PACE-GNC-SPEC-0067.pdf

Effective Date: November 19, 2018

Expiration Date: November 19, 2023

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

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

PACE-GNC-SPEC-0067, Revision -

Plankton, Aerosol, Cloud, ocean Ecosystem (PACE), Code 427

PACE Spacecraft Coarse Sun Sensor

Specification

Goddard Space Flight Center

Greenbelt, Maryland

National Aeronautics and

Space Administration

GSFC PACE CMO

11/19/2018

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

PACE CSS Specification PACE-GNC-SPEC-0067, Revision -i

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

PACE Spacecraft Component Standard Specification

Signature/Approval Page

Prepared By:

Eric Rogstad

Reviewed By:

Julian Ramirez-Brana

John Blackwood

Gary Davis

Beth Weinstein

Jack Sanders

Gary Won

Nikesha Davis

Tim Johnson

Toni Martinez

Dan McGuinness

Shavesha Rutledge

Raymond Ladbury

Jaime Eitnier

Craig Stevens

Dan Powers

Zach Boblitt

Approved By:

Andre’ Dress

Electronic Signatures available online at: https://ipdtdms.gsfc.nasa.gov/ ii

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

Preface

This document is under Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) Mission configuration control. Changes to this document require prior approval of the PACE

Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the PACE Configuration Management Office (CMO), along with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

iii

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Change History Log

Revision Effective Date Description of Changes

(Reference the SCoRe & Approval Date)

Revision - 11/19/2018 Baseline Release following the approval of PACE-CCR-0412 iv

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Table of TBDs/TBRs/TBSs

Action Item

No.

Location Summary Individual/

Organization

Actionee v

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Table of Contents

1.0 Introduction

1.1 General Information

1.2 Scope

2.0 Applicable Documents

3.0 Contract Description

3.1 Coarse sun sensor Description

3.2 Ground Support Equipment Description

4.0 Functional/Performance Requirements

4.1 CSS Flight Unit Functional/Performance Requirements

4.1.1 Performance

4.1.1.1 Current Output

4.1.1.2 Maximum Current

4.1.1.3 Calibrated Accuracy

4.1.1.4 Output Variability

4.2 Resource Allocations

4.2.1 NA

4.2.2 Nominal Power Allocation

4.3 NA

4.4 Electrical Grounding

4.4.1 NA

4.4.2 NA

4.4.3 NA

4.4.4 Mechanical Contact Resistance

4.4.5 Mating Method

4.4.6 NA

4.4.7 NA

4.4.8 Connector DC Resistance

5.0 Physical Requirements

5.1 Interface Documentation

5.1.1 Mechanical Interface

5.1.2 Electrical Interface

5.2 Mass Properties

5.2.1 Component Masses

5.2.2 Center of Mass Location

5.3 Physical Envelope

5.4 Mounting

5.5 Field of View

5.6 Alignment

5.6.1 Alignment Location/Orientation Accuracy

5.6.2 Alignment Knowledge Accuracy

5.6.3 Alignment Stability

5.6.4 Alignment Method

6.0 Environmental Requirements

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6.1 Mechanical Factors of Safety

