C.2 Attachment B_MSR-CCRS-ERM-SPEC-0002.pdf

Effective Date: September 08, 2021 Expiration Date: September 08, 2026

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

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

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

MSR-CCRS-ERM-SPEC-0002, Revision -

Mars Sample Return - Capture, Containment and Return System (MSR- CCRS) Project

NASA/GSFC Code 435

Earth Entry System (EES) Spin Eject Mechanism (SEM)

Technical Specification

MSR-CCRS CMO

September 08, 2021

RELEASED

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

EES SEM Technical Specification MSR-CCRS-ERM-SPEC-0002 Revision -

Effective Date: 09/08/2021 ii

Earth Entry System (EES) Spin Eject Mechanism (SEM) Technical Specification Signature/Approval Page

Prepared by:

[signature on file] Charlie Wang CCRS Spin Eject Mechanism RE Code 544

Reviewers/Approvers:

[signature on file] [signature on file] Chanel Duncan Brendan Feehan CCRS Chief Safety & Mission Assurance Officer

CCRS Project Systems Engineer Code 599

Code 383

[signature on file]

Charles Bacon Julian Ramirez-Brana CCRS GSFC Flight Manager CCRS Electrical Systems PDL Code 435 Code 561

Gordon Casto Radford Perry CCRS Mechanical Systems Lead Code 543

CCRS Contamination Control and Planetary Protection

Code 546

Veronica Otero Russell Stein CCRS Thermal PDL Spin Eject Mechanism PDL Code 545 Code 540

[signature on file] Bruno Sarli CCRS Inspection and Release Phase Lead

Code 599

Approved By:

David Littmann CCRS Project Manager Code 435

*** Electronic signatures are available on-line at: https://ipdtdms.gsfc.nasa.gov*** iii

Preface

This document is a Mars Sample Return (MSR) Capture, Containment and Return System (CCRS) Project configuration control board (CCB) controlled document. Changes to this document require prior approval of the CCB Chairperson or designee. Proposed changes shall be submitted in the Technical Data Management System (TDMS) via a configuration change request (CCR) along with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

All of the requirements in this document assume the use of the word "shall" unless otherwise stated.

Questions or comments concerning this document should be addressed to:

CCRS Configuration Management Office Mail Stop: 435 Goddard Space Flight Center Greenbelt, Maryland 20771 GSFC-DL-CCRS-TDMSCM-Admin@mail.nasa.gov iv

Change History Log

Revision Effective Date Description of Changes (Reference the CCR and Approval Date)

Revision - 09/08/2021 Released per MSR-CCRS-CCR-0051; 09/08/2021 v

Table of TBDs/TBRs/TBCs

Location Summary Individual/ Organization Actionee

Proposed Closure Date

2.3.2.8 TBD: LV RF Emissions Limits Rick Schnurr / 564 TBD->TBC by Contractor PDR TBC->closed by Contractor CDR

2.4.3 TBC: SEM Natural Frequencies Son Ngo / 543 Contractor SRR

2.4.6 TBR: SEM MICD Contractor and NASA

GSFC

Contractor PDR

2.4.7 TBC: RTAS Installation Loads Justin Lin / JPL CCRS PDR

2.4.8 TBD: RTAS Installation Deflection Justin Lin / JPL Contractor PDR

2.6.6 TBC: Dead Time Rick Schnurr / 564 CCRS PDR

2.7.1 TBD: Thermal ICD Veronica Otero / 545 TBD->TBC by

Contractor PDR TBC->closed by Contractor CDR

2.7.2 TBC: Survival Temperature Veronica Otero / 545 Contractor SRR

2.7.3 TBC: Operational Temperature Veronica Otero / 545 Contractor SRR vi

Table of Contents 1 INTRODUCTION ................................................................................................................. 1-1

1.1 Purpose ............................................................................................................................. 1-1

1.2 Scope ................................................................................................................................ 1-1

1.3 Application ....................................................................................................................... 1-1

1.4 Basic Description ............................................................................................................. 1-4

1.5 Verb Application .............................................................................................................. 1-4

1.6 Changes ............................................................................................................................ 1-4

1.7 Related Documentation .................................................................................................... 1-5

1.7.1 Applicable Documents ............................................................................................. 1-5

1.7.2 Reference Documents ............................................................................................... 1-5

2 TECHNICAL REQUIREMENTS ......................................................................................... 2-1

2.1 GENERAL REQUIREMENTS ....................................................................................... 2-1

2.1.1 Coordinate Systems for the SEM ............................................................................. 2-1

2.1.2 Units of Measure ...................................................................................................... 2-1

2.1.2.1 Hardware ............................................................................................................. 2-1

2.1.2.2 Dual Dimensions ................................................................................................ 2-1

2.1.2.3 Units Identification ............................................................................................. 2-1

2.1.3 Documentation Date ................................................................................................. 2-1

2.1.4 Mission Life ............................................................................................................. 2-1

2.2 FUNCTIONAL AND PERFORMANCE REQUIREMENTS ........................................ 2-1

2.2.1 EES Release ............................................................................................................. 2-1

2.2.2 EES Release Velocity ............................................................................................... 2-2

2.2.3 EES Release Lateral Velocity .................................................................................. 2-2

2.2.4 EES Release Spin Rate ............................................................................................. 2-2

2.2.5 EES Release Tip-off Rate ......................................................................................... 2-2

2.2.6 EES Release Pointing ............................................................................................... 2-2

