MPCV_72634-_Approved_for_Release.pdf

MPCV 72634

REVISION DRAFT

National Aeronautics and RELEASE DATE: JUNE 19, 2019 Space Administration

ORION MULTI-PURPOSE CREW VEHICLE (MPCV)

PROGRAM ORION MAIN ENGINE PROJECT

TECHNICAL REQUIREMENTS SPECIFICATION

(PTRS)

Approved for Public Release

Revision: B Document No: MPCV 72634 Release Date: 06/19/2017 Page: 2 of 157 Title: Orion Multi-purpose Crew Vehicle (MPCV) Program Orion Main Engine Project Technical Requirements Specification (PTRS)

REVISION AND HISTORY PAGE

Status Revision Change No.

Description Release Date

MPCV 72634 OME PTRS Revision History

Draft Draft Draft for release 06/19/19

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TABLE OF CONTENTS

SECTION PAGE

1.0 INTRODUCTION

1.1 PURPOSE AND SCOPE

1.2 REQUIREMENTS NOMENCLATURE

1.3 RELATIONSHIP TO PROGRAM REQUIREMENTS

1.4 REQUIREMENT NUMBERING

1.5 RESPONSIBILITY AND CHANGE AUTHORITY

1.6 MEASUREMENT UNITS

1.6.1 Tolerances and Conversions

2.0 APPLICABLE AND REFERENCE DOCUMENTS

2.1 APPLICABLE DOCUMENTS

2.2 REFERENCE DOCUMENTS

2.3 ORDER OF PRECEDENCE

3.0 REQUIREMENTS

3.1 MISSION OVERVIEW

3.2 SYSTEM OVERVIEW

3.3 CHARACTERISTICS

3.3.1 Functional/Performance Requirements

3.3.2 Physical Requirements

3.3.3 Environmental Requirements

3.3.4 Reliability and Maintainability

3.3.5 Life 39

3.3.6 Safety Requirements

3.4 DESIGN AND CONSTRUCTION

3.4.1 Materials and Processes

3.4.2 Structural Design

3.4.3 Electromagnetic Environments

3.4.4 Nameplates or Product Markings

3.4.6 Interchangeability

3.4.8 Human Engineering

3.4.9 Reusability

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3.4.10 Wire, Harness and Connectors

3.4.11 Standards and Other Standardized Requirements

3.5 LOGISTICS

3.6 GSE 60

3.7 INTERFACE REQUIREMENTS

3.8 SOFTWARE

4.0 PREPARATION FOR DELIVERY

4.1 PROTECTIVE PACKAGING

4.2 PACKAGE MARKING

5.0 CUSTOMER IMPOSED VERIFICATION REQUIREMENTS

5.1 DEFINITIONS

5.2 VERIFICATION REQUIREMENTS

5.2.1 Overview

5.2.2 Methods

5.2.3 Unique Test Requirements

5.3 ACCEPTANCE AND QUALIFICATION

5.3.1 General Requirements and Guidelines for Test

5.3.1.1 Test Baseline Definition and Tailoring

5.3.1.2 Acceptance Testing

5.3.1.3 Qualification Testing

5.3.1.4 Development Testing

5.3.1.5 Testing in Support of Troubleshooting

5.3.1.6 Instrumentation and DATA Acquisition and Processing

5.3.1.7 Inspection of Hardware Pre and Post Test

5.3.1.8 TEST PROBLEM REPORTING

5.3.1.9 TEST DOCUMENTS

5.3.1.10 TEST TOLERANCES

5.3.1.11 Alternative Strategies to the Standard Qualification and Acceptance Test Program

5.3.2 Unit Test Baseline

5.3.2.1 Test Requirements and Sequencing

5.3.2.2 UNIT Functional AND Performance Tests

5.3.2.3 Unit Leak Test

5.3.2.4 Unit Shock Test

5.3.2.5 Unit Random Vibration Test

5.3.2.6 Unit Acoustic Vibration Test

5.3.2.7 Unit Sinusoidal Vibration Test

5.3.2.8 Unit Acceleration Test

5.3.2.9 Unit Thermal Cycle Test

5.3.2.10 Unit Thermal Vacuum Test

5.3.2.11 Unit Thermal Gradient Test

5.3.2.12 Unit Depressurization/Repressurization Test

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5.3.2.13 Climatic Tests

5.3.2.14 Explosive Atmosphere

5.3.2.15 Unit Oxygen Compatibility Acceptance Test

5.3.2.16 Unit Life Test

5.3.3 Major Assembly Test Baseline

5.3.4 Requirements and Guidelines for Retest

5.4 RESERVED

APPENDIX A ACRONYMS, ABBREVIATIONS AND GLOSARY OF TERMS

A1.0 ACRONYMS AND ABBREVIATIONS

A2.0 GLOSSARY OF TERMS

APPENDIX B OPEN WORK

B1.0 TO BE DETERMINED

B2.0 TO BE RESOLVED

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LIST OF TABLES PAGE

3.3.2-1 OME CENTER OF GRAVITY

3.3.3-1 TRANSPORTATION LOADS IN THE USA AND BETWEEN EUROPE AND THE USA

3.3.3-2 TRANSPORTATION LOADS IN EUROPE

3.3.3-3 HANDLING LOADS

3.3.3-4 THERMAL TRANSPORTATION AND STORAGE ENVIRONMENT

3.3.3-5 ENGINE QUASI-STATIC LOADS (LIMIT LOADS)

3.3.3-6 ENGINE LOW FREQUENCY SINE ENVIRONMENT (LIMIT LOADS)

3.3.3-7 ENGINE ACOUSTIC ENVIRONMENT

3.3.3-8 ENGINE SHOCK ENVIRONMENT

3.3.3-9 ESM INTERNAL THERMAL CONDITIONS

3.3.3-10 GIMBAL OPERATIONAL LOADS

3.3.5-1 ENGINE LIFE BY PHASE

3.4.1-1 FLUIDS COMPATIBILITY

3.4.2-1 MINIMUM FACTORS OF SAFETY FOR PRESSURIZED HARDWARE

3.4.2-2 MINIMUM FACTORS OF SAFETY FOR PRESSURIZED HARDWARE (APPLICABLE

TO PRIMARY AND SECONDARY STRUCTURES ONLY)

