TDM-25-5N-SPEC-0001 .pdf

TDM-25-5N-SPEC-0001, Revision (-)

5N Non-Toxic Thruster Specification

National Aeronautics and Space Administration

5N/22N Thruster Project Reviewed Not Subject To Export Control

Thruster Spec TDM-25-5N-SPEC-0001 Rev (-) i Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Preface This document is under configuration control. Changes to this document require prior approval of the Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

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

Change History Log

Revision Effective Date

Description of Changes (Reference the CCR & Approval Date)

(-) 04/29/2025 Baseline Revision iii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this

Record of Deviations/Waivers

Deviation/Waiver Approved Date Description of Deviations/Waivers iv Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this

Table of Contents

1.0 Introduction

1.1 General Information

1.2 Scope

2.0 Applicable Documents

3.0 Thruster Description

4.0 Configuration/Performance/Functional Requirements

4.1 Thruster Configuration Requirements

4.1.1 Major Components List

4.1.2 Nozzle Configuration

4.1.3 Valve Configuration

4.1.3.1 Redundancy

4.1.3.2 Thruster Valve Fail-Safe

4.1.3.3 Inlet Filter Rating

4.1.4 Interface Tube Location

4.1.5 Valve Heater Configuration

4.1.6 Valve Thermal Sensor Configuration

4.1.7 Catalyst Bed/Reactor Heater Configuration

4.1.8 Catalyst Bed/Reactor Thermal Sensor Configuration

4.2 Thruster Performance Requirements

4.2.1 Thrust

4.2.2 Thrust Repeatability

4.2.3 Thrust Alignment

4.2.4 Steady State Specific Impulse

4.2.5 Steady-State Firing Duration

4.2.6 Pulsed Specific Impulse

4.2.7 Thruster Duty Cycle

4.2.8 Impulse Bit

4.2.9 Propellant Throughput

4.2.10 Emergency Start Capability

4.2.11 Roughness

4.2.12 Thrust Degradation

4.2.13 Propellant and Pressurant Requirements

4.2.13.1 Propellant Type

4.2.13.2 Pressurant Type

4.2.13.3 Helium-Saturated Propellants

4.2.13.4 Propellant Temperature

4.2.13.5 Gas Ingestion

4.2.14 Deep Thermal Cycles

4.2.15 Thruster Warm Pulses

4.2.16 Heat Soakback

4.3 Thruster functional requirements

4.3.1 Pressure Requirements

4.3.1.1 Maximum Expected Operating Pressure/Exposure Pressure v Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this

4.3.1.2 Proof Pressure

4.3.1.3 Burst Pressure

4.3.1.4 Surge Pressure

4.3.1.5 Interface Tube Pressure Requirements

4.3.2 Leakage Requirements

4.3.2.1 Internal Leakage

4.3.2.2 Internal Leakage During Random Vibration

4.3.2.3 External Leakage

4.4 Valve Specific requirements

4.4.1 Valve Cycles

4.4.2 Valve Inductance

4.4.3 Valve Heater Sizing

4.5 Catalyst bed/reactor heater requirements

4.5.1 Catalyst Bed / Reactor Heater Duration to Minimum Temperature

4.5.2 Catalyst Bed / Reactor Turn-On Cycles

4.6 Power

4.6.1 Power Allocations

4.6.1.1 Power Allocation – Thruster Valve

4.6.1.2 Power Allocation – Catalyst Bed / Reactor Heater

4.6.2 Primary (Unregulated) Power Input Requirements

4.6.2.1 Operating Voltage Range

4.6.2.2 Abnormal Voltages

4.6.2.3 Sudden Removal of Power

4.6.3 Load Induced Noise Requirements

4.6.3.1 Turn-on Current Transients

4.6.3.2 Operational Current Transients (One Time)

4.6.3.3 Low Frequency Repetitive Transients (during nominal operation)

4.7 Electrical Grounding

4.7.1 Primary Power DC Isolation

4.7.2 Internally Generated Secondary Returns

4.7.3 Internally Generated Secondary to Primary DC Isolation

4.7.4 Mechanical Contact Resistance

4.7.5 Mating Method

5.0 Physical Requirements

5.1 Mass Properties

5.1.1 Mass Allocation

