Attachment J-3 Specifications 10 May 2021.pdf

ER13-SPEC-0001

George C. Marshall Space Flight Center Revision: 03 Marshall Space Flight Center, Alabama 35812 EFFECTIVE DATE: 04-28-2021

Mars Ascent Vehicle (MAV) Reaction Control

System

SPECIFICATION

FOR

5N THRUSTER

ENGINEERING DIRECTORATE

Propulsion Systems Department

Engine Components Development & Technology Branch

ER13

Mars Ascent Vehicle Reaction Control System

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Document No: ER13-SPEC-0001 Revision: Rev.03

Effective Date: 04-28-2021 Page 3 of 29

APPROVALS

Prepared by: Date: _____________________

David B. (Brad) Bullard

ER13 Combustion Devices Team Lead

Reviewed: Date: _____________________

Matthew W. Marsh

ER13 Branch Chief

Approved by: Date: _____________________

Frankie R. Jernigan

ER11 MAV RCS Propulsion System Lead

Benjamin D. Davis

ER62 MAV Propulsion System Lead

Mark G. Dagostino

EE05 MAV Chief Engineer

Angela L. Jackman

ST20 MAV Project Manager

DAVID

BULLARD

Digitally signed by DAVID

BULLARD

Date: 2021.04.28 10:22:14 -05'00'

MATTHEW

MARSH

Digitally signed by

MATTHEW MARSH

Date: 2021.04.28 10:32:04 -05'00'

FRANKIE

JERNIGAN

Digitally signed by FRANKIE

JERNIGAN

Date: 2021.04.28 11:32:41 -05'00'

BENJAMIN

DAVIS

Digitally signed by

BENJAMIN DAVIS

Date: 2021.04.28 13:01:16 -05'00'

MARK

D'AGOSTINO

Digitally signed by MARK

D'AGOSTINO

Date: 2021.04.28 13:31:55 -05'00'

ANGELA

JACKMAN

Digitally signed by ANGELA

JACKMAN

Date: 2021.04.29 10:26:59 -05'00'

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Document No: ER13-SPEC-0001 Revision: Rev.03

Effective Date: 04-28-2021 Page 4 of 29

Table of Contents

1 SCOPE

2 DOCUMENTS

2.1 ORDER OF PRECEDENCE

2.2 APPLICABLE DOCUMENTS

Structural, Mechanical, and Material Fluids Electrical Other

2.3 DEFINITIONS AND TERMS

2.4 ABBREVIATIONS AND ACRONYMS

3 COMPONENT REQUIREMENTS

3.1 ITEM DESCRIPTION

3.2 PERFORMANCE REQUIREMENTS

Functional Operation Operating Fluid Thrust Steady State Specific Impulse (Isp) Startup Response Shutdown Decay Response Roughness Pulse Mode Performance Valve Electrical Actuation Operational Life Ancillary Electrical Components Leakage Physical Characteristics Length of Service

3.3 COMPONENT INTERFACES

Fluid Interfaces Structural Mounting Interface Nozzle Exit to Vehicle Interface Electrical Interfaces Nozzle Centerline Axis Orientation

3.4 ENVIRONMENTS

Propellant Inlet Temperature Range Operating Temperature Non-Operating Temperature External Pressure Humidity Random Vibration Shock Inertial Loads

3.5 DESIGN AND CONSTRUCTION

Material and Construction Standards Strength and Life Assessments Structural Design and Test Factors of Safety Threaded Joints Mechanisms Electrical Bonding

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Document No: ER13-SPEC-0001 Revision: Rev.03

Effective Date: 04-28-2021 Page 5 of 29

EEE Parts Fasteners External Cleanliness Internal Cleanliness Fluid Compatibility Workmanship Interchangeability

3.6 PART MARKING AND IDENTIFICATION

3.7 TRACEABILITY

Serialization Records

3.8 RELIABILITY

4 VERIFICATION

4.1 METHODS OF VERIFICATION

Inspection (I) Demonstration (D) Analysis (A) Test (T)

4.2 VERIFICATION MATRIX

5 PREPARATION FOR DELIVERY

5.1 METHODS OF PRESERVATION AND PACKAGING

Contamination Control Inlet Tube Thruster Orientation Retention of Cleanliness Storage Lead Wires

TABLES

Table 1. Random Vibration Levels Table 2. Shock Loads at Thruster Mounting Plate Table 3. Sample Acceptance Hot-Fire Test Matrix (TBR) Table 4: Verification Cross Reference Matrix Requirements

FIGURES

Figure 1. Thruster Envelope (center position) Figure 2. Thruster Envelope (right position) Figure 3. Thruster Envelope (left position) Figure 4. Thruster Mounting Configuration

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Document No: ER13-SPEC-0001 Revision: Rev.03

Effective Date: 04-28-2021 Page 6 of 29

1 SCOPE

This specification establishes the requirements for the performance, design, test, qualification, acceptance, and delivery of a 1 lbf (5 N) class hydrazine thruster. The thruster will be assembled into a blow-down system and operated in both steady state and pulsed modes. Within this document, the term “thruster” is synonymous with “Rocket Engine Assembly (REA),” which is frequently found in the propulsion industry.

