Attachment B RST-PROP-SPEC-0121-.pdf

Effective Date: August 31, 2020

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

RST-PROP-SPEC-0121, Revision -

Roman Space Telescope (RST), Code 448

Attitude Control System (ACS) Thruster Specification

GSFC RST CMO

August 31, 2020

Released

RST ACS Thruster Spec RST-PROP-SPEC-0121, Revision – ii

ACS Thruster Specification

Review/Signature/Approval Page

Prepared by:

Alison Rao

Approved by:

Jason Hylan

Electronic Approval available on-line in the RST CM Tool file://gs448share/WFIRST/01.0%20Project%20Management/1.4%20Configuration-Data%20Mgmt/Templates/Document%20Templates/TDMS/Templates%20(boilerplates)/Historical/in ii

Preface This document is a Roman Space Telescope (RST) Configuration Management (CM)-controlled document.

Note: Prior to May 20, 2020, the project name was Wide Field Infrared Survey Telescope

(WFIRST).

For the purposes of configuration management, the prefixes “WFIRST” and “RST” are completely interchangeable. For example, RST-MGMT-PROC-0024 is the same as WFIRST-

MGMT-PROC-0024.

Changes to this document require prior approval of the applicable Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the RST CM Office (CMO), along with supportive material justifying the proposed change.

In this document, a requirement is identified by “shall,” a good practice by “should,” permission by “may” or “can,” expectation by “will,” and descriptive material by “is.”

Questions or comments concerning this document should be addressed to:

RST Configuration Management Office Mail Stop 448 Goddard Space Flight Center Greenbelt, Maryland 20771 iii

Change History Log

Revision Effective Date Description of Changes (Reference the CCR & CCB/ERB Approval Date)

Revision - August 28, 2020 Initial Release per RST-CCR-0244 iv

Table of TBDs/TBRs/TBSs

Item No. Location Summary Individual/ Organization

Actionee

Due Date v

Table of Contents

1 INTRODUCTION

1.1 Purpose

1.2 Scope

1.3 Related Documentation

1.3.1 Applicable Documents and Forms

2 CONTRACT DESCRIPTION

2.1 ACS Thruster Description

2.2 Ground Support Equipment Description

3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

3.1 ACS Thruster Configuration Requirements

3.1.1 Major Components List

3.1.2 Nozzle Configuration

3.1.3 Valve Configuration

3.1.3.1 Redundancy

3.1.3.2 Independent Operation

3.1.3.3 Thruster Valve Fail-Safe

3.1.3.4 Inlet Filter Rating

3.1.4 Interface Tube Location

3.1.5 Valve Heater Configuration

3.1.6 Valve Thermostat Configuration

3.1.7 Valve Thermal Sensor Configuration

3.1.8 Catalyst Bed Heater Configuration

3.1.9 Catalyst Bed Thermal Sensor Configuration

3.2 ACS Thruster Functional Requirements

3.2.1 Pressure Requirements

3.2.1.1 Maximum Expected Operating Pressure/Exposure Pressure

3.2.1.2 Proof Pressure

Proof Pressure with Valves Open Proof Pressure with Valves Closed

3.2.1.3 Burst Pressure

Burst Pressure with Valves Open Burst Pressure with Valves Closed

3.2.1.4 Surge Pressure

3.2.1.5 Interface Tube Pressure Requirements

Interface Tube MEOP/Exposure Pressure Interface Tube Proof Pressure Interface Tube Burst Pressure

