Attachment_B_Specifications.pdf

80GSFC18R0053

Attachment B

Effective Date: May 7, 2018

Expiration Date: May 7, 2023

Check https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use

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

PACE-PROP-SPEC-0083, Revision -

Plankton, Aerosol, Cloud, ocean Ecosystem (PACE), Code 427

PACE Thruster Specification

Goddard Space Flight Center

Greenbelt, Maryland

National Aeronautics and

Space Administration

GSFC PACE CMO

05/07/2018

Released https://ipdtdms.gsfc.nasa.gov/

PACE Thruster Spec PACE-PROP-SPEC-0083, Revision -i

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

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

PACE Thruster Specification

Signature/Approval Page

Prepared By:

Jacob Stahl

Reviewed By:

Bobby Gheen

John Blackwood

Gary Davis

Beth Weinstein

Jack Sanders

Gary Won

Paul Mason

Dan McGuinness

Antonio Reyes

Shavesha Rutledge

Henry Mulkey

John Tota

Bob Kraeuter

Craig Stevens

Walt Ancarrow

Dave Sohl

Approved By:

Andre’ Dress

Electronic Signatures available online at: https://ipdtdms.gsfc.nasa.gov/ ii

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

Preface

This document is under Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) Mission configuration control. Changes to this document require prior approval of the PACE

Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the PACE Configuration Management Office (CMO), along with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

iii

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Change History Log

Revision Effective

Date

Description of Changes

(Reference the CCR & Approval Date)

Revision - 05/07/2018 Baseline Release following the approval of PACE-SCoRe-0257 iv

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Table of Contents

1.0 Introduction

1.1 General Information

1.2 Scope

2.0 Applicable Documents

3.0 Contract Description

3.1 Thruster Description

3.2 Ground Support Equipment Description

4.0 Functional/Performance Requirements

4.1 Thruster Flight Unit Functional/Performance Requirements

4.1.1 Thrust

4.1.2 Thrust Repeatability

4.1.3 Thrust Alignment

4.1.4 Specific Impulse

4.1.5 Pulsed Specific Impulse

4.1.6 Impulse Bit

4.1.7 Propellant Throughput

4.1.8 Thruster Duty Cycle

4.1.9 Catalyst Bed Heater

4.1.10 Inlet Pressure

4.1.11 Propellant Temperature

4.1.12 Propellant

4.1.13 Pressurant Gas

4.1.14 Leakage

4.1.14.1 External Leakage

4.1.14.2 Internal Leakage

4.1.14.3 Internal Leakage During Vibration

4.1.15 Firing Duration

4.1.16 Roughness

4.1.17 Proof and Burst Pressure

4.1.17.1 Thrusters

4.1.17.2 Propellant Inlet Tubes

4.1.18 Cycles

4.1.18.1 Thermal Cycles

4.1.18.2 Valve Cycles

4.1.19 Cold Starts

4.1.20 Thruster Sub-Components

4.1.20.1 Inlet Filter

4.1.20.2 Thruster Valves

4.1.20.3 Thruster Valve Heaters

4.1.20.4 Thruster Valve Thermostats

4.1.20.5 Thruster Valve Thermistors

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4.1.20.6 Catalyst Bed Heaters