5.1.2 Center of Mass Accuracy

5.1.3 Determination of Moments and Products of Inertia

5.2 Physical Envelope

5.3 Mounting

5.4 Alignment

5.4.1 Nozzle Direction

5.4.2 Nozzle Exit Plane

5.4.3 Position Alignment

5.4.4 Alignment Location/Orientation Accuracy

5.4.5 Alignment Measurement Tooling/Features

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

5.5 Fluid Interfaces

5.5.1 Interface Tube Material

5.5.2 Interface Tube Orientation

5.5.3 Interface Tube Length

5.5.4 Interface Tube Outer Diameter

5.5.5 Interface Tube Wall Thickness

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 Pressure

6.6.1 Operating External Pressure Range

6.6.2 Maximum Depressurization Rate

6.6.3 Launch Vehicle Environmental Control System Impingement Velocity

6.7 Ground Environments

6.7.1 Thermal Ground Environment

6.7.2 Humidity Ground Environment

6.8 Thermal Requirements

6.8.1 Flight Interface Design Temperature Limits

6.8.2 Valve Temperature Limits

6.8.3 Thermal Interfaces

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

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

7.1.3.3 Transportation Cleanliness

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

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

8.1.2.2 Material Limitations for Debris Casualty Area

8.1.3 Material Compatibility

8.1.3.1 Fluid Compatibility

8.1.3.2 Thruster Dewpoint

8.1.3.3 Passivation of Wetted Materials

8.2 electrical

8.2.1 Interface Requirements

8.2.2 Wires

8.2.2.1 Wire Specification

8.2.2.2 Minimum Wire Size

8.2.2.3 Wire Derating

8.2.2.4 Wire Composition

8.2.2.5 Wire Twists

8.2.2.6 Wire Insulation Resistance

8.3 Thermal Design

8.3.1 Thermal Conduction

8.3.2 Thermal Radiation

8.4 Identification and marking

8.5 workmanship

8.5.1 Workmanship Standards

8.5.2 Welding

8.5.3 Welding Procedures

8.5.4 Material Joining

8.6 Reliablity and mission life

8.6.1 Mission Life

8.6.2 Shelf Life

8.6.3 Reliability

8.7 Ground Handling

8.7.1 GSE Cleanliness

8.7.2 Test Harness

8.8 Interchangability

9.0 Mechanical Design Requirements

9.1 Structural Requirements

9.1.1 Component Fatigue

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

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

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

Appendix A Abbreviations and Acronyms ix Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this

List of Figures

Figure Page

Figure 6-1 Shock Envelope

List of Tables

Table Page

Table 2-1 Applicable Documents 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 6-5 Qualification Level Shock Response Spectrum Table 6-6 Depressurization Profile Table 6-7 Valve Temperature Levels Table 8-1 Limited materials for debris casualty area

5N Thruster Spec TDM-25-5N-SPEC-0001 Rev (-)

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

1.0 INTRODUCTION

1.1 GENERAL INFORMATION

The Space Technology Mission Directorate (STMD) Technology Demonstration Office (TMD) is investigating low toxicity propulsion in the 5N thrust levels for direct applications to NASA Science Mission Directorate (SMD) mission applications.

1.2 SCOPE

This specification describes the performance, design, development, physical characteristics, electrical, mechanical, environmental, and verification testing requirements for a low toxicity 5N thruster maturation program. Further the program will demonstrate a robust qualification program to meet a NASA SMD Mission within the bounds of the requirements contained within this specification. The low toxicity 5N thruster will herein be referred to as the thruster.