2.2.7 EES Release Performance Limits ............................................................................. 2-2

2.3 ENVIRONMENTAL REQUIREMENTS ....................................................................... 2-2

2.3.1 Transportation, Handling, and Storage Environments ............................................. 2-2

2.3.1.1 Maximum Loads on Flight Hardware ................................................................. 2-2

2.3.1.2 Pressure ............................................................................................................... 2-3

2.3.2 Launch and Ascent Environments ............................................................................ 2-3

2.3.2.1 Spacecraft Tip-Off Rates .................................................................................... 2-3

2.3.2.2 Static-Equivalent Accelerations .......................................................................... 2-3

2.3.2.3 Sinusoidal Vibration ........................................................................................... 2-5

2.3.2.4 Acoustic Vibration .............................................................................................. 2-5

2.3.2.5 Random Vibration .............................................................................................. 2-6

2.3.2.6 Mechanical Shock ............................................................................................... 2-7

2.3.2.7 Pressure ............................................................................................................... 2-7

2.3.2.8 Launch Site Electromagnetic Environment ........................................................ 2-7

2.3.3 In-Space Environments ............................................................................................ 2-8

2.3.3.1 Mechanical Shock ............................................................................................... 2-8

2.3.3.2 On-orbit Dynamic Load ...................................................................................... 2-8 vii

2.3.3.3 Magnetic compatibility with Orbiting Sample ................................................... 2-9

2.3.3.4 On-orbit Electromagnetic Compatibility ............................................................ 2-9

2.3.4 SEM Environmental Verification ............................................................................. 2-9

2.3.4.1 SEM System Level Environmental Verification ................................................ 2-9

2.3.4.2 SEM Verification Test Matrix ............................................................................ 2-9

2.4 Mechanical Requirements .............................................................................................. 2-10

2.4.1 Deliverables ............................................................................................................ 2-10

2.4.2 SEM Design Envelope ........................................................................................... 2-11

2.4.3 SEM Natural Frequencies ...................................................................................... 2-11

2.4.4 Lift and Handling points for EES I&T ................................................................... 2-11

2.4.5 Hold Down Release Mechanisms (HDRMs) ......................................................... 2-11

2.4.5.1 HDRM Usage and Life ..................................................................................... 2-11

2.4.5.2 HDRM Quantity ............................................................................................... 2-11

2.4.6 Interfaces ................................................................................................................ 2-11

2.4.7 RTAS Installation Loads (No margin applied) ...................................................... 2-11

2.4.8 RTAS Installation Deflection ................................................................................. 2-11

2.4.9 EES Fittings Structure Integrity ............................................................................. 2-12

2.5 Mass Requirements ........................................................................................................ 2-12

2.5.1 SEM Mass .............................................................................................................. 2-12

2.5.2 SEM Mass on the EES ........................................................................................... 2-12

2.5.3 EES Mass Properties .............................................................................................. 2-12

2.6 Electrical Requirements ................................................................................................. 2-13

2.6.1 Polarity Reversal Protection ................................................................................... 2-13

2.6.2 SEM Firing Pulses .................................................................................................. 2-13

2.6.3 SEM Firing Current ................................................................................................ 2-13

2.6.4 SEM Activation ...................................................................................................... 2-13

2.6.5 Pulse Width ............................................................................................................ 2-13

2.6.6 Dead Time .............................................................................................................. 2-13

2.6.7 Maximum Spread Current ...................................................................................... 2-13

2.6.8 Grounding ............................................................................................................... 2-13

2.6.9 Ground Straps ......................................................................................................... 2-13

2.6.10 Interface Electromechanical Design ....................................................................... 2-13

2.6.11 ESD Grounding ...................................................................................................... 2-13

2.7 Thermal Requirements ................................................................................................... 2-14

2.7.1 Thermal Hardware Mounting ................................................................................. 2-14

2.7.2 Survival (Non-Operational) Temperature .............................................................. 2-14

2.7.3 Operational Temperature ........................................................................................ 2-14

2.8 Fault Tolerance and Reliability Requirements .............................................................. 2-14

2.8.1 EES Release Reliability .......................................................................................... 2-14

2.8.2 Dual-Fault Tolerance - EES Release ...................................................................... 2-14

2.9 Terrestrial Contamination Control Requirements .......................................................... 2-14

2.9.1 Terrestrial External Surface Cleanliness ................................................................ 2-14

2.9.2 Terrestrial Internal Surface Particulate Cleanliness ............................................... 2-14

2.9.3 Internal Surface Molecular Cleanliness ................................................................. 2-14

2.9.4 Terrestrial Outgassing ............................................................................................ 2-14

2.9.5 GSE cleanliness ...................................................................................................... 2-15 viii

2.9.6 GSE outgassing ...................................................................................................... 2-15

2.9.7 Bioassay .................................................................................................................. 2-15

2.9.8 Compatibility with alcohol cleaning ...................................................................... 2-15

Appendix A SEM MICD .......................................................................................................... A-1 Appendix B Abbreviations and Acronyms ............................................................................... B-1

List of Figures Figure 1-1: CCRS Functional Description (Reference Only) ..................................................... 1-2 Figure 1-2: EES Major Components and Current Architecture (Reference Only) ...................... 1-3 Figure 1-3: EES Assembly Description (Reference Only) .......................................................... 1-3 Figure 1-4: EES-SEM Interface (Reference Only) ...................................................................... 1-4 Figure 2-1 Generic Mass-Acceleration Curve (MAC) – Flight Limit ......................................... 2-3 Figure 2-2: EES CG Location .................................................................................................... 2-12