4.1.20.7 Catalyst Bed PRT

4.1.21 Gas Ingestion

4.2 Resource Allocations

4.2.1 Mass Allocation

4.2.2 NA

4.2.3 Peak Power Allocation

4.3 Power

4.3.1 NA

4.3.1.1 Operating Voltage Range

4.3.1.2 Abnormal Voltages

4.4 Electrical Grounding

4.4.1 Primary Power DC Isolation from Component Chassis

4.4.2 Internally Generated Secondary Returns

4.4.3 Internally Generated Secondary to Primary DC Isolation

4.4.4 Mechanical Contact Resistance

4.4.5 Mating Method

4.5 NA

4.5.1 NA

4.5.2 NA

4.5.2.1 NA

4.5.2.2 NA

4.5.2.3 Passive Linearizing Resistor

4.5.2.4 Thermistor Resistor Conditioning

4.5.3 NA

4.5.4 NA

4.5.5 NA

4.5.5.1 NA

4.5.5.2 NA

4.5.5.3 Wire Pairing

5.0 Physical Requirements

5.1 Interface Documentation

5.1.1 Interface Drawings

5.2 Mass Properties

5.2.1 Component Masses

5.2.2 Center of Mass Location

5.2.3 Center of Mass Accuracy

5.3 Physical Envelope

5.4 Mounting

5.4.1 Mounting Flange

5.5 NA

5.6 Alignment

5.6.1 Alignment Location/Orientation Accuracy

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5.6.2 NA

5.6.3 NA

5.6.4 NA

5.6.5 Alignment Measurement Tooling/Features

5.7 Fluid Interfaces

6.0 Environmental Requirements

6.1 Mechanical Factors of Safety

6.2 Quasi-Static Acceleration

6.3 Frequency Requirement

6.3.1 Fundamental Launch Frequencies

6.4 Vibration

6.4.1 Sinusoidal Vibration

6.4.2 Random Vibration

6.5 Shock

6.6 NA

6.7 Transportation

6.7.1 Transportation Cleanliness

6.8 Pressure

6.8.1 Operating Pressure Range

6.8.2 NA

6.8.3 Launch Vehicle Environmental Control System Impingement Velocity

6.9 NA

6.10 Ground Environments

6.11 Thermal Requirements

6.11.1 Flight Interface Design Temperature Limits

6.11.2 Valve Temperature Limits

6.11.3 Thermal Interfaces

6.11.3.1 Conduction

6.11.3.2 Radiation

6.12 NA

6.13 Atomic Oxygen Fluence

6.13.1 Atomic Oxygen Analysis

6.13.2 Atomic Oxygen Testing

7.0 Cleanliness

7.1 Surface Contamination

7.1.1 Surface Contamination Levels at Delivery

7.1.1.1 Particulate Contamination

7.1.1.2 Molecular Contamination – Exposed Surfaces

7.1.1.3 Molecular Contamination – Covered Surfaces

7.1.1.4 External Cleanliness

7.1.1.5 Internal Cleanliness

7.1.2 Surface Contamination Generation

7.1.2.1 Particulate Generation

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

7.1.2.2 Molecular Generation

7.1.3 Cleanability

7.1.3.1 Cleanability – Sensitive Surfaces

7.1.3.2 Cleanability – Sensitive Surface Cleaning Methods

7.2 Electrostatic Cleanliness

7.2.1 NA

7.2.2 Conductive Surface Resistivity

8.0 Design & Construction Requirements

8.1 Parts, Materials & Processes (PMP)

8.1.1 EEE Parts

8.1.2 Materials

8.1.2.1 Material Conductivity

8.1.2.2 Material Limitations for Debris Casualty Area

8.1.2.3 Welding

8.1.2.4 Bi-Metallic Joints

8.1.2.5 Thin Films

8.1.2.6 Lubricants

8.1.2.7 Fluid Compatibility

8.1.2.8 Thruster Dewpoint

8.2 Electrical

8.2.1 NA

8.2.2 Harness Requirements

8.2.2.1 Connector Selection

8.2.2.2 Thermal Circuit Connections

8.2.2.3 NA

8.2.2.4 NA

8.2.2.5 NA

8.2.2.6 Harness Wires

8.2.3 NA

8.2.3.1 NA

8.2.3.2 Floating Conductors

8.3 NA

8.4 NA

8.5 Identification and Marking

8.6 Workmanship

8.6.1 Workmanship Standards

8.6.2 Welding Procedures

8.7 Reliability and Mission Lifetime

8.7.1 Mission Life

8.7.2 Reliability

8.8 Ground Handling

8.8.1 NA

8.8.2 NA

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

8.8.3 NA

8.8.4 NA

8.8.5 NA

8.8.6 NA

8.8.7 GSE Cleanliness

8.8.8 NA

8.8.9 Test Harness

9.0 Mechanical Design Requirements

9.1 Structural Requirements

9.1.1 Component Fatigue

9.1.2 Fracture Control Requirements

9.2 Fastening Systems

9.2.1 Fastener Performance Analysis

9.2.1.1 Factors of Safety

9.2.1.2 Supplemental Factor

9.2.1.3 Ultimate Design Loads

9.2.1.4 Yield Design Loads

9.2.1.5 Design Separation Load

9.2.2 Fastener Locking and Retention

9.2.2.1 Locking Features

9.2.2.2 Verification

9.2.2.3 Locking Features

9.2.2.4 Installation Torque Specification and Control

9.2.3 Fastened Joints Criteria

9.2.3.1 Minimum and Maximum Preload

9.2.3.2 Analysis Addressing Potential Rupture

9.2.3.3 Ultimate Strength Analysis

9.2.3.4 Applied Shear Loading

9.2.3.5 Shear Loading

9.2.3.6 Simultaneous Applied Tensile and Shear Loads

9.2.3.7 Allowable Yield Tensile Load

9.2.3.8 Separation Analysis

9.2.3.9 Seal Analysis

10.0 Logistics

10.1 NA

10.2 Ground Support Equipment

11.0 Verification Requirements

11.1 Verification Methods

11.1.1 Inspection

11.1.2 Analysis

11.1.3 Test

11.2 Inspection Requirements

11.2.1 Visual Inspection

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11.2.2 Physical Measurement

11.2.3 Documentation Search

11.3 Analysis Requirements

11.4 Test Requirements

11.4.1 Definitions

11.4.2 Test Factors

11.4.3 Test Tolerances

11.4.4 Test Restrictions

11.4.4.1 Failure During Tests

11.4.4.2 Modification of Hardware

11.4.4.3 External Adjustment

11.4.4.4 Re-Test Requirements

11.5 Required Tests

11.5.1 Performance Tests

11.5.1.1 NA

11.5.1.2 Functional Tests

11.5.1.3 NA

11.5.1.4 NA

11.5.1.5 Hot Fire Test

11.5.1.6 Gas Flow Impedance

11.5.1.7 Alignment

11.5.2 Mass Properties and Envelope Measurement

11.5.3 Static Loads/Strength Test

11.5.3.1 Sine Burst

11.5.3.2 Static Pull

11.5.4 Sine Sweep Survey

11.5.5 Sine Vibration

11.5.6 Random Vibration

11.5.7 NA

11.5.8 Proof Pressure Test

11.5.9 Shock

11.5.10 NA

11.5.11 NA

11.5.12 NA

11.5.13 Thermal Cycle Testing

11.5.14 Harness Tests

11.5.14.1 Continuity/Hi-Pot Tests

11.5.14.2 Harness Bake-Out

11.5.15 NA

11.5.16 Cleanliness

Appendix A Abbreviations and Acronyms x

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List of Figures

Figure Page

Figure 6-1 Sine Vibration Environment

Figure 6-3 Shock Envelope

Figure 8-1: Thermostat and Heater Wiring

List of Tables

Table Page

Table 2-1 Applicable Documents

Table 4-1 -3σ Worst Case Steady-State Specific Impulse

Table 4-2: Thermostat Set points

Table 6-1 Factors of Safety

Table 6-2 Thruster Design Limit Loads

Table 6-3 Thruster Sine Vibration Environment

Table 6-4 Thruster Random Vibration Environment, (22.7 kg, or less)

Table 6-6 Qualification Level Shock Response Spectrum

Table 6-7 Transportation Loads

Table 6-8: Valve Temperature Levels

Table 8-1 Limited materials for debris casualty area

Table 11-1 Test Factors and Durations

Table 11-2 Test Tolerances

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

1.0 INTRODUCTION

1.1 GENERAL INFORMATION

The Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) mission is a strategic climate continuity mission that will extend the high quality ocean ecological, ocean biogeochemical, cloud, and aerosol particle data records begun by NASA in the 1990s. The mission will be capable of collecting radiometric and polarimetric measurements of the ocean and atmosphere, from which these biological, biogeochemical, and physical properties will be determined. PACE data products will not only add to existing critical climate and Earth system records, but also answer new and emerging advanced science questions related to Earth’s changing climate.

