Attachment B - MSRCCRS-VS-SPEC.pdf

Effective Date: August 07, 2022 Expiration Date: August 07, 2027

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-VS-SPEC-0036, Revision -

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

NASA/GSFC Code 435

CCRS Capture-Enclosure (CE) Camera

Specification

MSR-CCRS CMO (CK)

August 07, 2022

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

CE Camera Specification MSR-CCRS-VS-SPEC-0036 Revision-

Effective Date: 08/07/2022 ii

Signature/Approval Page

Prepared by:

Ron Shiri

Reviewers/Approvers:

R. Ferrer 7/28/2022 R. Montgomery 7/27/2022 D. Wilson 7/27/2022 K. Cleveland 7/27/2022 G. Casto 7/27/2022 C. Bacon 7/27/2022 R. Schnurr 7/26/2022 B. Yang 7/26/2022 K. Grello 7/26/2022 B. Bos 7/26/2022 G. Cataldo 7/26/2022 K. Wrenn 7/20/2022 R. Shiri 7/19/2022 B. Rizk 7/19/2022 W. Posey 7/15/2022 D. Donovan 7/13/2022 T. Feehan 7/11/2022 J. Capone 7/06/2022 D. Donovan 7/06/2022

Approved by:

Keith Walyus MSRCCRS CCB Chairman 7/27/2022

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

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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 will 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 iv

Change History Log

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

Revision- July 28, 2022 Released per MSRSCCRS-CCR-00117, 07/27/2022 v

Table of TBDs/TBRs/TBSs

Action Item No.

Location Summary Individual/ Organization

Actionee vi

Table of Contents

1 INTRODUCTION

1.1 Purpose

1.2 Scope

1.3 Related Documentation

1.4 Applicable Documents

2 REQUIREMENTS

2.1 Design Heritage Requirement

2.2 Imaging and Performance Requirements

2.3 Mechanical Requirements

2.4 Magnetic Requirements

2.5 Environmental Requirements

2.6 Transportation Requirements

2.7 Pressure Requirements

2.8 On-Orbit Dynamic Requirements

2.9 Ground Environmental Requirements

2.10 Thermal Requirements

2.11 Contamination Requirements

2.12 Radiation Requirements

2.13 Electrical Requirements

2.14 EMI/EMC Requirements and Environment

2.15 Command and Telemetry Format

2.16 Command Interface

2.17 Imaging Commands

2.18 Image Generation

2.19 Lifetime Requirements

3 QUALITY ASSURANCE PROVISIONS

3.1 Quality Assurance Requirements

4 PREPARATION FOR DELIVERY

4.1 Delivery Requirements

ABBREVIATIONS AND ACRONYMS

vii

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List of Figures

Figure 1. Object Volume and Camera Volume Allocation Figure 2. Illustration of 2.2.5 pertaining to image saturation caused by overexposure Figure 3. Camera Design Limit Loads Figure 4. Dose in Silicon as a function of spherical aluminum shielding for 2027 launch worst-case scenario Figure 5. NIEL equivalent proton fluence in silicon at the center of solid aluminum spheres for 2027 launch worst-case scenario Figure 6. Active period-proton spectra to use for proton induced SEE computations Figure 7. Integral mission cumulative solar proton spectrum for total mission cruise duration for 2027 launch worst-case scenario Figure 8. CREME96 integral flux as a function of LET for four levels of activity for a component shielded by 1 g/cm2; peak 5-minute (worst case peak), worst day and worst week (worst case SPE) and quiet (GCR)

Figure 9. Proposed camera grounding scheme in which camera chassis is isolated from spacecraft structure to maintain a single-point ground near the Jettison Avionics assembly viii

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List of Tables

Table 1. Mechanical Factors of Safety Table 2. Mass-Acceleration Curve (MAC) Table 3. Camera Sine Vibration Environment (TBD) Table 4. Camera Random Vibration Environment, (22.7 kg, or less) Table 5. Qualification Level Shock Response Spectrum (Q=10) Table 6. Flight/Acceptance Level Acoustic Environments (acceptance/protoflight test duration = 1 minute; qualification test duration = 2 minutes) Table 7. Transportation Loads Table 8. CE Radiation Environment Table 9. Dose in silicon as a function of spherical aluminum shielding for 2027 launch worst-case scenario Table 10. Total mission non-ionizing dose in silicon as a function of spherical aluminum shielding for the 2027 launch worst-case scenario Table 11. Displacement damage equivalent 10-MeV proton fluence in silicon as a function of spherical aluminum shielding for 2027 launch worst-case scenario Table 12. Displacement damage equivalent 50-MeV proton fluence as a function of spherical aluminum shielding for 2027 launch worst-case scenario Table 13. Proton non-ionizing energy loss (NIEL) values used to calculate TNID in silicon as a function of spherical aluminum shielding Table 14. Proton differential spectra for proton induced SEE computations Table 15. Integral cumulative solar proton fluence for total mission duration for the

2027 launch worst-case scenario Table 16. Differential cumulative solar proton fluence for total mission duration for the

2027 launch worst-case scenario Table 17. CREME96 integral flux as a function of LET for four levels of activity for a component shielded by 1 g/cm2; peak 5-minutes (worst-case peak), worst day and worst week (worst case SPE) and quiet (GCR)

Table 18. EMI-EMC test levels and limits Table 19. Temperature and Humidity Storage

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1 Introduction

1.1 Purpose

The National Aeronautics and Space Administration (NASA) is developing the Capture, Containment and Return System (CCRS) as part of NASA’s Mars Sample Return campaign.