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

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 5N thruster Statement of Work (TDM-25-5N-SOW), 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

TDM-25-5N-SOW Thruster Statement of Work GSFC-S-311-P-079 Procurement Specification for Thermofoil Heater

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-DTL-32539 Performance Specification Propellant, ASCENT 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

ASTM-A511 Standard Specification for Seamless Stainless Steel Mechanical Tubing

AMS-QQ-S-763 Pressure Tubing Specification, Stainless Steel

SAE-AMS-2244C Tolerances Titanium and Titanium Alloy Tubing

AMS-2488D Anodic Treatment - Titanium and Titanium Alloys Solution pH 13 or Higher

AWS D17.1

Specification for Fusion Welding for Aerospace Applications

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

Document Number Title

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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3.0 THRUSTER DESCRIPTION

The thruster is a 5N (1-lbf) class engine that is capable of generating thrust by the catalytic decomposition or a reactor combustion (i.e. non-catalytic) scheme using a low toxicity propellant. The major components of the thruster include the flow control valve, catalyst bed for propellant decomposition and/or reaction chamber for propellant combustion, a nozzle, and thermal hardware for heating and measuring necessary components. The thruster valve is a normally closed valve that requires power to open the valve, and then return to closed in the absence of power.

This specification applies to the thrusters, the associated ground support equipment (GSE), and the paperwork for the 5N/22N Thruster Project. Please refer to the Statement of Work (SOW, TDM-5N-SOW-0001) and the Deliverable Items List and Schedule (TDM-5N-SOW-0001, Appendix A) for a complete list of all the hardware, verification testing, test reports, design reports, and analyses reports that should be delivered with the thrusters.

The 5N thruster design will be qualified for space flight to the qualification levels listed in this document. This qualification program is encompassing; however, not a blanket qualification wholesale for every mission. Mission requirements must be evaluated to determine if the hardware is qualified to meet the objectives.

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4.0 CONFIGURATION/PERFORMANCE/FUNCTIONAL REQUIREMENTS

This Section defines the configurational, performance and functional requirements for the thruster.

4.1 THRUSTER CONFIGURATION REQUIREMENTS

4.1.1 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

• Catalyst bed/Reactor heaters

• Catalyst bed/Reactor temperature sensor

4.1.2 Nozzle Configuration

The thruster shall contain a straight-line nozzle with 0-degree cant.

4.1.3 Valve Configuration

4.1.3.1 Redundancy

The thruster shall contain a single seat flow control valve at minimum.

If a series-redundant flow control valve is implemented in the thruster design, the upstream and downstream valves shall open and close independently.

4.1.3.2 Thruster Valve Fail-Safe

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

4.1.3.3 Inlet Filter Rating

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 not reduce thruster performance.

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4.1.4 Interface Tube Location

The thruster interface tube shall comply with clocking shown in Figure 4.1. Refer to Section 4.6 for interface tube dimensions. Other configurations can be accepted with NASA COR approval.

Figure 4-1 Thruster Tube Interface

4.1.5 Valve Heater Configuration

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

4.1.6 Valve Thermal Sensor Configuration

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

4.1.7 Catalyst Bed/Reactor Heater Configuration

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

4.1.8 Catalyst Bed/Reactor Thermal Sensor Configuration

The thruster shall contain redundant thermal sensors for sensing catalyst bed/reactor temperature.

4.2 THRUSTER PERFORMANCE REQUIREMENTS

4.2.1 Thrust

a) The thruster(s) shall be capable of operating in a blowdown configuration with a maximum inlet pressure of 400 psia (27.6 bar) at beginning of life (BOL), and 75 psia (5.17 bar) at end of life (EOL).

b) The thruster shall have a minimum thrust of 1-lbf (4.45 N) at an inlet pressure of 400 psia

(27.6 bar).

c) The thruster shall have a minimum thrust of 0.225-lbf (1.0 N) at an inlet pressure of 75 psia (5.17 bar).

Inlet Tube

Nozzle NozzleValve

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4.2.2 Thrust Repeatability

The thruster shall produce steady state thrust that is predictable to within 7% at all conditions detailed in Section 4.2.1.

4.2.3 Thrust Alignment

The thruster thrust vector shall be perpendicular to the mounting flange plane to within ±0.25°.

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

4.2.4 Steady State Specific Impulse

The minimum vacuum steady-state specific impulse provided by the thruster shall have a minimum of 250 seconds for all operating conditions.

4.2.5 Steady-State Firing Duration

The thruster shall be capable of continuous steady-state firing durations of up to and including 3600 seconds (60 minutes).

