within each volumesupplied.

3.2.3.4.2 RCS Pressure Sensors

The RCS shall provide a means of determining pressure in the RCS unit which includes pressure measurements within each volume supplied.

3.2.3.4.3 RCS Sensor Voltages

All sensors shall operate using the input voltages specified in 3.2.3.2.

3.2.4 VALVES

3.2.4.1 Valve Shuttle Time, Opening

Each RCS valve shall actuate with an opening response time of less than 10 ms.

3.2.4.2 Valve Shuttle Time, Closing

Each RCS valve shall actuate with a closing response time of less than 10 ms.

3.2.4.3 Cycle Life

Each RCS valve shall be capable of 250,000 cycles.

3.2.4.3 Delivered Cycle Life

95% of design cycle life shall be available at the time of delivery for the flight unit.

3.2.4.4 Force Margin

The force margin on each valve shall be as specified in NASA-STD-5017.

3.2.4.5 Leakage

Each RCS valve shall have a maximum valve leakage rate of 3.0 scch of GHe during exposure to the temperature extremes specified in paragraph 3.3.3.1.

3.2.5 SYSTEM PRESSURES

3.2.5.1 Maximum Design Pressure

The RCS shall have all pressurized hardware designed, tested, and analyzed for a MDP of 100 psia.

3.2.5.2 Proof Pressure

The RCS shall have all pressurized hardware designed, tested, and analyzed for a proof pressure of at least 1.5 times MDP as per SLS-RQMT-216.

3.2.5.3 Safety Factors

The RCS shall have all pressurized hardware designed and analyzed to comply with SLS-RQMT-216 for Ultimate and Yield. A summary of ultimate safety factors provided in SLS-RQMT-216, BASELINE are summarized in Table 1:

Table 1: Factors of Safety for Design of Pressure Systems

3.2.6 PROPELLANT SELECTION

3.2.6.1 Propellant Type

The propellant shall be R236fa or equivalent.

3.2.7 SYSTEM PERFORMANCE

3.2.7.1 Axial Thrusters

3.2.7.1.1 Number of Thrusters

There shall be two thrusters to provide –Z axis translation with the effective thrust in the –Z direction.

3.2.7.1.2 Location of Axial Thrusters

The –Z translational thrusters shall be located at the following position in the coordinate system specified in 3.1 [x y z]. Figure 8Figure 8 depicts the locations of the two axial (-Z) thrusters and are labeled using the numbering in the list below.

1. [59.17, 52.5, 366] mm ±0.2 mm Axial thruster located to the right of center on the NEA Scout forward face.

[-59.17, 52.5, 366] mm ±0.2 mm Axial thruster located to the left of center on the NEA Scout forward face.

3.2.7.1.3 Orientation of Axial Thrusters

The axial thrusters shall be oriented so that the thrust acts along the –Z direction, and the exhaust gases are expelled in the +Z direction. The list below provides the orientation for the axial thrusters using the numbering defined in Figure 8Figure 8.

1. Thruster shall be aligned with the +Z-axis ±0.5 deg Axial thruster located to the right of center on the NEA Scout forward face.

2. Thruster shall be aligned with the +Z-axis ±0.5 deg Axial thruster located to the left of center on the NEA Scout forward face.

Figure 8: Notional design of the RCS system depicting the locations of the Axial thrusters (2) and RCS thrusters (4).

3.2.7.2 RCS Thrusters

3.2.7.2.1 Number of Thrusters

There shall be four thrusters to perform attitude control.

3.2.7.2.2 Location of Thrusters

The RCS thrusters shall be located at the following position in the coordinate system specified in 3.1 [x y z]. Figure 8Figure 8 depicts the locations of the four RCS thrusters and are labeled using the numbering in the list below.

1. [101.92, 88.12, 366] mm ±0.2 mm RCS thruster located in the top-right on the NEA Scout forward face.

[101.92, 16.88, 366] mm ±0.2 mm RCS thruster located in the bottom-right on the NEA Scout forward face.

[-101.92, 16.88, 366] mm ±0.2 mm RCS thruster located in the bottom-left on the NEA Scout forward face.

[-101.92, 88.12, 366] mm ±0.2 mm RCS thruster located in the top-left on the NEA Scout forward face.

3.2.7.2.3 Orientation of Thrusters

The thruster orientation shall be as depicted in Figure 9Figure 9.