2 DOCUMENTS

2.1 Order of Precedence

For purposes of interpretation or construction in case of any omissions, conflicts, ambiguities or inconsistencies, the following documents shall rank in the following order of precedence:

1. Purchase Agreement

2. Statement of Work

3. This Specification

4. Applicable Documents listed in Section 2.2 herein.

It should be noted that nothing in this document supersedes applicable laws and regulations unless a specific exemption has been obtained.

2.2 Applicable Documents

The documents listed below, of the exact issue shown or the latest revision if not specified, form a part of this specification to the extent specified herein.

Structural, Mechanical, and Material MSFC-STD-486 Standard, Threaded Fasteners, Torque Limits for NASA-STD-5012 Strength and Life Assessments Requirements for Liquid-Fueled Space

Propulsion System Engines NASA-STD-5017 Design and Development Requirements for Mechanisms NASA-STD-5020 Requirements for Threaded Fastening Systems in Spaceflight Hardware NASA-STD-6008 NASA Fastener Procurement, Receiving Inspection, and Storage Practices for Spaceflight Hardware NASA-STD-6016 Standard Materials and Processes Requirements for Spacecraft

Fluids MIL-PRF-26536 High Purity Monopropellant Hydrazine (N2H4) MIL-PRF-27401 Propellant Pressurizing Agent, Nitrogen MIL-PRF-27407 Propellant Pressurizing Agent, Helium MIL-PRF-27415C Propellant Pressurizing Agent, Argon TT-I-735 Isopropyl Alcohol

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

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Effective Date: 04-28-2021 Page 7 of 29

IPC J-STD-001S Space and Military Applications Electronic Hardware Addendum to IPC

J-STD-001

MSFC-STD-3012 Electrical, Electronic, and Electromechanical (EEE) Parts Management and Control Requirements for MSFC Space Flight Hardware

NASA-STD-4003 Electrical Bonding for NASA Launch Vehicles, Spacecraft, Payloads, and Flight Equipment

NASA-STD-8739.4 Workmanship Standard for Crimping, Interconnecting Cables, Harnesses, and Wiring

Other IEST-STD-CC1246 Product Cleanliness Levels – Applications, Requirements, and

Determination MIL-STD-130 Identification Marking of U.S. Military Property SMC-S-016 Test Requirements for Launch, Upper-Stage and Space Vehicles

2.3 Definitions and Terms

Shall: Used to indicate a requirement which must be implemented and its implementation verified.

To Be Determined (TBD): A value that is not yet known and will be provided by the purchaser.

To Be Resolved (TBR): An approximate value is known, but requires review by the purchaser.

To Be Supplied (TBS): A value that will be provided by the supplier and agreed to by the purchaser.

Will: Used to indicate a statement of fact and is not verified.

2.4 Abbreviations and Acronyms

“A” Analysis (used in Table 4) Ar Argon gas ASTM American Society for Testing and Materials °C Degrees Celsius “D” Demonstration (used in Table 4) DC Duty Cycle ECF Environmental Correction Factor EEE Electrical, Electronic, Electromechanical EMF Electromotive Force EPW Electrical Pulse Width °F Degrees Fahrenheit FS Factor of Safety G, g Acceleration due to gravity GHe Gaseous Helium GN2 Gaseous Nitrogen GN&C Guidance, Navigation, and Control

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Effective Date: 04-28-2021 Page 8 of 29

Grms Root mean square of acceleration, as measured in multiples of G Hz Hertz (cycles/second) “I” Inspection (used in Table 4) IEST Institute of Environmental Science and Technology in Inch Isp Specific Impulse (lbf-sec/lbm) lbf Pound (force) lbm Pound (mass) LN2 Liquid Nitrogen MAV Mars Ascent Vehicle MDC Mission Duty Cycle MDP Maximum Design Pressure MEOP Maximum Expected Operating Pressure MS Margin of Safety MSFC Marshall Space Flight Center N2H4 Hydrazine (monopropellant) NASA National Aeronautics and Space Administration N Newton (unit of force) PP Planetary Protection PSD Power Spectral Density psia Pounds per Square Inch (absolute) RCS Reaction Control System REA Rocket Engine Assembly RTD Resistance Temperature Detector sccs Standard Cubic Centimeters per Second “T” Test (used in Table 4) TBD To Be Determined TBR To Be Resolved (with the thruster manufacturer) TBS To Be Supplied (by the thruster manufacturer) TC Thermocouple VC Visibly Clean VDC Voltage, Direct Current

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Document No: ER13-SPEC-0001 Revision: Rev.03

Effective Date: 04-28-2021 Page 9 of 29

3 COMPONENT REQUIREMENTS

3.1 Item Description

The MAV RCS monoprop thrusters will provide roll control for first stage ascent from the Martian surface, then attitude control for the coast period between first stage main engine shutdown and upper stage separation.

