-_Spec.pdf

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400-FORM-0002 (4/16/2014)

DRAFT

LANDER-PROP-SPEC-0002, Revision - EUROPA Lander Propulsion, Code 597

Engine Specification

EAR ECCN - [9E515]

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Goddard Space Flight Center Greenbelt, Maryland

National Aeronautics and Space Administration https://ipdtdms.gsfc.nasa.gov/

Engine SPEC LANDER-PROP-SPEC-0002, Revision - Effective Date: 08/08/2018 ii

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Engine Specification Signature/Approval Page

Revision - Revision A Prepared by:

Eric Cardiff

Reviewers/Approvers:

TBD

Approved by:

Eric Cardiff

*** Electronic signatures are available on-line at: https://ipdtdms.gsfc.nasa.gov*** https://ipdtdms.gsfc.nasa.gov/ iii

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Preface This document is a Europa Lander Propulsion project signature-controlled document. Changes to this document require prior approval of the applicable Product Design Lead (PDL) or designee. Proposed changes shall be submitted in the EUROPA Propulsion Technical Data Management System (TDMS) via a Signature Control Request (SCoRe) along with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

All of the requirements in this document assume the use of the word "shall" unless otherwise stated.

Questions or comments concerning this document should be addressed to:

Europa Lander Propulsion Configuration Management Office Mail Stop: 597 Goddard Space Flight Center Greenbelt, Maryland 20771 iv

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Change History Log

Revision Effective Date Description of Changes (Reference the SCoRe Approval Date)

Revision - TBD Released following approval of LANDER-SCoRe-TBD v

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Table of Contents

1. Introduction

1.1 General Information

1.2 Scope

1.3 Engine Definition and Description

1.4 Valve Definition

1.5 Filter Definition

1.6 Filter Capacity

1.7 Identification and Marking

1.8 Interchangeability

2. Applicable Documents

3. Interface Requirements

3.1 PROPELLANTS and Pressurant

3.2 TEST FLUIDS

3.3 ELECTRICAL INTERFACES

Safe closure Wires Ignition Proof Voltage Range ........................................... Error! Bookmark not defined.

Open Voltage Hold Open Voltage Drop-Out Voltage

3.4 MECHANICAL INTERFACE

Mounting Flange Valve Inlet Tube

3.5 THERMAL INTERFACES

Conduction

4. Functional Requirements

4.1 Life

ENGINE LIFE

Valve Life Cycle Number of Pulses Dry Cycles Engine Cold Starts

4.2 Nominal Operation

vi

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4.3 Operational Ranges

4.4 Gas Ingestion

4.5 Helium Saturated Propellants

4.6 Propellant Depletion

4.7 Surge Pressure

4.8 Steady-State Firing Duration

4.9 Thermal Functional

PROPELLANT TEMPERATURE

Engine Temperatures Engine Restart Heat Soakback

4.10 Leakage

EXTERNAL LEAKAGE

Internal Leakage

4.11 Electrical Functional

Minimum Electrical Pulse Width (EPW) Coil Resistance Coil Inductance Insulation Resistance “Hit and Hold” Valve Operation Valve No-Flow Operation Survival Voltage Grounding

4.12 Mass

4.13 Alignment

5. performance Requirements

5.0 Thrust

5.1 SPECIFIC IMPULSE

5.2 Propellant Throughput

5.3 Parameter Variability Due To Propellant Throughput

5.4 Response Times

5.5 Steady-State Firing Variability

Thrust Variability Propellant Flowrate Variability Specific Impulse Variability Mixture Ratio Variability vii

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5.6 Combustion Roughness

5.7 PULSE MODE OPERATION

Minimum Impulse Bit (IBIT) Pulse Mode Specific Impulse Pulse to Pulse Repeatability

5.7.3.1 Small EPW Isolated Pulses

5.7.3.2 Other Pulse Trains

Following the initial thermal transient as defined in Paragraph 4.3.5 the pulse to pulse Ibit repeatability for the pulses in any pulse train shall be ± 0.50%

6. Environmental Requirements

6.1 Handling Environments

6.2 Mechanical Factors of Safety

6.3 Quasi-Static Acceleration

6.4 Frequency Requirement

6.5 Random and Sine Vibration

6.6 Shock Requirements

6.7 Humidity

6.8 ESD Survival

6.9 Flight Interface Design Temperature Limits

7. Cleanliness

7.0 External Cleanliness

7.1 Internal Cleanliness

7.2 Assembly Outgassing

Appendix A Abbreviations and Acronyms (TBR) viii

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List of Tables Table Page Table 1 Applicable Documents

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1. INTRODUCTION

1.1 GENERAL INFORMATION

The Europa Lander mission is being formulated and implemented by a joint partnership between the Jet Propulsion Laboratory and the Applied Physics Laboratory, with NASA GSFC building the propulsion system for the Applied Physics Laboratory.