3.2.2 Leakage Requirements

3.2.2.1 Internal Leakage

3.2.2.2 Internal Leakage during Random Vibration

3.2.2.3 External Leakage

3.3 ACS Thruster Performance Requirements

3.3.1 Thrust

vi

3.3.1.1 Thrust at 340 psia Inlet Pressure

3.3.1.2 Thrust at 100 psia Inlet Pressure

3.3.1.3 Thrust Repeatability

3.3.1.4 Roughness

Spiking Reporting of Thruster Limitations

3.3.1.5 Steady-State Thrust Variability

3.3.1.6 Thrust Degradation

3.3.2 Inlet Pressure

3.3.3 Specific Impulse

3.3.4 Steady-State Firing Duration

3.3.5 Pulsed Performance Requirements

3.3.5.1 Duty Cycles

3.3.5.2 Minimum Pulse Duration

3.3.5.3 Minimum Impulse Bit

3.3.5.4 Ibit Repeatability, > 1 Second

3.3.5.5 Ibit Repeatability, < 1 Second

3.3.6 Throughput

3.3.7 Cold Start Capability

3.3.8 Propellant and Pressurant Requirements

3.3.8.1 Propellant Type

3.3.8.2 Pressurant Type

3.3.8.3 Helium-Saturated Propellants

3.3.8.4 Propellant Temperature

3.3.8.5 Gas Ingestion

3.3.9 Number of Maneuvers

3.3.10 Thruster Warm Pulses

3.3.11 Heat Soakback

3.4 Valve-Specific Requirements

3.4.1 Valve Cycles

3.4.2 Valve Inductance

3.4.3 Valve Heater Sizing

3.5 Catalyst Bed Heater-Specific Requirements

3.5.1 Catalyst Bed Heater Duration to Minimum Temperature

3.5.2 Catalyst Bed Heater Turn-On Cycles

3.6 Power

3.6.1 Power Allocations

3.6.1.1 Power Allocation – Thruster Valve

3.6.1.2 Power Allocation – Catalyst Bed Heater

3.6.2 Primary (Unregulated) Power Input Requirements

3.6.2.1 Operating Voltage Range

3.6.2.2 Abnormal Voltages

3.6.2.3 Sudden Removal of Power

3.6.3 Load Induced Noise Requirements

3.6.3.1 Turn-on Current Transients

3.6.3.2 Operational Current Transients (one time)

vii

3.6.3.3 Low Frequency Repetitive Transients (during nominal operation)

3.7 Electrical Grounding

3.7.1 Primary Power DC Isolation

3.7.2 Mechanical Contact Resistance

3.7.3 Mating Method

4 PHYSICAL REQUIREMENTS

4.1 Use of Metric Units

4.2 Mass Properties

4.2.1 Mass Allocation

4.2.2 Center of Mass Location

4.2.3 Center of Mass Accuracy

4.2.4 Determination of Moments and Products of Inertia

4.3 Physical Envelope

4.4 Mounting

4.5 Alignment

4.5.1 Angular Alignment

4.5.2 Position Alignment

4.5.3 Alignment GSE Size

4.6 Interface Tube Physical Requirements

4.6.1 Interface Tube Material

4.6.2 Interface Tube Length

4.6.3 Interface Tube Outer Diameter

4.6.4 Interface Tube Wall Thickness

5 ENVIRONMENTAL REQUIREMENTS

5.1 Quasi-Static Acceleration

5.2 Frequency Requirement

5.3 Vibration

5.3.1 Sinusoidal Vibration

5.3.2 Random Vibration

5.4 Shock

5.5 Acoustics

5.6 Pressure

5.6.1 Operating External Pressure Range

5.6.2 Maximum Depressurization Rate

5.7 Ground Environments

5.7.1 Thermal Ground Environment

5.7.2 Humidity Ground Environment

5.8 Thermal Requirements

5.8.1 Flight Interface Design Temperature Limits, Operational

5.8.2 Flight Interface Design Temperature Limits, Acceptance

5.8.3 Flight Interface Design Temperature Limits, Qualification

5.8.4 Flight Interface Design Temperature Limits, Survival

6 CLEANLINESS

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

viii

6.1.1.1 External Cleanliness

6.1.1.2 Internal Cleanliness

6.1.2 Surface Contamination Generation

6.1.2.1 Particulate Generation

6.1.2.2 Molecular Generation

Material Selection Material Selection - Silicones

6.1.3 Cleanability

6.1.3.1 Cleanability – Sensitive Surfaces

6.1.3.2 Cleanability – Sensitive Surface Cleaning Methods

6.1.4 Transportation Cleanliness

7 DESIGN & CONSTRUCTION REQUIREMENTS

7.1 Parts, Materials & Processes (PMP)

7.1.1 EEE Parts

7.1.2 Materials

7.1.2.1 Material Conductivity

7.1.3 Material Compatibility

7.1.3.1 Fluid Compatibility

7.1.3.2 Passivation of Wetted Materials

7.2 Electrical

7.2.1 Interface Requirements

7.2.2 Wires

7.2.2.1 Wire Specifications

7.2.2.2 Minimum Wire Size

7.2.2.3 Wire Derating

7.2.2.4 Wire Composition

7.2.2.5 Wire Twists

7.2.2.6 Wire Insulation Resistance

7.3 Thermal Design

7.3.1 Thermal Conduction

7.3.2 Thermal Radiation

7.4 Identification and Marking

7.5 Workmanship

7.5.1 Workmanship Standards

7.5.2 Welding

7.6 Reliability and Mission Lifetime

7.6.1 Mission Life

7.6.2 Shelf Life

7.7 GSE Cleanliness

7.8 Interchangeability

8 MECHANICAL DESIGN REQUIREMENTS

8.1 Mechanical Factors of Safety

8.2 Fracture Control Requirements

8.3 Materials

8.3.1 Dissimilar Metals

8.3.2 Stress Corrosion Cracking

ix

8.3.3 Fastened Joints

8.3.3.1 Fastener Locking

8.3.3.2 Fastened Joint Margin of Safety

8.3.3.3 Installation Torque Documentation

8.3.3.4 Seal Analysis

8.3.3.5 Critical Fasteners

8.4 Mechanism Design

8.4.1 Torque Margin Formula

APPENDIX A ABBREVIATIONS AND ACRONYMS

APPENDIX B THRUSTER DRIVING CIRCUIT SCHEMATIC

List of Figures Figure 3-1 Thruster Tube Interface

List of Tables

Table 5-1 Thruster Design Limit Loads Table 5-2 Sine Vibration Limit Levels Table 5-3 Random Vibration Limit Levels Table 5-4 Minimum Workmanship ASD Level Table 5-5 Shock Limit Levels Table 5-6 Depressurization Profile Table 5-7 Temperature Limits Table 5-8 Propellant-wetted In-flight Software Temperature Limits Table 8-1 Design Factors of Safety Table 8-2 Torque Margin Calculation Factors of Safety

1 INTRODUCTION

1.1 Purpose

The Roman Space Telescope (RST) is a mission responding to the 2010 National Research Council New Worlds, New Horizons (NWNH) Astronomy and Astrophysics Decadal Survey top priority recommendation in the large space mission category. The science program includes two dedicated investigations to tackle outstanding questions in dark energy research and exoplanet exploration, and includes a substantial General Observer program to enable targeted investigations of astrophysical phenomena to advance other goals from the Decadal Survey. A coronagraph instrument is included in the payload for purposes of advancing the present state of the art of coronagraph technology. This document defines the requirements for the ACS thruster design.

1.2 Scope

This specification describes the electrical, mechanical, and environmental requirements for a space-qualified ACS thruster for the NASA Goddard Space Flight Center (GSFC) RST Mission.

1.3 Related Documentation

RST documents can be obtained from the RST CM Tool.

1.3.1 Applicable Documents and Forms

The following documents and drawings in effect on the day this specification was signed (except where noted) apply to the fabrication and to the electrical, mechanical, and environmental requirements of the ACS 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 RST ACS Thruster Statement of Work (RST-PROP-SOW-0042), in which case the Statement of Work takes precedence.