3.4.2-3 LOAD COMBINATION CRITERIA FOR COMPONENTS

3.4.2-4 MATERIAL THICKNESS FOR STRESS CALCULATIONS

3.4.2-5 FASTENER REQUIREMENTS HERITAGE AND TRACEABILITY

3.6-1 FACTORS OF SAFETY FOR PROOF LOAD

3.7-1 DYNATUBE FITTING PARTS AND HARDWARE OWNERSHIP

3.7-2 AVAILABLE COMMAND INTERFACES

3.7-3 AVAILABLE COMMAND INTERFACE DETAILS

3.7-4 AVAILABLE INSTRUMENTATION INTERFACES

3.7-5 POSITION INDICATOR SWITCH ELECTRICAL INTERFACE DETAILS

3.7-6 TYPE 1 PRESSURE SENSOR ELECTRICAL INTERFACE DETAILS

3.7-7 TYPE 2 PRESSURE SENSOR ELECTRICAL INTERFACE DETAILS

3.7-8 TYPE 2 PRESSURE SENSOR VARIATIONS

3.7-9 TYPE 1 TEMPERATURE SENSOR ELECTRICAL INTERFACE DETAILS

3.7-10 TYPE 2 TEMPERATURE SENSOR ELECTRICAL INTERFACE DETAILS

3.7-11 LVDT ELECTRICAL INTERFACE DETAILS

3.7-12 HEATER ELECTRICAL INTERFACE DETAILS

5.3.1-1 TEST CONDITION TOLERANCES

5.3.2-1 UNIT QUALIFICATION TEST BASELINE

5.3.2-2 UNIT ACCEPTANCE TEST BASELINE

5.3.2-3 UNIT QUALIFICATION RECOMMENDED TEST SEQUENCE

5.3.2-4 UNIT ACCEPTANCE RECOMMENDED TEST SEQUENCE

5.3.2-5 MINIMUM UNIT AVT WORKMANSHIP SCREENING LEVEL

A1.0-1 ACRONYMS AND ABBREVIATIONS

A2.0-1 GLOSSARY OF TERMS

B1.0-1 TO BE DETERMINED ISSUES

B2.0-1 TO BE RESOLVED ISSUES

B2.0-2 VERIFICATION CROSS REFERENCE MATRIX

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LIST OF FIGURES PAGE

1.3-1 ORION MAIN ENGINE REQUIREMENT FLOW

3.1-1 EXAMPLE ORION MISSION

3.2-1 ORION-SYSTEM

3.2-2 ORION-HERITAGE OMS-E (REF)

3.3.1-1 ENGINE INLET PRESSURES AT START (ENVELOPE)

3.3.1-2 ENGINE INLET PRESSURES AT START (CORNER POINTS)

3.3.1-3 ENGINE OPERATING PRESSURES ENVELOPE

3.3.1-4 ENGINE OPERATING PRESSURES (CORNER POINTS)

3.3.1-5 ENGINE GIMBAL DEFLECTION RANGE

3.3.2-1 ESM COORDINATES

3.3.3-1 ESM EXTERNAL THERMAL ENVIRONMENT WITH MAIN ENGINE IN STRAIGHT

POSITION

3.3.3-2 ESM EXTERNAL THERMAL ENVIRONMENT WITH MAIN ENGINE IN CANTED

POSITION

3.7-1 VOLUME ALLOCATION FOR OME (INCLUDING ENGINE GIMBALING) (DIMENSIONS

IN INCHES)

3.7-2 VOLUME ALLOCATION FOR OME WITH DETAIL FOR THRUST MOUNT INTERFACES

(DIMENSIONS IN INCHES)

3.7-3 ENGINE MOUNT PAD GEOMETRY (DIMENSIONS IN INCHES)

3.7-4 FIGURE TITLE PITCH AND YAW ACTUATOR MOUNTING GEOMETRY. ENGINE

GIMBAL POINT

3.7-5 PITCH AND YAW ACTUATOR ENGINE ATTACHMENT GEOMETRY (DIMENSIONS IN

INCHES)

3.7-6 GAS AND FLUID INTERFACE LOCATIONS (DIMENSIONS IN INCHES)

3.7-7 FUEL AND OXIDIZER INTERFACE PANEL GEOMETRY, ESM SIDE. FOR

REFERENCE ONLY

3.7-8 GAS AND FLUID INTERFACE LOCATIONS (DIMENSIONS IN INCHES)

3.7-9 GAS, FLUID, AND ELECTRICAL CONNECTOR LOCATIONS

3.7-10 FUEL AND OXIDIZER INLET LINE CONNECTION GEOMETRY

3.7-11 FUEL AND OXIDIZER INLET LINE CONNECTION GEOMETRY

3.7-12 DYNATUBE FITTING LOCATIONS ON INTERFACE BRACKETS

3.7-13 DYNATUBE FITTING DRAWING

3.7-14 THRUST VECTOR MISALIGNMENT AND OFFSET MOMENT

5.3.2-1 UNIT MINIMUM WORKMANSHIP VIBRATION TEST SPECTRUM

5.3.2-2 QUALIFICATION THERMAL CYCLE TEST

5.3.2-3 ACCEPTANCE THERMAL CYCLE TEST

5.3.2-4 QUALIFICATION THERMAL VACUUM TEST

5.3.2-5 SAND, DUST, OR REGOLITH TEST REQUIREMENTS

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1.0 INTRODUCTION

The Orion Main Engine (OME) project is an effort to procure a replacement Orion main engine replacing for the legacy Space Shuttle Orbital Maneuvering System-Engine (OMSE). The OME is being designed and delivered under the Orion Multi-Purpose Crew Vehicle (MPCV) Program, delivered as Government Furnished Equipment (GFE) to the European Space Agency Orion Prime Contractor and subsequently to the MPCV Orion Prime Contractor. The OME project will design, build and deliver a system that will provide propulsive capability for the Orion spacecraft in order to meet desired mission objectives

1.1 PURPOSE AND SCOPE

This specification defines the OME project’s generic and unique design and performance requirements pertaining to its application on the Orion MPCV. This specification defines OME system requirements pertaining to performance, design, construction, and handling of the hardware and delineates the methods by which the OME developer will develop, test, and certify the hardware. Detailed facility, acceptance, and qualification test plans and procedures will be generated separately during the development process, as required. Interface requirements between the OME project and the Orion European Service Module (ESM) are contained herein. The verification requirements for the OME system hardware are contained in section 5, customer imposed verification requirements.