1.2 SCOPE

This specification describes the electrical, mechanical, environmental and performance requirements for a space-qualified Propulsion Subsystem Engine for the Europa Lander mission.

1.3 ENGINE DEFINITION AND DESCRIPTION

The Propulsion Subsystem engine is nominally a 20 lbf engine. The subsystem will use 16 engines, consisting of 8 primary engines and a secondary redundant string of 8 engines. The engine will be pressure fed and operate with monomethylhydrazine (MMH) as the fuel and nitrogen tetroxide (MON-3) as the oxidizer at a nominal mixture ratio of 1.63 TBR.

1.4 VALVE DEFINITION

The engine shall use dual seat solenoid valves to control the flow of propellant to the engine.

1.5 FILTER DEFINITION

The valve shall incorporate a 25 micron absolute titanium filter at the valve inlet.

1.6 FILTER CAPACITY

The filter shall have a capacity of >100 mg TBR.

1.7 IDENTIFICATION AND MARKING

The unit shall be marked with the part number and a unique sequential serial number in the area designated on the ICD in a manner to be approved by the COR.

1.8 INTERCHANGEABILITY

The engine shall be directly interchangeable in form, fit, and function with other engines of the same part number.

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2. APPLICABLE DOCUMENTS

The following documents and drawings shall apply to the fabrication and to the electrical, mechanical, and environmental requirements of the engine to the extent specified herein. In the event of conflict between this specification and any referenced document, this specification will govern, with the exception of the Europa Lander Engine Statement of Work (LANDER-PROP-SOW-002), in which case the Statement of Work takes precedence.

The following is a list of the applicable specifications and publications. The vendor shall use the current version of the document as of the award of the contract.

Table 1 Applicable Documents TBR Pending Rick Pfister

Document Number Title

NASA Std. 8719.24, Rev. A Range Safety User Requirements

SAE-AMS-2244C Tolerances (R) Titanium and Titanium Alloy Tubing

IEST-STD-CC1246D Product Cleanliness Levels And Contamination Control Program

ASTM E-595-07 Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment

MIL-C-5541 Military Specification, Chemical Conversion Coatings on Aluminum and Aluminum Alloys

AMS 2488 Anodic Treatment - Titanium and Titanium Alloys Solution Ph 13 Or Higher

EEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, and Derating

EUROPA-PROP-SPEC-0001 Subsystem Specification

MIL-PRF-27407B Propellant, Helium, Pressurizing Agent

MIL-STD-889B Dissimilar Metals

MIL-PRF-26539E Propellant, Nitrogen Tetroxide (MON 3, Low Iron)

MIL-PRF-27404C Propellant, Monomethylhydrazine

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JSC-SPEC-C-20C Water, High Purity, Specification for

TT-I-735A Isopropyl Alcohol

J-STD-001ES Requirements for Soldered Electrical and Electronic Assemblies

LANDER-PROP-SOW-0002 Europa Lander Engine Statement of Work

JPL-D-80TBD, Rev. - Environmental Requirements Document

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3. INTERFACE REQUIREMENTS

3.1 PROPELLANTS AND PRESSURANT

The propellants and pressurant gas used in the engine shall be:

• fuel: monomethylhydrazine (MMH) per MIL-PRF27407C

• oxidizer: Nitrogen Tetroxide (MON 3) per MIL-PRF-26539E

• Pressurant: Gaseous helium per MIL-PRF-27407B

3.2 TEST FLUIDS

The engine shall be capable of being flushed with test fluids for flow testing, propellant flushing, and calibration. The only flushing fluids to be used shall be distilled deionized or demineralized water per JSC-SPEC-C-20TBR and isopropyl alcohol (IPA) per TT-I- 735A.

3.3 ELECTRICAL INTERFACES

Safe closure

In the event of an electrical power failure or component failure, the engine valves shall close safely. If power fails and the valves are closed they shall remain closed. If power fails and the valves are open they shall return to the closed position. Electrical power shall be required only during engine operation.

Wires

Each valve shall have a separate set of lead wires.

Ignition Proof

Electrical components of the engine shall not cause ignition of any explosive mixtures surrounding or in contact with the engine.