CCRS will be the primary payload of the European Space Agency’s Earth Return Orbiter which will launch into space on an Ariane 6 no earlier than 2027. The CCRS will capture an Orbiting Sample (OS) container in Mars orbit, process and prepare the OS for a return to Earth and deliver the Martian surface samples to the Utah Test and Training Range (UTTR) operated by the United States Air Force. The Capture Enclosure of CCRS provides the OS capture and OS processing functionality for the system.

To carry out the CCRS tasks, after the Orbiting Sample (OS) is captured and stationary, the Capture-Enclosure (CE) cameras will collect and transfer to the Avionics sub-system at least one image of the OS endcap to determine its orientation as described by 1) the OS Interface Control Document (ICD) and 2) Flight Subsystems to Vision System Interface Control Document (ICD).

The Capture Enclosure (CE) architecture requires two cameras to acquire these data sets using vantage points inside the CE. The cameras will be mounted to fixed position inside CE. The illumination for the CE imaging will be provided by an array of LDs illuminating the Orbiting Sample (OS) after CCRS capture with an illuminance and areal coverage to enable acquisition of an image by the Vision System camera to allow an observer to determine the correct end of OS facing the CCRS interior.

This document’s purpose is to define the CE camera specifications for the flight model as well as the engineering development units. The specification along with the CCRS Capture-Enclosure (CE) Camera Statement of Work together scope the technical breadth of the work necessary for a successful camera delivery.

1.2 Scope

Except where specifically noted in this document or the statement of work, all the specifications described in this document apply to the camera components delivered to NASA by the contractor.

1.2.1 A complete flight CE camera will consist of two cameras. Additional flight-spare and engineering model unit will also be required as described in the statement of work

(SOW).

1.2.2 The CE cameras will document the post-capture condition and orientation of the OS during processing for return to Earth.

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1.3 Related Documentation

The following documents form a part of this specification to the extent specified herein.

1.3.1 This specification will be considered the superseding requirements in case of a conflict with any of the documents listed in this section.

1.3.2 The contractor will immediately report all such conflicts to NASA/GSFC for resolution.

Applicable Documents are documents with material relevant to requirement compliance.

Reference Documents are documents included to provide additional information and background content for clarification purposes.

1.4 Applicable Documents

MSR-CCRS-VS-SOW-0008 CCRS CE Camera Statement of Work

GSFC-STD-7000 General Environmental Standard (GEVS)

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

MIL-STD-461 Requirements for the Control of Electromagnetic Interference Characteristics of Subsystems and Equipment

MSR-CCRS-SYS-SPEC-0004 CCRS Mechanical and Mechanism Design Specification

ECSS-Q-ST-70-55 Microbial examination of flight hardware and cleanrooms

NASA-STD-8719.24 NASA Payload Safety Requirements

1.4.1 Reference Documents

Reserved

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2 Requirements

2.1 Design Heritage Requirement

2.1.1 [VS-C-001] The CE camera design shall have previously successfully operated as part of a spaceflight mission in an environment relevant to the conditions described in this specification or shall be based on hardware that has completed a full environmental test campaign relevant to the conditions described in this specification.

Note: Minor modifications to the hardware and software of previously flown camera are acceptable unless new technologies are required for the system to function.

Rationale: Due to the short delivery schedule, there is not time for the contractor to develop and mature new technology with acceptable risk.

2.2 Imaging and Performance Requirements

2.2.1 [VS-C-002] The CE camera shall generate color image information via incorporation of a

Bayer filter on the camera detector array.

Note: Bayer filters are described in U.S. Patent No. 3,971,065.

2.2.2 [VS-C-003] When mounted within the camera envelope allocation specified by 2.3.12 the delivered CE camera shall view and resolve with a contrast ≥ 0.6 a 3 mm or larger object positioned anywhere within the object volume described in Figure 1.

Note: This contrast value applies to the entire camera. It includes the optical as well as detector effects and any other effect that degrades the camera’s point spread function (PSF). It applies to all image orientations on the camera’s detector. This requirement sets the object-space depth of field for the camera as well as the image contrast.

Figure 1. Object Volume and Camera Volume Allocation

2.2.3 [VS-C-004] In a single exposure the CE camera shall image a scene with a luminance of

7.8 lumens/m2/steradian (7.8 candela/m2) with a signal-to-noise ratio (SNR) ≥ 10 using an exposure time ≤ 1 s Rationale: The stated luminance value accounts for the minimum possible illuminance from the illumination system, the lowest BRDF of the orbiting sample and the worst-case contamination expected in the capture enclosure after the orbiting sample capture.

Note: The selected exposure time freezes out motion. The luminance values are set by the reflectivity of Capture Enclosure hardware and the illuminance provided by the illumination system.