Each thruster will contain a dual-seat valve assembly, a catalyst bed, a chamber and nozzle, two valve heaters (one per valve), a catalyst bed heater, a propellant inlet filter, and instrumentation to monitor temperature in the catalyst bed and both valves in the assembly.

3.2 Performance Requirements

Because performance differences will exist due to the need for multiple nozzle configurations (see Section 3.2.13), the values provided within this specification are written to envelope all configurations.

All performance values are to be reported for an external environment corresponding to vacuum conditions.

Functional Operation The thruster shall provide thrust when electrical power is applied to the propellant valve.

Operating Fluid The thruster shall meet the performance and life requirements of this specification when supplied with monopropellant grade hydrazine per MIL-PRF-26536 “High Purity Monopropellant Hydrazine (N2H4).”

3.2.2.1 Propellant Conditions

In addition to the pressure condition stated in this section, the thruster shall meet all performance requirements for the full range of temperatures specified in Sections 3.4.1, 3.4.2, and 3.4.3 that are applicable to the mission phase (operational/non-operational).

3.2.2.1.1 Operating Pressure Range

The thruster shall be operational over a propellant feed pressure range of 500 to 100 psia.

3.2.2.1.2 Maximum Expected Operating Pressure (MEOP) at Thruster Inlet The maximum expected operating pressure at the thruster valve inlet is 500 psia.

3.2.2.1.3 Maximum Design Pressure (MDP) at Thruster Inlet

With the valve closed, the thruster shall withstand a Maximum Design Pressure (MDP) of 1000 psia at the propellant inlet without damage. This MDP is the maximum feedline pressure and is equal to MEOP + transient surge pressures due to water hammer.

3.2.2.1.4 Proof Pressure

The thruster shall meet the performance requirements of Section 3.2 after exposure to proof pressure, as specified in NASA-STD-5012 “Strength and Life Assessments Requirements for Liquid-Fueled Space Propulsion System Engines,” for a minimum duration of five (5) minutes

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Effective Date: 04-28-2021 Page 10 of 29 without detrimental yielding. Proof factors are to be applied to the relevant internal MDP for each pressurized sub-component.

3.2.2.1.5 Burst Pressure

The thruster shall not rupture when exposed to burst pressure, as specified in NASA-STD-5012 “Strength and Life Assessments Requirements for Liquid-Fueled Space Propulsion System Engines,” for a minimum of one (1) minute. Permanent deformation is acceptable. Burst factors are to be applied to the relevant internal MDP for each pressurized sub-component.

3.2.2.1.6 Reference Conditions

For the purposes of this document, reference conditions shall indicate 350 psia propellant inlet pressure and 70 °F propellant inlet temperature.

Rationale: Define a single design reference point for performance.

Thrust Vacuum steady-state thrust shall be 1.0 lbf ± 0.05 lbf when evaluated at reference propellant inlet conditions.

3.2.3.1 Performance Variation with Inlet Pressure

Nominal vacuum steady state thrust and Isp versus inlet pressure over the operational range, including error bands, shall be as specified in Figure TBS when evaluated at the reference temperature provided in Section 3.2.2.1.6.

Rationale: Supplier to provide F and Isp vs. Pf curves with error bands.

3.2.3.2 Thrust Repeatability

The total variation in steady-state thrust between thrusters shall not exceed ± 5% under identical propellant supply conditions.

3.2.3.3 Thrust Degradation

The steady state thrust degradation at the end of life shall be less than 5% compared to the beginning of life.

Steady State Specific Impulse (Isp) The minimum steady state specific impulse at reference conditions shall be no less than 228 seconds over the operational life of the thruster.

Startup Response The startup response to 90% of steady state chamber pressure shall not exceed 100 (TBR) milliseconds with the thruster at equilibrium bed temperature.

Shutdown Decay Response The shutdown decay response to 10% of steady state chamber pressure shall not exceed 150 (TBR) milliseconds with the thruster at equilibrium bed temperature.

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Effective Date: 04-28-2021 Page 11 of 29

Roughness Chamber pressure roughness shall not exceed 20% peak-to-peak of the mean as taken over a one second moving average (after 1 second of ON time) over the operational life of the thruster.

Pulse Mode Performance The thruster shall be designed to allow for operation in pulse mode.

3.2.8.1 Duty Cycles (DC)

The thruster shall have no known duty cycle limitations, as demonstrated during required testing.

3.2.8.2 Minimum Electrical Pulse Width (EPW)

The thruster shall provide for pulsed operation with a 20 millisecond minimum commanded pulse width.