1.2 REQUIREMENTS NOMENCLATURE

The following definitions differentiate between requirements and other statements.

Shall: This is the only verb used for the binding requirements.

Should/May: These verbs are used for stating non-mandatory goals.

Will: This verb is used for stating facts or declaration of purpose.

Rationale is provided for requirements. The rationale is not binding and is only used to provide additional information or examples in order to clarify intent of the requirement. In the event of an inconsistency between the requirement and the rationale, the requirement is binding and takes precedence.

In some cases, the values of quantities included in a requirement have not been determined and are designated as either “To Be Resolved” (TBR) or “To Be Determined” (TBD). Where approximate values of such quantities are known and provide useful guides for development, these values are shown along with the TBR notation. Where no value is known, a TBD will be used as a placeholder.

1.3 RELATIONSHIP TO PROGRAM REQUIREMENTS

The OME PTRS requirements are derived/flow down from higher level requirements documents representing the MPCV Program System Requirements Document and flowing to the Prime Contractor Integrated vehicle requirements, and ESM requirements. This document in conjunction with the CEV-T- 031420, TVC2 Specification and the OME/TVC2ESM ICD (RIBRE-TBD) represent primary governing documents for the integrated TVC/OME/ESM system.

The relationship between MPCV 72634, the OME PTRS, and the other MPCV program Requirements documents is shown in Figure 1.3-1.

CEV T 031206

Subsystem Spec

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MPCV SRD

MPCV-72000

ESM SRD

MPCV-72512

OME/TVC2/ESM

ICD (RIBRE-00yy

Program Documents

Prime Contractor

TVC2 Documents

OME Documents

MPCV 72536

MPCV to ESM

Prop and GN&C

IRD

MPCV-RIBRE-SPE-0001

ESM System Specification AB

MPCV-RIBRE-SPE-0004

MPCV-SM Prop Subsystem

Specification AB

CM Prop LM

CEV-T-031210

Subsystem Spec

GN&C LM

CEV-T-031570

CEV System Requirements

Specification - MPCV System Requirements LM

MPCV 72634

OME PTRS NASA

CEV-T-031206

Subsystem Spec Prop (RFA-0067)

CEV T 031420

TVC2 Spec LM

AB

NASA

NASA

NASA

FIGURE 1.3-1 - ORION MAIN ENGINE REQUIREMENT FLOW

1.4 REQUIREMENT NUMBERING

In an effort to keep requirement traceability more clear and achievable (particularly viewing the requirements traceability table and verification and validation matrices), each requirement contained in this specification is denoted by a unique identifier in the following format:

OME.XXXXX

Where legacy MPCV requirements are incorporated in this specification the numbering format will include the unique requirement number appended with OME. For example

OME.SDVR0005

represents the SDVR requirement from MPCV 70135 with the OME prefix.

1.5 RESPONSIBILITY AND CHANGE AUTHORITY

Proposed changes to this document shall be submitted by an Orion Multi-Purpose Crew Vehicle (MPCV) Program Change Request (CR) to the appropriate MPCV Board for consideration and disposition per the Program Technical Attributes Listing.

All such requests will adhere to the MPCV Program Configuration Management Change Process.

The appropriate NASA Office of Primary Responsibility (OPR) identified for this document is identified within the Program Technical Attributes Listing. The NASA OPR is the Crew and Service Module Office.

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1.6 MEASUREMENT UNITS

Units will be indicated in English units, followed by metric SI units in parenthesis to the maximum extent possible.

1.6.1 Tolerances and Conversions

The MPCV Program will utilize NIST SP811, Guide for the Use of the International System of Units (SI) for standardization and conversion of the units of measure. The document provides guidance on the use of SI units, including the notation and conversion.[

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2.0 APPLICABLE AND REFERENCE DOCUMENTS

2.1 APPLICABLE DOCUMENTS

Document Number Document Title

ANSI/AIAA S-080 Standard for Space Systems - Metallic Pressure Vessels, Pressurized Structures, and Pressure Components

ANSI/AIAA S-081 Standard for Space Systems - Composite Overwrapped Pressure Vessels (COPV)

CEV-T-025000 CEV Electromagnetic Compatibility Control and Verification Document.

CEV-T-027000 Electrical, Electronic, and Electromechanical Control, Management and Implementation Plan.

CPIA 655 Combustion Stability Specification and Verification Procedure

LM-ORN-0995 Multi-Purpose Crew Vehicle (MPCV) Vehicle Configuration and Performance (VCAP) Data Book

MPCV 70038 Orion Multi-Purpose Crew Vehicle (MPCV) Program: Hazard Analyses Requirements

MPCV 70135 Orion Multi-Purpose Crew Vehicle (MPCV) Program: Structural Design and Verification Requirements

MPCV 70156 Cross Program Fluid Procurement and Use Control Specification

MPCV 72506 Orion Multi-Purpose Crew Vehicle (MPCV) Standard for the Design and Fabrication of Ground Support Equipment

NASA-STD-5012 Strength And Life Assessment Requirements For Liquid Fueled Space Propulsion System Engines

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 NASA Standard Materials and Processes Requirements for Spacecraft

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Document Number Document Title

NASA-STD-(I)-6016 Standard Materials and Processes Requirements for Spacecraft

NSTS 08307A Criteria for Preloaded Bolts

SLS-PLAN-062 SLSP Integrated Vehicle Loads Control Plan

SMC-S-025 Evaluation and Test Requirements for Liquid Rocket Engines

2.2 REFERENCE DOCUMENTS

Document Number Document Title

MIL-STD-810F Test Method Standard for Environmental Engineering Considerations and Laboratory Tests, Method 508

MPCV 70043 Orion Multi-Purpose Crew Vehicle (MPCV) Program: Hardware Failure Modes and Effects Analysis/Critical Items List (FMEA/CIL) Requirements Document