1.2 SCOPE

This specification describes the performance, design, development, physical characteristics, electrical, mechanical, environmental, and verification testing requirements for a space-qualified

Propulsion Subsystem monopropellant hydrazine thrusters for the NASA Goddard Space Flight

Center (GSFC) PACE Mission, herein referred to as the thruster.

The propulsion system includes eight (8) thrusters, four (4) primary and four (4) redundant, for correcting launch dispersions, stationkeeping, attitude control, momentum unloading, and safe mission disposal at end of life. All eight thrusters will have straight nozzles, and one spare thruster will be required.

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

2.0 APPLICABLE DOCUMENTS

The following documents and drawings in effect on the day this specification was signed shall apply to the fabrication and to the electrical, mechanical, and environmental requirements of the thruster 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 PACE thruster

Statement of Work (PACE-PROP-SOW-0048), in which case the Statement of Work takes precedence.

The following is a list of the applicable specifications and publications.

Table 2-1 Applicable Documents

Document Number Title

PACE-PROP-SOW-0048 PACE Thruster Statement of Work

PACE-PROP-LIST-0030 PACE Thruster Deliverable Items List and Schedule

(DILS)

GSFC-S-311-P-079 Procurement Specification for Thermofoil Heater

GSFC-S-311-641

Switch, Thermostatic, Bimetallic, SPST, Corrosion

Resistant Steel, Hermetic, Detail Specification for

GSFC-S-311-P-18

Thermistor, (Thermally Sensitive Resistor), Insulated and Uninsulated, Negative Temperature

Coefficient, Specification for

NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety

Requirements

NASA-STD-5001B Structural Design And Test Factors Of Safety For

Spaceflight Hardware

NASA-STD-6016 Standard Materials and Processes Requirements for

Spacecraft

NASA-HDBK-7005 Dynamic Environment Criteria

NASA-STD-7001 Payload Vibroacoustic Test Criteria

GSFC-STD-7000A General Environmental Verification Standard

(GEVS)

NASA-STD-5019A Fracture Control Requirements for Spaceflight

Hardware

NASA-STD-5020 Requirements for Threaded Fastening Systems in

Spaceflight Hardware

MIL-PRF-26536G Performance Specification Propellant, Hydrazine

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

FED-STD-TT-I-735A Isopropyl Alcohol

MIL-PRF-27401G Propellant, Nitrogen, Pressurizing Agent

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

MIL-PRF-27415C Propellant Pressurizing Agent, Argon

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

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

Document Number Title

ASTM-A511

Standard Specification for Seamless Stainless Steel

Mechanical Tubing

AMS-QQ-S-763

Pressure Tubing Specification, Stainless

Steel

AMS-2488D

Anodic Treatment - Titanium and Titanium Alloys

Solution pH 13 or Higher

AWS D17.1

Specification for Fusion Welding for Aerospace

Applications

KSC-KTI-5212D Material Selection List for Plastic Films, Foams, and

Adhesive Tapes

IPC J-STD-001F Requirements for Soldered Electrical and Electronic

Assemblies

IPC J-STD-001FS WAM1 Space Applications Electronic Hardware Addendum

IEST-STD-CC-1246E Product Cleanliness Levels – Applications, Requirements, and Determination

ASTM E-595-07 Standard Test Method for Total Mass Loss and

Collected Volatile Condensable Materials from

Outgassing in a Vacuum Environment

MIL-DTL-5541 Chemical Conversion Coatings on Aluminum and

Aluminum Alloys

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

NASM 33540 Safety Wiring, Safety Cabling, Cotter Pinning, General Practices for

SAE AS567 Safety Cable, Safety Wire, Key Washers, and Cotter

Pins for Propulsion Systems, General Practices for

Use of

541-WI-5330.1.41 Fastener Locking Using Arathane 5753

MSFC-STD-3029A Guidelines for the Selection of Metallic Materials for

Stress Corrosion Cracking Resistance in Sodium

Chloride Environments

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

3.0 CONTRACT DESCRIPTION

3.1 THRUSTER DESCRIPTION

The thruster is a 22N (5-lbf) class engine that is capable of generating thrust by the catalytic decomposition of hydrazine. The major components of the thruster include the thrust chamber, catalyst bed, nozzle, thruster valves, mounting bracket, catalyst bed heaters, thruster valve heaters, thruster valve thermostat, and thermal sensors.

This specification applies to the thrusters, the associated GSE, and the paperwork for the PACE mission. Please refer to the Statement of Work (SOW, PACE-PROP-SOW-0048) and the

Deliverable Items List and Schedule (DILS, PACE-PROP-LIST-0030) for a complete list of all the hardware, verification testing, test reports, design reports, and analyses reports that should be delivered with the thrusters.

Included in this specification are specific requirements for the thrust and specific impulse, under a variety of conditions and duty cycles, throughout the mission life, using the specified propellant, high-purity grade hydrazine per MIL-PRF-26536G. Other principle requirements include the ability to have knowledge of the angle and repeatability of the thrust vector, the mass and power allocations, the inclusion of a filter, and the inclusion of heaters and thermostats on the valve. Also specified are the expected thermal and vibration environments, as well as the electrical inputs throughout which the thruster must operate and survive.

3.2 GROUND SUPPORT EQUIPMENT DESCRIPTION

The following ground support equipment (GSE) will be provided:

Nine (9) red protective covers to prevent incidental damage to the thrust chamber and nozzle during I&T.