2.2.4 [VS-C-005] The CE camera shall meet the contrast and SNR requirements specified above (2.2.2 and 2.2.3) when illuminated by ≤ 9500 lux from a light source located anywhere outside of the camera’s as-built field of view.

Rationale: This is the worst-case maximum illuminance that a single illumination module could produce based on the current capture enclosure design.

2.2.5 [VS-C-006] When imaging an object over-exposed by a factor of 100 or less with the

Capture Enclosure Camera, the image saturation caused by the over-exposure shall remain isolated to the portions of the image containing the over-exposed feature to within 10% of the maximum dimension of the image feature when it is properly exposed as shown in Figure 2.

Note: This means that if an object in the field of view subtends a diameter of 100 pixels when it is properly exposed, then when it is overexposed the saturated features caused by

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the overexposure will not increase the size of the saturated image by more than 10% (100 pixels to 110 pixels) if the over-exposure factor is 100 or less.

Rationale: This requirement controls how well the camera detector controls saturation and the optical system suppresses in-field stray light. It is necessary in case we have specularly-reflective objects in the field of view that cannot be made compliant with the BRDF (bi-directional reflectance distribution function) requirements.

Figure 2. Illustration of 2.2.5 pertaining to image saturation caused by overexposure

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2.2.6 [VS-C-007] The CE camera shall acquire images at a rate ≥ 1 Hz when requested by CE.

2.2.7 [VS-C-008] Any ghost images created by the CE camera at the camera’s detector plane shall have a maximum illuminance ≤ 0.03 the maximum illuminance generated by the original image at the detector plane.

2.3 Mechanical Requirements

2.3.1 [VS-C-009] Each CE camera shall have removable, non-flight protective covers that fit within the envelope and are removable along the camera axis.

2.3.2 [VS-C-010] The CE camera shall contain provisions for a grounding strap to be attached from the component chassis for connection to an external conductive structure as documented in a Mechanical Interface Control Drawing (MICD).

Note: The MICD will be developed between NASA and camera contractor.

2.3.3 [VS-C-011] The grounding lug locations on the CE camera component chassis or the tie points in contact with the ground strap shall have a minimum contact area of 80 mm².

2.3.4 [VS-C-012] The grounding lug locations on the CE camera component chassis shall remain free of any material finish that may affect the reliability of the ground connection.

2.3.5 [VS-C-013] The DC resistance between the component chassis and ground strap of the

CE camera components shall not exceed 2.5 mΩ.

2.3.6 [VS-C-014] The DC resistance between the component chassis and mounting surface of the CE camera components shall not exceed 2.5 mΩ.

2.3.7 [VS-C-015] The CE camera component connectors shall electrically connect to chassis with a DC resistance less than or equal to 10 mΩ.

2.3.8 [VS-C-016] The CE camera surface resistivity shall be less than or equal to 1010

Ω/square.

2.3.9 [VS-C-017] The CE camera mechanical interface shall match the description documented in the Mechanical Interface Control Drawing (MICD) which will be developed between the contractor and NASA GSFC.

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2.3.10 [VS-C-018] The CE camera electrical interface shall match the description documented in the Electrical Interface Control Document (EICD) which will be developed between the contractor and NASA GSFC.

2.3.11 [VS-C-019] Each CE camera mass, not including harnessing, shall not exceed 0.75 kg.

2.3.12 [VS-C-020] The CE camera shall not exceed an envelope 170 mm x ⌀ 100 mm in size, as shown in Figure 1.

2.3.13 [VS-C-021] The CE camera shall provide external, optical alignment references for each camera with surfaces at least 15 mm x 15 mm in size.

2.3.14 [VS-C-022] The CE camera external, optical alignment references shall offer two adjacent surfaces suitable for alignment measurements without affecting the basic integrity of the system.

2.3.15 [VS-C-023] Over the operational temperature range, the angular offset of the CE camera’ boresight relative to the mounting interface shall not exceed 5 milliradians.

2.3.16 [VS-C-024] The CE camera angular alignment knowledge shall relate each camera boresight axis with respect to its mounting interface to an accuracy of 2 milliradian (3-σ) with respect to each axis.

Note: The uncertainty calculation needs to include measurement accuracy, thermal effects and shifts due to the launch environment and gravity release.

2.3.17 [VS-C-025] The CE camera fastening systems shall comply with the CCRS Mechanical and Mechanism Design Specification MSRCCRS-SYS-SPEC-0004, Sections 3.1.

2.4 Magnetic Requirements

2.4.1 [VS-C-026] The CE camera shall not create a magnetic field greater than or equal to 0.05 mT

2.5 Environmental Requirements

2.5.1 [VS-C-027] The CE camera shall meet the performance requirements of Section 2.2 after exposure to the environments specified in this section.

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2.5.2 [VS-C-028] The CE Camera shall have a probability of failure <3.2% when subjected to the operating environments and mission operational lifetime requirements detailed within this specification.