Inrush Current

The inrush current shall be less than TBD A.

Open Voltage

The engine valves shall open when a voltage 12 - 18 VDC is applied across the fuel and oxidizer valves wired in series. TBR

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Hold Open Voltage

The engine valves shall remain open when the applied voltage is reduced to 6 VDC across the fuel and oxidizer valves wired in series.

Drop-Out Voltage

The drop-out voltage is defined as the voltage below which an open valve will close for a valve at 70F and pressurized to 300 psia. The drop-out voltage for the single seat valve as measured during the valve acceptance tests shall be ≥ 0.375 VDC and ≤ 1.70

VDC.

3.4 MECHANICAL INTERFACE

Mounting Flange

The engine mounting flange shall allow the engine to be aligned on the spacecraft.

Valve Inlet Tube

The valve inlet tube shall have a 0.250 (+0.004, -0.000) inch OD with a 0.028 ±0.002 inch wall thickness. The material shall be 6Al-4V Titanium. The inlet tube length shall be 1.25 ± 0.02 inches.

Engine Length

The length of the engine downstream of the mounting interface shall be less than TBD in.

3.5 THERMAL INTERFACES

The engine shall be thermally stable at all operational conditions specified herein and shall not experience thermal runaway or any other abnormal operational condition or limitation due to heating or other thermal conditions.

Conduction

The engine shall not use the spacecraft as a thermal sink to conduct heat from the engine in order to maintain a thermally-stable operating condition. When tested, the interface between the engine and the test stand shall simulate an adiabatic interface to demonstrate compliance with this requirement. TBR, how much flux is acceptable

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4. FUNCTIONAL REQUIREMENTS

4.1 LIFE

ENGINE LIFE

The engine shall be capable of meeting all specification requirements after:

a. Two (2) years of integration and test

b. One (1) year of ground storage

c. 11.3 years of mission life TBR.

Valve Life Cycle

Each propellant valve shall be qualified to 1,000,000 wet cycles.

Number of Pulses

The engine shall be capable of delivering a minimum of TBD pulses in pulse trains of any duration at any condition within the ranges given in Paragraph 4.1.1.

Dry Cycles

The valve shall meet all the requirements of this specification subsequent to being subjected to 200 dry cycles without fluid flow during checkout testing.

Engine Cold Starts

The engine shall have the demonstrated capability to perform a minimum of 300 starts, defined as going from a cold non-firing condition to thermal equilibrium at the standard operational condition.

4.2 NOMINAL OPERATION

The nominal operational conditions for the engine shall be:

Engine Inlet Pressure = 300 psia TBR

Engine Inlet Propellant Temperature = 70F Propellant Mixture Ratio = 1.65

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4.3 OPERATIONAL RANGES

The engine shall be capable of meeting all operational requirements when operated under the following conditions or any combination of these conditions:

Engine Inlet Pressure = 100 – 400 psia Propellant Mixture Ratio = 1.0 – 2.1 (Oxidizer/Fuel) Engine Inlet Propellant Temperature = -10C to 60C (14F – 140F) Engine Injector Initial Temperature ≥ -17C (1.4F) TBR

4.4 GAS INGESTION

The engine shall be capable of continuous or pulse mode operation when up to 150 cm3 (9.15 in3) of helium at standard temperature and pressure (STP) is ingested by the engine on either the fuel or oxidizer sides or both sides simultaneously either at the start of a firing or during a firing.

Ingestion of helium shall have no detrimental effect on the engine operation during or subsequent to the gas ingestion.

4.5 HELIUM SATURATED PROPELLANTS

The engine shall be capable of meeting all the requirements of this specification when the fuel and oxidizer propellants are fully saturated, partially saturated or not saturated with helium.

4.6 PROPELLANT DEPLETION

The engine shall withstand the depletion of either fuel or oxidizer during a firing without any adverse effects. Duration TBD. Delete?

4.7 SURGE PRESSURE

The engine shall meet all requirements of this specification following a transient surge pressure of 1500 psia.

4.8 STEADY-STATE FIRING DURATION

The mission requires the engine to have the capability to perform a 2 TBR hour steady-state burn. The engine shall have the demonstrated capability to perform a 4 hour steady-state burn at any condition within the ranges given in Paragraph 4.1. There shall be no time limitations on the duration of any steady-state firing for the operational conditions given in Paragraph 4.1.1.

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4.9 THERMAL FUNCTIONAL

PROPELLANT TEMPERATURE

The engine shall be capable of operating at any propellant temperature between -10C and 60C.