MPCV 70080 Cross Program Electromagnetic Environmental Effects (E3) Requirements Document

MPCV 70135 Orion Multi-Purpose Crew Vehicle (MPCV) Program: Structural Design and Verification Requirements

MPCV 72554 Orion Multi-Purpose Crew Vehicle (MPCV) Program: OMS-E and TVC Design Description and Interface Document

MPCV-AS-TN-0012 Preliminary data for auxiliary thrusters and OMS-E thermal environment

MSFC-DWG-20M02540 Assessment of Flexible Lines for Flow Induced Vibration

MSFC-PROC-404 Marshall Space Flight Center: Procedure, Gases, Drying and Preservation Cleanliness Level and Inspection Modes

MSFC-SPEC-626 Test Control Document for Assessment of Flexible Lines for Flow Induced Vibration

NASA-HDBK-7005 Dynamic Environmental Criteria

NASA-STD-5020 Requirements for Threaded Fastening Systems in Spaceflight Hardware

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Document Number Document Title

NASA-STD-6016 NASA Standard Materials and Processes Requirements for Spacecraft

NASA-STD-(I)-5019 Fracture Control Requirements for Spaceflight Hardware

NSTS 08123 Certification of Flex Hoses and Bellows for Flow Induced Vibration

NSTS 08307A Criteria for Preloaded Bolts

SLS-PLAN-062 SLSP Integrated Vehicle Loads Control Plan

SLS-SPEC-044-07 Cross Program Vehicle Design Environments Volume 7: Natural Environments

SMC-S-005 AFSC Space And Missile Systems Center Standard: Space Systems - Flight Pressurized Systems

2.3 ORDER OF PRECEDENCE

In the event of a conflict between the text of this document and the references cited herein, the text of this document takes precedence. Nothing in this document, however, supersedes applicable laws and regulations unless a specific exemption has been obtained.

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3.0 REQUIREMENTS

3.1 MISSION OVERVIEW

The Orion spacecraft is intended to carry astronauts to destinations beyond Low Earth Orbit (LEO). The vehicle will be capable of conducting regular in-space operations in conjunction with payloads delivered by the Space Launch System (SLS) for all missions. Orion’s active mission including transit to and from destinations may be up to 21 days in duration, but can be extended up to 210 days, which includes a time period of quiescent operations, while docked to a habitation element. While the Orion provides all services necessary to support as many as 4 crewmembers for an active mission of 21 days, actual mission crew complement and duration may vary based on mission design.

The European Service Module (ESM) serves as Orion’s primary source for power, life support consumables, and propulsion. The ESM propulsion system provides translation burns and attitude control maneuvers for Orion. Orion performs different types of translation burns to support the specific mission design, such as trans-lunar injection, outbound and return trajectory correction burns, lunar orbit insertion burns and departure burns. These translation burns may be done with the ESM OME or the ESM auxiliary thrusters and in limited circumstances the ESM Reaction Control System (RCS) thrusters.

Attitude control is accomplished with the ESM RCS thrusters. An example mission is illustrated in Figure 3.1-1.

Orion uses the ESM OME to conduct a deorbit burn (in the event translunar injection has not occurred) or a trans-earth injection burn to return the Orion spacecraft to Earth in the case of a mission abort while in space. The timing of these burns varies relative to the time of abort initiation. The OME may be used to achieve an alternate mission orbit rather than aborting when the destination orbit cannot be achieved,. In these scenarios one or more burns would be required to alter the trajectory and would include a deorbit burn at the end of the mission.

In addition to an On-orbit Abort, OME is also used during ascent in an abort scenario known as an abort-to-orbit (ATO). In an ATO, the ESM OME is used to achieve a safe orbit and later to conduct a deorbit burn to return the Orion to Earth.

FIGURE 3.1-1 - EXAMPLE ORION MISSION

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3.2 SYSTEM OVERVIEW

The Orion consists of the Crew Module (CM), Service Module (SM), and Launch Abort Systems (LAS) (Figure 3.2-1). The CM provides a habitable pressurized volume to support crewmembers and cargo during all elements of a given mission from Launch Operations to Earth Entry, Descent and Landing (EDL), and Recovery. The SM is comprised of 4 components, the Crew Module Adaptor (CMA), the European Service Module (ESM), the Spacecraft Adapter (SA), and the Spacecraft Adapter Jettisonable Fairings (SAJ). The SM provides services to the CM in the form of propulsion, consumables storage, heat rejection and power generation. The LAS provides an abort capability to safely transport the CM away from the launch vehicle stack in the event of an emergency on the launch pad or during ascent.

The ESM Propulsion system is the main Orion propulsion system and provides nominal 6DOF translational and rotational control of the Crew and Service Module (CSM) from separation with the SLS Upper stage until CM-SM separation prior to Earth reentry. For early Orion missions, the Shuttle heritage OMS-E provides the main propulsive force for long-duration on-orbit burns which require substantial delta-V. This heritage OMS-E, shown in Figure 3.2-2, will transition to the newly developed OME, providing the same capabilities and performance. The thrust vector is controlled by gimbaling the main engine in the yaw and pitch directions with the Thrust Vector Control (TVC) system. The TVC system is existing and the new OME will support the existing TVC interfaces.

The OME is a pressure-fed, hypergolic-reacting bipropellant, fixed thrust engine. The engine uses monomethylhydrazine (CH3NHNH2, also known as MMH) as the fuel and nitrogen tetroxide with 3% mixed oxides of nitrogen (N2O4, also known as MON-3) as the oxidizer. The OME TVC assembly (existing) gimbals the engine to provide two-axis (pitch and yaw) thrust vector control. Roll control is provided by the RCS system.

The OME is controlled by the Propulsion Drive Electronics (PDE), which serve as the central communication system/interface between the CM/CMA command and control systems and the SM propulsion system actuators and sensors. All guidance information calculated on the CM is distributed via the onboard data network (ODN), to the PDE, which translates the received commands into electrical pulses in order to operate valves (e.g. start/stop thrusters) and which scans propulsion system sensor values. PDEs are single fault tolerant at the box-level with each box containing two channels for controlling the pressurization system and downstream thrusters.