6.2 Quasi-Static Acceleration

6.3 Frequency Requirement

6.3.1 Fundamental Launch Frequencies

6.4 Vibration

6.4.1 Sinusoidal Vibration

6.4.2 Random Vibration

6.5 Shock

6.6 NA

6.7 Transportation

6.7.1 Transportation Cleanliness

6.8 Pressure

6.8.1 Operating Pressure Range

6.8.2 Maximum Depressurization Rate

6.8.3 Launch Vehicle Environmental Control System Impingement Velocity

6.9 On-Orbit Dynamic Environment

6.9.1.1 Dynamic Linear Acceleration

6.9.1.2 Dynamic Angular Acceleration

6.10 Ground Environments

6.11 Thermal Requirements

6.11.1 Flight Interface Design Temperature Limits

6.12 Charged Particle Radiation Requirements

6.12.1 Definitions

6.12.2 Total Ionizing Dose

6.12.2.1 Minimum TID Tolerance for EEE Parts and Materials

6.12.2.2 NA

6.12.3 Displacement Damage Dose

6.12.4 NA

6.12.5 Charging Environment

6.13 Atomic Oxygen Fluence

6.13.1 Atomic Oxygen Analysis

6.13.2 Atomic Oxygen Testing

7.0 Cleanliness

7.1 Surface Contamination

7.1.1 Surface Contamination Levels at Delivery

7.1.1.1 Particulate Contamination

7.1.1.2 Molecular Contamination – Exposed Surfaces

7.1.1.3 Molecular Contamination – Covered Surfaces

7.1.2 Surface Contamination Generation

7.1.2.1 Particulate Generation

7.1.2.2 Molecular Generation

7.1.2.3 Intentional and Unintentional Vents

7.1.2.4 Intentional and Unintentional Vents – Filters

7.1.2.5 Intentional Vents – Pressure Buildup Prevention

7.1.3 Cleanability

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7.1.3.1 Cleanability – Sensitive Surfaces