Engine Temperatures

The engine shall be capable of performing all mission operations with initial engine injector temperature as low as TBD C.

Engine Restart

The engine shall be capable of restarting and operating at any combination of engine, valve or propellant temperatures resulting from thermal soakback from either steady-state or pulse mode firings over the range of operational conditions given in 4.1.2.

Heat Soakback

Heat soakback from the engine for any steady-state or pulse mode firing shall not exceed the maximum allowable valve temperature or maximum allowable temperature for any other component on the engine.

4.10 LEAKAGE

EXTERNAL LEAKAGE

External leakage from the engine joints or assembly shall be < 1 x 10-6 scc/sec GHe.

Internal Leakage

The maximum allowable leakage across the valve seat is 3 scc/hr (8.3 x 10-4 scc/sec) of GHe.

4.11 ELECTRICAL FUNCTIONAL

Minimum Electrical Pulse Width (EPW)

The engine shall be capable of producing thrust at an EPW ≥ 8 msec.

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Coil Resistance

The resistance of each valve coil at 70F shall be TBD Ω ± 1 Ω. Single seat vs dual.

Coil Inductance

The inductance of each coil for an open valve at any temperature within the ranges given in Paragraph 4.12.9 shall be greater than TBD mH and less than TBD mH.

Insulation Resistance

The insulation resistance shall be greater than 100 MΩ when measured at 70°F with a potential of 500 ± 50 VDC applied between the shorted coil leads and the case.

“Hit and Hold” Valve Operation

The planned operational mode for any engine firing is to open the valve at the initial supply voltage given in Section 4.12.5 (the hit voltage) and after 35 - 45 msec, drop the voltage to a voltage in the 4 – 9 VDC range. This operational mode will be used for both steady-state and pulse mode engine firings. The valve shall have the capability to operate in this “hit and hold” mode at all operational conditions.

Valve No-Flow Operation

The valve shall have the capability to operate at the voltages given in Section 4.12.5 with no flow through the valve for 2.0 minutes without overheating the valve.

The valve shall have the capability to operate at the hold voltage of 4 - 9 VDC with no flow through the valve for 2.0 minutes without overheating the valve.

Survival Voltage

The valve shall survive without damage when subjected to a voltage of up to 40 VDC for one (1) minute with no flow.

Grounding

The DC resistance of a mechanical contact from the top of the engine mounting feet to the bottom of the engine valve mounting feet shall be less than or equal to 1 milliohm.

All conductors with areas greater than 1 mm2 within the engine shall have a bleed path to the feet or case of the engine, or to a lead wire, with a resistance less than or equal to 100 mega-ohm.

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The DC resistance of a mechanical contact from the shield on any lead wire to the bottom of the engine mounting feet shall be less than or equal to 2.5 milliohm.

The shield on each lead wire shall be terminated to the housing in a 360 degree manner.

4.12 MASS

The mass of the engine shall not exceed TBD kg (TBD lbm) including all tubing and valves. The mass shall be measured to 0.10 kg (0.22 lbm). The CG shall be calculated and indicated on the Interface Control Drawing (ICD).

4.13 ALIGNMENT

The centerline of the nozzle shall be perpendicular to the mounting flange to within ± 0.5 degrees. The position of the nozzle centerline with respect to the mounting interface centerline shall be less than ± 0.050 inches (± 1.27 mm). These requirements can be verified by the drawing tolerances.

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5. PERFORMANCE REQUIREMENTS MOVE FORWARD

5.1 THRUST

The engine thrust shall be 20 lbf (90 N) at nominal conditions, and no more than 28 lbf (125N).

5.2 SPECIFIC IMPULSE

The nominal vacuum specific impulse for the engine at a feed pressure of 270 psia with ambient temperature propellants for a steady-state firing of 30s or greater shall be ≥

320.0 s.

5.3 PROPELLANT THROUGHPUT

The engine shall be qualified for a propellant throughput of >2,250 kg TBR.

5.4 PARAMETER VARIABILITY DUE TO PROPELLANT THROUGHPUT

The engine shall not exhibit a change in thrust, mixture ratio, specific impulse or propellant flowrate due to the propellant throughput level.

5.5 RESPONSE TIMES

The thrust startup time for any steady-state or pulse mode firing at any operational condition as measured by T90 (time to 90% thrust) shall not exceed 15 msec. The thrust shutdown time as measured by T10 (time to 10% thrust) shall not exceed 10 msec.