Note on 3.2.7.2.3: The orientations of the four RCS thrusters are described using two angles: elevation and azimuth. First, the elevation angle is the positive angle measured from the RCS unit’s +Z face (X-Y plane) to the jet nozzle. For example, if the elevation angle is 90 degrees, the RCS jet exhaust is directed in the spacecraft positive Z-axis, while the thrust is in the spacecraft negative Z-axis. Second, the azimuth angle is the angle measured counter-clockwise (positive) from the NEA Scout X-axis to the jet nozzle.

Figure 9: Sketch to depict the RCS Jet angles as described in Section 3.2.8.2.

These angles should not to be confused with Euler rotations. These angles simply describe the directions the nozzles are pointed (exhaust direction), and not the direction of the thrust. In all cases, the jet exhaust is directed along the spacecraft’s positive Z-axis, and toward the spacecraft’s geometric center-line.

1. [30, 225] deg ±0.5 deg RCS thruster located in the top-right on the NEA Scout forward face.

[30, 135] deg ±0.5 deg RCS thruster located in the bottom-right on the NEA Scout forward face.

[30, 45] deg ± 0.5 deg RCS thruster located in the bottom-left on the NEA Scout forward face.

[30, 315] deg ±0.5 deg RCS thruster located in the top-left on the NEA Scout forward face.

3.2.7.3 Thrust Direction Tolerance

All thrusters shall have a thrust vector alignment within ±0.5 degrees relative to the nominal thruster direction under static conditions.

3.2.7.4 Thruster Operation

All thrusters shall be capable of a minimum of 250,000 firings.

3.2.7.5 Total System Impulse

The RCS unit shall be capable of providing a minimum of 500 N-s of total impulse.

3.2.7.6 Thuster Leakage Rate

The RCS thruster shall have a leakage rate no more than 3.0 scch of GHe.

3.2.7.7 Thrust

Each thruster shall be capable of providing a minimum of 25 mN +/- 1 mN of thrust.

3.2.7.8 Specific Impulse

Each thruster shall be capable of providing a specific impulse of at least 40 seconds at a minimum.

3.2.7.9 Start Response

The time to 90% of the steady-state thrust from the issuance of the thruster assembly ON command shall not exceed 15 ms.

3.2.7.10 Shutdown Response

The time to 10% of the steady-state thrust from the issuance of the thruster assembly OFF command shall not exceed 15 ms.

3.2.7.11 Minimum Impulse Bit

The minimum impulse bit for the RCS shall be a maximum of 0.5 mN-s.

3.2.7.12 Duration of Operation

The RCS shall be able to thrust continuously for at least 10 minutes at a minimum thrust level of 50 mN with the effective thrust in the direction of the –Z axis.

3.2.7.13 Minimum Usable Pressure

The RCS thrusters shall operate with a tank pressure greater than 5 psia.

3.2.7.14 Thruster Availablility

The RCS shall be able to proceed from an idle state to a thrusting state in 10 minutes.

3.2.7.15 Thrust Variablility

The thrust variation between the axial thrusters shall not exceed 5% .The thrust variation between the RCS thrusters shall not exceed 5%.

3.2.7.16 Independent Thruster Operation

Each thruster shall be capable of operating individually.

3.2.7.17 System External Leakage

The RCS shall have an external leakage rate no more than 0.0036 scch of GHe, not including valves.

3.2.7.18 Fault Tolerance To Leakage

The propulsion system shall maintain single-fault tolerance to propellant leakage prior to spacecraft deployment.

3.2.7.19 Fill Valve External Leakage

The fill valve shall have an external leakage rate no more than 1e-6 sccs of GHe.

3.2.8 OPERATIONAL LIFETIME

The post-launch operational life for the RCS shall be a minimum of 2.5 years excluding propellant depletion.

3.2.9 RADIATION TOLERANCE

3.2.9.1 Dose Depth

The propulsion system shall meet all safety, interface and performance requirements defined in this document during and after exposure to the radiation dose depth curve in Figure 10 for a 2.5-year (TBR) mission that matches the appropriate level of equivalent shielding of the electronics.

Figure 10: NEA Scout Dose Depth

3.2.9.2 Latch-up

All electronic parts used in the propulsion system shall utilize technologies or designs immune to latch-up below a linear energy transfer (LET) of 37 MeV-cm2/mg.

3.2.10 MATERIAL SELECTION

Materials shall be in accordance with NASA-STD-6016.

3.2.10.1 MATERIALS IDENTIFICATION AND USAGE LIST

The contractor shall provide a Materials Identification and Usage List with content as specified in NASA-STD-6016.