4.2.6 Pulsed Specific Impulse

The average pulsed specific impulse of the thruster shall be greater than 235 seconds at duty cycles greater than 60% post a number of warm-up pulses.

The number of warm-up pulses is defined as when steady-state performance is reached for that duty cycle operation.

4.2.7 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 4.2.5 at all inlet pressures and thermal environments specified herein.

4.2.8 Impulse Bit

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

b) The thruster shall have a minimum impulse bit (Ibit) of less than 0.2 N-s for a valve command pulse duration of 20 milliseconds. The minimum impulse bit is defined as the area under the thrust-time curve for a given valve open pulse width.

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c) 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.

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

4.2.9 Propellant Throughput

The thruster shall have a minimum qualified throughput capability of 220.5 lbs (100 kg).

4.2.10 Emergency Start Capability

The vendor shall demonstrate thruster emergency operational capability for a worst case cold thermal starting condition. No thrust degradation will result from this operation.

4.2.11 Roughness

a) Chamber pressure roughness shall be less than 10% during steady-state operation for inlet pressure between 75 and 400 psia.

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

c) The vendor shall provide hot-fire data that would indicate the presence of chamber pressure spiking (rapid isolated increases of chamber pressure caused by collection and detonation of un-decomposed/un-combusted propellant). While occasional spiking is allowed, data will be reviewed by vendor and NASA to determine whether any spiking is considered excessive. Spiking that is determined excessive shall be investigated.

4.2.12 Thrust Degradation

The thruster shall maintain steady-state thrust degradation of less than 5% throughout the qualification lifetime, as compared to beginning of life steady-state thrust.

4.2.13 Propellant and Pressurant Requirements

4.2.13.1 Propellant Type

The thruster shall use High Purity Advanced Space Craft Energetic Non-Toxic (ASCENT), formerly AF-M315E, propellant per MIL-DTL-32539.

4.2.13.2 Pressurant Type

The thruster shall use either helium pressurant gas per MIL-PRF-27407, Type 1 Grade B or nitrogen pressurant gas per MIL-PRF-27401G.

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4.2.13.3 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.

4.2.13.4 Propellant Temperature

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

4.2.13.5 Gas Ingestion

The thruster shall be capable of ingesting 0.5 grams of GHe at 400 psia (27.6 bar) at 20 deg-C before operation without any harm to the Catalyst Bed/Reactor, the thruster, or any degradation in performance during subsequent normal firing.

4.2.14 Deep Thermal Cycles

The thruster shall be qualified for a minimum of 1,000 deep thermal cycles.

A single deep thermal cycle is defined as a temperature excursion from ambient temperature to the thruster’s minimum pre-heating temperature (warm) to the maximum thruster operational temperature and back down to the ambient temperature.

4.2.15 Thruster Warm Pulses

The thruster shall be qualified for 200,000 pulses while at operational temperature.

4.2.16 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 operations.

4.3 THRUSTER FUNCTIONAL REQUIREMENTS

4.3.1 Pressure Requirements

4.3.1.1 Maximum Expected Operating Pressure/Exposure Pressure

a) The thruster shall operate at any expected operating inlet pressure from BOL 400 psia (27.6 bar) to EOL 75 psia (5.17 bar).

b) The thruster shall have a Maximum Expected Operating Pressure (MEOP) of 400 psia (27.6 bar), where the valve will be open when input voltage is applied as described herein.

c) The maximum exposure pressure to the thrusters in a non-operating condition with valves closed shall be 600 psia (41.37 bar). The exposure pressure is the sum of the MEOP and

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this the back-pressure relief cracking pressure (up to 200 psid). 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.

4.3.1.2 Proof Pressure

a) The thrusters shall have a proof pressure capability of not less than 600 psig (41.37 bar) with the valves open. (Maximum exposure pressure)

b) The thrusters shall have a proof pressure capability of not less than 900 psig (62.1 bar) with the valves closed. (1.5x maximum exposure pressure)

4.3.1.3 Burst Pressure

a) The thrusters shall have a burst pressure capability of not less than 1,500 psig (103.4 bar) with the valves open. (2.5x maximum exposure pressure)

b) The contractor shall demonstrate through a qualification program that the valve has a burst pressure of at least 1,500 psig (103.4 bar). (2.5x maximum exposure pressure)

4.3.1.4 Surge Pressure

The thruster shall withstand a peak surge (water-hammer) pressure of 1,500 psig at the closed thruster valve.