5.6 STEADY-STATE FIRING VARIABILITY

For the purposes of this paragraph the initial thermal transient is taken as the time required for the engine injector temperature as measure by the thermocouples on the injector backside to reach 70% of its steady-state value.

Thrust Variability

During any steady-state firing, after the initial thermal transient, for constant engine inlet conditions the thrust shall vary by no more than ± 2%.

Propellant Flowrate Variability

During any steady-state firing, after the initial thermal transient, for constant engine inlet conditions the fuel and oxidizer flowrates shall vary by no more than ± 2%.

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Specific Impulse Variability

During any steady-state firing, after the initial thermal transient, for constant engine inlet conditions the specific impulse shall vary by no more than ± 3%.

Mixture Ratio Variability

During any steady-state firing, after the initial thermal transient, for constant engine inlet conditions the mixture ratio shall vary by no more than ± 0.05.

5.7 COMBUSTION ROUGHNESS

Chamber pressure oscillations shall not cause any damage to the engine or performance shortfalls during any steady-state or pulse mode operation.

The engine shall not exhibit any occurrence of “spiking” (rapid isolated increases of chamber pressure), at any combination of feed pressure, temperature, mixture ratio, hardware temperature, pulse mode duty cycle or any other condition.

5.8 PULSE MODE OPERATION

Minimum Impulse Bit (IBIT)

The engine shall be capable of producing an isolated pulse minimum Ibit of TBD mN-s ± 15 mN-s at the nominal condition with an EPW ≥ 8 msec.

Pulse Mode Specific Impulse

The minimum specific impulse for any engine as a function of EPW with t(off) ≥ 60 s at a an engine inlet pressure of 270 psia and propellant at ambient temperature shall be as given in Table 3.

Table 3. Minimum Specific Impulse Requirement vs EPW Electrical Pulse Width (ms) Specific Impulse (s)

20 150 TBR

50 200 TBR

100 280 TBR

500 300 TBR

1000 320 TBR

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Pulse to Pulse Repeatability

5.8.3.1 Small EPW Isolated Pulses

Pulse-to-pulse Ibit repeatability for any pulse in a train of isolated pulses (toff ≥ 60 s) with EPW = 8 – 1000 msec at the nominal condition shall be ≤ ± 15 mN-s.

5.8.3.2 Other Pulse Trains

Following the initial thermal transient as defined in Paragraph 4.3.5 the pulse to pulse Ibit repeatability for the pulses in any pulse train shall be ± 0.50%.

Number of pulses

TBD.

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6. ENVIRONMENTAL REQUIREMENTS

Environmental design requirements for the engine are specified in this section. The spacecraft components shall be capable of meeting their performance requirements after exposure to the environments specified in this section.

6.1 HANDLING ENVIRONMENTS

Flight hardware shall be designed to survive without degradation in the thermal, pressure and relative humidity environments specified below in Table 5.

Table 4. Handling Environments

Control Parameter Low Limit High Limit

Air Temperature (Storage) Air Temperature (Operational) Temperature Change Rate Pressure (10,000 ft max. altitude) Relative Humidity

+5°C

+5°C -10°C/hr 6.9×104 N/m2 (520 Torr) 30%

+50°C

+40°C +10°C/hr 1×105 N/m2 (760 Torr) 70%

6.2 MECHANICAL FACTORS OF SAFETY

The engine shall demonstrate positive Margins of Safety under limit loads for all yield and ultimate failures using the Factors of Safety (FS) defined in Table 6.

Margin of Safety (MS) is defined as follows:

MS = Allowable Stress (or Load) Design Limit Stress (or Load)x FS

− 1

Table 5. Factors of Safety

Type of Hardware1 Static/Sine Random/Acoustic2 Tested Metallic Structure Yield 1.25 1.6

Tested Metallic Structure Ultimate 1.4 1.8

Stability Ultimate 1.4 1.8

Bonded Inserts/Joints Ultimate 1.5 1.9

Untested Flight Structure Yield- metallic only 2.0

Untested Flight Structure Ultimate - metallic only 2.6

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1 – Factors of safety for pressurized systems to be compliant with NASA-STD-9719.24, “Range Safety User Requirements.”

2 – Factors shown should be applied to statistically derived peak response based on RMS level. As a minimum, the peak response shall be calculated as a 3-sigma value.

6.3 QUASI-STATIC ACCELERATION

Quasi-static acceleration represents the combination of steady-state accelerations and the low frequency mechanically transmitted dynamic accelerations that occur during launch.