3.2.12 EEE Parts

The propulsion system shall meet the requirements of MSFC-STD-3012 Grade 4 at a minimum.

3.2.13 SLS Safety Requirements

The propulsion system shall meet the SLS safety requirements as defined in SLS-RQMT-216.

3.2.14 CLEANLINESS

3.2.14.1 System Cleanliness

Payloads shall comply with the GSDO-RQMT-1080, Cross-Program Contamination Control Requirements document for visibly clean standard level. (GDDO-RQMT-1080 is specified by IDRD SLS SPIE RQMT 018 Section 3.2.5).

3.2.14.2 Cleanliness Protection for Fluid Systems

A filter having an absolute rating of 10 microns shall be installed downstream of the propellant tank and upstream of any valves or thrusters.

3.3 ENIVRONMENTS

3.3.1 VIBRATION

The propulsion system shall meet requirements after exposure to vibration environments as defined below.

3.3.1.1 Random Vibration

The propulsion system shall meet requirements after exposure to the random vibration load factors as defined in Figure 11 and Table 2. The acceptance test levels in Table 2 shall be used as flight environments for analysis purposes.

Table 2 Secondary Payload Component Loads Due to Random Vibration Configuration 1a – 18.6 kg (41 lbm) Payload

Figure 11 Generalized Random Vibration Test Levels

3.3.1.2 Low Frequency Vibration

The propulsion system shall meet requirements during and after exposure to the low frequency load factors as defined in Table 3.

Table 3 Secondary Payload Component Loads Due to Low Frequency Loads

3.3.2 SHOCK

The propulsion system shall meet all safety and interface requirements defined in this document during and after exposure to the low frequency and random vibration load factors as defined in Table 3.

Figure 2: Dispenser Local Coordinate System

3.3.2.1 Pyrotechnic Shock

The propulsion system shall meet the performance requirements of this specification after exposure to the shock environments in Table 4.

Figure 3: Pyrotechnic Shock Spectrum Table 4: Test Shock Environment due to Spacecraft Separation

Note 1: Use Figure 9 with this table.

Note 2: The shock environment presented should be treated as “test to” environment and not “analyze to” environment

3.3.2.2 Flight Shock

The propulsion system shall meet requirements after exposure to flight shock environments as defined in Table 4.

3.3.2.3 Ground Shock

The propulsion system shall meet requirements after exposure to ground shock environments as defined in Table 5.

Table 5 Procedure for Shock Environments

3.3.3 THERMAL

3.3.3.1 Thermal Safety

The propulsion system shall meet all safety and interface requirements defined in this document during exposure to a temperature range of -24 deg C to 55 deg C. This thermal range encompasses the environment for launch processing and ascent. During this phase, the propulsion system will be unpowered.

3.3.3.2 Thermal Performance

The propulsion system shall meet performance requirements while exposed to a temperature range of -10 deg C to 55 deg C.

3.4 STORAGE

3.4.1 Storage Lifetime, Fueled

The RCS system shall meet the safety requirements as defined in SLS-RQMT-216 during and after storage in the unpowered configuration for a minimum of 36 months in the fueled configuration.

3.4.1.1 Leakage During Storage

The RCS thrusters shall have a maximum valve leakage rate of 3.0 scch of GHe after acceptance testing and during fueled storage.

3.4.2 Storage Lifetime, Unfueled

The RCS shall meet all requirements after a minimum of at least 2 years in storage while in an unfueled state.

3.4.3 Storage Environment

The valve shall meet the operating performance requirements specified herein after exposure to the following storage conditions.

2) Temperature From minus 23ºF to plus 120ºF

3) Humidity 8 to 100 percent relative humidity

3.5 QUALITY AND INSPECTION

3.5.1.1 SAE AS9100 Compliance

The hardware shall be compliant with SAE AS9100.

3.5.1.2 Certificate of Compliance

The supplier shall submit a certificate of Compliance (COC), typed and signed by responsible representative of the company, stating that all materials, processes, and/or finished components supplied under this purchase order meet the applicable specifications.

3.5.1.3 Part Numbers

The COC shall contain all applicable part/specifications numbers.

3.5.1.4 Packaging

The supplier shall package ESDS items to prevent damage during shipping and handling and the items must be identified as ESD sensitive with the proper marking on the shipping container as required in ANSI/ESD 20.20. Item 1 MSFC QA: MSFC QA will verify COC, ADP, receipt of quantity ordered, and receipt and compliance with the above requirements. Inspect for and document any shipping damage.