7.1.3.2 Cleanability – Sensitive Surface Cleaning Methods

7.2 Electrostatic Cleanliness

7.2.1 Conductive Surface Ground Path

7.2.2 Conductive Surface Resistivity

7.2.3 Closeout of Gaps and Apertures

7.2.3.1 Conductive Tape Surface Resistivity

7.2.3.2 Conductive Tape, Grounding

7.2.4 Exposed Harness Specific Requirements

8.0 Design & Construction Requirements

8.1 Parts, Materials & Processes (PMP)

8.1.1 EEE Parts

8.1.2 Materials

8.1.2.1 Material Conductivity

8.1.2.2 NA

8.2 Electrical

8.2.1 Test Sensors

8.2.2 Interface Requirements

8.2.2.1 Connector Selection

8.2.2.2 Signal Segregation

8.2.2.3 Test and Flight Signal Isolation

8.2.2.4 Test Interfaces

8.2.3 Mitigation of Internal Charging

8.2.3.1 Mitigation Strategies for Internal Charging

8.2.3.2 Floating Conductors

8.2.3.3 Dielectric Structures

8.3 Safety

8.4 NA

8.5 Identification and Marking

8.6 Workmanship

8.6.1 Workmanship Standards

8.6.2 Connector

8.6.2.1 GSE Cable Connectors

8.6.2.2 Prevention of Connector Mismating

8.6.2.3 Test Connectors

8.6.2.4 Connector Identification

8.6.2.5 Protection of Unused Test Connectors

8.6.2.6 Connector Savers

8.7 Reliability and Mission Lifetime

8.7.1 Mission Life

8.7.2 Operating Time

8.7.2.1 Total Time

8.7.2.2 Failure Free Time

8.8 Ground Handling

8.8.1 Ground Support Equipment (GSE) Design

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8.8.2 NA

8.8.3 NA

8.8.4 NA

8.8.5 NA

8.8.6 NA

8.8.7 GSE Cleanliness

8.8.8 GSE Bakeout

8.8.9 Test Harness

9.0 Mechanical Design Requirements

9.1 Structural Requirements

9.1.1 Component Fatigue

9.1.2 Fracture Control Requirements

9.2 Fastening Systems

9.2.1 Fastener Performance Analysis

9.2.1.1 Factors of Safety

9.2.1.2 Supplemental Factor

9.2.1.3 Ultimate Design Loads

9.2.1.4 Yield Design Loads

9.2.1.5 Design Separation Load

9.2.2 Fastener Locking and Retention

9.2.2.1 Locking Features

9.2.2.2 Verification

9.2.2.3 Locking Features

9.2.2.4 Installation Torque Specification and Control

9.2.3 Fastened Joints Criteria

9.2.3.1 Minimum and Maximum Preload

9.2.3.2 Analysis Addressing Potential Rupture

9.2.3.3 Ultimate Strength Analysis

9.2.3.4 Applied Shear Loading

9.2.3.5 Shear Loading

9.2.3.6 Simultaneous Applied Tensile and Shear Loads

9.2.3.7 Allowable Yield Tensile Load

9.2.3.8 Separation Analysis

9.2.3.9 Seal Analysis

10.0 Logistics

10.1 NA

10.2 Ground Support Equipment

10.3 Transportation Equipment

11.0 Verification Requirements

11.1 Verification Methods

11.1.1 Inspection

11.1.2 Analysis

11.1.3 Test

11.2 Inspection Requirements

11.2.1 Visual Inspection

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11.2.2 Physical Measurement

11.2.3 Documentation Search

11.3 Analysis Requirements

11.4 Test Requirements

11.4.1 Definitions

11.4.2 Test Factors

11.4.3 Test Tolerances

11.4.4 Test Restrictions

11.4.4.1 Failure During Tests

11.4.4.2 Modification of Hardware

11.4.4.3 External Adjustment

11.4.4.4 Re-Test Requirements

11.5 Required Tests

11.5.1 Performance Tests

11.5.1.1 Comprehensive Performance Test

11.5.1.2 Limited Performance Test

11.5.1.3 Abbreviated Functional Test

11.5.2 Mass Properties Measurement

11.5.3 Static Loads/Strength Test

11.5.3.1 Sine Burst

11.5.3.2 Static Pull

11.5.4 Sine Sweep Survey

11.5.5 Sine Vibration

11.5.6 Random Vibration

11.5.7 NA

11.5.8 NA

11.5.9 Shock

11.5.10 Thermal Vacuum Bake-out

11.5.11 Thermal Vacuum Test

11.5.11.1 Thermal Vacuum Test Parameters

11.5.11.2 Thermal Vacuum Test Profile

Appendix A Abbreviations and Acronyms x

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List of Figures

Figure Page

Figure 6-1. Sine Vibration Environment Figure 6-2. NA Figure 6-3. Shock Envelope

Figure 6-4. Total Ionizing Dose-Depth Curve (includes x2 margin) Figure 6-5. NA Figure 6-6. NA Figure 8-6. NA Figure 11-1. Thermal Vacuum Profile

List of Tables

Table Page

Table 2-1. Applicable Documents Table 6-1. Factors of Safety Table 6-2. CSS Design Limit Loads

Table 6-3. CSS Sine Vibration Environment Table 6-4. CSS Random Vibration Environment, (22.7 kg, or less)

Table 6-5. NA

Table 6-6. Qualification Level Shock Response Spectrum

Table 6-7. NA Table 6-8. NA

Table 6-9. Transportation Loads Table 6-10. Temperature Limits at Box Mounting Interface Table 6-11. Dose (including x2 margin) as a Function of Shielding

Table 6-12. NA Table 6-13. NA Table 8-1. NA Table 8-2. NA

Table 8-3. NA Table 8-4. NA Table 11-1. Test Factors and Durations

Table 11-2. Test Tolerances Table 11-3. Thermal Vacuum Test Parameters

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

1.0 INTRODUCTION

1.1 GENERAL INFORMATION

The Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) mission is a strategic climate continuity mission that will extend the high quality ocean ecological, ocean biogeochemical, cloud, and aerosol particle data records begun by NASA in the 1990s. The mission will be capable of collecting radiometric and polarimetric measurements of the ocean and atmosphere, from which these biological, biogeochemical, and physical properties will be determined. PACE data products will not only add to existing critical climate and Earth system records, but also answer new and emerging advanced science questions related to Earth’s changing climate.

1.2 SCOPE

This specification describes the electrical, mechanical, environmental, and verification testing requirements for a space-qualified Coarse Sun Sensor for the NASA Goddard Space Flight

Center (GSFC) PACE Mission.

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

2.0 APPLICABLE DOCUMENTS

The following documents and drawings in effect on the day this specification was signed shall apply to the fabrication and to the electrical, mechanical, and environmental requirements of the

Coarse Sun Sensor to the extent specified herein. In the event of conflict between this specification and any referenced document, this specification will govern, with the exception of the PACE Coarse Sun Sensor Statement of Work (PACE-GNC-SOW-0037), in which case the

Statement of Work takes precedence.