Document Number Title AMS5569 Steel, Corrosion and Heat Resistant, Seamless and Welded

Hydraulic Tubing 19Cr - 9.5Ni - 0.03C max Cold Drawn, 1/8 Hard Temper

AMS5647 Steel, Corrosion-Resistant, Bars, Wire, Forgings, Tubing, and Rings, 19Cr - 9.5Ni, Solution Heat Treated

ASTM C1683 Standard Practice for Size Scaling of Tensile Strengths Using Weibull Statistics for Advanced Ceramics

ASTM E595 Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment

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

IEST-STD-CC1246E Product Cleanliness Levels – Applications, Requirements, and Determination

JSC SPEC C20 Water, High Purity Specification for MIL-PRF-26536E Propellant, Hydrazine MIL-PRF-27401 Propellant Pressurizing Agent, Nitrogen

Document Number Title MIL-PRF-27407 Propellant Pressurizing Agent, Helium MIL-PRF-27415 Propellant Pressurizing Agent, Argon MIL-STD-889C Dissimilar Metals MSFC-STD-3029A MSFC Technical Standard: Guidelines for the Selection of

Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments

NASA-STD-5018 Strength Design and Verification Criteria for Glass, Ceramics, and Windows in Human Space Flight Applications

NASA-STD-5019 Fracture Control Requirements for Spaceflight Hardware NASA-STD-5020A, w/ Change 1

Requirements for Threaded Fastening Systems in Spaceflight Hardware

NASA-STD-6012 Corrosion Protection for Space Flight Hardware NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety

Requirements RST-PROP-SOW-0042 Attitude Control System Thruster Statement of Work S-311-641 NASA GSFC General Requirements for Thermostatic Switches S-311-P-079 NASA GSFC Procurement Specification for Thermofoil Heater S-311-P-18 NASA GSFC Specification for Negative Temperature

Coefficient, Insulated and Uninsulated Thermistor (Thermally Sensitive Resistor)

TT-I-735 Isopropyl Alcohol

2 CONTRACT DESCRIPTION

2.1 ACS Thruster Description

The ACS thruster is a monopropellant hydrazine thruster capable of providing fine thrust for attitude control of the RST spacecraft on orbit. The spacecraft contains a suite of ACS thrusters that are pulsed to provide control about the three spacecraft axes during delta-V maneuvers and to unload spacecraft momentum wheels when needed between maneuvers. The flight units will be individual thrusters consisting of flow control valves, catalyst beds for decomposing hydrazine, a nozzle, and thermal hardware for heating and measuring necessary components. The thruster valves are normally closed valves that require power to open the valve and then return to the closed position in the absence of power.

Each thruster design will either have been qualified for space flight or the vendor will designate one (1) unit as protoflight to test at the qualification levels listed in this document. Thruster designs that have been qualified previously will include qualification documentation as described in the ACS Thruster Statement of Work (SOW), RST-PROP-SOW-0042.

2.2 Ground Support Equipment Description

The ACS thruster procurement includes the following ground support equipment (GSE) to be delivered with the flight hardware.

• Test plugs – This GSE installs into the nozzle and facilitates an AN-type gas connection for ground testing, such as thruster valve leakage tests.

• Alignment hardware – This GSE installs into the thruster nozzle, self-aligning, and facilitates alignment error verification after thruster installation in the higher assembly.

• Nozzle covers – This GSE encompasses the thruster nozzle exit to prevent contamination from entering the nozzle.

RST ACS Thruster Spec RST-PROP-SPEC-0121, Revision -

3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the ACS thruster.

3.1 ACS Thruster Configuration Requirements

3.1.1 Major Components List

The thruster shall include the following components that meet the requirements herein.

• Thruster valve with inlet propellant filter

• Thrust chamber

• Nozzle

• Valve heaters

• Valve thermistors

• Thermostats

• Catalyst bed heaters

• Catalyst bed platinum resistance thermistors (PRT)

3.1.2 Nozzle Configuration

The thruster shall contain a nozzle with a 90-degree cant.

3.1.3 Valve Configuration

3.1.3.1 Redundancy

The thruster shall contain a series-redundant flow control valve.

3.1.3.2 Independent Operation

The upstream and downstream valves shall open and close independently.

3.1.3.3 Thruster Valve Fail-Safe

In the event of a loss of electrical power, the thruster valve shall remain in the closed position if closed and return to the closed position if open.

3.1.3.4 Inlet Filter Rating

The thruster valve shall include an inlet propellant filter with an absolute filtration rating of 30 microns or larger.

3.1.4 Interface Tube Location

The thruster interface tube shall comply with the clocking shown in Figure 3-1. Refer to Section 4.6 for interface tube dimensions.

Figure 3-1 Thruster Tube Interface

3.1.5 Valve Heater Configuration

The thruster shall contain a dual-element valve heater per S-311-P-079 with redundant heater circuits.

3.1.6 Valve Thermostat Configuration

The valve heater circuits shall employ a series-redundant philosophy with two thermostats per S-311-641 wired in series per circuit.

3.1.7 Valve Thermal Sensor Configuration

The thruster shall contain redundant thermistors per S-311-P-18 for sensing valve temperature.

3.1.8 Catalyst Bed Heater Configuration

The thruster shall contain a dual-element catalyst bed heater with redundant heater circuits.

3.1.9 Catalyst Bed Thermal Sensor Configuration

The thruster shall contain redundant PRTs for sensing catalyst bed temperature.

3.2 ACS Thruster Functional Requirements

3.2.1 Pressure Requirements

3.2.1.1 Maximum Expected Operating Pressure/Exposure Pressure The thruster shall have a Maximum Expected Operating Pressure (MEOP) of 400 psia, where the valve will open when input voltage is applied as described herein.

The thruster shall have an Exposure Pressure of 700 psia, where the valve will open when input voltage is applied as described herein. The exposure pressure is the sum of the maximum feed pressure (MEOP) and the back-pressure relief cracking pressure of two levels of latch valves (up to 150 psid each). Note that upon valve opening at this worst-case exposure pressure, the pressure will immediately drop to the operating pressure range described herein, so the thruster does not need to operate up to this exposure pressure. Due to the propulsion system launch configuration, exposure pressure is only a consideration on orbit.

3.2.1.2 Proof Pressure

Proof Pressure with Valves Open

The thruster shall have a proof pressure capability of not less than 1,050 psia with valve open.

Proof Pressure with Valves Closed The thruster shall have a proof pressure capability of not less than 1,050 psia with valve closed.

3.2.1.3 Burst Pressure

Burst Pressure with Valves Open

The thruster shall have a burst pressure of not less than 1,750 psia with valve open.