An alignment fixture for measuring the alignment of the thruster’s nozzle on the spacecraft. The fixture will make use of mirrors for optical measurement of the nozzle orientation to within ±0.1°. Eight (8) nozzle plugs will be used for alignment.

Two (2) nozzle plugs (with AN interface) will be used for valve testing purposes that seals against the nozzle such that valve leakage and flow testing may be performed.

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

4.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the thruster(s) as defined in

Section 3.1.

4.1 THRUSTER FLIGHT UNIT FUNCTIONAL/PERFORMANCE

REQUIREMENTS

4.1.1 Thrust

a) The thruster(s) shall be capable of operating in a blowdown configuration with a maximum inlet pressure of 400 psia (27.58 bar) at beginning of life (BOL), and 85 psia

(5.86 bar) at end of life (EOL).

b) The nominal thrust for each thruster at an inlet pressure of 400 psia (27.58 bar) shall be less than 7.87 lbf (35 N).

c) The nominal thrust for each thruster at an inlet of 85 psia (5.86 bar) shall be at least 2.25 lbf (10 N).

d) The contractor shall provide thrust data for both steady-state and pulse mode operation versus inlet pressure for nominal and ±3σ performance, including EOL.

e) If the statistical ±3σ cannot be provided, the contractor shall demonstrate by analysis or test what the statistical confidence is in the nominal steady state performance.

4.1.2 Thrust Repeatability

Steady-state thrust and specific impulse for each thruster shall be predictable within 7%. For each thruster, acceptance test data are to be provided to allow the prediction of variations in thrust and specific impulse as a function of inlet pressure and temperature.

4.1.3 Thrust Alignment

a) The thruster thrust vector shall be perpendicular to the thruster nozzle exit plane to within

±0.25°, before and after environment testing.

b) The position of the nozzle centerline as built shall be within ± 0.050 inches (1.3 mm) of the position as designed.

4.1.4 Specific Impulse

a) The minimum vacuum steady-state specific impulse provided by the thruster shall be as shown in Table 4-1 for the corresponding inlet pressures.

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

The contractor will provide specific impulse versus inlet pressure curves for nominal and worst case performance.

b) If the statistical -3σ performance cannot be provided, the contractor shall demonstrate by analysis or test what the statistical confidence is in the nominal steady-state performance.

Table 4-1 -3σ Worst Case Steady-State Specific Impulse

Inlet Pressure (psia, bar) -3σ Worst Case Isp (s)

85, 5.86 217.5

100, 6.9 218.0

150, 10.3 220.5

200, 13.8 221.5

250, 17.2 222.5

300, 20.7 223.0

350, 24.1 223.5

400, 27.6 223.5

4.1.5 Pulsed Specific Impulse

The average pulsed specific impulse of all of the thrusters shall be greater than 200 seconds at duty cycles greater than 60% post a number of warm-up pulses. The number of warm-up pulses will be agreed upon by the NASA GSFC COR and the vendor.

4.1.6 Impulse Bit

a) The thruster shall be capable of operating in a pulsed mode with valve command pulse durations of 20 milliseconds (ms). The minimum impulse bit is defined as the area under the thrust-time curve for a given valve open pulse width.

b) The impulse bit shall be repeatable to +/- 5% for any continuous firing duration of 1 second or greater with similar initial conditions.

c) For any firing duration less than 1 second, the impulse bit shall be repeatable to +/- 25%.

4.1.7 Propellant Throughput

All thrusters shall be qualified for a minimum throughput capability of 259.0 lbs (117.5 kg).

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4.1.8 Thruster Duty Cycle

Thruster operation shall not have any duty cycle restrictions for on times between 20 ms and the firing duration defined in Section 4.1.15 at all inlet pressures and thermal environments specified herein.

4.1.9 Catalyst Bed Heater

The catalyst bed heater shall be designed to achieve operational catalyst bed temperature within

45 minutes of enabling from the worst-case minimum temperature.

4.1.10 Inlet Pressure

a) The thruster shall operate at any expected operating inlet pressure from BOL 400 psia

(27.6 bar) to EOL 85 psia (5.86 bar).

b) The maximum exposure to the thrusters in a non-operating condition with valves closed shall be 600 psia (41.4 bar).

4.1.11 Propellant Temperature

The thruster shall operate at any propellant temperature between 12 and 40°C.

4.1.12 Propellant

The thruster shall be qualified to operate with high purity grade hydrazine per MIL-PRF-

26536G.

4.1.13 Pressurant Gas

The pressurant gas shall be gaseous helium per MIL-PRF-27407D.

4.1.14 Leakage

4.1.14.1 External Leakage

Each thruster shall have an external leak rate of less than or equal to 1 x 10^-6 scc/s GHe at the

MEOP of 400 psia (27.6 bar), with valves open and nozzle plugged.

4.1.14.2 Internal Leakage

All thruster valve seats shall have an internal leak rate of less than or equal to 5 scc/h GHe at test differential pressures of 25 psid (1.72 bar), 400 psid (27.6 bar), and the maximum differential pressure of 600 psid (41.4 bar), 400 psid MEOP + 200 psid latch valve back pressure relief.

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4.1.14.3 Internal Leakage During Vibration

a) The contractor shall provide data on the internal leak rate during vibration testing.

b) All thruster valve seats shall have an internal leak rate of less than or equal to 5 scc/h

GHe at a test differential pressure of 600 psid (41.4 bar) when exposed to the vibration environment defined in Section 6.4. Measurement of the leak rate during vibration testing is not required if compliance can be demonstrated with prior test data and if approval is received from the NASA/GSFC COR.

4.1.15 Firing Duration

The thruster shall be capable of continuous steady-state firing durations of up to and including

2700 seconds.