3.2.8.3 Minimum Impulse Bit (MIB)

Impulse bit, when evaluated at the minimum electrical pulse width defined in Section 3.2.8.2, shall be no greater than 0.022 (TBR) lbf-sec when supplied with propellant at the reference conditions defined in Section 3.2.2.1.6.

3.2.8.3.1 MIB Variation with Inlet Pressure

The variation of impulse bit with propellant inlet pressure, when evaluated at the minimum electrical pulse width defined in Section 3.2.8.2, shall be as shown in Figure TBS for the operational pressure range shown in Section 3.2.2.1.1 when supplied with propellant at the reference temperature defined in Section 3.2.2.1.6.

Rationale: MIB vs. Pf to be provided by the supplier.

3.2.8.4 Impulse Bit Repeatability

Impulse bit repeatability at a 100 millisecond or greater pulse duration shall not exceed ± 15% of the nominal pulse train average at equilibrium conditions and with constant inlet pressure, temperature, and voltage.

3.2.8.5 Impulse Bit Variability

The total variation in impulse bit between thrusters shall not exceed ± 5% under identical propellant supply conditions.

3.2.8.6 Pulse Mode Specific Impulse

The thruster shall provide Isp, including error bands, as a function of electrical ON-time as specified in Figure TBS when operated over the range of required operating conditions. Hot start-up pulses, cold start-up pulses, and ambient starts are to be indicated.

Rationale: Isp vs. EPW to be provided by the supplier.

3.2.8.7 Impulse Bit Variation with Electrical Pulse Width

The thruster shall provide impulse bit as a function of electrical ON-time as specified in Figure TBS when operated over the range of required operating conditions. Hot start-up pulses, cold start-up pulses, and ambient starts are to be indicated.

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Effective Date: 04-28-2021 Page 12 of 29

Rationale: Ibit vs. EPW to be provided by the supplier.

Valve Electrical Actuation The thruster shall meet the performance requirements of Section 3.2 with an electrical actuation as described in this section over the full range of operating conditions defined within this specification. These requirements apply to each valve, unless specifically stated otherwise.

3.2.9.1 Operating Voltage

The thruster shall be designed for operation 28 +6/-4 VDC nominal power.

3.2.9.2 Back-EMF Suppression Diodes

All electrical power equipment used to open and close the propellant valves shall have back- EMF suppression diodes at a voltage level of 60 VDC to protect the valve coils from over-voltage surges during valve closing.

3.2.9.3 Valve Actuation Frequency

The valve shall be designed to allow for operation at an actuation frequency up to 20 Hz at the minimum EPW, as specified in Section 3.2.8.2.

3.2.9.4 Pull-in Voltage

The valve pull-in voltage shall meet or exceed opening force margin requirements, as defined by Section 3.5.5.

3.2.9.5 Drop-Out Voltage

The valve drop-out voltage shall meet or exceed closing force margin requirements, as defined by Section 3.5.5.

3.2.9.6 Power Consumption

The total power draw for both valves shall not exceed 8.25 Watts at a temperature of 70 °F.

3.2.9.7 Dielectric Strength

The valve shall have no electrical breakdown, flashover, or current leakage greater than 2 mA with a voltage of 600 +50/-0 VAC applied for a duration of one minute minimum between the lead wires and valve mounting interface.

3.2.9.8 Insulation Resistance

-0 VDC applied between the lead wires and valve mounting interface.

Operational Life

3.2.10.1 Total Impulse

The thruster shall provide a total impulse of 50 lbf-sec.

3.2.10.2 Number of Hot-Fire Pulses

The thruster shall provide a minimum of 2,000 pulses at a variety of duty cycles.

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Effective Date: 04-28-2021 Page 13 of 29

3.2.10.3 Valve Dry Cycles

The valve shall provide a minimum of 1500 dry cycles.

3.2.10.4 Valve Wet Cycles

The valve shall provide a minimum of 4,000 cycles when wetted with propellant.

3.2.10.5 Cold Start Capability

The thruster shall meet steady state performance requirements after 5 cold starts, with a catalyst bed temperature less than 60 °F.

Ancillary Electrical Components

3.2.11.1 Catalyst Bed Heater

The thruster shall include an independently-commanded, dual-element catalyst bed heater with a total power draw (both elements energized) of TBS Watts at 70 °F.

Rationale: Power requirements for heater to be resolved with supplier based on environments.

Catalyst bed heaters will be thermostatically controlled by the vehicle to maintain temperature above the lower threshold value shown in Section 3.4.2.

3.2.11.2 Valve Heater

The thruster shall contain one independently-commanded propellant valve heater for each valve in the dual-seat assembly. Total power draw (both units energized) shall not exceed TBS Watts at 70 °F.

Rationale: Power requirements for heater to be resolved with supplier based on environments.

Valve heaters will be thermostatically controlled by the vehicle to maintain temperature above the lower threshold value shown in Section 3.4.2.

3.2.11.3 Thruster Catalyst Bed Temperature Monitoring

The catalyst bed shall be equipped with a temperature probe, appropriately placed to indicate average catalyst bed temperature.