There are many applicable documents (standards, procedures, etc.) throughout this specification document. The expectation is suppliers will be compliant with the requirements within those documents. However, we recognize often comparable documents may be used in lieu of those listed herein. The alternative documents will be evaluated for compliance and approval is required prior to usage.

The following is a list of the applicable specifications and publications.

Table 2-1. Applicable Documents

Document Number Title

PACE-GNC-SOW-0037 PACE CSS Statement of Work

NFPA 70 National Fire Protection Association National

Electric Code NASA-STD-5001B Structural Design And Test Factors Of Safety For

Spaceflight Hardware

NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety

Requirements

NASA-STD-6016 Standard Materials and Processes Requirements for

Spacecraft

NASA-HDBK-7005 Dynamic Environment Criteria

NASA-STD-7001 Payload Vibroacoustic Test Criteria

IEST-STD-CC-1246E Product Cleanliness Levels And Contamination

Control Program

ASTM E-595-07 Standard Test Method for Total Mass Loss and

Collected Volatile Condensable Materials from

Outgassing in a Vacuum Environment

MIL-DTL-5541 Chemical Conversion Coatings on Aluminum and

Aluminum Alloys

MIL-A-8625F Anodic Coatings for Aluminum and Aluminum

Alloys

EEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, and Derating

MIL-STD-461F Military Standard, Electromagnetic Emission And

Susceptibility Requirements For The Control Of

Electromagnetic Interference (EMI)

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

Document Number Title

GSFC-STD-7000A General Environmental Verification Standard

(GEVS)

NASA-STD-5019A Fracture Control Requirements for Spaceflight

Hardware

NASA-STD-5020 Requirements for Threaded Fastening Systems in

Spaceflight Hardware

NASA-STD-5017A Design and Development Requirements for

Mechanisms

FAA AC 20-71 Federal Aviation Administration Advisory Circular

(AC) 20-71, “Dual Locking Devices on Fasteners".

NASM 33540 Safety Wiring, Safety Cabling, Cotter Pinning, General Practices for

SAE AS567 Safety Cable, Safety Wire, Key Washers, and Cotter

Pins for Propulsion Systems, General Practices for

Use of

541-WI-5330.1.41 Fastener Locking Using Arathane 5753

MSFC-STD-3029A Guidelines for the Selection of Metallic Materials for

Stress Corrosion Cracking Resistance in Sodium

Chloride Environments

MIL-STD-462, Notice 2 Electromagnetic Interference Characteristics, Measurement of, 1 May 1970

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

3.0 CONTRACT DESCRIPTION

3.1 COARSE SUN SENSOR DESCRIPTION

Twelve Coarse Sun Sensors (CSS) will be used by the PACE Guidance, Navigation and Control

(GN&C) subsystem. These Coarse Sun Sensors will be mounted upon the exterior of the PACE observatory.

The Coarse Sun Sensors will be used to provide sun position information as a key part of the

PACE attitude control system.

3.2 GROUND SUPPORT EQUIPMENT DESCRIPTION

Ground support equipment (GSE) will enable aliveness testing of the CSS by illuminating the

CSS to generate a current output from the CSS during spacecraft Integration and Testing. The

GSE will include twelve CSS stimulators and a control panel, which shall meet the requirements as defined in Section 10.2.

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

4.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the Coarse Sun Sensor as defined in Section 3.1.

4.1 CSS FLIGHT UNIT FUNCTIONAL/PERFORMANCE REQUIREMENTS

4.1.1 Performance

4.1.1.1 Current Output

The output of each CSS shall be a current that is approximately proportional to the cosine of the angle of the sun with respect to the CSS boresight axis.

4.1.1.2 Maximum Current

Each CSS optic shall not exceed a peak output of 2.0 mA corresponding to full solar illumination on the boresight axis.

4.1.1.3 Calibrated Accuracy

The maximum error between the estimated angle (calibration applied to the sensor output) and the actual incident angle for each CSS shall be less than or equal to 5 degrees over the entire field of view.