4.1.16 Roughness

a) Chamber pressure oscillations shall not cause any damage to thruster or performance shortfalls during all modes of operation.

b) The thruster shall not exhibit any occurrence of “spiking” (rapid isolated increases of chamber pressure caused by collection and detonation of undecomposed propellant) at any condition of feed pressure, temperature, etc. specified herein.

4.1.17 Proof and Burst Pressure

4.1.17.1 Thrusters

a) The thrusters shall have a proof pressure capability of not less than 1,500 psia (103.4 bar) and a burst pressure of not less than 2,200 psia (151.7 bar) with the valves open.

b) The thruster valves, when closed, shall be capable of withstanding a proof pressure of

1,500 psia (103.4 bar).

c) The contractor shall demonstrate through a qualification program that each valve has a burst pressure of at least 2,200 psia (151.7 bar).

4.1.17.2 Propellant Inlet Tubes

a) The interface tubes shall be capable of withstanding a proof pressure of at least 900 psia

(62 bar) (1.5x MEOP).

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b) The contractor shall demonstrate through a qualification program that each valve’s interface tubes have burst pressures of 2,400 psia (165.5 bar) (4x maximum exposure pressure) at a minimum.

4.1.18 Cycles

4.1.18.1 Thermal Cycles

a) All thrusters shall be qualified for two times the number of nominal expected thermal cycles required for the PACE mission (160 cycles). A single thermal cycle is defined as a temperature excursion from at or below the thruster’s minimum pre-heating temperature to the nominal firing operation temperature and back down to the minimum operating temperature.

b) The catalyst bed heater design shall be qualified to 160 cycles (TBR).

4.1.18.2 Valve Cycles

All thruster valves shall be qualified for 40,000 cycles (TBR).

4.1.19 Cold Starts

a) The contractor shall provide data on the number of demonstrated cold starts and the thrust degradation resulting from cold starts.

b) The cold start catalyst bed temperature shall be defined by the contractor and provided as part of the cold start data.

4.1.20 Thruster Sub-Components

4.1.20.1 Inlet Filter

a) The thruster shall incorporate a filter to protect the valve seats from contamination.

b) The filter shall have an absolute rating of 25 micrometers or larger.

c) The filter capacity shall be at least 0.4 mg of AC Coarse Test Dust (ACCTD). This allows sufficient capacity to not reduce thruster performance.

4.1.20.2 Thruster Valves

a) All thrusters shall have a dual-seat (series-redundant) propellant valve, with each seat able to be independently verified for leakage.

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

b) Each seat shall be electrically operated independently and normally closed when unpowered.

4.1.20.3 Thruster Valve Heaters

a) There shall be one dual-element heater per thruster (per GSFC specification S-311-P-079).

This allows for series-redundant thermostatic control (Operational, and Survival), as shown in Figure 8-1 in section 8.2.2.2.

b) The thruster valve heater element shall consume, at most, 4 W of power at 27 VDC. This corresponds to a maximum of 6.35 W at 34 VDC.

c) Each valve heater watt density shall be less than 1.5 W/cm2 @ 34 VDC to prevent local overheating.

d) The heaters shall be sized for the worst-case cold thermal condition and shall have a maximum duty cycle of 70% at 27 V, according to thruster thermal environments in section 6.11.

e) All thermal hardware shall be installed prior to acceptance testing.

4.1.20.4 Thruster Valve Thermostats

a) There shall be four thermostats mounted on the thruster valves per GSFC specification S-

311-641. This setup (two operational and two survival thermostats) provides series-redundant thermostatic control, as shown in Figure 8-1in section 8.2.2.2.

b) The set points of the thermostats shall be per Table 4-3, where the tolerance is ±2°C (which is allowed by the specification for non-standard tolerances).

Table 4-2: Thermostat Set points

ON Set point (°C) OFF Set point (°C)

Operational 16±2 24±2

Survival 10±2 19±2

4.1.20.5 Thruster Valve Thermistors

a) There shall be one (1) thermistor per thruster valve, at a minimum, that has a resistance of

2252 Ω at 25°C.

b) They shall be qualified IAW GSFC specification S-311-P-18.

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4.1.20.6 Catalyst Bed Heaters

a) The thrusters shall include heaters to preheat the catalyst bed prior to thruster operation.

b) The heaters shall have prime and redundant circuits.

c) Each heater shall be operated independently.

d) Each catalyst bed shall have a primary and redundant heater with a power of less than 6.5

W, per element, at 27 VDC.

4.1.20.7 Catalyst Bed PRT

a) The temperature sensor for the catalyst bed shall have a resistance of 2000Ω at 0°C or

500Ω at 0°C and comply with Wire Specification listed in Section 8.2.2.6.1of this document.

b) There shall be one PRTs per thruster catalyst bed.

4.1.21 Gas Ingestion

The thruster must be capable of ingesting 0.5 g of GHe at 400 psia (24.1 bar) before operation without any harm to the catalyst bed, the thruster, or any degradation in performance during subsequent normal firing.

4.2 RESOURCE ALLOCATIONS

4.2.1 Mass Allocation

a) Each thruster shall have a mass of less than or equal to 1.76 lbm (0.8 kg) including harness mass.

b) The harness length for the thruster shall be at least 3.28 ft (1 m).

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4.2.3 Peak Power Allocation

All thruster hardware shall be able to operate at the peak power values listed below for the full 90 minutes required to preheat the catalyst bed then operate the thrusters for the longest steady state fire, per Sections 4.1.9, 4.1.15

a) All thruster valves shall be designed with a maximum current draw of 1.3 A over the specified operating voltage range in section 4.3.1.

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

b) The maximum current draw shall be the total for both valve coils wired in parallel.

c) The thruster valve power shall have a peak (instantaneous) power consumption of less than or equal to 24.5 W at 28 VDC.

d) The catalyst bed heater(s) for the thruster shall have a peak (instantaneous) power consumption of less than or equal to 13.2 W at 28 VDC.

e) The valve heater for the thruster shall have a peak (instantaneous) power consumption of a total of 4 W, at 28 VDC.