List of Tables

Table 1-1: Applicable Documents ............................................................................................... 1-5 Table 1-2: Reference Documents ................................................................................................ 1-5 Table 2-1: SEM Release Performance Limits ............................................................................. 2-2 Table 2-2: Maximum Loads During Ground Handling ............................................................... 2-3 Table 2-3: Maximum Loads During Transportation .................................................................... 2-3 Table 2-4: MAC Design Limit Loads .......................................................................................... 2-4 Table 2-5: Interface Sine Sweep Vibration Levels ...................................................................... 2-5 Table 2-6: Acoustic Environment ................................................................................................ 2-5 Table 2-7: Random Vibration Levels (Less than 50lbs) .............................................................. 2-6 Table 2-8: Shock Response Spectrum (Q=10) ............................................................................ 2-7 Table 2-9: LV RF Emissions Limits ............................................................................................ 2-7 Table 2-10: Shock Response Spectrum (Q=10) .......................................................................... 2-8 Table 2-11: On-orbit Electromagnetic Compatibility .................................................................. 2-9 Table 2-12: SEM Verification Test Matrix .................................................................................. 2-9 Table 2-13: RTAS Installation Loads (No margin applied) ...................................................... 2-11 Table 2-13: EES Mass Properties (relative to CG) .................................................................... 2-12

1-1

1 INTRODUCTION

1.1 Purpose

This technical specification establishes the requirements for the Mars Sample Return (MSR) Capture, Containment and Return System (CCRS) Earth Entry System (EES) Spin Eject Mechanism (SEM).

1.2 Scope

This specification establishes the functional, performance, environmental, mechanical/electrical interface, contamination control and verification testing requirements of the SEM for the MSR program.

The scope of work to be performed for the design, development, fabrication, testing, and delivery of the SEM is provided in the accompanying Statement of Work (SOW) MSR-CCRS-ERM-SOW- 0001.

1.3 Application

The Spin Eject Mechanism (SEM) is an integral part of the Capture, Containment and Return System (CCRS) Earth Entry System (EES). It is responsible for restraining the Earth Entry System (EES) during launch, cruise, and Capture and Containment (CCM) robotic operations while in Martian orbit. To facilitate robotic operations, the SEM requires a hollow inner diameter for robotic access to install the Contained Orbiting Sample (C-OS) SCV Lid, and EES Lid onto the EES Aeroshell. The completed system ready for transport back to Earth is shown in Figure 1-3.

The SEM is responsible for restraining the EES during transport back to Earth, and finally ejecting the EES safely back to Earth.

Images detailing the CCRS and EES architecture are shown in Figure 1-1 and Figure 1-2 . The SEM is part of the primary load path into the EES for all mission stages.

1-2

Figure 1-1: CCRS Functional Description (Reference Only)

1-3

Figure 1-2: EES Major Components and Current Architecture (Reference Only)

Figure 1-3: EES Assembly Description (Reference Only)

1-4

1.4 Basic Description

In its most basic description, the SEM can be imagined as a hollow cylinder with that provides a deployable structural load path between the CCRS Module and the EES. This interface needs to be strong enough to properly address launch loads while also remaining capable of providing precise and controlled ejection energy into the EES during the payload ejection event to return the payload to Earth.

Figure 1-4: EES-SEM Interface (Reference Only)

The design envelope for these components is controlled by Mechanical Interface Control documents as specified in Section 2.4.2 and 2.4.6.

Note: Images shown in Figure 1-1 through Figure 1-4 are notional design concepts for reference only. These do not represent a flight proven design.

1.5 Verb Application

Statements containing the verb “shall” are binding requirements regardless of location within this document. The verbs “should” and “may” are used for stating non-mandatory goals or denoting a statement of best practice. The verb “will” is used in a statement of fact, expected occurrence, or declaration of purpose.

1.6 Changes

All revisions, deletions and other changes to this document and any referenced documents must be authorized in writing by the Buyer’s Contracting Officer.

1-5

1.7 Related Documentation

The latest revision of the applicable and reference documents at the time of contract execution shall be used if different from that shown in Table 1-1and Table 1-2 unless otherwise specified. In the event of conflict between this specification and any referenced document, this specification will govern, with the exception of the Statement of Work, MSR-CCRS-ERM-SOW-0001, in which case the SOW shall take precedence.

1.7.1 Applicable Documents

The following documents listed in Table 1-1 shall apply to the development of the SEM in the situations where they are specifically referenced in this document.

Table 1-1: Applicable Documents

Document No. Title

MSR-CCRS-ERM-SOW-0001 Rev.

Mars Sample Return (MSR) Capture, Containment and Return System (CCRS) Spin Eject Mechanism (SEM) Statement of Work

MSR-CCRS-SYS-REQ-0002 Rev. - CCRS Environmental Requirements Document

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

ASTM E2900 Standard Practice for Spacecraft Hardware Thermal Vacuum Bakeout

ECSS‐Q‐ST‐70‐55 Microbiological examination of flight hardware and cleanrooms

1.7.2 Reference Documents

The following documents listed in Table 1-2 may contain information pertinent to the performance of the Contract and should be referred to for clarification purposes only. The Applicable Documents and other Government produced contract documentation shall take precedence over the reference documentation.