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FIGURE 3.2-1 - ORION-SYSTEM

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FIGURE 3.2-2 - ORION-HERITAGE OMS-E (REF)

3.3 CHARACTERISTICS

3.3.1 Functional/Performance Requirements

OME.0001 Fuel and Oxidizer

The engine shall perform as specified herein using monomethylhydrazine (MMH) fuel and nitrogen-tetroxide (MON-3) oxidizer per MPCV 70156 Cross Program Fluid Procurement and Use Control Specification.

Rationale: This requirement defines the physical and chemical characteristics of the propellant and oxidizer that are used in the ESM and therefore will be used for the ESM main engine.

OME.0003 Propellant Saturation

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The engine shall meet its performance requirements with propellants saturated from 0 to 100 percent with helium per MPCV 70156 Cross Program Fluid Procurement and Use Control Specification.

Rationale: Helium gas is used to pressurize the propellant tanks. The helium gas may saturate the propellants.

OME.0004 Propellant Inlet Pressure

The engine shall be capable of operating at pressure levels at the engine inlet as shown in Figure Engine Inlet at Start (envelope) and Figure Engine Inlet at Start (corner points) for the start box and Figure Engine operating Pressures (Envelope) and Figure Engine Operating Pressures (Corner Points) the operating box.

Rationale: This requirement is derived from the existing ESM performance capability and mission analysis, and therefore the OME must function within these ESM system parameters. The ESM system includes a common propellant storage element supplying 3 different engine types which must function as a system, with this operating range analyzed for the OME.

FIGURE 3.3.1-1 - ENGINE INLET PRESSURES AT START (ENVELOPE)

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FIGURE 3.3.1-2 - ENGINE INLET PRESSURES AT START (CORNER POINTS)

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FIGURE 3.3.1-3 - ENGINE OPERATING PRESSURES ENVELOPE

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FIGURE 3.3.1-4 - ENGINE OPERATING PRESSURES (CORNER POINTS)

OME.0005 Propellant Inlet Temperature

The engine shall be capable of operating at propellant temperature levels at the engine inlet as shown below, excluding the effect of engine operation on post shutdown heat soakback.

nominal:

70 ⁰F (21.1 ⁰C)

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30 ⁰F (-1.12 ⁰C) to 125 ⁰F (51.6 ⁰C) steady- state:

70 ⁰F +/- 30 ⁰F (21.1 ⁰C +/- 16.6 ⁰C)

Rationale: This requirement is derived from the existing ESM performance capability, and therefore the OME must function within the ESM system parameters.

OME.0006 Propellant Inlet Temperature Difference

The engine shall operate with fuel delivered to the engine inlet within +/- 10 ⁰F (+/- 5.5 ⁰C) of the oxidizer temperature during steady state operation.

Rationale: This requirement is derived from the existing ESM performance capability, and therefore the OME must function within the ESM system parameters.

OME.0007 Operational Altitude

The engine shall operate from altitudes of 300,000 ft (91.4 km) to space vacuum.

Rationale: This altitude represents the actual minimum altitude for flight operations of approximately 300,000 ft. Test conditions may vary below 300,000 ft as necessary to accommodate test facility limitations.

OME.0008 Propellant Valve Seal Leakage

The internal gaseous Helium leakage shall not exceed 70 SCCH with any differential pressure, between 0.0 psi (0.0 bar) to MEOP, applied across any propellant valve.

Rationale: This requirement is derived from the existing ESM performance capability, and therefore the OME must function within the ESM system parameters.

OME.0009 Actuation System Leakage

The internal leakage shall not exceed 10 scch of gaseous Nitrogen with any differential pressure between 0.0 psi (0.0 bar) and MEOP applied across any engine valve pneumatic actuation system component.

Rationale: This requirement is derived from the existing ESM performance capability, and therefore the OME must function within the ESM system parameters.

OME.0010 External Leakage

External leakage from any component, component/ interconnecting hardware, flange joint and metallurgical joints shall not exceed 1x 10-4 (SCCS) of gaseous Helium with any internal pressure from 0.0 psig (0.0 bar) to MEOP at any temperature specified herein.

Rationale: This is tied to standard leak requirements for mechanical joints.

OME.0011 Vacuum Thrust

The engine shall have a steady-state vacuum thrust of 6000 lbf (26689 N) with a tolerance of +200/-170 lbf (+889N/-756 N) at normal operating pressure as shown in Figure Engine Operating Pressures (Envelope)

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Rationale: This requirement is derived from the existing ESM performance capability and mission performance analysis.

OME.0012 Specific Impulse

The engine shall have a specific impulse greater than 315 s at normal operating conditions per Figures: Engine operating Pressures (Envelope) and Engine Operating Pressures (Corner Points).

Rationale: This requirement is derived from the existing ESM performance capability and mission performance analysis.

OME.0013 Mixture Ratio

The engine shall operate with a nominal oxidizer to fuel mixture ratio (by weight) of 1.65 with a tolerance of +/- 0.03 at normal operating pressure as shown in Figure: Engine Operating Pressures (Envelope).

Rationale: This requirement is derived from the existing ESM performance capability and propellant system sizing for an overall 1.65 mixture ratio.

OME.0014 Start Up Transient

The engine shall develop 90 percent steady state thrust within 0.450 +/- 0.100 seconds of electrical on-signal.

Rationale: This requirement is derived from the existing ESM performance capability, vehicle controllability characteristics, and interaction with propellant feedsystem during start.

OME.0015 Engine Firing Duration

The engine shall be capable of a continuous firing from 2 secs minimum to 1030 seconds maximum.

Rationale: The maximum burn duration value assumes all usable propellant passes through the main engine in a single burn. The minimum burn duration is based on established operations for minimum engine firing - to provide some delta-V overlap to protect for loss of auxiliary engines.

Engine start is defined by sending of electrical ON signal.

OME.0017 Minimum Engine Off Time

The engine shall allow a restart no later than 240 seconds after receipt of a shutdown signal.

Rationale: The minimum engine off time is based on established operations to provide the capability to perform OME burn as soon as 240 seconds after receipt of a shutdown signal from a previous burn (note in the case of EM-1/2 limited to 10 seconds ON for the 240 seconds OFF time)

OME.0018 Gas Ingestion

The engine shall operate nominally, with no engine damage, subsequent to ingestion of a maximum of 24 [in3] ( 390 [cm3]) of undissolved helium at nominal inlet conditions in either or both propellants during start-up and at any operating mode.