4.3 POWER

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4.3.1.1 Operating Voltage Range

The thruster shall be designed to operate over the bus voltage range of +24 to +35 VDC at their primary power inputs.

The thruster shall meet all performance requirements within this range.

4.3.1.2 Abnormal Voltages

The thruster shall survive without performance degradation after indefinite exposure to an anomalous voltage range of 0 to +40 VDC for at least 500ms. This requirement is to be verified by analysis or test of a non-flight unit.

4.4 ELECTRICAL GROUNDING

4.4.1 Primary Power DC Isolation from Component Chassis

At the thruster’s primary power interfaces, primary power and primary power returns shall be isolated from signal grounds and from the component chassis by a DC resistance of greater than or equal to 1 MΩ.

4.4.2 Internally Generated Secondary Returns

Secondary returns (power and signal grounds) shall be referenced to the component chassis ground by connecting them at one or more places.

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4.4.3 Internally Generated Secondary to Primary DC Isolation

Secondary power (or signal) inputs shall be isolated from primary power by a DC resistance of greater than 1 MΩ.

4.4.4 Mechanical Contact Resistance

The DC resistance of the mechanical contact between two conductive mating surfaces (internal to the component) shall be less than or equal to 2.5 mΩ DC resistance.

4.4.5 Mating Method

The primary mating method for a PACE component is metal-to-metal contact between component mounting feet (or base plate) and the spacecraft structure. When the use of this method is not possible as determined by the Interface Control Document (ICD), the use of a ground strap is necessary.

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4.5.2.3 Passive Linearizing Resistor

The linearizing resistor for the thermistors, if required, will be located with the temperature sensor conditioning circuit in the Spacecraft avionics.

4.5.2.4 Thermistor Resistor Conditioning

Thermistor resistor temperature sensors will be conditioned to provide a voltage range which represents a temperature range that includes a margin of greater than or equal to 5 degrees C beyond the unit qualification or non-operational survival limit, whichever is worse.

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4.5.5.3 Wire Pairing

Each positive wire shall be twisted with the associated return wire. All outputs should be kept as close as possible to their return.

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5.0 PHYSICAL REQUIREMENTS

5.1 INTERFACE DOCUMENTATION

The contractor shall use metric units when interfacing with NASA GSFC including any drawings, documents, models, except for the following cases:

Heritage Component: Components that have been previously qualified, or of similar design heritage, may be specified in English units where use of metric equivalents would lead to additional cost to the program.

Fasteners: Although bolt patterns will be defined using metric dimensioning, use of

English fasteners (with hole dimensioning and tolerancing) is permitted.

Angular Measurement: Angular measurement may be expressed in degree of arc or in an appropriate subdivision of degree of arc such as second of arc (arc-sec) when advantageous to application.

5.1.1 Interface Drawings

An Interface Control Drawing (ICD) will be developed between the contractor and NASA GSFC which shall define both the mounting interface and the electrical interface.

5.2 MASS PROPERTIES

5.2.1 Component Masses

See Section 4.2.1 for total mass.

The mass of the thruster shall be measured to within 50 g.

5.2.2 Center of Mass Location

The contractor shall define the center of mass in the ICD.

5.2.3 Center of Mass Accuracy

The center of mass shall be calculated using a high fidelity CAD model and labeled in the thruster interface drawing.

5.3 PHYSICAL ENVELOPE

a) The thruster outside dimensions, including mounting flanges and connectors, shall not exceed the volume envelope of 7.87in x 3.93in x 3.54in (200 mm length X 100 mm x 90

mm) as defined in the ICD.

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

b) This envelope shall not apply to the interface tube.

5.4 MOUNTING

a) The thruster shall interface with the flight system via tubing, according to Figure 5-1.

b) The length of the thruster from the base of the spacecraft mounting interface to the end of the nozzle shall be less than 4.0 in (100 mm).

c) Mounting interface flatness, and co-planarity requirements for the component side of the interface (including brackets, if any, and shims) shall be as defined in the ICD.

Figure 5-1: Thruster Tube Interface

5.4.1 Mounting Flange

The thruster shall incorporate a mounting flange to attach the engine to the spacecraft mounting bracket, as per the ICD provided by the contractor and approved by NASA/GSFC.

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5.6 ALIGNMENT

5.6.1 Alignment Location/Orientation Accuracy

See section 4.1.3 for thrust alignment.

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5.6.5 Alignment Measurement Tooling/Features

a) The interchangeable alignment plugs shall be made such that they can be used for any of the delivered thrusters.

b) The optical reference surfaces shall be a cube of at least 0.6 in3 (9800 mm3).

c) Optical reference surfaces shall be capable of producing an autocollimation indication.

5.7 FLUID INTERFACES

a) The interface tubes shall be 304L stainless steel per ASTM-A511 or AMS-QQ-S-763.

b) The interface tubes shall have the following dimensions: 0.25 inches ± 0.004 inches (6.35 mm ±0.1016 mm) OD, 0.028 inches ±0.004 inches (0.7112 mm ±0.1016 mm) wall thickness, as shown in Figure 5-1.

c) The interface tube shall be perpendicular to the centerline of the thruster valve body.

d) The interface tube shall have a minimum length of 2 inches (50 mm) from the thruster body centerline.

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6.0 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements for the spacecraft components are specified in this section.

The thrusters shall meet their performance requirements in Section 4.0 during and after exposure to the environments specified in this section.