3.2.11.4 Thruster Valve Temperature Monitoring

Each thruster valve in the dual-seat assembly shall be equipped with a temperature probe, appropriately placed to indicate average valve temperature.

Leakage

3.2.12.1 Thruster External Leakage

The thruster shall have an external leakage rate of less than or equal to 1 x 10-3 sccs GHe over the operating pressures in Section 3.2.2.1.1.

3.2.12.2 Thruster Valve Internal Leakage

The thruster valve shall have an internal leakage rate, when the valve is closed, of less than or equal to 1 x 10-3 sccs GHe over the operating pressures in Section 3.2.2.1.1.

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Document No: ER13-SPEC-0001 Revision: Rev.03

Effective Date: 04-28-2021 Page 14 of 29

Physical Characteristics Each thruster shall consist of: (1) a serially-configured, dual-seat, normally-closed, direct-acting propellant valve, (2) a catalyst bed, (3) a chamber and nozzle, (4) two valve heaters, (5) a dual-element catalyst bed heater, (6) a propellant inlet filter, and (7) three temperature probes, one each for independently monitoring temperature of the catalyst bed and both valves, as indicated in Sections 3.2.11.3 and 3.2.11.4, respectively. The thruster valve assembly shall be aligned along the axis of the catalyst bed and the exit nozzle shall be at a right angle to the axis of the catalyst bed.

Multiple nozzle configurations will be required to accommodate location-specific placement on the vehicle and contour matching of the vehicle outer mold line, as indicated in Sections 3.2 and 3.2.13.1.

3.2.13.1 Envelope Dimensions

The thruster shall fit within the envelopes defined by Figure 1, Figure 2, and Figure 3.

Figure 1. Thruster Envelope (center position)

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Figure 2. Thruster Envelope (right position)

Figure 3. Thruster Envelope (left position)

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Effective Date: 04-28-2021 Page 16 of 29

3.2.13.2 Inlet Tube

The propellant inlet tube shall be made from 304L stainless steel, aligned parallel to (preferably concentric with) the valve axis, and conform to the requirements identified in Figure 1, Figure 2, and Figure 3.

3.2.13.3 Propellant Filter

The thruster shall include a propellant filter at the inlet to the valve assembly with sufficient filtration to prevent loss of valve function and/or performance due to ingestion of contaminants.

3.2.13.4 Mass

The thruster shall have a mass of less than or equal to 0.926 lbm (420 grams), including lead wires.

3.2.13.5 Lead Wires

Valve and heater lead wires shall be a minimum of 18 inches long, as measured from the top of the valve containing the propellant interface to the vehicle.

3.2.13.5.1 Valve Wiring Configuration

Valve solenoids for each of the two valves in the assembly shall be wired in a parallel.

Length of Service

3.2.14.1 Shelf Life

The thruster shall meet the performance requirements of Section 3.2 for a minimum of ten (10) years after hardware delivery and prior to removal from packaging.

3.2.14.2 Service Life

The thruster shall meet the performance requirements of Section 3.2 for a minimum of ten (10) years after removal from packaging.

3.2.14.3 Pressure Exposure

The thruster shall meet the performance requirements of Section 3.2 when exposed to MEOP conditions, as stated in Section 3.2.2.1.2, for a minimum of 1 year.

3.2.14.4 Propellant Exposure

The thruster shall meet the performance requirements of Section 3.2 when exposed to propellant, as stated in Section 3.2.2, for a minimum of 1 year.

3.3 Component Interfaces

Fluid Interfaces

3.3.1.1 Vehicle Propellant Interface

The propellant interface to the vehicle shall be as defined in Figure 1, Figure 2, and Figure 3.

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Effective Date: 04-28-2021 Page 17 of 29

Structural Mounting Interface The thruster shall include a flange with through holes to enable hard-mounting to a mechanical surface on the spacecraft structure, as shown in Figure 4. The 3X bolt hole diameters and bolt circle diameter shown may be treated as TBR to accommodate heritage flange design.

Figure 4. Thruster Mounting Configuration

Nozzle Exit to Vehicle Interface The nozzle will interface to the vehicle outer mold line through a compliant gasket (TBR) to minimize loads transferred to the thruster.

Electrical Interfaces

3.3.4.1 Wiring

Thruster electrical wiring interfaces shall terminate with flying leads (TBR).

3.3.4.2 Electrical Schematics

The thruster electrical schematic shall be as defined in Figure TBS.

Rationale: Supplier to provide thruster electrical schematic.

Nozzle Centerline Axis The thruster nozzle centerline axis shall be oriented parallel to within ±0.5 degrees of the mounting surface.

Orientation The thruster will be mounted to the vehicle with the thruster valve positioned above the catalyst bed.

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Effective Date: 04-28-2021 Page 18 of 29

3.4 Environments

Propellant Inlet Temperature Range The thruster shall be operational over a propellant temperature range of +57.2 °F to +104 °F (+14 °C to +40 °C).