2.5.2 [VS-C-029] The CE camera shall demonstrate positive Margins of Safety under limit loads for all yield and ultimate failures using the Factors of Safety (FS) defined in Table 1. Margin of Safety (MS) is defined as follows:

MS = (Allowable Stress (or Load) / (Applied Limit Stress (or Load) x FS)) -1

Table 1. Mechanical Factors of Safety

2.5.3 [VS-C-030] Any CE camera primary or secondary structure comprised of composite materials, Beryllium, bonded joints and/or bonded inserts shall receive proof testing to

1.25 x Limit Load.

Note: Actual flight component testing is preferred but testing of representative sets of hardware with a similarity qualification argument can be used if approved by the NASA/GSFC Contracting Officer Representative. Qualification by analysis only is not acceptable.

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2.5.4 [VS-C-031] The CE camera design shall withstand the quasi-static design limit loads defined in the mass-acceleration curve (MAC) shown in Figure 3 and Table 2 without damage or degradation of performance.

Note: Quasi-static acceleration represents the combination of steady-state accelerations and the low frequency mechanically transmitted dynamic accelerations that occur during launch. The design loads shown below will be updated based on the results of coupled loads analysis. Linear interpolation should be used between breakpoints to determine the appropriate limit load as a function of Camera weight. Note that these design limit loads are intended to cover only the low frequency launch environment and must be used in conjunction with the random vibration environments to assess structural margins.

Usage Notes:

• MAC loads should be applied in each axis independently Usage Notes:

• MAC loads should be applied in each axis independently

• Hardware should be designed to show positive margin for limit load using factors of safety defined this document.

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

Table 2. Mass-Acceleration Curve (MAC)

Figure 3. Camera Design Limit Loads.

2.5.5 [VS-C-032] Any CE camera component shall have a fundamental frequency greater than

100 Hz when hard mounted at its structure interface.

Note: If a component fails to meet the specified fundamental frequency, the vendor will supply a detailed finite element model of the component, correlated to modal survey test results up to 50 Hz, to be used in coupled loads analyses. Requirements for the submitted finite element model are documented in the Statement of Work (SOW).

2.5.6 [VS-C-033] The CE camera shall undergo qualification, protoflight or acceptance (level depends on qualification status of unit) sine vibration testing on all three axes at the levels shown in Table 3.

Note: Stiff components showing a first mode greater than 100Hz can be exempted from sine vibration testing upon approval by the NASA/GSFC COR. A generic sine vibration specification is provided for Protoflight (PF), Qualification (QT) and Acceptance (AT) levels applied at the Capture Enclosure to camera interface.

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Table 3. Camera Sine Vibration Environment (TBD)

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

Usage Notes:

• Levels may be notched to not exceed 1.25 times the design limit load. Peak levels at the low end of the frequency range (5-20 Hz typically) may be ramped up as needed to accommodate shaker table displacement limitations.

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

• Qualification level sweep rate shall be 2 oct/min

• Vibration levels are defined at the hardware mounting interface.

2.5.7 [VS-C-034] The CE camera shall demonstrate its ability to meet its performance requirements after being subjected to the random vibration environment in Table 4, applied at the mounting interface.

Note: This environment will be updated with random vibration analysis. For a lightweight camera, the highest design loads may be from this random vibration environment. These levels are applicable to hardware weighing less than 50 lbs and having resonant frequencies greater than 80 Hz.

Note: During the test, the test input level may be reduced (notched) at critical frequencies, if required, to limit the random vibration loads and/or acceleration responses to 3 dB above design limit levels.

Table 4. Camera Random Vibration Environment, (22.7 kg, or less)

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 0.016 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 Usage Notes:

• Acceptance/protoflight level random test duration = 1 min

• Qualification level random test duration = 2 min

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2.5.8 [VS-C-035] Any notching required by the contractor for the random vibration test shall comply with the random vibration notching requirement in the CCRS Capture Enclosure (CE) Camera Statement of Work.

2.5.9 [VS-C-036] The CE camera shall meet its performance requirements after being subjected to the shock environment in Table 5, applied at the camera mechanical interface.

Note: GSFC may defer the CE camera shock testing to the level of assembly that allows for actuation of the actual shock producing device, though a supplier shock sensitivity assessment would still be required in lieu of a component level test.

Table 5. Qualification Level Shock Response Spectrum (Q=10)

Frequency (Hz) Acceptance Level (g) Protoflight/Qualification (g) 100 160 224 630 1000 1400

10000 1000 1400

2.5.10 [VS-C-037] The CE camera design shall withstand, without any damage or degradation of performance, the equivalent Acceptance/Protoflight acoustic levels shown in Table 6.

Note: Contractor acoustic testing is currently not required for this procurement due to camera being located inside the structure.

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Table 6. Flight/Acceptance Level Acoustic Environments (acceptance/protoflight test duration = 1 minute; qualification test duration = 2 minutes)

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 80 132.5 135.5 100 133.0 136.0 125 133.8 136.8 160 133.7 136.7 200 133.0 136.0 250 133.0 136.0 315 133.0 136.0 400 131.0 134.0 500 129.0 132.0 630 128.4 131.4 800 128.6 131.6 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

2.6 Transportation Requirements

2.6.1 [VS-C-038] The maximum loads that CE camera will experience during ground operations or transportation shall be less than the values in Table 7 without damage or degradation of performance.