6.1 MECHANICAL FACTORS OF SAFETY

a) The thruster as well as Mechanical Ground Support Equipment (MGSE) shall demonstrate positive Margins of Safety under limit loads for all yield and ultimate failures using the Factors of Safety (FS) defined in Table 6-1. Margin of Safety (MS) is defined as follows:

MS = (Allowable Stress (or Load) / (Applied Limit Stress (or Load) x FS)) -1

Table 6-1 Factors of Safety

b) Primary and secondary structure comprised of composite materials, Beryllium, bonded joints and/or bonded inserts shall be proof tested to 1.25 x Limit Load; qualification by analysis only is not acceptable. Actual flight component testing is preferred, but testing of

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representative sets of hardware with a similarity qualification argument can be used if approved by the NASA/GSFC COR.

6.2 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 thruster shall be designed to withstand the quasi-static design limit loads defined in the mass-acceleration curve (MAC) shown in Table 6-2 without damage or degradation of performance. The design loads shown below will be updated based on the results of coupled loads analysis.

Linear interpolation should be used between breakpoints to determine the appropriate limit load as a function of thruster weight. Note that these design limit loads are intended to cover only the low frequency launch environment and must be used in conjunction with the random vibration environments to assess structural margins.

Table 6-2 Thruster Design Limit Loads

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6.3 FREQUENCY REQUIREMENT

6.3.1 Fundamental Launch Frequencies

The thruster shall have a fundamental frequency greater than 100 Hz when hard mounted at its spacecraft interface. Any component, which fails to meet the specified fundamental frequency, must supply a finite element model, correlated to modal survey test results up to 50 Hz, to be used in coupled loads analyses. Requirements for the submitted finite element model are in the

SOW and discussed in Section 11.5.4.

6.4 VIBRATION

6.4.1 Sinusoidal Vibration

The thruster shall undergo qualification, protoflight or acceptance (level depends on qualification status of unit) sine vibration testing on all three axes at the levels shown in Table

6-3. Stiff components showing a first mode greater than 150Hz can be exempted from sine vibration testing upon approval by the NASA/GSFC COR. A generic sine vibration specification is provided for Protoflight (PFT), Qualification (QT) and Acceptance (AT) levels applied at the PACE to thruster interface. See Section 11.4.1 for definitions of Protoflight, Qualification, and Acceptance.

Table 6-3 Thruster Sine Vibration Environment

Frequency Protoflight, Qual Level Flight, Acceptance Level

5-20 Hz 0.63 in (double amplitude) 0.50 in (double amplitude)

20-100 Hz 12.5 G 10.0 G

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Figure 6-1 Sine Vibration Environment

Levels may be notched to not exceed 1.25 times the design limit load outlined in section 6.2.

Peak levels at the low end of the frequency range (5 – 20 Hz typically) may be ramped up as needed to accommodate shaker table displacement limitations.

6.4.2 Random Vibration

a) The thruster shall demonstrate its ability to meet its performance requirements after being subjected to the random vibration environment in Table 6-4, for components weighing

22.7 kg (50 lb) or less, applied at the Spacecraft to thruster interface.

Table 6-4 Thruster Random Vibration Environment, (22.7 kg, or less)

Frequency ASD Level (g2/Hz)

(Hz) Qualification Acceptance

20-50

50-800

800-2000

0.026

+6 dB/oct

0.16

-6 dB/oct

0.026

0.013

+6 dB/oct

0.08

-6 dB/oct

0.013

Overall 14.1 Grms 10.0 Grms

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This environment will be updated with random vibration analysis. Note for lightweight thruster, the highest design loads may be from this random vibration environment.

b) The contractor shall provide a random vibration analysis along with static loads analysis.

Please see NASA-HDBK-7005 and NASA-STD-7001 for more information.

During the test, the test input level will be reduced (notched) at critical frequencies, if required, to limit the random vibration loads and/or acceleration responses to 3 dB above design limit levels.

c) Notching shall be limited to -12 dB of the original input and to a bandwidth of less than

100 Hz to limit the random vibration responses to 3dB above design limit levels.

Notching beyond these limits will require NASA/GSFC COR approval.

6.5 SHOCK

a) The thruster shall be designed to meet its performance requirements after being subjected to the shock environment in Table 6-5, applied at the thruster interface to the PACE spacecraft structure.

Table 6-5 Qualification Level Shock Response Spectrum

Freq (Hz) SRS (G)

100 100

800 750

8000 750

10000 1000

b) A shock susceptibility and attenuation assessment shall be performed on all Spacecraft components. If the flight shock environment as shown on a Shock Response Spectra

(SRS) plot (Q=10) is enveloped by the curve shown below (Figure 6-2), then the shock environment can be considered benign and there is low risk in deferring the shock test to the Observatory level.

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Figure 6-2 Shock Envelope

c) Analysis supporting this conclusion (i.e. deferral of shock testing) shall be provided to the NASA/GSFC COR for review.

d) Any component determined to be susceptible to the shock environment (e.g., shock levels are above the curve) shall have shock testing performed at the component level

(preferably on a qualification unit).

e) Component self-induced shock testing shall be accomplished by two actuations at the component level for each self-induced shock source (in order to account for the scatter associated with the actuation of the device) for the first flight unit, and a single actuation on subsequent units.

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6.7 TRANSPORTATION

In addition to the launch loads shown above, the thruster shall also be designed to withstand the maximum transportation loads shown in Table 6-6 without damage or degradation of performance.

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Table 6-6 Transportation Loads

6.7.1 Transportation Cleanliness

Materials and enclosure used for transportation and storage shall not generate molecular or particle contaminants or degrade the surface cleanliness of the item or adjacent items

6.8 PRESSURE

6.8.1 Operating Pressure Range

The thruster shall be designed to meet all performance requirements while operating over a pressure range of 1.08 x 105 N/m2 (813 Torr) to 1.3 x 10-12 N/m2 (1 x 10-14 Torr).

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6.8.3 Launch Vehicle Environmental Control System Impingement Velocity

The thruster exterior surfaces shall not suffer damage or degradation when exposed to the

Launch Vehicle (LV) Environmental Control Systems (ECS) airflow velocity of 10 m/sec.