Operating Temperature The thruster shall meet the performance requirements of Section 3.2 during and after exposure to an operating environment temperature range of +41 °F to +250 °F (+5 °C to +121 °C).

Non-Operating Temperature The thruster shall meet the performance requirements of Section 3.2 after exposure to a non-operating (non-wetted) environment temperature range of -40 °F to +250 °F (-40 °C to +121 °C).

External Pressure The thruster shall permit operation at external pressures ranging from 0 psia to 16 psia (TBR) without incurring damage due to nozzle flow separation or other atmospheric pressure influences.

Humidity The thruster shall meet all performance requirements of Section 3.2 after exposure to humidity levels between 30 and 50%.

Random Vibration The thruster shall meet the leakage, structural, and operational requirements defined within this specification during and after exposure to the random vibration levels specified in Table 1 for a minimum of 6 minutes per axis for qualification and 1 minute per axis minimum for acceptance.

Table 1. Random Vibration Levels

Acceptance Qualification Frequency PSD PSD

(Hz) (G^2/Hz) (G^2/Hz) 20 0.01 0.02 60 0.63 2.50

215 0.63 2.50 260 0.25 1.00 360 0.25 1.00 450 0.04 0.07 650 0.04 0.07

2000 0.01 0.02 Grms 13.84 26.01

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Shock The thruster shall meet the internal leakage requirement specified in Section 3.2.12.2 after exposure to the shock levels specified in Table 2 in each of the three orthogonal axes. The thruster does not have to meet operational performance requirements after shock exposure.

Table 2. Shock Loads at Thruster Mounting Plate

Qualification Frequency Acceleration

(Hz) (g) 100 10

4100 8900 10000 8900

Inertial Loads The thruster shall meet the structural requirements specified in Sections 3.5.2 and 3.5.3 during exposure to inertial loads of ± 59.3 g’s.

3.5 Design and Construction

Material and Construction Standards The thruster shall comply with the requirements of NASA-STD-6016 “Standard Materials and Processes Requirements for Spacecraft” regarding materials and construction for a non-human-rated mission.

Strength and Life Assessments The thruster shall comply with the requirements of NASA-STD-5012 “Strength and Life Assessments Requirements for Liquid-Fueled Space Propulsion System Engines.”

Structural Design and Test Factors of Safety The thruster shall have a positive Margin of Safety (MS) for all yield and ultimate limit loads using the Factors of Safety (FS) defined in NASA-STD-5012 “Strength and Life Assessments Requirements for Liquid-Fueled Space Propulsion System Engines.”

Threaded Joints Threaded joints within the thruster shall meet the requirements of NASA-STD-5020 “Requirements for Threaded Fastening Systems in Spaceflight Hardware.”

Mechanisms The thruster shall meet the requirements of NASA-STD-5017 “Design and Development Requirements for Mechanisms.”

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Electrical Bonding The thruster shall meet NASA-STD-4003 “Electrical Bonding for NASA Launch Vehicles, Spacecraft, Payloads, and Flight Equipment,” Class H and S, measured between the thruster and mounting interface.

EEE Parts EEE parts used in the thruster shall be Grade 1 minimum per MSFC-STD-3012 “Electrical, Electronic, Electromechanical (EEE) Parts Management and Control Requirements for MSFC Space Flight Hardware.”

Fasteners Fasteners shall comply with NASA-STD-6008 “NASA Fastener Procurement, Receiving Inspection, and Storage Practices for Spaceflight Hardware” and MSFC-STD-486 “Standard, Threaded Fasteners, Torque Limits for.”

External Cleanliness The thruster external cleanliness shall meet the requirements of IEST-STD-CC1246 “Product Cleanliness Levels – Applications, Requirements, and Determination,” Level VC-0.5-1000 Visibly Clean (VC).

Internal Cleanliness The thruster internal cleanliness shall meet the requirements of IEST-STD-CC1246 “Product Cleanliness Levels – Applications, Requirements, and Determination,” Level 100 R1 Particulate Cleanliness, with the exception of areas designed for exposure to catalyst.

Fluid Compatibility The thruster shall be compatible with the following purge gases and cleaning fluids:

GN2 per MIL-PRF-27401 “Propellant Pressurizing Agent, Nitrogen” GHe per MIL-PRF-27407 “Propellant Pressurizing Agent, Helium,” Type 1, Grade A AR per MIL-PRF-27415C “Propellant Pressurizing Agent, Argon” Isopropyl Alcohol per TT-I-735 “Isopropyl Alcohol,” Grade A

Workmanship

3.5.12.1 Soldered Electrical Connections

Soldered electrical connections shall be in accordance with IPC J-STD-001S “Space and Military Applications Electronic Hardware Addendum to IPC J-STD-001.”