Table 7. Transportation Loads

Method Fore/Aft (g) Lateral (g) Vertical (g) Air -2.0 +2.0 -2.0 +2.0 -5.0 +5.0 Road -2.0 +2.0 -2.0 +2.0 -3.0 +3.0

Notes:

- Minus sign denotes downward for vertical load. Vertical Loads include gravity (-1g)

- Vertical loads to be applied independently. In-plane loads to be combined with 1g vertical loading

2.6.2 [VS-C-039] Materials and enclosures used for the CE camera transportation and storage shall not generate molecular or particulate contaminants that would make the system non-compliant with the contamination requirements in Section 2.10.

2.7 Pressure Requirements

2.7.1 [VS-C-040] The CE camera shall meet all performance requirements while operating over a pressure range of 1.08 x 105 N/m2 (813 Torr) to 1.3 x 10-12 N/m2 (1 x 10-14 Torr).

2.7.2 [VS-C-041] Pressure experienced by the CE camera 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.

2.7.3 [VS-C-042] The CE camera shall meet all performance requirements after exposure to a peak depressurization rate of less than -5.0 kPa/Sec during launch and ascent.

2.8 On-Orbit Dynamic Requirements

2.8.1 [VS-C-043] The CE camera shall survive all linear and angular accelerations to the stated requirement levels at the same time as described in 2.7.2 and 2.7.3.

Note: The linear and angular acceleration are not mutually exclusive.

2.8.2 [VS-C-044] The CE camera shall survive on-orbit when subjected to the maximum linear acceleration rate due to nominal thruster firing.

Note the linear acceleration rate will be mutually agreed upon between the contractor and NASA after selection.

2.9 Ground Environmental Requirements

2.9.1 [VS-C-045] The CE camera shall meet all of the performance requirements during exposure to air temperatures between +5° and +25° C and relative humidity between 30% and 70%.

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2.9.2 [VS-C-046] After being powered “OFF” and exposed to air temperatures of +5° to +30° C and relative humidity of 0 to 70%, the CE camera shall meet all of its performance requirements.

2.10 Thermal Requirements

2.10.1 [VS-C-047] Each camera shall contain at least one internal temperature sensor with resistance greater than 2000 Ohms located on or near the camera’s detector that provides an electrical signal proportional to temperature.

Note: Alternate resistance values needs to be cleared with GSFC

2.10.2 [VS-C-048] The internal camera temperature sensors shall measure temperature over a range from -50° C to +70° C with an accuracy of ±1° C.

2.10.3 [VS-C-049] The camera shall operate and meet all performance requirements over a temperature range of at least -25° C to +25° C.

2.10.4 [VS-C-050] The camera shall meet all performance requirements after extended exposure to temperatures ranging from -40° to +55° C.

2.10.5 [VS-C-051] The camera peak power output shall not exceed 8.8 W for any operational mode.

2.10.6 [VS-C-052] Each CE camera housing, lens barrel, and baffle outer surface treatments shall have a minimum normal emissivity of 0.8 over the wavelength range of 4 to 40 microns.

NOTE: Treatment options include tapes, multi-layer insulation, and metallic coatings or anodize, etc. Preferred option will be settled after camera vendor selection.

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2.11 Contamination Requirements

2.11.1 [VS-C-053] The CE camera shall be maintained at VC-0.5-1000+UV per IEST-STD-

CC1246E.

2.11.2 [VS-C-054] The surface particulate cleanliness of the CE camera shall be less than 0.02% area coverage particulate and < 0.33 µg/cm2 molecular (NVR) upon delivery to GSFC.

2.11.3 [VS-C-055] The CE camera shall be compatible with damp assays using procedures described in ECSS-Q-ST-70-55.

2.11.4 [VS-C-056] The CE camera shall be compatible with alcohol cleaning (70% IPA or ethanol).

2.12 Radiation Requirements

2.12.1 [VS-C-057] The CE camera shall support the worst case 2895-day mission radiation environment as defined in Table 8 with full functional integrity.

Table 8. CE Radiation Environment

2027 launch worst case scenario

• Earth-Mars cruise: 708 days

• Mars-Orbital Phase: 1519 days

• Mars-Earth cruise: 668 days

• Total: 2895 days

• Radiation Model(s):

o Trapped: AE9/AP9 o Solar proton model: ESP at 95% confidence level

2.12.2 [VS-C-058] The CE camera shall survive a total ionizing dose for the worst-case environment.

Note: As reference points, the dose behind 2.5 mm Al solid spherical shielding is 30.4 krad(Si); the worst-case dose behind a 25-mm composite aluminum honeycomb panel structure with solid aluminum equivalent density of 0.05 g/cm3 is 146 krad (Si) (0.47 mm equivalent spherical aluminum). Additional calculation(s) will be necessary for other shield geometries. Total ionizing dose as a function of shielding thickness is calculated for the 95% confidence level so no additional margin is required. Figure 4 gives the top-level total ionizing dose requirement for the CCRS environment exposures. Tabulated data for this figure are shown in Table 9.