Burst Pressure with Valves Closed The thruster shall have a burst pressure of not less than 1,750 psia with valve closed.

3.2.1.4 Surge Pressure

The thruster shall withstand a peak surge (water-hammer) pressure of 1,050 psi at the closed thruster valve.

3.2.1.5 Interface Tube Pressure Requirements

Interface Tube MEOP/Exposure Pressure

The thruster interface tube shall have a MEOP of not less than 400 psia.

The thruster interface tube shall have an on-orbit Exposure Pressure of not less than 700 psia.

Interface Tube Proof Pressure The thruster interface tube shall have a proof pressure capability of not less than 1,050 psia.

Interface Tube Burst Pressure The thruster interface tube shall have a burst pressure of not less than 2,800 psia.

3.2.2 Leakage Requirements

3.2.2.1 Internal Leakage

Each thruster valve seat shall maintain an internal leakage of less than 5 standard cubic centimeters per hour (scch) of gaseous helium in the closed position with inlet pressures of both 25 psia and MEOP. Valve leakage is defined as leakage through a single closed valve seat and out the nozzle.

3.2.2.2 Internal Leakage during Random Vibration

Each thruster valve seat shall maintain an internal leakage of less than 5 scch of gaseous helium in the closed position with an inlet pressure of 25 psia while subjected to the acceptance random vibration environment stated herein. This requirement may be satisfied through similarity if applicable qualification data exist.

3.2.2.3 External Leakage

The thruster shall maintain an external leakage of less than 1 x 10^-6 standard cubic centimeters per second (sccs) of gaseous helium at MEOP with both thruster valve seats open and the nozzle plugged.

3.3 ACS Thruster Performance Requirements

3.3.1 Thrust

3.3.1.1 Thrust at 340 psia Inlet Pressure

The thruster shall have a minimum thrust of 4.3 N at an inlet pressure of 340 psia. Thrust below this threshold due to repeatability shall be reviewed for acceptance by the NASA/GSFC Contracting Officer's Representative (COR).

3.3.1.2 Thrust at 100 psia Inlet Pressure

The thruster shall have a minimum thrust of 1.5 N at an inlet pressure of 100 psia.

3.3.1.3 Thrust Repeatability

The thruster shall produce thrust that is predictable to within 7% for a given duty cycle, including steady state.

3.3.1.4 Roughness

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

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

Reporting of Thruster Limitations If there are limitations on thruster operation with respect to roughness, the contractor shall describe the limitations, the technical basis for the limitations, and the risks associated with thruster operation at the limiting conditions.

3.3.1.5 Steady-State Thrust Variability

The steady-state thrust shall vary by no more than +/-2% for constant engine inlet conditions after the initial thermal transient.

3.3.1.6 Thrust Degradation

The thruster shall maintain steady-state thrust degradation of less than 5% throughout the mission life defined by Section 7.6.1, as compared to beginning of mission steady-state thrust.

3.3.2 Inlet Pressure

The thruster shall operate over an inlet pressure range of 80-400 psia.

3.3.3 Specific Impulse

The thruster shall have a minimum steady-state specific impulse of 210 seconds for all operating conditions.

3.3.4 Steady-State Firing Duration

The thruster shall be capable of continuous steady-state firing durations of up to 60 minutes.

3.3.5 Pulsed Performance Requirements

3.3.5.1 Duty Cycles

The thruster shall not have any duty cycle operation restrictions for on times between the minimum on time in Section 3.3.5.2 and the firing duration defined in Section 3.3.4 at all inlet pressures and thermal environments specified herein.

3.3.5.2 Minimum Pulse Duration

The thruster shall be qualified for a minimum firing time of 20 milliseconds.

3.3.5.3 Minimum Impulse Bit

The thruster shall have a minimum impulse bit (Ibit) of less than 0.2 N-s for a pulse of 20 milliseconds for all duty cycles and feed pressures.

3.3.5.4 Ibit Repeatability, > 1 Second

The thruster shall have Ibit performance that is repeatable to +/- 5% for any continuous firing duration of 1 second or greater with similar initial conditions.

3.3.5.5 Ibit Repeatability, < 1 Second

The thruster shall have Ibit performance that is repeatable to +/- 25% for any firing duration less than 1 second with similar initial conditions.

3.3.6 Throughput

The thruster shall have a minimum qualified throughput of 163 kg.

3.3.7 Cold Start Capability

The thruster shall be capable of surviving 3 cold starts with no more than 10% total degradation in thrust. A cold start is defined as starting thruster operation with a catalyst bed temperature of 12 deg-C or lower.

3.3.8 Propellant and Pressurant Requirements

3.3.8.1 Propellant Type

The thruster shall use High-Purity Grade hydrazine per MIL-PRF-26536E (with Amendment 1) as the propellant.

3.3.8.2 Pressurant Type

The thruster shall use helium pressurant gas per MIL-PRF-27407, Type I Grade B.

3.3.8.3 Helium-Saturated Propellants

The thruster shall be capable of meeting all the requirements of this specification when the propellant is fully saturated, partially saturated, or not saturated with helium. Testing may be performed using unsaturated propellant with customer approval.

3.3.8.4 Propellant Temperature

The thruster shall operate at any propellant temperature between 10 and 50 deg-C.

3.3.8.5 Gas Ingestion

The thruster shall be capable of ingesting 0.195 g of gaseous helium at 354 psia at 20 deg-C before operation without any harm to the catalyst bed, the thruster, or any degradation in performance during subsequent normal firings.

3.3.9 Number of Maneuvers

The thruster shall be qualified for a minimum of 2,075 maneuvers. A single maneuver is defined as a temperature excursion from below the thruster’s minimum operational start temperature to the maximum thruster operational temperature and back down below the minimum operational start temperature.

3.3.10 Thruster Warm Pulses

The thruster shall be qualified for 160,000 pulses while at operational temperature.

3.3.11 Heat Soakback

The thruster shall not exceed the maximum allowable temperature of any component in the thruster due to heat soakback for any steady-state or pulse-mode firing.