Rationale: This requirement is derived from the existing ESM performance capability and system operating characteristics

OME.0019 Starts

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The engine shall be capable of performing all operations specified herein for a minimum of 10 starts per mission.

Rationale: This requirement is derived from the existing ESM performance capability and mission performance analysis.

OME.0022 Peak Thrust

The peak thrust shall not exceed 150% of the mean steady state thrust during the start transient.

Rationale: This requirement is derived from the existing ESM performance capability and induced load limits on the vehicle.

OME.0023 Shutdown Response Time

The engine shall decay to less than 10% nominal thrust within 2.3 seconds from the electrical-off-signal.

Rationale: This requirement is derived from the existing ESM performance capability, mission analysis, and system shutdown characteristics. Tail off impulse shall be documented and provided for guidance modeling.

OME.0024 Max Gimbal Deflection

The engine shall be capable of being gimbaled during operation with the maximum hardstop angles specified in Figure: Engine Gimbal Deflection Range.

Rationale: The engine gimbal ring must accommodate the full range of motion of the TVC and provide for margin against the minimum required gimbal range.

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FIGURE 3.3.1-5 - ENGINE GIMBAL DEFLECTION RANGE

OME.0025 Dynamic Stability

The engine shall be designed to achieve stable combustion throughout its operational range defined in Figure Engine Operating pressures (Envelope).

Rationale: Stable combustion is defined in ref: CPIA 655. This contains the state-of-the-practice guidelines for assessing combustion stability of liquid rocket engines.

OME.0026 Roughness

The engine thrust roughness shall be less than +/- 5% of average steady state thrust at normal operating conditions.

Rationale: Thrust roughness will impact attitude and attitude rate motion. Thrust roughness is defined as the total peak-to-peak chamber pressure deviation divided by the nominal chamber pressure.

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OME.0085 Heat Shield

The engine shall provide a heat shield to protect the ESM from radiant heat from the nozzle.

Rationale: The ESM is an existing system that has been designed to withstand the environment of the legacy OMS-E which included heat shields. Any new engine must incorporate similar thermal protection features.

3.3.2 Physical Requirements

OME.0027 Mass

The dry engine assembly weight shall not exceed 278 lbm.(126.1 kg).

Rationale: This requirement is derived from the existing ESM performance capability and mass allocation to the engine (including nozzle, heat shield, and all required components).

OME.0103 Center of Gravity

The OME nominal Center of Gravity (CG), including nozzle and heatshield, relative to the OME assembly coordinate system defined in figure ESM Coordinate System shall be in accordance with the values in table OME Center of Gravity.

Rationale: in order to maintain overall vehicle CG within currently defined envelope the OME CG must be within the limits defined.

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FIGURE 3.3.2-1 - ESM COORDINATES

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TABLE 3.3.2-1 - OME CENTER OF GRAVITY

Case Nominal Center of Gravity

X in (cm)

Y in (cm)

Z in (cm)

Total Engine Dry -8.2 +3.0/-1.0 0.0 +/-3.0 0.0 +/-3.0

3.3.3 Environmental Requirements

OME.0028 Transportation and Handling Loads

The engine shall be capable of meeting the operating performance requirements specified herein after exposure to the transportation loads given in Table Transportation Loads in the USA and between Europe and the USA, Table Transportation Loads in Europe, and Table Handling Loads.

Rationale: On the longitudinal direction, a positive acceleration is in the force direction and hence a negative one is in the aft direction. For the vertical direction, a positive acceleration is in the up direction and therefore a negative acceleration is in the down direction.

Remark 1: The values given in the tables correspond to the acceleration of the carrier plus the effect of gravity in the vertical axis (Z). To compute the interface loads between the transported cargo and the carrier, an inertial loading on the cargo shall be applied in the opposite direction.

Remark 2: Tables Transportation Loads in the USA and between Europe and the USA and Table Handling Loads are extracted from MPCV-RIBRE-SPE-0011, section 4.1.23.

Remark 3: Table Transportation Loads in Europe preliminarily taken from MPCV-RIBRE-CP-0088 to account for the change of the aircraft.

TABLE 3.3.3-1 - TRANSPORTATION LOADS IN THE USA AND BETWEEN EUROPE AND

THE USA

Transportation Limit-Load Factors1

MODE LONGITUDINAL (g) LATERAL (g) VERTICAL (g)

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Water + 0.75 + 2.5 + 2.5, - 0.5

NASA Barge (MAF to KSC)2 + 0.75 + 1.0 + 2.25, - 0.25

NASA Barge (Inland Waterway) + 0.5 + 0.5 + 1.4, -0.6

Air3 + 3.0 + 1.5 + 3.0, - 1.0

Super Guppy Case #1 + 1.0 + 2.5, - 1.0

Super Guppy Case #2 + 0.5 + 1.0 + 1.6

C-5, C-17 +3.0, -1.5 + 1.5 + 2.0, - 1.0

Crash Landing4 +3.0, -1.5 + 1.5 + 4.5, - 2.0

Ground

Truck + 2.0 + 2.0 + 3.0, - 1.0

Rail (humping shocks)5 + 30.0 + 5.0 + 15.0

Rail (rolling) + 3.0 + 1.5 + 3.0, - 1.0

Slow-moving dolly + 1.0 + 0.75 + 2.0

Slow-moving dolly (smooth surface)6 + 0.5 + 0.5 + 1.0 + 0.5

1 The load factors in this table apply to the transport vehicle axes in g’s [gravity].

2 MAF = Michoud Assembly Facility, KSC = Kennedy Space Center

3 Air limit load factors provided envelope Super Guppy, C5, C17 operational loads.

4 Crash loads are to be assessed independently in the three orthogonal directions except gravity;

Vertical gravity load of 1.0 g shall be applied simultaneously with longitudinal and lateral crash loads.

The crash load case is an ultimate load case.

5 These are shock conditions and should not be treated as quasi-steady accelerations. Event specific analyses shall be performed for shock loads or inputs likely to occur at the element/module fundamental frequency.