Figure 4. Dose in Silicon as a function of spherical aluminum shielding for 2027 launch worst-case scenario.

Table 9. Dose in silicon as a function of spherical aluminum shielding for 2027 launch worst-case scenario

2.12.3 [VS-C-059] The CE camera shall survive a total non-ionizing dose (TNID) for the worst-case environment.

Note: As reference points, the TNID behind 2.5 mm Al solid spherical shielding is 6.31x108 MeV/g(Si); the worst-case dose behind a 25-mm composite aluminum honeycomb panel structure with solid aluminum equivalent density of 0.05 g/cm3 is 2.83x109 MeV/g(Si) (0.47 mm equivalent spherical aluminum). Additional calculation(s) will be necessary for other shield geometries. Total non-ionizing dose as a function of shielding thickness is calculated for the 95% confidence level so no additional margin is required. Table 10 gives the top-level total non-ionizing dose requirement for the CCRS environment exposures. Table 11 and Table 12 give the non-ionizing energy loss (NIEL)

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equivalent proton fluence for silicon at the 95% confidence level, for 10 MeV protons and 50 MeV protons. For example, the 10-MeV equivalent proton fluence necessary to deposit the TNID behind 2.5 mm solid Al shielding is 6.70x1010 cm-2; the 10 MeV equivalent fluence for the TNID behind 0.47 mm of solid aluminum shielding is 3.0x1011 cm-2. Figure 5 shows the 10-MeV equivalent proton fluence as a function of shielding (values from Table 11).

[VS-C-060] Table 13 shall be used to determine equivalent proton fluence for other proton energies as needed.

Figure 5. NIEL equivalent proton fluence in silicon at the center of solid aluminum spheres for 2027 launch worst-case scenario.

Table 10. Total mission non-ionizing dose in silicon as a function of spherical aluminum shielding for the 2027 launch worst-case scenario.

Table 11. Displacement damage equivalent 10-MeV proton fluence in silicon as a function of spherical aluminum shielding for 2027 launch worst-case scenario.

Table 12. Displacement damage equivalent 50-MeV proton fluence as a function of spherical aluminum shielding for 2027 launch worst-case scenario.

Table 13. Proton non-ionizing energy loss (NIEL) values used to calculate TNID in silicon as a function of spherical aluminum shielding

2.12.4 [VS-C-061] All CE camera electrical, electronic and electromechanical components shall be designed to avoid or tolerate errors and not suffer nor induce permanent loss of function due to non-destructive single event effects (SEE).

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2.12.5 [VS-C-062] CE camera electrical, electronic and electromechanical components with a non-destructive linear energy transfer (LET) threshold less than 20 MeV•cm2/mg shall be evaluated for both heavy ion and proton sensitivity.

Note: All single-event test data should be gathered such that the ion’s range is at least 100 micrometers – independent of LET.

2.12.6 [VS-C-063] CE camera electrical, electronic and electromechanical components with a non-destructive linear energy transfer (LET) threshold between 20 and 75 MeV•cm2/mg shall be evaluated for heavy ion sensitivity.

Note: All single-event test data should be gathered such that the ion’s range is at least 100 micrometers – independent of LET.

2.12.7 [VS-C-064] CE camera electrical, electronic and electromechanical components with a non-destructive single event effect linear energy transfer (LET) threshold of 75 MeV•cm2/mg shall require no further evaluation.

2.12.8 [VS-C-065] For destructive SEE, such as single-event latch-up (SEL), the minimum LET for all CE camera electrical, electronic and electromechanical components shall be 75 MeV-cm2/mg with a test ion range of at least 100 micrometers independent of LET.

Note: Deviations in the minimum LET and test ion range for components may be considered on a case-by-case basis by the Parts Control Board (PCB).

2.12.9 [VS-C-066] CE camera power MOSFETs, bipolar transistors, and Schottky diodes shall be de-rated into their SEE safe operating areas to prevent single-event gate rupture (SEGR) and single-event burnout (SEB).

2.12.10 [VS-C-067] All CE camera parts shall be reviewed by the contractor to ensure that the ion range is sufficient to result in a continually increasing LET through the sensitive layer(s) (typically the epilayer(s) of the device).

2.12.11 [VS-C-068] All Schottky diodes used in the CE camera shall be derated to 50% of rated breakdown voltage if no SEB test data exist.

2.12.12 [VS-C-069] The CE camera shall meet requirements after solar events that generate charged particles.

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2.12.13 [VS-C-070] Analysis of SEE rates shall be performed by the contractor using the proton and heavy-ion spectra given in the sub-sections below.

Note: Charged particle interference during solar events is of concern because it can impact the observation times of CCDs or CMOS detectors as well as other sensitive components.

Solar Proton Spectra

The differential flux for the Worst Week, Worst Day, and 5 Minute Peak during a solar event, based on the CREME96 October 1989 model, is given in Figure 6 and Table 14. The maximum proton transient throughout which the CCRS camera is expected to meet performance the 5 Minute Peak during a solar event. Values behind 1 g/cm2 (3.7 mm aluminum) shielding are included in the table.

The mission cumulative proton fluence spectra (integral and differential) are given in Table 15 and Table 16, with the integral spectra plotted in Figure 7.