Table 1-2: Reference Documents

1-6

Document Number Title

MSR-CCRS-ERM-CDRL-0002

Rev. -

Mars Sample Return (MSR) Capture, Containment and Return System (CCRS) Spin Eject Mechanism (SEM) Deliverable Items List and Schedule (DILS)

MSR-CCRS-SMA-REQ-0003

Rev. -

Mars Sample Return (MSR) Capture, Containment and Return System (CCRS) Mission Assurance Requirements (MAR)

MSR-CCRS-SYS-SPEC-0004

Rev. -

CCRS Mechanical and Mechanism Design Specification

MSR-CCRS-SYS-SPEC-0010

Rev. -

CCRS Thermal Model Exchange Guidelines

NASA-STD-5017A Revalidated w-Change 1 Design and Development Requirements for Mechanisms

GSFC-STD-7000B General Environmental Verification Standard (GEVS) for GSFC Flight Programs and Projects

Ariane 6 User’s Manual Issue 2 Revision 0 February 2021

2-1

2 TECHNICAL REQUIREMENTS

2.1 GENERAL REQUIREMENTS

2.1.1 Coordinate Systems for the SEM

SEM Reference Coordinate System shall be oriented such that the Y-Z plane aligns with the Spacecraft (CCRS) interface plane and the positive X-axis projects away from the Spacecraft and through the EES geometric centerline between the heatshield stagnation point and the center of the backplate of the backshell. This will be the same orientation as the CCRS coordinate system.

The positive Y-axis projects parallel toward the Capture Module Lid (As shown in Figure 1-2).

The physical origin of the system will need to be located in reference with a physical piece of hardware.

2.1.2 Units of Measure

SEM design shall use the International System of Units (SI) Metric Standard, with the following exceptions and clarifications:

2.1.2.1 Hardware

English hardware shall be used. SI hardware may be used for inseparable assembly interfaces, or other locations if deemed essential to the design, with permission from the Government Technical Representative or designee and authorized in writing by the Buyer’s Contracting Officer.

2.1.2.2 Dual Dimensions

Interface control documents and design drawings shall use dual dimensions with SI as primary.

2.1.2.3 Units Identification

Documentation shall clearly identify the specific units to be used.

2.1.3 Documentation Date

Dates in documentation, drawings, and specifications shall be used in the format DD MMM YYYY, where DD specifies the day in two digits, MMM specifies the month in a three letter abbreviation and YYYY specifies the year in four digits, with a space, a dash, or a slash as separator. The accepted abbreviations for the month are Jan, Feb, Mar, Apr, May, Jun, Jul, Aug, Sep, Oct, Nov, Dec. An all caps spelling is also acceptable.

2.1.4 Mission Life

SEM shall meet all requirements up to 8 years post-launch.

2.2 FUNCTIONAL AND PERFORMANCE REQUIREMENTS

The following requirements apply with the EES characteristics as defined in Section 2.5.3.

2.2.1 EES Release

SEM shall eject the EES such that it does not re-contact elements of itself (i.e. hardware left behind) and remains in the as-designed/as-specified deployed envelope.

2-2

2.2.2 EES Release Velocity

SEM shall eject the EES with velocity magnitude of 350 ±11.5mm/s at 99.7% probability and 95% confidence in the axial direction.

2.2.3 EES Release Lateral Velocity

SEM shall deploy the EES with lateral velocity of less than 4 mm/s at 99.7% probability and 95% confidence.

2.2.4 EES Release Spin Rate

SEM shall release the EES with the spin rate of 5.3 RPM ±0.2 RPM at 99.7% probability and 95% confidence.

2.2.5 EES Release Tip-off Rate

SEM shall release the EES with a nominal tip-off rate of less than 1.93 deg/s at 99.7% probability and 95% confidence.

2.2.6 EES Release Pointing

SEM shall release the EES such that the principal spin axis shall be parallel ±0.5° to the X-axis defined in Section 2.1.1.

2.2.7 EES Release Performance Limits

Analysis of SEM deployment performance at 99.9999% probability and 95% confidence shall not exceed the limits in Table 2-1.

Table 2-1: SEM Release Performance Limits

Parameter Minimum Maximum Release Velocity (mm/s) 306.0 361.5 Lateral Velocity (mm/s) 0.0 16.7 Spin Rate (RPM) 4.6 5.5 Tip-off Rate (deg/s) 0.0 11.0

2.3 ENVIRONMENTAL REQUIREMENTS

SEM shall be designed to achieve its specified performance requirements after exposure to any single or reasonable combination of natural and induced environments from pre-launch storage, transportation, test, launch, ascent, and on-orbit operations.

2.3.1 Transportation, Handling, and Storage Environments

SEM flight hardware shall be protected during manufacturing, assembly, handling, transportation, storage and other pre-launch activities such that none of the acceptance test environments defined herein is exceeded.

2.3.1.1 Maximum Loads on Flight Hardware

Maximum loads that SEM will experience during ground operations or transportation shall be less than the values in Table 2-2 and Table 2-3.

2-3

Table 2-2: Maximum Loads During Ground Handling

Table 2-3: Maximum Loads During Transportation

2.3.1.2 Pressure

Pressure experienced by SEM during ground handling and transportation shall be less than 19.3 kPa with a maximum de-compression rate of -1.79 kPa/sec.

Note: The average pressure change rate experienced during normal air transport is equivalent to a 3.8-7.6 m/sec change in altitude.

2.3.2 Launch and Ascent Environments

Launch environments derive from the specific launch vehicle on which the ERO Spacecraft (which hosts SEM) is launched, and its ascent profile. The primary and secondary launch vehicles have been selected for the ERO mission, but details are not yet available from the launch provider.

Generic environments are included in this requirements document from GSFC-STD-7000B. Ariane 6 User’s Manual can be used as a reference.

2.3.2.1 Spacecraft Tip-Off Rates

The rotational rates once the launch vehicle has separated from the space vehicle are not available at this time. SEM shall be designed using loads presented in Section 2.3.2.2, 2.3.2.3, and 2.3.2.5 until updated requirements are available.