4.3.1.5 Interface Tube Pressure Requirements

4.3.1.5.1 MEOP/Exposure Pressure

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

The thruster interface tube shall have an Exposure Pressure of not less than 600 psia.

4.3.1.5.2 Proof Pressure

The interface tubes shall be capable of withstanding a proof pressure of not less than 900 psig (62.1 bar) (1.5x maximum exposure pressure).

4.3.1.5.3 Burst Pressure

The interface tubes shall be capable of withstanding a proof pressure of not less than 2,400 psig (165.5 bar) (4.0x maximum exposure pressure).

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4.3.2 Leakage Requirements

4.3.2.1 Internal Leakage

The thruster valve seat(s) shall have an internal leak rate of less than or equal to 5 standard cubic centimeters per hour (scch) of gaseous helium (GHe) at test differential pressures of 25 psid (1.72 bar), 400 psid (27.6 bar), and the maximum differential pressure of 600 psid (41.37 bar), 400 psid MEOP + 200 psid latch valve back pressure relief.

4.3.2.2 Internal Leakage During Random Vibration

The thruster valve seat(s) shall have an internal leak rate of less than or equal to 5 scch 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, or by analysis, and if approval is received from the NASA.

4.3.2.3 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 of 400 psi (27.6 bar) with both thruster valve seats open and the nozzle plugged.

4.4 VALVE SPECIFIC REQUIREMENTS

4.4.1 Valve Cycles

The thruster valve shall be qualified for 400,000 open/close cycles for each valve seat.

(2x Thruster Warm Cycles)

4.4.2 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) valve(s) are wired in parallel (by the customer).

4.4.3 Valve Heater Sizing

The thruster valve heater shall be sized to maintain allowable valve temperatures with a maximum heater duty cycle of 70% at 27 VDC input voltage under worst-case cold thermal conditions, assuming only one heater circuit is active.

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4.5 CATALYST BED/REACTOR HEATER REQUIREMENTS

4.5.1 Catalyst Bed / Reactor Heater Duration to Minimum Temperature

The catalyst bed / reactor 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.

4.5.2 Catalyst Bed / Reactor Turn-On Cycles

The Catalyst Bed / Reactor heater shall be qualified to 2,000 turn-on cycles.

(2x Deep Thermal Cycles)

4.6 POWER

4.6.1 Power Allocations

4.6.1.1 Power Allocation – Thruster Valve

The thruster shall have a nominal (continuous) power consumption of less than or equal to 15 W across the operating voltage range in Section 4.6.2.1 and the protoflight/qualification temperature range in Table 6-7.

This power allocation applies when the upstream and downstream valve circuits are wired in parallel, if applicable.

4.6.1.2 Power Allocation – Catalyst Bed / Reactor Heater

The Catalyst Bed / Reactor heater shall have a nominal (continuous) power consumption of less than or equal to 35W across the operating voltage range in Section 4.6.2.1 and the protoflight/ qualification temperature range in Table 6-7. This power allocation applies to a single catalyst bed heater circuit.

4.6.2 Primary (Unregulated) Power Input Requirements

4.6.2.1 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 for all expected load conditions (except when turned off).

The thruster shall meet all performance requirements within this range.

4.6.2.2 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.

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4.6.2.3 Sudden Removal of Power

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

4.6.3 Load Induced Noise Requirements

4.6.3.1 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).

4.6.3.2 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.

4.6.3.3 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.

4.7 ELECTRICAL GROUNDING

4.7.1 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 MΩ.

4.7.2 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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4.7.3 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.

4.7.4 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.

4.7.5 Mating Method

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

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5.0 PHYSICAL REQUIREMENTS

5.1 MASS PROPERTIES

5.1.1 Mass Allocation

The mass of the thruster shall be less than or equal to 1.10 lbm (0.5 kg) including harness mass.

The mass of the thruster shall be measured to within 0.055 lbm (0.025 kg).