3.4 Valve-Specific Requirements

3.4.1 Valve Cycles

The thruster valve shall be qualified for 152,606 open/close cycles for each valve seat.

3.4.2 Valve Inductance

The thruster valve shall have a total maximum inductance of 300 mH. This total maximum value is for the flight configuration in which the upstream and downstream valves are wired in parallel (by the customer).

3.4.3 Valve Heater Sizing

The thruster valve heater shall be sized to maintain allowable valve temperatures with a maximum heater duty cycle of 70% at 30 VDC input voltage under worst-case cold thermal conditions, assuming only one heater circuit is active.

3.5 Catalyst Bed Heater-Specific Requirements

3.5.1 Catalyst Bed Heater Duration to Minimum Temperature

The catalyst bed heater shall be designed to achieve the minimum operational start temperature within 45 minutes of enabling from the worst-case minimum temperature with only a single catalyst bed heater active.

3.5.2 Catalyst Bed Heater Turn-On Cycles

The catalyst bed heater shall be qualified to 2,002 turn-on cycles.

3.6 Power

3.6.1 Power Allocations

3.6.1.1 Power Allocation – Thruster Valve

The thruster shall have a nominal (continuous) power consumption of less than or equal to

13.6 W across the operating voltage range in section 3.6.2.1 and the protoflight/qualification temperature range in Table 5-7. This power allocation applies when the upstream and downstream valve circuits are wired in parallel.

3.6.1.2 Power Allocation – Catalyst Bed Heater

The catalyst bed heater shall have a nominal (continuous) power consumption of less than or equal to 10.4 W across the operating voltage range in section 3.6.2.1 and the protoflight/ qualification temperature range in Table 5-7. This power allocation applies to a single catalyst bed heater circuit.

3.6.2 Primary (Unregulated) Power Input Requirements

3.6.2.1 Operating Voltage Range

The thruster components shall be designed to operate over the bus voltage range of +24 to +35 VDC at their primary power inputs during all mission phases and for all expected load conditions (except when turned off).

3.6.2.2 Abnormal Voltages

The thruster shall survive without electrical overstress after exposure to an anomalous voltage range of 0 to +40 VDC and input power short or open. This requirement is to be verified by analysis or testing of a non-flight unit.

3.6.2.3 Sudden Removal of Power

The thruster shall meet its performance requirements without degradation after exposure to an abrupt, unannounced removal of power.

3.6.3 Load Induced Noise Requirements

3.6.3.1 Turn-on Current Transients

The amplitude and duration of in-rush currents, at both initial application of prime power and subsequent component turn-on, shall be per the following:

• <10 microseconds: 10A

• Between 10 microseconds and 10 milliseconds: 3 times of max steady state or 3A whichever is larger

• >10 milliseconds: not to exceed max steady state

The compliance of this requirement can be demonstrated using a solid-state power switch with the similar turn on characteristics as the flight power control switch (optically coupled solid state relay). A schematic of the thruster driving circuit is provided in Appendix B.

3.6.3.2 Operational Current Transients (one time)

a. The thruster operational current transients shall be less than or equal to 5A for a period of less than 1 millisecond.

b. The rate of change of the current shall be less than or equal to 20 mA/millisecond.

3.6.3.3 Low Frequency Repetitive Transients (during nominal operation) Any repetitive current transients, in the frequency range between 1 Hz to 200 Hz, shall not exceed 0.8 A peak-to-peak.

This measurement will be performed with an oscilloscope current probe in time domain.

3.7 Electrical Grounding

3.7.1 Primary Power DC Isolation

At the thruster 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 Mohm.

3.7.2 Mechanical Contact Resistance

The DC resistance of the mechanical contact between two current-carrying conductive mating surfaces used for electronic components shall be less than or equal to 2.5 milliohm.

3.7.3 Mating Method

The mating method for the thruster will use thermal isolating spacers between the thruster mounting plate and the spacecraft structure.

4 PHYSICAL REQUIREMENTS

4.1 Use of Metric Units

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

• SI units with parenthetical English units are permitted for engineering and manufacturing drawings.

• Angular measurement may be expressed in degree of arc or in an appropriate subdivision of degree of arc such as minute or second of arc (arc-min or arc-sec).

• Hardware that has been previously built, or new hardware of similar design heritage, may be specified in English units where use of metric equivalents would lead to additional risk and cost to the program, but interfaces with that hardware must use metric with parenthetical English units.

• Although bolt patterns will be defined using metric dimensioning, use of English fasteners (with hole dimensioning and tolerancing) is permitted.

4.2 Mass Properties

4.2.1 Mass Allocation

The mass of the thruster shall be less than or equal to 0.5 kg.

4.2.2 Center of Mass Location

The contractor shall define the center of mass in the Mechanical Interface Control Document

(MICD).

4.2.3 Center of Mass Accuracy

The center of mass of the thruster shall be determined to within ± 5 mm relative to an external reference.

4.2.4 Determination of Moments and Products of Inertia

The final moments of inertia of the thruster shall be calculated to within 1%. The final products of inertia of the thruster shall be calculated to within 5%.

4.3 Physical Envelope

The thruster external dimensions, including mounting provisions, shall be within an envelope of 170 mm x 86 mm x 51 mm, excluding the interface tube.

4.4 Mounting

The thruster shall accommodate being hard-mounted with mechanical fasteners on a single mechanical surface of the spacecraft structure.

4.5 Alignment

4.5.1 Angular Alignment

The centerline of the nozzle shall be perpendicular with respect to the mounting flange to within ± 0.25 degrees.

4.5.2 Position Alignment

The position of the nozzle centerline with respect to the mounting interface centerline shall be within ± 1.5 mm.

4.5.3 Alignment GSE Size

Optical reference surfaces on the alignment GSE shall be at least 1.52 x 1.52 cm.

4.6 Interface Tube Physical Requirements

4.6.1 Interface Tube Material

The interface tube shall be 304L stainless steel per AMS5647 and AMS5569, or equivalent specifications with customer approval.

4.6.2 Interface Tube Length

The interface tube shall have a length of 1.5 +/- 0.06 inches.