2.3.2.2 Static-Equivalent Accelerations

The mass-acceleration curve (MAC) shown in Figure 2-1 provides a set of quasi-static limit loads that shall be used for preliminary hardware design. The MAC covers the peak dynamic response of the payload hardware to the low-frequency launch environment.

Figure 2-1 Generic Mass-Acceleration Curve (MAC) – Flight Limit

Method Ground Handling (Dolly) -0.25 0.25 -0.25 0.25 -1.60 0.00 Forklifting -1.00 1.00 -0.75 0.75 -1.50 0.50 Hoisting 0.00 0.00 0.00 0.00 -1.50 0.00 Notes :

- Minus s ign denotes downward for vertica l load. Vertica l Loads include gravi ty (-1 G)

- Vertica l loads to be appl ied independently. In-plane loads to be combined with 1 G vertica l loading

Fore/Aft (G) Lateral (G) Vertical (G)

Method Air (C5) -2.00 1.00 -1.00 1.00 -3.60 1.60 Road -1.00 1.00 -0.75 0.75 -2.50 0.50 Notes :

- Minus s ign denotes downward for vertica l load. Vertica l Loads include gravi ty (-1 G)

- Vertica l loads to be appl ied independently. In-plane loads to be combined with 1 G vertica l loading

Fore/Aft (G) Lateral (G) Vertical (G)

2-4

The breakpoints for the MAC curve are given in Table 2-4.

MAC loads are applicable to all components and subsystems regardless of their fundamental frequency. However, the mass that should be used for deriving loads using the MAC is a function of the fundamental frequency of the component or the fundamental frequency of the structure on which the component is mounted. One should use the component mass down to the level of assembly which either has its fundamental frequency below 80 Hz or is driven by structure which has fundamental frequency below 80 Hz. For example, an electronics box with a first mode of 120 Hz which mounts to spacecraft structure with a first mode below 80 Hz would be designed based on a MAC load using the weight of the electronics box. Everything within the electronics box would be designed for the same quasi-static MAC load as the electronics box. The MAC curve should not be applied to items within the electronics box in this example.

Table 2-4: MAC Design Limit Loads

Hardware Mass (Kg)

Limit Load (G, any

Direction) 1 or less 68.0

5 49.0 10 39.8 20 31.2 40 23.8 60 20.2 80 17.8 100 16.2 125 14.7 150 13.5 175 12.6

2-5

Usage Notes:

• MAC loads shall be applied in each axis independently

• Hardware shall be designed to show positive margin for limit load using factors of safety defined in MSR-CCRS-SYS-REQ-0002.

• Linear interpolation may be used between breakpoints to determine limit load

2.3.2.3 Sinusoidal Vibration

Table 2-5 provides a generic sine environment for the preliminary design of components and subsystems. The sine sweep vibration levels shown in Table 2-5 are defined at the hardware mounting interface.

Table 2-5: Interface Sine Sweep Vibration Levels

Frequency Flight Level Protoflight/Qual Level 5 to 20 Hz 0.5 in. (double amplitude) 0.63 in (double amplitude) 20 to 100 Hz4 10.0 g 12.5 g

• Flight/Protoflight level sweep rate shall be 4 oct/min

• Qualification level sweep rate shall be 2 oct/min

2.3.2.4 Acoustic Vibration

The acoustic environment SEM will experience is shown in Table 2-6. While all hardware should be assessed for sensitivity to direct acoustic impingement, unless the component or subsystem has structure which is light-weight and has large surface area (typically a surface to weight ratio of < 150 in2/lb), it is expected that the random environment will be the dominant high-frequency loading condition rather than the acoustic environment.

Table 2-6: Acoustic Environment

One-Third Octave Center Frequency (Hz)

Flight Level (dB)

Qual/Protoflight Level (dB)

20 126.5 129.5 25 127.7 130.7

31.5 127.0 130.0 40 128.5 131.5 50 130.0 133.0 63 131.5 134.5 80 132.5 135.5 100 133.0 136.0

200 or Greater 12.0

2-6

125 133.8 136.8 160 133.7 136.7 200 133.0 136.0 250 133.0 136.0 315 133.0 136.0 400 131.0 134.0 500 129.0 132.0 630 128.4 131.4 800 128.6 131.6 1000 126.9 129.9 1250 123.1 126.1 1600 118.3 121.3 2000 116.5 119.5 2500 115.0 118.0 3150 113.1 116.1 4000 112.5 115.5 5000 111.8 114.8 6300 111.0 114.0 8000 110.0 113.0 10000 109.1 112.1

OASPL 143.7 146.7

Usage Notes:

• Acceptance/protoflight test duration shall be 1 min

• Qualification test duration shall be 2 min

2.3.2.5 Random Vibration

Table 2-7 provides a set of random vibration levels for hardware design. The levels shown are the generalized random vibration environment defined in ref A1. These levels are applicable to hardware weighing less than 50 lbs and having resonant frequencies greater than 80 Hz. The levels for hardware weighing more than 50 lbs may be reduced using the method specified in GEVS.

Hardware with resonant frequencies below 80 Hz may be designed using only the MAC design loads specified in Section 2.3.2.3 as the MAC loads include the effect of mechanically transmitted random vibration up to 80 Hz.

Table 2-7: Random Vibration Levels (Less than 50lbs)

Frequency (Hz) Acceptance Protoflight/Qualification 20 0.013 g2/Hz 0.026 g2/Hz

20 – 50 +6 dB/Oct +6 dB/Oct 50 – 800 0.080 g2/Hz 0.16 g2/Hz

800 – 2000 -6 dB/Oct -6 dB/Oct 2000 0.013 g2/Hz 0.026 g2/Hz

Overall 10.0 grms 14.1 grms

2-7

• Acceptance/protoflight level random test duration shall be 1 min

• Qualification level random test duration shall be 2 min

2.3.2.6 Mechanical Shock

Table 2-8 provides a generic shock environment that may be used for hardware design until the mission specific shock environments can be defined. Hardware shall be designed to withstand the shock spectrum environment shown in Table 2-8 without any damage or degradation of performance.