5.1.2 Center of Mass Accuracy

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

5.1.3 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%.

5.2 PHYSICAL ENVELOPE

a) The thruster outside dimensions, including mounting flanges and connectors, shall be within an envelope of 7.874 in x 3.937 in x 2.362 in (200 mm length X 100 mm x 60 mm)

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

5.3 MOUNTING

a) The thruster shall accommodate being hard mounted with mechanical fasteners on a single mechanical surface (i.e. Spacecraft Structure).

b) 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.

5.4 ALIGNMENT

5.4.1 Nozzle Direction

The centerline of the nozzle (i.e. thrust vector) shall be perpendicular with respect to the mounting flange as shown in Figure 4-1 in Section 4.1.4.

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5.4.2 Nozzle Exit Plane

The nozzle exit plane shall be perpendicular with respect to the nozzle centerline to within ±0.25 degrees, before and after environment testing.

5.4.3 Position Alignment

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

5.4.4 Alignment Location/Orientation Accuracy

See Section 4.2.3 for thrust alignment.

5.4.5 Alignment Measurement Tooling/Features

a) The interchangeable alignment plugs shall be made such that they can be used for any of the 5N thruster.

b) The optical reference surfaces shall be a cube of at least 0.5 in3 (8194 mm3).

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

5.5 FLUID INTERFACES

5.5.1 Interface Tube Material

The interface tube shall be 6Al-4V Titanium per SAE-AMS-2244C.

5.5.2 Interface Tube Orientation

The interface tube shall be perpendicular to the centerline of the thruster valve body as shown in Figure 4-1 in Section 4.1.4. Other configurations can be accepted with NASA COR approval.

5.5.3 Interface Tube Length

The interface tube shall have a uniform straight length of 1.50 +/- 0.05 inches.

5.5.4 Interface Tube Outer Diameter

The interface tube shall have an outer diameter of 0.25 +0.004/-0.000 inches along the length in Section 5.5.3.

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5.5.5 Interface Tube Wall Thickness

The interface tube shall have a wall thickness of 0.028 +0.004/-0.000 inches along the length described in Section 5.5.3. The thickness tolerance may be verified for only the first 0.8 +/-0.05 inches of tube length starting at the interface end of the tube.

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6.0 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements are specified in this Section.

The thruster 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

The thruster 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 (see NASA-STD-5001A for more information on other materials [e.g. glass]). 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

Flight Hardware Design/Analysis Factors of Safety Applied to Limit Loads1,2 Type of Hardware Static Sine Random/Acoustic4,5

Metallic Yield 1.253 1.25 1.6 Metallic Ultimate 1.43 1.4 1.8 Stability Ultimate 1.4 1.4 1.8 Beryllium Yield 1.4 1.4 1.8 Beryllium Ultimate 1.6 1.6 2.0 Composite Ultimate 1.5 1.5 1.9 Bonded Inserts/Joints Ultimate 1.5 1.5 1.9

1 – Factors of safety for pressurized systems to be compliant with NASA-STD-8719.24 (with Annex) (NASA Expendable Launch Vehicle Payload Safety Requirements).

2 – Factors of safety for glass and structural glass bonds specified in NASA-STD-5001.

3 – If qualified by analysis only, positive margin must be shown for factors of safety of 2.0 on yield and 2.6 on ultimate. See Section 2.4.1.1.1 of GSFC-STD-7000A 4 – Factors shown should be applied to statistically derived peak response based on RMS level. As a minimum, the peak response shall be calculated as a 3-sigma value.

5 – Factors shown assume that qualification/protoflight testing is performed at acceptance level plus 3dB. If the difference between acceptance and qualification levels is less than 3dB, then the above factors may be applied to qualification level minus 3dB.

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

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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 (DLL) defined in the mass-acceleration curve (MAC) shown in Table 6-2 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.

3. 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

Thruster Mass (kg)

Limit Load (g, any direction)

0.5 or less 69.9 1 67.9

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 interface. Requirements are met assuming rigid stiffness to restrained Degree of Freedom (DOF).

If the thruster is unable to meet the specified fundamental frequency, a finite element model must be supplied, 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 (TDM-25-5N- SOW-0001, Section 3.5).