3.5.12.2 Crimping, Interconnecting Cables, Harness, & Wiring

Any crimping, interconnecting cables, and wiring shall be in accordance with NASA-STD-

8739.4 “Workmanship Standard for Crimping, Interconnecting Cables, Harnesses, and Wiring.”

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

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Effective Date: 04-28-2021 Page 21 of 29

Interchangeability The thruster shall be directly interchangeable in form, fit, and function with other thrusters of the same part number. Each of the three thruster nozzle configurations should be identified by a unique part number or dash number.

3.6 Part Marking and Identification

The thruster shall be identified per MIL-STD-130 “Identification Marking of U.S. Military Property” and, at a minimum, include the following:

Manufacturer Manufacturer Part Number Serial Number

3.7 Traceability

Serialization All non-standard parts shall be serialized with a unique serial number.

Records Records shall provide capability for backwards traceability to identify fabrication, inspection, processing, test, and operating records with traceability to the procurement documents.

3.8 Reliability

The thruster shall have a Mean Time To Failure (MTTF) of TBS hours.

4 VERIFICATION

Supplier shall conduct a verification program that demonstrates the hardware design is qualified and meets all requirements contained in this document. Supplier shall provide documentation showing conformance to this specification.

4.1 Methods of Verification

Verification methods define the approach to be utilized to show compliance with a design or performance parameter. Requirements defined within this specification shall be verified by the methods described in Sections 4.1.1 through 4.1.4.

Inspection (I) Verification by inspection may include visual inspection of the physical hardware, physical measurement of a property of the hardware, or a documentation search demonstrating hardware of an identical design has demonstrated fulfillment of a requirement.

Demonstration (D) Verification by demonstration is the exercise of the hardware, software or operation to assure that qualitatively specified functions can be performed. Demonstration generally does not entail special equipment or sophisticated measuring instrumentation.

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Document No: ER13-SPEC-0001 Revision: Rev.03

Effective Date: 04-28-2021 Page 22 of 29

Analysis (A) Verification by analysis involves the use of engineering analysis, qualitative assessment, computer modeling and/or simulation to ensure compliance to requirement(s). If necessary, analysis combined with test results may be used to provide this evidence.

Test (T) Verification by test (e.g., functional, environmental) is the actual operation to ensure that performance characteristics of the hardware are in accordance with the requirement(s). Testing will also include methods to verify requirements such as material compatibility, which may or may not be performed with an end item. All acceptance and qualification testing shall be performed in accordance with NASA-STD-5012 “Strength and Life Assessments Requirements for Liquid-Fueled Space Propulsion System Engines,” NASA-STD-5017 “Design and Development Requirements for Mechanisms,” and SMC-S-016 “Test Requirements for Launch, Upper-Stage and Space Vehicles,” using the bounding test criteria for each test.

Hot-fire testing should be performed in a test cell environment simulating vacuum conditions to the lowest pressure possible.

4.1.4.1 Test Restrictions

4.1.4.1.1 Test Equipment Failure during Testing

The test will be stopped if test equipment fails during testing in cases where this failure will result in damage to the thruster. A complete record shall be maintained and included in the test report.

4.1.4.1.2 Thruster Failure during Testing

In the event of a thruster failure during test, the test will be stopped, the configuration made safe and maintained to the extent practical, and the purchaser notified. A complete record shall be maintained and included in the test report.

4.1.4.1.3 Modification of Thruster

Modification of the thruster shall not be permitted once acceptance and/or qualification testing has started.

4.1.4.1.4 Restriction on Freon Use

Freon or other halogenated solvent shall not be used in the vicinity of the thruster, as it can poison the catalyst bed.

4.1.4.1.5 Lubricant Use

The use of fluorinated greases shall be avoided or minimized.

4.1.4.1.6 Catalyst and Valve Seat Temperature Exposure

The catalyst shall not be exposed to temperatures above TBS °F in air, nor shall the valve be exposed to temperatures above TBS °F in air.

Rationale: Supplier to provide temperatures at which the catalyst and valve, respectively, become susceptible to oxidation or other detrimental effects.

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Effective Date: 04-28-2021 Page 24 of 29

4.1.4.2.2.2 Thermal Testing

Thermal testing shall be conducted in accordance with SMC-S-016 “Test Requirements for Launch, Upper-Stage and Space Vehicles.”

For the thruster valve, thermal cycle tests shall be performed in lieu of thermal vacuum testing.

Non-operating temperature, one cycle, to be performed prior to operating temperature testing.

Hot and cold temperature plateaus shall be held for a minimum of four (4) hours at the acceptance temperature levels. The “hot and cold temperature plateaus” are 19.8 °F (11 °C) beyond the operating and non-operating temperatures shown in Sections 3.4.2 and 3.4.3.

Pressurization of the valve when not performing leakage tests is not required.

For the thruster assembly, thermal vacuum requirements may be satisfied with hot-fire testing.