Table 2-8: Shock Response Spectrum (Q=10)

Frequency (Hz) Acceptance Level (g) Protoflight/Qualification (g)

100 20 28 1000 500 700 10000 500 700

• Shock testing performed by firing the actual device shall consist of 2 actuations for protoflight/qualification testing and 1 actuation for acceptance testing.

• Simulated shock testing shall be performed to the levels specified in Table 2-8 with 2 tests per axis for qualification testing and 1 test per axis for acceptance/protoflight testing.

All components shall be assessed for damage due to shock based on shock sensitivity or proximity to shock sources. Components not considered susceptible to the shock environment may have shock testing deferred to the level of assembly that allows for actuation of the actual shock producing device. Any component considered to be susceptible to the shock environment should consider shock testing at the component level to demonstrate shock qualification.

2.3.2.7 Pressure

SEM shall be designed to be compatible with the peak depressurization rate of 5.0 kPa/s (50 mbar/s).

2.3.2.8 Launch Site Electromagnetic Environment

The system shall demonstrate by analysis that the delivered and installed unit, as well as on-orbit performance, will not be damaged or degraded by exposure to the launch vehicle and launch site RF emissions less than values in Table 2-9.

Table 2-9: LV RF Emissions Limits

Frequency Range E-Field (V/m) (no margin)

200 MHz – 18 GHz (except of peaks below)* TBD

2.200 GHz to 2.3 GHz (LV S-Band) TBD

2-8

5.4 GHz to 5.9 GHz (C-Band Range Radar) TBD

1262.19 to 1345.04 MHz (Launch Site Radars)** TBD

* Consistent with LV EM field strength near the separation plane ** SC Radiated Susceptibility (RS) limit is 5 V/m; LS radar might need mask/cover considerations

Note: Launch vehicle and launch site values are presently under development. SEM shall assume values from comparable systems such as Ariane 5 or Falcon 9 until such values are explicitly known and provided via specification update.

2.3.3 In-Space Environments

SEM will be exposed to the environment described in this section, based on the concept of operations of the ERO and CCRS mission.

2.3.3.1 Mechanical Shock

ERO spacecraft will separate a major module during in-space operations. This results in a shock environment similar to that experienced during launch vehicle separation.

Table 2-10 provides a generic shock environment that may be used for hardware design until the mission specific shock environments can be defined. Hardware shall be designed to withstand the shock spectrum environment shown in Table 2-10 without any damage or degradation of performance.

Table 2-10: Shock Response Spectrum (Q=10)

Frequency (Hz) Acceptance Level (g) Protoflight/Qualification (g)

100 160 224 630 1000 1400

10000 1000 1400

Usage Notes:

• Shock testing performed by firing the actual device shall consist of 2 actuations for protoflight/qualification testing and 1 actuation for acceptance testing.

• Simulated shock testing shall be performed to the levels specified in Table 2-10 with 2 tests per axis for qualification testing and 1 test per axis for acceptance/protoflight testing.

All components shall be assessed for damage due to shock based on shock sensitivity or proximity to shock sources. Components not considered susceptible to the shock environment may have shock testing deferred to the level of assembly that allows for actuation of the actual shock producing device. Any component considered to be susceptible to the shock environment should consider shock testing at the component level to demonstrate shock qualification.

2.3.3.2 On-orbit Dynamic Load

SEM shall be compatible with the following dynamic environment:

• Linear Acceleration: 0 m/s²

• Angular Acceleration: 0.015 deg/s² around X, 0.024 deg/s² around Y, 0.1 deg/s² around Z

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• Angular Rate: 2 deg/s, all axes (Pure torque configuration. Some residual forces can appear but are considered low.)

2.3.3.3 Magnetic compatibility with Orbiting Sample

SEM shall not generate magnetic fields greater than 0.2 mT at the surface of the unit.

2.3.3.4 On-orbit Electromagnetic Compatibility

SEM shall meet the requirements as shown in Table 2-11.

Table 2-11: On-orbit Electromagnetic Compatibility

Radiated Susceptibility (E Field)

The unit shall not show any malfunction or deviation from the specified performance when irradiated with the following E-fields:

• 30 MHz to 18 GHz: 2 V/m rms The radiated Eƒ-field shall be amplitude modulated by a sine wave at 1 kHz with a modulation depth of 50 % For non RF units which do not feature high frequency clocks, the upper frequency may be limited to 10 times its highest frequency or 1 GHz whichever is higher.

Radiated Susceptibility (E Field)

The unit shall not show any malfunction or deviation from the specified performance when irradiated with the E-fields as listed below

2.3.4 SEM Environmental Verification

SEM environmental verification shall meet the requirements in Section 4 of the CCRS Environmental Requirements Document MSR-CCRS-SYS-REQ-0002.

2.3.4.1 SEM System Level Environmental Verification

SEM mechanisms and components shall be capable of a minimum of 50 deployments without requiring replacement or refurbishment.

2.3.4.2 SEM Verification Test Matrix

Table 2-12 presents the minimum expected testing that shall be performed as validation and verification of the requirements as specified herein. Specified testing of design elements not applicable to the as-proposed configuration may be removed and alternate testing strategy proposed in its place. Formal test plan shall be reviewed and approved at SRR. Tests listed in this section shall not override the requirements in Section 2.3.4.