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6.10 GROUND ENVIRONMENTS

The thruster shall meet all of their performance requirements during exposure to air temperature between +5 and +40 degrees C and relative humidity between 30% and 70%.

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6.11 THERMAL REQUIREMENTS

6.11.1 Flight Interface Design Temperature Limits

The thruster shall be capable of operating in vacuum at the thruster/Spacecraft interface with temperatures ranging from -20°C to +50°C.

6.11.2 Valve Temperature Limits

The thruster shall be capable of operating when its valve temperatures are within the temperature limits shown in Error! Reference source not found. without damage or permanent performance degradation.

Table 6-7: Valve Temperature Levels

Minimum Temperature (ºC) Maximum Temperature (ºC)

Valve Open +10 +40

Valve Closed +10 +100

Qualification +4 +100

6.11.3 Thermal Interfaces

NASA/GSFC COR will provide representative thermal environments for the contractor to incorporate with their thermal model of the thruster assembly at the Kick-off Meeting discussed in Section 3.4.1 of the SOW. This shall include both Radiative and Conductive interfaces. The models representative of the thruster thermal environment will be provided in Thermal Desktop format. Alternative formats may be proposed by the contractor if necessary.

6.11.3.1 Conduction

The maximum allowable conduction from the thrusters to the spacecraft shall be provided to the contractor by the NASA/GSFC COR at the Kick-off Meeting.

6.11.3.2 Radiation

The maximum allowable radiation from the thrusters to the spacecraft shall be provided to the contractor by the NASA/GSFC COR at the Kick-off Meeting. The spacecraft blanket temperatures in the Radiative Field of View of the thruster assembly shall be maintained below

250°C.

The contractor shall provide analysis and documentation demonstrating compliance with the blanket requirement stated above. The inclusion or exclusion of a heat shield shall be the contractor’s responsibility.

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6.13 ATOMIC OXYGEN FLUENCE

When a component has surfaces that will be exposed to the external space environment, all external materials shall survive an atomic oxygen fluence (Observatory orbit velocity direction) of 2.0E+20 atoms/cm2 without loss of structural integrity or loss of critical performance criteria.

6.13.1 Atomic Oxygen Analysis

An analysis shall be performed for all external materials and finishes to verify the compatibility with the AO environment. The analysis should show that any AO degradation does not pose a contamination hazard for other components on the spacecraft (source of particles, molecular films, debris, etc.).

6.13.2 Atomic Oxygen Testing

If no data exists for the proposed material and finishes, a test shall be performed, exposing a representative sample to AO and verifying no loss of structural integrity or loss of critical performance criteria.

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7.0 CLEANLINESS

The requirements in this section ensure the cleanliness of the thruster at delivery, so as not to adversely affect its own performance, as well as not be a source of contamination to other items, including not generating contaminants following delivery in excess of that permitted below by virtue of its design, materials of construction, or operation.

7.1 SURFACE CONTAMINATION

7.1.1 Surface Contamination Levels at Delivery

7.1.1.1 Particulate Contamination

The thruster shall meet IEST-STD-CC1246E VC-0.5-1000 + UV, or equivalent, when inspected with both UV and white light in a darkened room.

7.1.1.2 Molecular Contamination – Exposed Surfaces

If the thruster has surfaces exposed to the space environment, all such surfaces shall meet a molecular surface cleanliness level of IEST-STD-CC1246E R3.3E-1 on all external and critical surfaces, when tested in accordance with IEST-STD-CC1246E VC-0.5-1000 + UV, or equivalent.

7.1.1.3 Molecular Contamination – Covered Surfaces

All thruster surfaces not exposed to the space environment shall meet IEST-STD-CC1246E VC-

0.5-1000 + UV, or equivalent, when inspected in a darkened room with white and UV light. The surfaces shall be free of molecular contamination – for example films, spots, or other.

7.1.1.4 External Cleanliness

All hardware surfaces shall be verified to be “visibly clean, highly sensitive” per JSC-SN-C-

0005D, prior to delivery to NASA/GSFC. This is accomplished by maintaining all hardware in a double bag for storage.

7.1.1.5 Internal Cleanliness

The propulsion components, lines, and fittings shall be cleaned, and verified internally clean, to level 100R1 per IEST-STD-CC1246E prior to integration in the system.

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7.1.2 Surface Contamination Generation

7.1.2.1 Particulate Generation

The contractor shall not employ any of the following particle generating materials or processes into the thruster design or construction without prior approval by the NASA/GSFC COR:

Paints prone to shedding due to large paint pigment molecules, overspray, poor adhesion, etc.

Dry lubricants (e.g. molybdenum disulfide).

Surfaces prone to corrosion or oxides because of a lack of corrosion protection or dissimilar metals in close contact.

Fabrics with brittle constituents (e.g., composites, graphite or glass).

Perforated materials when material is highly susceptible to tear propagation (e.g., multi-layer insulation (MLI)).

Metal oxides (bare [untreated] aluminum and magnesium, iron, non-corrosion resistant steel, etc.).

Braided metallic or synthetic wires, ropes, slings, etc. unless measures have been taken to contain any broken filaments or fibers (sheathing, sealing with polymers, covering, etc.).

Woven materials especially cut or unfinished ends (metal braid, EMI shielding, lacing cord, expando sleeving), unless measures have been taken to prevent fraying or generation of particles (cut with a hot knife, seal with polymer, bag, etc.).

Materials with thin films known to erode or crack or flake when subjected to normal handling (e.g., indium tin oxide [ITO] or other rigid or brittle semiconductor or ceramic coating on flexible substrates, Teflon, MLI, etc.).

Foams, highly textured materials.

Trapped debris in holes.

7.1.2.2 Molecular Generation

7.1.2.2.1 Material Selection

The thruster’s materials shall have a total mass loss (TML) less than 1.00%…