4.1.4.2.3 Random Vibration

Leakage shall be monitored during random vibration testing, with the maximum expected operating pressure (MEOP), as specified in Section 3.2.2.1.2, supplied at the valve inlet.

Catalyst attrition shall be determined by comparing pre and post-test thruster weight. Structural margins shall be determined analytically at predicted operating temperatures. Hot fire testing shall follow random vibration testing to verify that thruster performance meets the requirements of Section 3.2 following vibration.

4.1.4.3 Qualification Testing

Perform qualification testing on each thruster per Section 4.1.4, with tailoring and clarification to test requirements defined in the sections below.

Testing of subcomponent such as valves, heaters, and instrumentation can be performed at a lower level, as appropriate.

A minimum of one (TBR) thruster shall undergo qualification testing. Should additional units be required, each unit shall be identical in design, materials, and manufacture to the flight-rated design.

Rationale: NASA-STD-5012B specifies a minimum of one qualification unit for small, pressure-fed engines. The required number of qualification units will be resolved with input from the supplier, based on current hardware capability relative to the requirements being imposed within this document.

4.1.4.3.1 Performance Variation across Range of Operating Conditions Where Section 3 requirements specify performance across the operational range of the thruster as TBS, the supplier shall provide the information identified to the customer based on measured data obtained during testing and shall include error bands as appropriate.

4.1.4.3.2 Proof and Burst Pressure

Proof and burst pressure tests, identified in Sections 3.2.2.1.4 and 3.2.2.1.5, respectively, shall include the required environmental correction factor (ECF) if not conducted at planned operating temperature, as specified in NASA-STD-5012 “Strength and Life Assessments Requirements for Liquid-Fueled Space Propulsion System Engines.”

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Effective Date: 04-28-2021 Page 25 of 29

4.1.4.3.3 Thermal

4.1.4.3.3.1 Thermal Test Margin

Temperatures indicated in Sections 3.4.2 and 3.4.3 reflect expected conditions. A 37.8 °F (21 °C) test margin shall be applied to both the upper and lower limits for qualification testing, per SMC-S-016 “Test Requirements for Launch, Upper-Stage and Space Vehicles,” unless otherwise indicated in the verification requirement.

4.1.4.3.3.2 Thermal Testing

Thermal testing shall be conducted in accordance with SMC-S-016 “Test Requirements for Launch, Upper-Stage and Space Vehicles.”

For the thruster valve, thermal cycle tests shall be performed in lieu of thermal vacuum testing.

Non-operating temperature, one cycle, to be performed prior to operating temperature testing.

Hot and cold temperature plateaus shall be held for a minimum of four (4) hours at the qualification temperature levels. The “hot and cold temperature plateaus” are 37.8 °F (21 °C) beyond the operating and non-operating temperatures shown in Sections 3.4.2 and 3.4.3. Flight configuration leakage testing shall be performed at the first, third and last thermal cycles as part of the minimum and maximum temperature dwells in order to detect performance degradation.

Pressurization of the valve when not performing leakage tests is not required.

Thruster thermal vacuum requirement may be satisfied with hot-fire testing.

4.1.4.3.4 Random Vibration

Leakage shall be monitored during random vibration testing, with the maximum expected operating pressure (MEOP), as specified in Section 3.2.2.1.2, supplied at the valve inlet.

Catalyst attrition shall be determined by comparing pre and post-test thruster weight. Structural margins shall be determined analytically at predicted operating temperatures. Hot fire testing shall follow random vibration testing to verify that thruster performance meets the requirements of Section 3.2 following vibration testing.

4.1.4.3.5 Shock Exposure

The thruster propellant inlet shall be pressurized to MEOP (TBR) during shock exposure to quantify any leakage past the propellant valves, during or after the shock event, that could lead to hydrazine contamination of the surrounding environment.

4.1.4.3.6 External Pressure

The thruster will be required to operate in ambient pressure environments ranging from vacuum conditions to 16 psia without suffering detrimental effects. This is to ensure that demonstration flights can be conducted within earth’s atmosphere (without suffering damage due to excessive nozzle separation, for example) prior to mission launch.

4.1.4.3.7 Impulse Life

At least one qualification thruster unit shall demonstrate 2X mission life during qualification testing, while meeting the performance requirements provided in Section 3.2.

Specification for Mars Ascent Vehicle Reaction Control System 5N Thruster

Document No: ER13-SPEC-0001 Revision: Rev.03

Effective Date: 04-28-2021 Page 29 of 29

An identification tag shall be placed between the inner and outer bags displaying the statement 'Open in Clean Room Environment only.'

A humidity indicator shall be placed between the inner and outer bags.

Storage The thruster shall be designed for storage for a period of time equal to or exceeding the durations stated in Sections 3.2.14.1 and 3.2.14.2 without requiring repair, maintenance, or retesting at the end of storage.

Lead Wires Exercise caution to prevent damage to lead wires. Avoid unnecessary flexing of the leads during thruster handling.