Figure 6. Active period-proton spectra to use for proton induced SEE computations

Table 14. Proton differential spectra for proton induced SEE computations

Figure 7. Integral mission cumulative solar proton spectrum for total mission cruise duration for 2027 launch worst-case scenario.

Table 15. Integral cumulative solar proton fluence for total mission duration for the 2027 launch worst-case scenario

Table 16. Differential cumulative solar proton fluence for total mission duration for the

2027 launch worst-case scenario

Heavy-Ion Spectra

The CREME96 solar event flux (based on the October 1989 event) and the quiet galactic comic ray flux for the elements hydrogen through uranium were used to calculate worst-case LET spectra behind 1 g/cm2 (3.7 mm aluminum) shielding. Figure 8 provides the integral LET spectra for the CCRS environment for SEE rate analyses. Tabulated data for this figure are shown in Table 17.

Figure 8. CREME96 integral flux as a function of LET for four levels of activity for a component shielded by 1 g/cm2; peak 5-minute (worst case peak), worst day and worst week (worst case SPE) and quiet (GCR)

Table 17. CREME96 integral flux as a function of LET for four levels of activity for a component shielded by 1 g/cm2; peak 5-minutes (worst-case peak), worst day and worst week (worst case SPE) and quiet (GCR)

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2.13 Electrical Requirements

2.13.1 [VS-C-071] The CE camera shall operate with input voltages ranging from 4.5 V to 5.5

V.

2.13.2 [VS-C-072] The CE camera shall survive an overvoltage application of up to 6.875 V.

2.13.3 [VS-C-073] The CE camera shall survive a sudden removal of power indefinitely.

Note: The CE camera will be powered every time the spacecraft avionics LVPC is powered. Each camera head is powered from a separate shared switch.

2.13.4 [VS-C-074] The CE camera primary power isolation shall exceed 10 MΩ and 0.015 nF from the chassis.

2.13.5 [VS-C-075] The CE camera shall support a data interface connector supporting SpaceWire up to 100 Mbps.

Note: The standard 9-pin SpW interface may be bundled with manufacturer-specific inputs and outputs at the camera, including split connectors, flying leads and dedicated power input or sensor outputs. We will negotiate the precise details of this Camera/cable connection configuration as part of ICD activity.

2.13.6 [VS-C-076] Any low-voltage differential signaling (LVDS) interface between the camera and the electronics box shall operate at a signal clock rate of 100 MHz or less.

2.13.7 [VS-C-077] All low-voltage differential signaling (LVDS) and SpaceWire interfaces shall operate over #28 LVDS cables with an impedance of 100 ± 10 Ohm with a maximum length of 15 feet

2.13.8 [VS-C-078] The CE camera signal grounding shall be enabled through the SpW cable back to a single-point ground back at the spacecraft avionics.

Rationale: In order to prevent ground loops, grounding within the camera will reference a single-point ground at the Jettison Avionics box.

2.13.9 [VS-C-079] Camera grounding architecture shall conform to Figure 9 in which the camera chassis is isolated from the spacecraft structure.

2.13.10 [VS-C-080] A CE camera shall dissipate 4 W maximum while imaging.

2.13.11 [VS-C-081] The CE camera shall provide access to the camera FPGA JTAG pin-out for the purposes of re-programming the camera FPGA, if necessary, during pre-launch activities.

2.13.12 [VS-C-082] The CE camera shall have an equivalent capacitance across its power supply input lines less than 470 µF.

Figure 9. Proposed camera grounding scheme in which camera chassis is isolated from spacecraft structure to maintain a single-point ground near the Jettison Avionics assembly

2.14 EMI/EMC Requirements and Environment

2.14.1 [VS-C-083] The CE camera shall successfully complete EMI-EMC testing following

MIL-STD-461G and GSFC-STD-7000 tailored to the levels and limits shown in Table 18.

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Table 18. EMI-EMC test levels and limits

Inrush [VS-C-084] The Camera electronics inrush current shall not exceed dI/dt < 2 A/μs

Voltage Transients (excluding with On/Off operations)

[VS-C-085] Conducted voltage transients on the primary power bus, appearing during nominal mode switching (excluding ON / OFF) or appearing during periodic < 1 Hz or aperiodic activities shall be ≤ 0.1 Vpp when measured with at least 50 MHz bandwidth with the standard MIL-STD-461 version G LISN.

Radiated Emission (E field)

[VS-C-086] The unit shall not exceed the specified limits in the range 2 MHz - 18GHz. For non RF units which do not feature high frequency clocks, the upper frequency can be limited to 1 GHz or the 10th harmonic of the highest clock frequency (whatever is greater), as long as the 5th - 10th Harmonics are at least 10 dB below the limit.

The limits are given in the table below. Data for notches in the 390-404 MHz will be taken for information only.

Radiated Emission (H field)

[VS-C-087] The Radiated H field generated by Units, shall not exceed the limits specified in the table below when measurement are done at 7cm of the source.

Susceptibility (E Field)

[VS-C-088] 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 [VS-C-089] 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)

[VS-C-090] The unit shall not show any malfunction or deviation from the specified performance when irradiated with the E-fields as listed in the table below applies only to units powered during launch.