4.6.3 Interface Tube Outer Diameter

The interface tube shall have an outer diameter of 0.25 +/- 0.003 inches along the length in Section 4.6.2.

4.6.4 Interface Tube Wall Thickness

The interface tube shall have a wall thickness of 0.028 +/- 0.0028 inches along the length in Section 4.6.2.

5 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements are specified in this section.

The thruster will meet its performance requirements in Section 3 during and after exposure to the environments specified in this section.

5.1 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 (DLL) defined in the Mass-Acceleration Curve (MAC) shown in Table 5-1 without damage or degradation of performance.

Note:

1. For masses within the breakpoints, linear interpolation is used to determine load.

2. Loads are considered to act in any direction, one axis at a time.

Table 5-1 Thruster Design Limit Loads

Component Mass (kg) Limit Load (g)

0.5 35.9 1 35

5.2 Frequency Requirement

The thruster shall have a fundamental frequency greater than 100 Hz when hard mounted at its interface. Requirements are met assuming rigid stiffness to restrained Degree of Freedom (DOF).

5.3 Vibration

5.3.1 Sinusoidal Vibration

The thruster shall withstand the sinusoidal vibration levels defined in Table 5-2.

Note:

1. Levels defined are acceptance levels. Analysis is conducted against acceptance with the safety factors defined in Table 8-1. Testing factors are defined in RST-PROP-SOW- 0042.

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

3. The sine sweep vibration levels shown in Table 5-2 are defined at the mounting interface of the component.

4. Input levels may be notched to limit the test article’s center of gravity (CG) response to

1.25 times its DLL outlined in Table 5-1.

5. Test verification is required over the range 5 to 50 Hz. Analytical verification is required from 50 to 100 Hz.

6. Components with a first fundamental frequency equal to or greater than 75 Hz can forgo sine vibration testing from 5 to 50 Hz, while verification through analysis of the 50 to 100 Hz is still required.

7. Components with a fundamental frequency equal to or greater than 150 Hz can forgo the sine vibration environment until higher levels of assembly upon approval by the

NASA/GSFC COR.

Table 5-2 Sine Vibration Limit Levels

Frequency (Hz) Limit Level (g)

5 to 100 10.0

5.3.2 Random Vibration

The thruster shall withstand the random vibration environment in Table 5-3 applied at the mounting interface.

The random vibration inputs may be notched to limit the test article’s CG response to 1.25 times DLL. Notched inputs must envelope minimum workmanship levels in order to screen for design or manufacturing flaws. A delicate optic, detector, sensor, etc., which is shown analytically to have negative margin at minimum workmanship levels may notch further at those frequency bands with project concurrence.

Table 5-3 shows the limit level random vibration environment for components with mass less than

22.7 kg. Levels defined are acceptance levels. Analysis is conducted against acceptance with the safety factors defined in Table 8-1. Testing factors are defined in RST-PROP-SOW-0042.

Table 5-4 defines minimum workmanship for components weighing less than 45.4 kg. Random vibration specifications may be updated based on acoustic analysis.

During testing, force limiting control is recommended and the control method must be approved by the NASA/GSFC COR.

Table 5-3 Random Vibration Limit Levels

Frequency (Hz) ASD Limit Level (g2/Hz)

20 0.013 20 – 50 +6 dB/oct

50 – 800 0.08 800 – 2000 -6 dB/oct

2000 0.013 Overall 10.0 grms

Table 5-4 Minimum Workmanship ASD Level

Frequency (Hz) ASD Level (g2/Hz)

20 0.01 20 – 80 +3 dB/oct

80 – 500 0.04 500 – 2000 -3 dB/oct

2000 0.01 Overall 6.8 grms

5.4 Shock

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

Levels defined are acceptance levels. Testing factors are defined in RST-PROP-SOW-0042.

Table 5-5 Shock Limit Levels Frequency (Hz) Limit Load (g)

100 75 100-738 6.5 dB/oct.

738 858 10000 858

Note: A shock susceptibility and attenuation assessment is to be performed on the thruster design.

This assessment can be based on past shock tests or other relevant information.

5.5 Acoustics

N/A

5.6 Pressure

5.6.1 Operating External Pressure Range

The thruster shall be designed to withstand 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).

5.6.2 Maximum Depressurization Rate

The thruster shall be designed to meet all performance requirements after exposure to a depressurization profile shown in Table 5-6 during launch and ascent.

Table 5-6 Depressurization Profile

Transonic Late Early Sec psia sec psia sec psia

0.0 14.70 0.0 14.70 0.0 14.70

10.0 14.20 10.0 14.40 10.0 14.00

15.0 13.70 15.0 14.00 15.0 13.40

20.0 12.95 20.0 13.35 20.0 12.55

25.0 11.95 25.0 12.45 25.0 11.45

41.2 7.90 44.8 7.90 28.0 10.64

43.2 7.70 46.8 7.80 30.2 9.75

45.2 6.24 48.8 6.50 32.2 9.00

46.6 5.40 50.4 6.00 33.5 8.50

48.7 4.30 52.1 5.50 35.0 8.00

50.9 3.30 53.8 5.00 36.4 7.50

54.0 2.40 55.3 4.60 37.9 7.00

58.0 1.60 56.8 4.20 39.4 6.50

65.0 1.00 58.4 3.80 41.0 6.00

70.0 0.80 60.0 3.40 42.7 5.50

80.0 0.50 61.4 3.10 44.4 5.00

110.0 0.30 62.8 2.80 46.3 4.50

150.0 5.00 64.2 2.50 48.0 4.05

65.5 2.25 50.0 3.75

66.9 2.00 52.0 2.88

68.0 1.80 57.0 2.00

68.8 1.68 65.0 1.30

75.0 1.10 80.0 0.60

85.0 0.60 110.0 0.30

110.0 0.30 150.0 0.05

150.0 0.05

5.7 Ground Environments

5.7.1 Thermal Ground Environment

The thruster shall meet all performance requirements after exposure to air temperature between +5 and +30 deg-C.

5.7.2 Humidity Ground Environment

The thruster shall meet all performance requirements after exposure to relative humidity between 0% and 70%.