6 Transport shall adhere to Slow-Moving Dolly (Smooth Surface) Operational Constraints

TABLE 3.3.3-2 - TRANSPORTATION LOADS IN EUROPE

TRANSPORTATION LILMIT-LOAD FACTORS

MODE

LONGITUDINAL

(g)

LATERAL

(g)

VERTICAL

(g)

Handling

Hoisting 0.0 0.0 +1.4,-0.7

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Mating ±0.5 ±0.5 +2.0, 0.0

Road: quasistatic ±1.1 ±1.2 +3.0,-1.0

Air

Antonov +2.3,-1.5 ±1.5 +2.5,-2.0

Sea

Slamming 0.0 0.0 +1.8,-0.2

Waves ±0.3 ±0.5 +1.6,-0.4

TABLE 3.3.3-3 - HANDLING LOADS

Handling Load Factors

MODE LONGITUDINAL (g) LATERAL (g) VERTICAL (g)

Jacking + 0.5 + 0.5 + 2.0

Forklift + 1.0 + 0.5 + 2.0

Matting and Erecting + 0.5 + 0.5 + 2.0

Hoisting (Land Ops) + 1.33 applied upward in any direction within 20o of vertical

Hoisting (Water Ops) + 2.67 applied upward in any direction within 20o of vertical

OME.0029 Ground Temperature, Humidity, and Pressure Environments

The engine shall be compatible with the ground environments specified in Table Thermal Transportation and Storage Environment.

Rationale: The engine may experience these environments during storage and transportation.

TABLE 3.3.3-4 - THERMAL TRANSPORTATION AND STORAGE ENVIRONMENT

Phase Temperature Hygrometry Pressure

ESM integration in Bremen 22 oC + 2 oC 40% to 55% 950 to 1050 hPa

Transport in Europe and in US (aircraft excluded)

22 oC + TBD oC < 55% at 16 oC 950 to 1050 hPa

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Aircraft transport 22 oC + TBD oC < 55% at 16 oC 150 to 1050 hPa

MPCV integration and test in US 22 oC + 2 oC 40% to 55% 950 to 1050 hPa

Launch preparation operations 20 oC + 2 oC < 5% 950 to 1050 hPa

OME.0031 Acceleration

The engine shall meet performance requirements after exposure to the quasi-static loads given in Table Engine Quasi- Static Loads (Limit Loads) at the center of gravity of the engine.

Rationale: The relevant flight environment must be respected by the engine. The loads refer to the global axes of the MPCV-ESM system.

TABLE 3.3.3-5 - ENGINE QUASI-STATIC LOADS (LIMIT LOADS)

Load Case QSLx (g) QSLy (g) QSLz (g)

1 -2.6 +2.7 +3.6

2 +5.1 +2.7 +3.6

3 -1.9 +5.7 +3.6

4 +4.5 +5.7 +3.6

5 -1.9 +2.7 +6.6

6 +4.5 +2.7 +6.6

OME.0032 Low Frequency Vibration

The engine shall meet its functional requirements after exposure to the frequency spectrum as defined in Table Engine Low Frequency Sine Environment (Limit Loads).

Rationale: The relevant flight environment must be respected by the engine. The loads refer to the global axes of the MPCV-ESM system. These vibration spectrum will be used to develop the qualification test envelope.

TABLE 3.3.3-6 - ENGINE LOW FREQUENCY SINE ENVIRONMENT (LIMIT LOADS)

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Frequency (Hz) Level (g)

X 5 - 60

60 - 200

2.00

1.00

Y 5 – 30

30 – 200

3.00

1.00

Z 5 – 30

30 – 200

2.00

1.00

OME.0033 Random Vibration

The engine shall meet functional requirements after exposure to the random vibration loads applied at the mounting interfaces as defined in Table Engine Random Vibration Environment (Limit Loads).

Rationale: The relevant flight vibration environment must be respected by the engine.

ENGINE RANDOM VIBRATION ENVIRONMENT (LIMIT LOADS)

OME Limit Loads

Frequency (Hz) PSD (g²/Hz)

20 3.50E-02

45 0.225

65 0.225

70 0.15

160 0.15

200 0.06

800 0.06

2000 1.00E-02

OME.0035 Gateway Vibration Environment

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When the Orion is docked with the Gateway, the engine shall comply with the Gateway requirements for 0.2g load factors acting in any direction.

OME.0036 Acoustic Loads

The engine shall meet its functional requirements after exposure to the acoustic environment given in Table Engine Acoustic Environment.

Rationale: The relevant flight environment must be respected by the engine. The loads refer to the global axes of the MPCV-ESM system.

TABLE 3.3.3-7 - ENGINE ACOUSTIC ENVIRONMENT

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OME.0037 Shock

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The engine shall meet its functional requirements after exposure to the limit (MEE) shock loads given in Table Engine Shock Environment at the mounting points of the engine.

Rationale: The relevant flight environment must be respected by the engine. The loads refer to the global axes of the MPCV-ESM system.

TABLE 3.3.3-8 - ENGINE SHOCK ENVIRONMENT

Frequency (Hz) SRS MEE (g)

100 1.6

4150 330

10000 330

OME.0038 ESM External Thermal Environment

The engine shall meet its functional requirements after exposure to the external thermal environment as specified in Figure ESM External Thermal Environment with main Main Engine in Straight Position and Figure ESM External Thermal Environment with main Main Engine in Canted Position. Each figure is divided into 3 thermal zones, which are shown as concentric circles about the ESM centerline in the y-z plane, at the aft of the ESM. The radius of each zone is shown on the figures a R. Temperatures for the zone are listed as is the radiative emmissivity (e).

Rationale: This defines the thermal loads induced on the engine from the aft part of the ESM.

There are two thermal environments for the Main Engine, Straight and Canted. Each figure is divided into 3 thermal zones, which are shown as concentric circles about the ESM centerline in the y-z plane, at the aft of the ESM. The radius of each zone is shown on the figures as R.

Temperatures for the zone are listed as is the radiative emissivity (e).