The engine shall demonstrate its ability to meet its performance requirements after being subjected to a net limit load of 60 g’s at the center of mass of the valve.

6.4 FREQUENCY REQUIREMENT

The engine shall have a fundamental frequency greater than 80Hz when hard mounted at its spacecraft interface.

6.5 RANDOM AND SINE VIBRATION

The random vibration and sine vibration requirements for the engine are given in Tables 6 and 7 and are shown in Figures 2 and 3. Vibration requirements are applied at the engine level.

6.6 SHOCK REQUIREMENTS

The shock requirements for acceptance and qualification testing are given in Table 8 and shown in Figure 4. Shock requirements are applied at the engine level.

TBR for environment

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Frequency (Hz) Protoflight/Qual Level Acceptance Level

Hz

20 - 80 + 5.9 dB/oct + 5.9 dB/oct

80 - 150 0.30 g2 / Hz 0.15 g2 / Hz

150 - 200 - 13.8 dB/oct - 13.8 dB/oct

200 - 500 0.08 g2 / Hz 0.04 g2 / Hz

500 - 2000 - 3.0 dB/oct - 3.0 dB/oct

Over All 10.8 grms 7.6 grms

Table 6. Random Vibration Levels

Hz g g

5 0.7 0.5

7 1.5 1.1

20 1.5 1.1

40 4.0 2.8

80 4.0 2.8

Table 7. Sine Vibration Levels

Hz g g

100 100 70

1000 1400 1000

2000 2300 1650

10000 2300 1650

Table 8. Shock Levels

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Figure 2. Random Vibration Spectrum

Figure 3. Sine Vibration Spectrum

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Figure 4. Shock Spectrum

6.7 HUMIDITY

The engine shall be able to meet performance requirements after exposure to relative humidity levels of 0% to 70% for the defined mission life.

6.8 ESD SURVIVAL

All EEE Parts within the valve shall survive repeated exposure to ESD pulses up to 1000 V Human Body Model per MIL-STD-883G.

6.9 FLIGHT INTERFACE DESIGN TEMPERATURE LIMITS

The engine shall be capable of surviving indefinitely without damage or permanent performance degradation when its temperatures are within the qualification survival limits shown in Table 10.

Table 9. Qualification Survival Temperature Levels at Mounting Interface Minimum

Temperature (ºC) Maximum

Temperature (ºC) Allowable Flight Temperature 0 +33 Flight Acceptance -5 +38 Protoflight or Qualification -15 +55

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7. CLEANLINESS

7.0 EXTERNAL CLEANLINESS

The COPV external cleanliness shall meet IEST-VC-0.5-1000 per IEST-STD-CC1246, prior to delivery to NASA/GSFC.

7.1 INTERNAL CLEANLINESS

The COPV internal cleanliness shall meet level 100A per IEST-STD-CC1246 (as modified by the following: no metal particles allowed above 25 um) prior to integration in the system.

7.2 ASSEMBLY OUTGASSING

When measured in a vacuum of 10-6 Torr at 50°C, the engine outgassing shall not exceed 2E-10 g/sec per kg of unit under test of mass that is condensable on a Quartz Crystal Monitor (QCM) that is operated at -20°C.

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APPENDIX A ABBREVIATIONS AND ACRONYMS (TBR)

Abbreviation/ Acronym Definition

AC Alternate Current CCB Configuration Control Board CCR Configuration Change Request CLA Coupled-Loads Analysis CM Configuration Management CMO Configuration Management Office COR Contracting Office Representative CVCM Collected Volatile Condensable Mass DC Direct Current EEE Electrical, Electronic, and Electromechanical EMI Electromagnetic Interference EOL End of Life FS Factor of Safety GSFC Goddard Space Flight Center I&T Integration and Test MS Margin of Safety MEOP Maximum Expected Operating Pressure M Mega, 106 MGSE Mechanical Ground Support Equipment µ Micro, 10-6 m Milli, 10-3 Mohms Megaohms MS Margin of Safety N/A Not Applicable NASA National Aeronautics and Space Administration OD Outer Diameter PSIA Pounds per Square Inch – Absolute PSID Pounds per Square Inch – Differential QCM Quartz Crystal Monitor RE Radiated Emissions RS Radiated Susceptibility SC Spacecraft SOW Statement of Work STP Solar Terrestrial Probe TBD To Be Determined TBR To Be Reviewed TML Total Mass Loss

5.8.4 Number of pulses

Appendix A Abbreviations and Acronyms (TBR)