Susceptibility (H Field)

[VS-C-091] Units shall not show any malfunction or deviation from the specified performance when irradiated along any axis with an H-field of :

• 3 mT in the frequency range DC - 10 Hz

• 120 dBpT in the frequency range 50 Hz - 50 kHz.

Radiated Susceptibility (H Field)

[VS-C-092] In the case where a unit uses technologies which are known to be sensitive to AC or DC magnetic fields (e.g. USO's, motors, etc.), this sensitivity shall be reported together with an estimate of the maximum allowed magnetic field.

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2.14.2 [VS-C-093] The CE camera shall meet all of its requirements after exposure to launch vehicle and launch site radio frequency (RF) emissions.

Rationale: Launch vehicle and launch site radio frequency (RF) emissions are expected in the 200 MHz – 18 GHz, 2.200 GHz - 2.3 GHz (S-band), 5.4 GHz - 5.9 GHz (C-band range radar) and 1262.19 – 1345.04 MHz (launch site radar) radio bands. The camera needs to survive this environment.

Note: The peak electric field emission levels will be defined in an interface requirements document or interface control document in consultation with NASA after contract award.

Verification by analysis is an acceptable verification approach for this requirement.

2.15 Command and Telemetry Format

2.15.1 [VS-C-094] The CE camera shall output uncompressed images or images with lossless compression.

Note: Examples of image file formats: RAW, FITS, JPEG-2000

2.16 Command Interface

2.16.1 [VS-C-095] The CE camera shall transmit command and telemetry to and from the system via SpaceWire.

2.17 Imaging Commands

2.17.1 [VS-C-096] The CE camera shall accept imaging commands to control, at a minimum:

camera exposure time, imaging rate, sub-framing, pixel binning, image compression type and image compression ratio.

Note: The camera may also support other imaging parameters controllable by command.

2.18 Image Generation

2.18.1 [VS-C-097] The CE camera image transfer rate shall change via command to rates agreed upon with NASA.

2.18.2 [VS-C-098] At any single readout, the image size output by the CE camera shall be less than 32 Mega Bytes.

Note: Larger images can be downlinked in sections.

2.18.3 [VS-C-099] The CE camera shall generate uncompressed images when commanded.

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2.19 Lifetime Requirements

2.19.1 [VS-C-100] The CE camera shall operate within specification after four (4) years of dormancy in the operational environment.

2.19.2 [VS-C-101] The CE camera shall operate within specification with no maintenance or calibration required for up to 5 years of storage under the conditions specified in Table 19.

Table 19. Temperature and Humidity Storage

Facility Relative Humidity Temperature Storage, up to 5 years 30 to 60% 18 to 24 °C

[VS-C-102] The CE camera shall operate within specification with no maintenance or calibration during the component flight mission life, defined to be 4.5 years herein with a 10% operational duty cycle.

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3 Quality Assurance Provisions

3.1 Quality Assurance Requirements

3.1.1 [VS-C-103] The CE camera contractor shall design, develop, manufacture and deliver the

CE camera following the quality assurance requirements specified in the CCRS CE Camera Statement of Work (MSRCCRS-VS-SOW-0008).

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4 Preparation for Delivery

4.1 Delivery Requirements

4.1.1 [VS-C-104] The contractor shall design and prepare packaging and shipping arrangements to safely deliver the CE camera consistent with the requirements specified in the Statement of Work.

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Abbreviations and Acronyms

ATLO Assembly, Test and Launch Operations BRDF Bi-directional Reflectance Distribution Function C&DH Command & Data Handling CE Capture Enclosure CMN Common Mode Noise CCRS Capture, Containment and Return System CS Conducted Susceptibility DDD Displacement Damage Dose DPC Discrete Power Controller ECC Error Correction Code EDAC Error Detection and Correction EEE Electrical, Electronic and Electromechanical ELDRS Enhanced Low Dose Rate Sensitivity EM Electrical Model EMI/EMC Electromagnetic Interface/Electromagnetic compatibility ERO Earth Return Orbiter ETU Engineering Test Unit FOV Field of View FSW Flight Software GSFC Goddard Space Flight Center ICD Interface Control Document IDP Instrument Dependent protocol IP Internet Protocol ITAR International Trade in Arms Regulation LET Linear Energy Transfer MICD Mechanical Interface Control Drawing MPP Manufacturing Process Plans NASA National Aeronautical and Space Administration NIEL Non-Ionizing Energy Loss NSPAR Non-Standard Part Approval Request OS Orbiting Sample PF Protoflight PS Pyrotechnic Shock RDM Radiation design Margin RE Radiated Emissions RF Radio Frequency RS Radiated Susceptibility RV Random Vibration SBU Sensitive But Unclassified SEE Single Event Effects SEU Single Event Upsets SMRF Structural Mechanical Reference Frame SOW Statement of Work

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SP Science Protocol TBC To Be Confirmed TBD To Be Determined TBS To Be Scheduled TID Total Ionizing Dose TNID Total Non-Ionizing Dose TV Thermal Vacuum UDP User Datagram Protocol