6.4 VIBRATION

6.4.1 Sinusoidal Vibration

The thruster shall withstand acceptance and qualification sine vibration testing on all three axes at the levels shown in Table 6-3.

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Note:

1. Levels defined are acceptance levels. Analysis is conducted against acceptance with the safety factors defined in Table 6-1. Testing factors are defined in SOW (TDM-25-5N- SOW, Section 5.6).

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 6-3 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 6-1.

5. 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.

Table 6-3 Thruster Sine Vibration Environment

Frequency Acceptance Level Qualification Level 5-20 Hz 0.50 in (double amplitude) 0.63 in (double amplitude)

20-100 Hz 10.0 G 12.5 G

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) Acceptance Qualification

20-50 50-800

800-2000

0.013 +6 dB/oct

0.08 -6 dB/oct

0.013

0.026 +6 dB/oct

0.16 -6 dB/oct

0.026 Overall 10.0 Grms 14.1 Grms

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 DLL.

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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 DLL. Notching beyond these limits will require NASA 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 6-5, applied at the thruster interface.

Table 6-5 Qualification Level Shock Response Spectrum

Frequency (Hz) Level (Q=10) 100 50 g

100 to 1500 10.0 dB/Octave 1500 to 3000 1500 g

3000 to 10,000 3000 g Three Mutually Perpendicular Axes

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.

Figure 6-1 Shock Envelope

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

c) 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).

d) 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.

6.6 PRESSURE

6.6.1 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).

6.6.2 Maximum Depressurization Rate

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

Table 6-6 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

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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

6.6.3 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.

6.7 GROUND ENVIRONMENTS

6.7.1 Thermal Ground Environment

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

6.7.2 Humidity Ground Environment

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

6.8 THERMAL REQUIREMENTS

6.8.1 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.

6.8.2 Valve Temperature Limits

The thruster shall be capable of operating when its valve temperatures are within the temperature limits shown in Table 6-7 without damage or permanent performance degradation.

Table 6-7 Valve Temperature Levels Minimum Temperature (ºC) Maximum Temperature (ºC)

Valve Open +10 +65 Valve Closed +10 +100 Qualification +4 +100

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6.8.3 Thermal Interfaces

NASA 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.

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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 R1 on all external and critical surfaces.

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. This is accomplished by maintaining all hardware in a double bag for storage.

7.1.1.5 Internal Cleanliness

The thruster 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.

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:

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• 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% and a collected volatile condensable mass (CVCM) less than 0.10%, when measured in accordance with ASTM E-595 unless a materials usage agreement has been generated and approved by the NASA COR.

7.1.2.2.2 Material Selection – Silicones

Silicones on external surfaces shall not be exposed to the space environment unless approved by the NASA 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.

7.1.3 Cleanability

7.1.3.1 Cleanability – Sensitive Surfaces

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

7.1.3.2 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.

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7.1.3.3 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.

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8.0 DESIGN & CONSTRUCTION REQUIREMENTS

8.1 PARTS, MATERIALS & PROCESSES (PMP)

8.1.1 EEE Parts

The contractor’s Quality Assurance system for EEE parts will be in accordance with the requirements in SOW (TDM-25-5N-SOW). For additional details on EEE parts compliance specific to harness wires, see Section 8.2.2 in this document.

8.1.2 Materials

The thruster assembly will be comprised of materials and processes in accordance with the requirements in the SOW (TDM-25-5N-SOW).

The thrusters shall function within specification after exposure to propellant for a 5-year duration.

8.1.2.1 Material Conductivity

All mounting surfaces should be conductive as defined in Section 4.7.

8.1.2.2 Material Limitations for Debris Casualty Area

The contractor shall report the mass, dimensions, and use of materials shown in Table 8-1.

Table 8-1 Limited materials for debris casualty area Material Quantity

Titanium Mass > 50g Steel Mass > 50g Inconel Mass > 50g Invar Mass > 50g Materials with melting points above 1200° C (1473 K)

Mass > 50g

8.1.3 Material Compatibility

8.1.3.1 Fluid Compatibility

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

• ASCENT…