5.8 Thermal Requirements

5.8.1 Flight Interface Design Temperature Limits, Operational The thruster shall meet all performance requirements within the Operational limits shown in Table 5-7.

5.8.2 Flight Interface Design Temperature Limits, Acceptance

Acceptance temperatures shown in Table 5-7 are test temperatures referenced in RST-PROP-

SOW-0042.

5.8.3 Flight Interface Design Temperature Limits, Qualification The thruster shall meet all performance requirements within the Protoflight/Qualification limits shown in Table 5-7.

5.8.4 Flight Interface Design Temperature Limits, Survival

When powered “OFF”, the thruster shall be capable of surviving indefinitely when its temperatures are within the survival limits shown in Table 5-7 without damage or permanent performance degradation.

Table 5-7 Temperature Limits

Minimum Temperature (deg-C) Maximum Temperature (deg-C) Operational +10 +50 Acceptance Test +5 +55 Protoflight/Qualification Test

0 +60

Survival (Unpowered) -5 +120 Note: The hardware must meet all performance specifications within the qualification temperatures. Operation is not required at survival temperature.

Table 5-8 Propellant-wetted In-flight Software Temperature Limits

Software Temperature Limit

Minimum Temperature (deg-C) Maximum Temperature (deg-C)

Green +10 +50 Yellow +6 +60 Red +5 +120

Note: This table is used for RST Software purposes only. Verification at these temperatures is not required, unless the temperature coincides with acceptance/qualification temperatures.

6 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.

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 External Cleanliness

The thruster external surfaces shall meet IEST-STD-CC1246E Particulate Cleanliness Level 300, with visual inspection meeting Level VC-0.5-1000 + UV, or equivalent, when inspected with both UV and white light in a darkened room.

6.1.1.2 Internal Cleanliness

The thrusters shall be cleaned, and verified internally clean, to level 50R1 per IEST-STD- CC1246E (as modified by the following: no metal particles allowed above 25 um) prior to delivery.

6.1.2 Surface Contamination Generation

6.1.2.1 Particulate Generation

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

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

• 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.

6.1.2.2 Molecular Generation

Material Selection

The thruster materials shall have a total mass loss (TML) less than 1.00% and a collected volatile condensable mass (CVCM) less than 0.10%, when measured in accordance with ASTM E595 unless a materials usage agreement has been generated and approved by the NASA/GSFC

COR.

Material Selection - Silicones Silicones on external surfaces shall not be exposed to the space environment unless approved by the NASA/GSFC COR. Silicones should be avoided or minimized. It is highly recommended that silicones be baked out at a high temperature prior to integration into the system to prevent extended bakeouts of the entire assembly.

6.1.3 Cleanability

6.1.3.1 Cleanability – Sensitive Surfaces

If any surfaces are not cleanable with Isopropyl alcohol and polyester wipes or light vacuuming, they shall be identified on the MICD.

6.1.3.2 Cleanability – Sensitive Surface Cleaning Methods

Alternate cleaning methods shall be identified and appropriate documentation provided for any surfaces that are not cleanable with Isopropyl alcohol.

6.1.4 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.

7 DESIGN & CONSTRUCTION REQUIREMENTS

7.1 Parts, Materials & Processes (PMP)

7.1.1 EEE Parts

The thruster contractor’s Quality Assurance system for EEE parts will be in accordance with the requirements in the SOW, RST-PROP-SOW-0042.

7.1.2 Materials

The thruster will be comprised of materials and processes in accordance with the requirements in the SOW, RST-PROP-SOW-0042.

7.1.2.1 Material Conductivity

All mounting surfaces should be conductive as defined in Section 3.7.

7.1.3 Material Compatibility

7.1.3.1 Fluid Compatibility

The thrusters shall be fabricated with materials that are compatible with all fluids described below:

• Hydrazine per MIL-P-26536E (with amendment 1), High Purity Grade*

• Deionized water per JSC SPEC C20

• Isopropyl alcohol per TT-I-735, Grade A

• Nitrogen per MIL-PRF-27401, Grade B

• Helium per MIL-PRF-27407

• Argon per MIL-PRF-27415 (with amendment 1)

• Xenon, research grade

* Hydrazine compatibility is required for wetted components only.

7.1.3.2 Passivation of Wetted Materials

All material surfaces that are wetted with propellant shall be passivated.

7.2 Electrical

7.2.1 Interface Requirements

The thruster shall be delivered with a one (1) meter minimum flying lead harness length for the entire bundle.

7.2.2 Wires

7.2.2.1 Wire Specifications

All internal wire types shall meet the Wire and Cable Requirements in Section W1 of EEE-

INST-002.

7.2.2.2 Minimum Wire Size

No wire smaller than AWG #26 shall be used.

7.2.2.3 Wire Derating

The current carrying capacity of the wires shall be derated for continuous operation at the required current levels in a vacuum, as defined in Section W1 of EEE-INST-002.

7.2.2.4 Wire Composition

Wire conductors shall be copper or a copper alloy, plated with suitable materials except nickel or other magnetic materials.

7.2.2.5 Wire Twists

Power wires shall be twisted with their power return wires and signal wires with their signal return wires.

7.2.2.6 Wire Insulation Resistance

The insulation resistance shall be greater than 100 Mohms when measured at 21 +/- 10 deg-C with a potential of 500 +/- 50 VDC applied between the shorted leads and the case.

7.3 Thermal Design

7.3.1 Thermal Conduction

The maximum allowable conduction from the thrusters to the spacecraft will be provided to the contractor by the NASA/GSFC COR.

7.3.2 Thermal Radiation

The maximum allowable radiation from the thrusters to the spacecraft will be provided to the contractor by the NASA/GSFC COR. The inclusion or exclusion of a heat shield will be the contractor’s responsibility.

7.4 Identification and Marking

Each unit shall be permanently marked with the vendor part number, vendor name, and a unique sequential serial number in the area designated on the interface control drawing in a manner to be approved by the NASA/GSFC COR. Permanent marking will use alcohol-proof ink, engraving, or laser etching.