Table 2-12: SEM Verification Test Matrix

Article Minimum Required Testing Prototype - Testing as required to validate proposed design is capable of meeting deployment parameters

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Engineering Development Unit

(EDU)

- Kick-off spring characterization

- Kick-off spring mission life performance validation

- Bearing frictional performance and drag characterization

- Characterization of other parasitics and failure modes

- Ambient Deployment with weighted mockup

- Ambient Deployment with zero-gravity or gravity negation

Engineering Test Unit (ETU) / Qualification Unit

- Kick-off spring characterization and screening

- Ambient Deployment with weighted mockup

- Ambient Deployment with zero-gravity or gravity negation

- Random Vibration and post-vibration Ambient Deployment with zero-gravity or gravity negation

- Shock and post-shock Ambient Deployment with zero-gravity or gravity negation

- Thermal Vacuum with zero-gravity or gravity negation deployment

- Thermal Vacuum and post-thermal Ambient Deployment with zero-gravity or gravity negation Assembly, Integration, & Test Unit (AM) / Structural – Thermal Model

- Validation of mass properties

- Validation of thermal properties

Flight Model (FM) - Kick-off spring characterization and screening

- Ambient Deployment with weighted mockup

- Ambient Deployment with zero-gravity or gravity negation

- Random Vibration and post-vibration Ambient Deployment with first motion

- Shock and post-shock Ambient Deployment with first motion

- Thermal Vacuum with first motion deployment

- Thermal Vacuum with post-thermal ambient first motion deployment

- Contamination Bake-out as-required

2.4 Mechanical Requirements

2.4.1 Deliverables

SEM deliverables shall be comprised of two separate elements:

1. Fixed Chassis: This fixed part of the SEM is rigidly mounted to the CCRS Module primary structure and remains with CCRS Module after the ejection event. All powered components shall be on this side of the system attached to the CCRS.

2. EES Fitting(s): This part of the SEM remains with the EES after the ejection event and shall consist of passive non-electrical components as there are no electrical provisions crossing over this SEM-EES interface. These fittings should be as compact and

2-11 lightweight as possible to minimize disruptions to EES mass properties and aerodynamics. This fitting(s) could be a single fitting or multiple subassemblies.

2.4.2 SEM Design Envelope

SEM shall be compatible with the EES dimensional envelope as defined in the SEM Mechanical Interface Control Document (MICD) in Section 2.4.6, which will be developed between the contractor and NASA GSFC. The dimensions provided in Appendix A shall be for reference only and are superseded by the final released drawing.

2.4.3 SEM Natural Frequencies

SEM mounted on a rigid interface including EES mass properties shall have its first main frequencies as greater than 60 Hz (TBC) lateral and greater than 60 Hz (TBC) axial.

2.4.4 Lift and Handling points for EES I&T

SEM shall provide GSE interface points to allow all Integration & Test activities, safe-movement to environmental testing and handling.

2.4.5 Hold Down Release Mechanisms (HDRMs)

2.4.5.1 HDRM Usage and Life

SEM HDRMs shall be field/customer resettable (not permanent) during ground tests. HDRM shall be capable of no less than 50 resets prior to requiring factory replacement or refurbishment.

2.4.5.2 HDRM Quantity

SEM shall utilize no greater than four (4) HDRM each with a primary and a secondary firing drive. Preferred configuration utilizes no more than three (3) HDRMs.

2.4.6 Interfaces

SEM shall interface with EES and CCRS Module structure per SEM MICD (Appendix A)

(TBR).

2.4.7 RTAS Installation Loads (No margin applied)

SEM shall withstand Robotic Transfer Assembly System (RTAS) installation loads without releasing EES or negatively affecting SEM release performance. Loads shall be no greater than as specified in Table 2-13 when applied along the EES central axis.

Table 2-13: RTAS Installation Loads (No margin applied)

Load Type Operational (TBC) Limit (TBC) Axial (N) 473 1380

Lateral (N) 120 155 Torque (N-m) 0.5 1.0 Moment (N-m) 52 65

2.4.8 RTAS Installation Deflection

SEM shall prevent unacceptable deflections at hold down points during installation of EES closeout lids using robotic arm (RTAS). Permissible deflections to be determined after assessment of proposed HDRM configuration and provided by SRR. (TBD)

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2.4.9 EES Fittings Structure Integrity

EES Fittings as described in Section 2.4.1 item 2 shall maintain structural integrity through entry, decent, and landing, and shall feature an interface point to be used for EES retrieval at the landing site. Entry, descent, and landing loading and environmental conditions will be provided prior to Contractor SRR. Interface points and design shall be via mutual agreement and defined on the MICD.

2.5 Mass Requirements

2.5.1 SEM Mass

SEM total mass shall not exceed 14kg.

2.5.2 SEM Mass on the EES

The portion of SEM mass that remains on the EES after separation (see Section 2.4.1, EES Fitting(s)) shall not exceed 1.81 kg.

2.5.3 EES Mass Properties

EES shall be assumed to have mass properties as shown in Table 2-13. Figure 2-2 shows the Center of Gravity (CG) location reference of the EES.

Table 2-14: EES Mass Properties (relative to CG)

Property Nominal Range Unit Mass 79.4 ±5% kg CG from EES Nose 293 ±3 mm CG Radial Offset 0 [0 5] mm Ixx 9.17 ±5% kg-m2 Iyy 5.75 ±5% kg-m2 Izz 5.75 ±5% kg-m2

Figure 2-2: EES CG Location

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2.6 Electrical Requirements

2.6.1 Polarity Reversal Protection

SEM shall have built-in protection to prevent damage due to polarity reversal at the power inputs.