SECTION 4.0 VERIFCATION

4.1 QUALITY CONFORMANCE

The verification method and the relationship to Section 3.0 requirements shall be as defined in the verification matrix. Each performance and design requirement specified in Section 3.0 of this specification shall be verified by inspection, analysis, and/or test.

4.1.1 METHODS OF VERIFICATION

4.1.1.1 Inspection

Inspections are physical examinations used to verify that the unit is in compliance with the design, manufacturing, and procurement requirements as defined in the drawings and specifications. Inspections of process paperwork will be used to ensure that all quality assurance provisions have been implemented in the test program procedures. Inspections may include visual or visual with aid of NDE processes using standard quality control methods. All test equipment used in the test program will be inspected to ensure compliance with the test program plans and procedures. Requirements that will be verified by inspection may be performed during procurement, fabrication, or assembly and test of the unit.

4.1.1.2 Analysis

Verification by analysis involves the use of engineering analysis, qualitative assessment, computer modeling and/or simulation to ensure compliance to requirement(s). If necessary, analysis combined with test results may be used to provide this evidence. Analyses shall be traceable to the requirements of this specification and the specific revision level of the design drawings and related documentation.

4.1.1.2.1 Analysis Methods

Analysis methods and procedures shall be included in the NASA review and acceptance of analysis used to satisfy requirements.

4.1.1.2.2 Analysis Reporting

Sufficient information on methods of analysis used to satisfy requirements shall be provided to NASA.

4.1.1.3 Test

Verification by test (e.g., functional, environmental) is the actual operation to ensure that the performance characteristics of the unit are in accordance with the requirement(s). Testing will also include methods to verify requirements such as material compatibility, which may or may not be performed with an end item. The analysis of data derived from tests is an integral part of the test program and should not be confused with analysis defined in paragraph 4.1.1.2. Tests shall be used to determine quantitative compliance to requirements and produce quantitative results.

4.1.2 VERIFICATION MATRIX

The verification matrix provides the minimum test, analysis, and inspection activities needed to verify that the design meets the requirements defined in Section 3.0. This table, combined with the values in the requirements in Section 3.0, forms the basis of a verification plan that shall be submitted by the seller for approval by NASA. Columns annotated with "I" indicate that inspection is required. Columns annotated with "T" indicate that testing is required. Columns annotated with "A" indicate that analysis is required. Columns annotated with "X" indicate the paragraph is a header, for information only, is a target guideline, or is a statement of fact signified by “will”.

Table 6 Verification Matrix (TBR)

4.1.3 TEST CONDITIONS

4.1.3.1 Standard Test Conditions

If not called out otherwise, test conditions shall follow SMC-S-016 at the discretion of NASA.

4.1.3.2 Test Tolerances

If not called out otherwise, test tolerances shall follow SMC-S-016 at the discretion of NASA.

4.1.4 Test Responsibility

Propulsion subsystem qualification and acceptance testing and all other verification activities shall be completed prior to acceptance of hardware and the delivery of the hardware for integration into the NEA Scout.

4.1.5 Verification Phasing

Verification shall occur at each phase of the development and delivery cycle, including qualification and acceptance as described in Table 7.

4.1.6 Development

This phase of the verification coincides with the design development of the item prior to the Critical Design Review (CDR). This verification will establish feasibility of the proposed concepts to satisfy the design prior to authorization to make the flight qualification unit and begin qualification testing. The CDR package will constitute a large portion of the verification data for this phase.

4.1.7 Qualification

This phase of verification will coincide with the fabrication and testing of the qualification unit, which will verify suitability of the post CDR baselined design with the requirements of this specification under the comprehensive environments and loads defined for the preflight ground and flight operations. The purpose here is to verify that the design meets the requirements of the specification. Prior to the commencement of the qualification testing the qualification unit will be subjected to a full acceptance verification set. The qualification unit acceptance and the qualification test reports will constitute a large portion of this verification package.

4.1.7.1 Qualification Testing

Qualification testing of the propulsion subsystem shall include inspection, performance, leakage, shock, vibration, thermal vacuum, burst factor pressure testing, and life per SMC-S-016. Other qualification tests necessary to satisfy the verification requirements in Table 7 Verification Matrix will also be conducted. Performance tests will be conducted before environmental testing and after each environmental test.

4.1.8 Acceptance

The acceptance phase of verification will be performed on all articles fabricated in accordance with the approved design baseline and verification procedures. The acceptance verification will be a set of verification activities preformed prior to the qualification verification for the qualification item(s) and after completion of manufacturing of all production items to assure compliance with the design baseline.