7.5 Workmanship

7.5.1 Workmanship Standards

The workmanship standards and processes outlined in the SOW will be used.

7.5.2 Welding

All welds shall be performed and inspected according to a vendor procedure approved by the

NASA/GSFC COR.

7.6 Reliability and Mission Lifetime

7.6.1 Mission Life

The thruster shall meet all performance specifications through five (5) years of operation in space, following a three (3) month commissioning period.

7.6.2 Shelf Life

The thruster shall not suffer any degradation in performance when stored for five (5) years either on the S/C or in bonded storage.

7.7 GSE Cleanliness

All ground and test support equipment provided with the thruster shall be compatible with the flight component and the cleanroom environment where the flight component or test component will reside.

7.8 Interchangeability

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

8 MECHANICAL DESIGN REQUIREMENTS

For additional guidance on the design and analysis of threaded fastening systems in NASA spaceflight hardware, consult NASA-STD-5020A, “Requirements for Threaded Fastening Systems in Spaceflight Hardware”.

8.1 Mechanical Factors of Safety

Positive Margin of Safety (MS) shall be demonstrated analytically for the thruster design, including interfaces to mechanical GSE, using the MS formula below with the appropriate Factors of Safety (FS) defined in Table 8-1 applied to flight limit loads.

Note:

1. MS is defined as follows:

𝑀𝑀𝑀𝑀 =

𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴 𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿 𝐿𝐿𝐴𝐴𝐴𝐴𝐿𝐿 𝐴𝐴𝑜𝑜 𝑀𝑀𝐿𝐿𝑜𝑜𝐴𝐴𝑆𝑆𝑆𝑆

𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐿𝐿𝐴𝐴𝐿𝐿 𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿 𝐿𝐿𝐴𝐴𝐴𝐴𝐿𝐿 𝐴𝐴𝑜𝑜 𝑀𝑀𝐿𝐿𝑜𝑜𝐴𝐴𝑆𝑆𝑆𝑆 ∗ 𝐹𝐹𝑀𝑀

− 1

2. Positive MS in joints that are both bonded and fastened is demonstrated assuming only the bond carries the load, only the bolt carries the load, or if the capability of the joint requires both the bond and fasteners, a detailed analysis is performed.

a. The adhesive portion of the joint is usually the stiffest load path and thus carries most of the load. In this scenario, the full strength of the bolted interface is not realized until after the failure of the adhesive.

3. Positive MS is demonstrated for adhesively bonded joints for all potential failure modes under all loading conditions and temperatures.

a. Failure modes may include cohesive failure in the adhesive, failure at the adhesive to adhered interface, and adherent failure.

b. Interaction between adhesively bonded joints in close proximity should be included when analyzing bonded joints.

c. Temperature stresses in bonded joints should consider bond and consolidation cure temperatures.

d. Adhesive bond fillets and bondline thickness should be considered in bonded joint design and analysis under temperature loading.

Table 8-1 Design Factors of Safety

Type of Hardware Static /Sine Random 2

Tested Metallic Yield 1.251 1.6

Tested Metallic Ultimate 1.41 1.8

Stability/Buckling Ultimate 1.4 1.8 Bonded inserts/Joints Ultimate 1.5 1.9

(1) For qualification by analysis only, positive margin must be shown with FS of 2.0 and 2.6 for metallic yield and ultimate, respectively.

(2) Factors shown are applied to statistically derived peak response based on limit/acceptance RMS level.

As a minimum, the peak response must be calculated as a 3-sigma value.

8.2 Fracture Control Requirements

Thrusters must be evaluated for fracture criticality. Thrusters are included under tier 4 of a four-tier system used to evaluate fracture criticality of hardware; tier 4 has associated mitigation requirements that shall be followed, as specified below.

Tier 4 fracture classification: Fracture control requirements (per NASA-STD-5019) shall apply to the following elements only: (1) Pressure vessels, dewars, lines, fittings and other pressurized components (per NASA-STD 8719.24); (2) Castings (unless hot isostatically pressed and the flight article is proof tested to 1.25 times limit load); (3) Weldments; (4) Parts made of materials in Tables II or III of MSFC-SPEC-3029A if under sustained tensile stress; (5) Parts made of materials susceptible to cracking during quenching; (6) Nonredundant, mission-critical preloaded springs loaded to greater than 25 percent of ultimate strength.

Notes for fracture/fatigue:

1. Fracture analysis is performed assuming an initial crack size. NDE or proof testing is required to verify this initial crack size assumption. If proof testing, the proof tests must be at load sufficient enough to screen for any crack size above the assumed. This often results in over-test compared to DLL.

2. If NASGRO is used for brittle component fracture analysis, the NASGL module is used.

3. A fail-safe component is a structural component in a redundant load path which after the loss of the component, the remaining load path(s) has sufficient structural capability to withstand the redistributed loads, and the loss of the component will not cause a catastrophic hazard.

Examples of redundant load paths are multiple fasteners in bolted joints, stitched welded structures, and riveted and bolted structures. A factor of safety of 1.0 shall be applied to the remaining load path(s) when demonstrating fail-safe.

4. Critical fastened joints are joints whose failure results in mission-ending failure. Class-C fasteners that are contained within Class-C hardware are not critical fasteners.

5. Mission cycles include ground testing and handling, transportation, launch, and on orbit loading. The load spectrum will include the load level and the accompanying number of cycles and duration (for static loading) at each level during the hardware service life. Mission cycle load spectrum is included in the Fracture Control Plan.

6. In cases where fracture toughness of a metal is not known, contact RST project for help in fracture classification.

8.3 Materials

8.3.1 Dissimilar Metals

Dissimilar materials in direct contact with each other shall be compatible as defined in Table I of MIL-STD-889C, unless approved by the NASA/GSFC COR. Use of similar materials sealed per NASA-STD-6012 may be acceptable in some instances.

8.3.2 Stress Corrosion Cracking

All metallic materials shall be selected from Table I of MSFC-STD-3029, unless otherwise approved by Parts, Materials, and Processes Control Board.

8.3.3 Fastened Joints

8.3.3.1…