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637 °C (1178 °F) ε = 0.63

537 °C (998 °F) ε = 0.68

357 °C (674 °F) ε = 0.76

ESM CL

R = 15.75 in (400 mm)

R = 39.37 in (1000 mm)

R = 75.51 in (1918 mm)

FIGURE 3.3.3-1 - ESM EXTERNAL THERMAL ENVIRONMENT WITH MAIN ENGINE IN

STRAIGHT POSITION

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777 °C (1430 °F) ε = 0.53

507 °C (944 °F) ε = 0.67

417 °C (782 °F) ε = 0.71

ESM CL

R = 15.75 in (400 mm)

R = 39.37 in (1000 mm)

R = 75.51 in (1918 mm)

FIGURE 3.3.3-2 - ESM EXTERNAL THERMAL ENVIRONMENT WITH MAIN ENGINE IN

CANTED POSITION

OME.0039 Plume Fluxes

The nozzle extension shall withstand a 10 min 55 sec exposure to the plume heating thermal environment defined in TBD-634-001, with plume fluxes as defined in TBD-634-001.

Rationale: The plume fluxes have been obtained using the IEMC version 1.0 code. A 1.3 margin factor has been applied. This 30% margin value is based on AD&S internal heritage. The duration corresponds to an Abort-to-Orbit.

OME.0040 ESM Internal Thermal Environment

The engine shall be compliant with the ESM internal thermal flight environment as defined in Table ESM Internal Thermal Conditions.

Rationale: The relevant flight environment must be respected by the engine. MPCV-AS-TN-0012 provides the analysis of the thermal environments at the ESM aft side.

TABLE 3.3.3-9 - ESM INTERNAL THERMAL CONDITIONS

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Calculated Marged

Hot Case Enclosure Temperature 57.9 oF (14.4 oC) 88.5 oF (31.4 oC)

Cold Case Enclosure Temperature 48.8 oF (9.4 oC)

Hot and Cold Case Enclosure Emissivity

0.8

OME.0041 Gimbal Loads

Engine gimballing shall induce loads less than those given in Table Gimbal Operational Loads

Rationale: Gimbal loads provided by NASA in NASA-ESM-Prop-2016-015.

TABLE 3.3.3-10 - GIMBAL OPERATIONAL LOADS

Type Load Operational Load

Ground Checkout Flight Non-Firing Flight Firing Breakaway (Maximum)

Inertia (for per unit acceleration)

1.5 lbf/(in/sec^2) (262 N/(m/sec^2))

1.5 lbf/(in/sec^2) (262 N/(m/sec^2))

1.5 lbf/(in/sec^2) (262 N/(m/sec^2)) -

Friction 0-40 lbf (1-178 N)

12-100 lbf (53-445 N)

14-135 lbf (62-601 N)

275 lbf (1223 N)

Thrust and Gravity

80 lbf (356 N)

0 lbf (0 N)

1 to +/- 60 lbf (4 to +/- 267 N)

60 lbf (267 N)

Spring 6 lbf/in (1.1 N/mm)

12 lbf/in (2.1 N/mm)

12 lbf/in (2.1 N/mm)

12 lbf/in (2.1 N/mm)

OME.0086 Natural Environments

The engine shall meet its requirements during and after exposure to the environments defined in the SLS-SPEC-044-07 Cross Program Vehicle Design Environments Volume 7: Natural Environments.

Rationale: The engine will be exposed to a variety of natural environments that could make it unable to meet its requirements, potentially in combination with induced environments. This requirement defines the limits of these effects and assures that they will be mitigated by the design.

3.3.4 Reliability and Maintainability

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OME.0042 Reliability

The engine shall fulfill its main functions during its life duration under the specified environmental conditions with a Probability of success of R>0.99862 and at a minimum confidence level of 95%

Rationale: The intent of this requirement is to capture the reliability of the heritage hardware for the MPCV missions.

3.3.5 Life

OME.0054 Engine Life

The engine, including all movable parts, instrumentation and wiring shall have a lifetime that meets or exceeds 4815 days, represented in Table X before replacement or additional servicing is necessary, beginning when the OME is transferred to the government.

Rationale: Due to unpredictable timeframes between construction and launch, it may be necessary for the OME to remain in a controlled storage from acceptance to installation for 10 years. The OME is delivered to ESA for installation & ESM integration, which could take up to 183 days. Following ESA Integration, the ESM is delivered to the Operations & Check-out facility at KSC to complete CM - SM Mate and Spacecraft processing. Once, Orion Spacecraft has completed assembly and vehicle level DD-250, Orion is delivered to KSC Exploration Ground Systems for 500 days of processing, including 100 days of integrated to the launch vehicle. A maximum of 68 days of launch pad time is assumed and included in the 100 days of integrated launch vehicle operations.The remaining 232 days allows the vehicle to meet its full mission time-line including 21 days of active mission, quiescent operations and post mission recovery. The OME is actively used for a total duration of 21 days.

TABLE 3.3.5-1 - ENGINE LIFE BY PHASE

Phase of Life Duration Storage Life 3650 days (or 10 years) ESA Integration 183 days KSC O&C Processing (Pre-Spacecraft DD- 250)

250 days

KSC Ground Operations 400 days Integrated launch vehicle Operations

• Launch Pad 100 days

• 68 days Full Mission Duration

• OME Expected Operation Timeframe

232 days

• 21 days

Total Engine Life 4815 days

3.3.6 Safety Requirements

OME.0043 Inadvertant Firing

The engine shall be failure tolerant to inadvertent operation.

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Rationale: This requirement mitigates hazard causes listed in Hazard Report MPCV-FLT-002.

The purpose of this requirement is intended to preclude inadvertent engine firings as well as if a purge system exists, inadvertent operation of the purge system.

OME.0045 Propagation of Failures

The engine shall prevent the propagation of failures that could result in a critical or catastrophic hazard.

Rationale: Failures must be prevented from propagating to minimize harm and prevent loss of critical resources, and to ensure reconfiguration capability for vehicle recovery. Faults and failures can be identified via the FMEA/CIL, among other techniques.

OME.0079 Failure Tolerance to Catastrophic Events

The engine shall provide failure tolerance to catastrophic events, with specific levels of failure tolerance and implementation (similar or dissimilar redundancy) derived via an integration of the design and safety analysis without the use of EVA,…