Each article produced will have a unique acceptance verification data package though the content will be similar or identical for each.

4.1.8.1 Acceptance Testing

Acceptance testing of the propulsion subsystem shall include inspection, wear-in, performance, leakage, vibration, and proof pressure testing per SMC-S-016. Other acceptance tests necessary to satisfy the verification requirements in Table 7 Verification Matrix will also be conducted. Performance tests will be conducted before and after each environmental test.

0. Valve Force Margins Valve force margins shall be verified at during the acceptance phase. Method may be a combination of analysis and test, but shall be anchored in acceptance test data. (e.g. pull in/ drop out voltage test in combination with a magnetic analysis to correlate force output to voltage input.)

4.1.9 Test Reporting Test reporting shall include sufficient information on instrumentation, process, environment and other factors to allow accurate assessments of data quality.

4.5 HARDWARE DELIVERABLES

4.5.1 FLIGHT UNIT

One flight unit shall be delivered per the Statement of Work.

4.5.2 QUALIFICATION UNIT

One qualification unit shall be delivered per the Statement of Work.

4.5.3 NON-FLIGHT HARDWARE/TEMPORARY INSTALLATIONS

The Offeror shall develop and deliver any non-flight hardware and temporary installations, such as pull tabs, thruster plugs, and protective covers that are required by design to protect and/or provide other functions to the hardware prior to installation and/or launch. All non-flight hardware and temporary installations shall be identified and written documentation explaining their fit/form/function provided.

4.5.4 GROUND SUPPORT EQUIPMENT

The Offeror shall provide ground support equipment as required for integration and checkout of the propulsion system

4.5.4.1 Hardware Check-out

The Offeror shall provide procedures for the check-out of hardware. The check-out test will allow the evaluation of whether the equipment still meets requirements after shipping, environmental testing, or unforeseen events. This includes the following requirements:

Requirements 3.2.3.3.1-4 regarding voltages.

Requirements 3.2.3.4.1-3 regarding sensors.

Requirements 3.2.4.1, 2, 4, and 5 regarding valves Requirement 3.2.7.1 regarding system leakage Requirement 3.2.7.6 regarding thruster leakage

4.5.5 RCS INTERFACE

The Offeror shall provide a means to interface with the RCS electronically, i.e. a control board with correct firmware or similar interface.

SECTION 5.0 PREPARATION FOR DELIVERY

5.1 PRESERVATION REQUIREMENTS

Preservation methods include the application of protective measures and materials to maintain the item within specified cleanliness levels and to prevent damage or degradation in reliability or performance of the item when exposed to the natural and induced ground based transportation and ground handling environments. The requirements for preservation methods and levels of protection shall be in accordance with NPR6000.1.

5.2 PACKING REQUIREMENTS

Packing may include special design containers used for the transportation and storage of the item. Packing may be used in conjunction with the appropriate preservation methods, handling procedures, and methods of transport to prevent damage or degradation in reliability or performance of the item when exposed to the natural and induced ground based transportation and ground handling environment. Packing requirements shall be in accordance with NPR6000.1 and supplemented by the following paragraphs.

5.2.1 MONITORING DEVICES

Instrumentation for monitoring or recording in-transit environments (e.g... shock, vibration, temperature, etc.), shall be specified and approved by NASA prior to implementation.

5.2.2 TEMPORARILY INSTALLED HARDWARE IDENTIFICATION

All temporarily installed devices such as caps, plugs, covers, support brackets, protective plates, etc., shall be highly visible, red in color or shall have attached highly visible red colored streamers to ensure that they are easily identified under casual observation.

5.2.3 MARKING FOR SHIPMENT

Interior and exterior containers shall be marked and labeled in accordance with NPR6000.1 and MIL-STD-129 including precautionary markings necessary to ensure safety of personnel and facilities and to ensure safe handling, transport, and storage. NASA Critical Item Label (NASA Form 1368) shall be applied in accordance with NPR6000.1. The following identification information on interior and exterior containers shall include as a minimum:

a) NASA Contract Number

b) Item Name

c) Seller’s Part Number

d) Serial Number

e) Seller’s Name

f) Seller’s Cage Code

g) Quantity

h) Cleaning Marking – Add the words “THIS PART HAS BEEN CLEANED TO LEVEL 100A OF IEST-STD-CC1246.”

i) Gross Weight image1.gif image2.emf image3.png image4.png image5.png image6.png image7.png image8.png image9.png image10.png image11.png image12.png image13.jpeg image14.jpeg