2.6.2 SEM Firing Pulses

SEM shall have access to up to four primary and four secondary switched pulsed services.

2.6.3 SEM Firing Current

SEM HDRM services shall supply 17V at 7.0A at no more than 2.4 ohms per HDRM.

2.6.4 SEM Activation

SEM HDRMs shall be energized one at a time. If the SEM design features more than one HDRM, timing of release shall not be required to be synchronous. All deployment characteristics shall be met when, and if, each HDRM is released independently.

2.6.5 Pulse Width

SEM HDRMs maximum pulsed duration shall not exceed 500ms.

2.6.6 Dead Time

SEM HDRMs shall have a dead time between pulses not less than 10ms. TBC

2.6.7 Maximum Spread Current

SEM shall be designed to achieve its specified performance with Max current spread within +/- 20%.

2.6.8 Grounding

The system shall have a DC Impedance across the units deployment interface (separation plane) of less than one ohm measured at the mechanical mounting interface and the ground strap locations.

2.6.9 Ground Straps

The system shall provide for a minimum of two ground strap mounting locations on both sides of the deployed interface.

2.6.10 Interface Electromechanical Design

Mechanical coatings and surface treatments at the deployment interface (separation plane) shall be compatible with meeting the grounding requirement.

2.6.11 ESD Grounding

All metal components in the unit shall be grounded to one or the other side of the separation plane mounting surfaces/ground strap locations with less than a 10 ohm impedance.

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2.7 Thermal Requirements

2.7.1 Thermal Hardware Mounting

SEM shall provide mounting surface to heaters, thermostats and thermistors per the EES Thermal ICD (TBD), which will be developed between the contractor and NASA GSFC.

2.7.2 Survival (Non-Operational) Temperature

SEM shall meet survival temperatures: - 30 to +50 (TBC) °C

2.7.3 Operational Temperature

SEM shall be designed to achieve its specified nominal and off-nominal release performance at temperature +18 to +28 (TBC) °C. CCRS will bring SEM to this temperature range for EES release.

NOTE: EES must remain significantly colder than SEM at time of release. Thermal distortions and differentials must be accounted for in the design and performance analysis. Thermal gradients are currently unknown but both CCRS structure and EES interfaces are expected to be up to 50C colder than the SEM Operational Temperature. SEM Operational Temperature was selected as desired for mechanism operation and may be reduced closer to expected interface temperatures by mutual agreement to reduce thermal induced errors and better meet deployment performance. EES temperature at release will be provided at Contractor SRR.

2.8 Fault Tolerance and Reliability Requirements

2.8.1 EES Release Reliability

SEM shall have a probability of failure <0.1% when subjected to the operating environments and mission lifetime requirements detailed within this specification.

2.8.2 Dual-Fault Tolerance - EES Release

No two failures of the SEM subcomponents shall lead to the inadvertent release of the EES.

2.9 Terrestrial Contamination Control Requirements

2.9.1 Terrestrial External Surface Cleanliness

The external surface cleanliness of the SEM shall be maintained at VC-0.5-1000+UV per IEST-

STD-CC1246E.

2.9.2 Terrestrial Internal Surface Particulate Cleanliness

The internal surface particulate cleanliness of the SEM shall be less than 0.1% area coverage at delivery.

2.9.3 Internal Surface Molecular Cleanliness

The internal surface molecular cleanliness of the SEM shall be less than 0.5 µg/cm2 at delivery.

2.9.4 Terrestrial Outgassing

SEM shall demonstrate a terrestrial outgassing stability of <1.0%/hour (per ASTM E-2900, Method B) at 50C and measured by a QCM at -40C.

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2.9.5 GSE cleanliness

All GSE shall be cleaned to VC-0.5-1000 (per IEST-STD-CC1246E) prior to entering a controlled environment or cleanroom.

2.9.6 GSE outgassing

GSE used in a vacuum environment shall demonstrate an outgassing stability of <3.0%/hour (per ASTM E-2900, Method B) at 50°C and measured by a QCM at -40°C.

2.9.7 Bioassay

SEM shall be compatible with damp assays using procedures described in ECSS-Q-ST-70-55.

2.9.8 Compatibility with alcohol cleaning

SEM shall be compatible with alcohol cleaning (70% IPA or ethanol).

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Appendix A SEM MICD

B-1

Appendix B Abbreviations and Acronyms AI&T Assembly, Integration and Test AIAA American Institute of Aeronautics and Astronautics AM AI&T Model ATP Acceptance Test Procedure BTC Break-The-Chain CAD Computer Aided Design CAM Containment Assurance Module CBE Current Best Estimate CCB Configuration Control Board CCM Capture and Containment Module CCRS Capture, Containment and Return System CDR Critical Design Review CDRL Contract Data Requirements List CG Center of Gravity CPT Comprehensive Performance Testing C-OS Contained OS (OS + PCV +SCV) COTS Commercial Off-the-Shelf DFMR Design For Minimum Risk DID Data Item Deliverables DILS Deliverable Items List and Schedule EAR Export Administration Regulations EAC Estimate At Completion EDL Entry, Descent, and Landing EDP EES Delivery Phase (EDP) EDU Engineering Development Unit EES Earth Entry System (EES Aeroshell + C-OS + CAM lid; system that separates for Earth return) EIDP End Item Data Package EM Engineering Model ERM Earth Return Module ERO Earth Return Orbiter ESA European Space Agency ETU Engineering Test Unit GEVS General Environmental Verification Standard GSE Ground Support Equipment HDRM Hold-Down…