4.1.1.4 Output Variability

Each CSS shall have a Beginning of Life (BOL) nominal output within ±5 % of the average of all 12 CSS corresponding to full solar illumination on the boresight axis.

4.2 RESOURCE ALLOCATIONS

4.2.1 NA

4.2.2 Nominal Power Allocation

a. The CSS shall be a current source.

b. The CSS shall not require or consume any power.

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4.3 NA

4.4 ELECTRICAL GROUNDING

4.4.1 NA

4.4.2 NA

4.4.3 NA

4.4.4 Mechanical Contact Resistance

The DC resistance of the mechanical contact between two conductive mating surfaces (internal to the component) shall be less than or equal to 2.5 mΩ DC resistance.

4.4.5 Mating Method

The primary mating method for a PACE component is metal-to-metal contact between component mounting feet (or base plate) and the spacecraft structure. When the use of this method is not possible as determined by the Mechanical Interface Control Document (MICD), the use of a ground strap is necessary.

4.4.6 NA

4.4.7 NA

4.4.8 Connector DC Resistance

Component connectors shall be electrically connected to chassis with a DC resistance less than or equal to 10 mΩ.

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

5.0 PHYSICAL REQUIREMENTS

5.1 INTERFACE DOCUMENTATION

The contractor shall use metric units when interfacing with NASA GSFC including any drawings, documents, models, except for the following cases:

Heritage Component: Components that has been previously qualified, or of similar design heritage, may be specified in English units where use of metric equivalents would lead to additional cost to the program.

Fasteners: Although bolt patterns will be defined using metric dimensioning, use of

English fasteners (with hole dimensioning and tolerancing) is permitted.

Angular Measurement: Angular measurement may be expressed in degree of arc or in an appropriate subdivision of degree of arc such as second of arc (arc-sec) when advantageous to application.

5.1.1 Mechanical Interface

The mounting interface shall be defined in the Mechanical Interface Control Drawing (MICD), which will be developed between the contractor and NASA GSFC.

5.1.2 Electrical Interface

The electrical interface shall be defined in the Electrical Interface Control Document (EICD), which will be developed between the contractor and NASA GSFC.

5.2 MASS PROPERTIES

5.2.1 Component Masses

a. Each Coarse Sun Sensor, including baffle, shall have a mass less than or equal to 18 grams (g).

b. The mass of the CSS shall be measured to within 2 g.

5.2.2 Center of Mass Location

The contractor shall define the center of mass in the Mechanical ICD.

5.3 PHYSICAL ENVELOPE

The CSS outside dimensions, including baffle, mounting flanges, and connectors, shall not exceed the volume envelope of 3.5 cm length x 3.5 cm width x 2.0 cm height as defined in the

Mechanical ICD.

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

5.4 MOUNTING

a. Mounting interface flatness, and co-planarity requirements for the component side of the interface (including brackets, if any, and shims) shall be as defined in the MICD.

b. The CSS shall have a flatness of 0.125 mm from any 2 points on the mounting interface for a component <0.3 m2.

5.5 FIELD OF VIEW

a. The CSS glint-free Field of View shall be 74.75 degrees from the sensor boresight axis.

b. The Contractor shall provide baffles, as required, to achieve the glint-free Field of View requirement.

c. The CSS Field of View shall be described in the MICD.

5.6 ALIGNMENT

5.6.1 Alignment Location/Orientation Accuracy

The boresight axis of the CSS shall be perpendicular to the CSS mounting surface within 0.2 degrees.

5.6.2 Alignment Knowledge Accuracy

Knowledge of the CSS boresight axis alignment to the mechanical reference shall be 0.1 degrees or better.

5.6.3 Alignment Stability

Throughout environments, the CSS boresight axis shall not shift with respect to the mechanical reference by more than 0.1 degrees.

5.6.4 Alignment Method

The CSS units will be hard-mounted on a mechanical surface on both the solar array and the spacecraft structure.

a. Proper fit/alignment of the CSS to the structure shall be inherent in its design, fabrication, and assembly to the structure, through the use of close dimensional control in the location of mounting holes and the use of correct mounting hardware.

b. The contractor shall provide a drill template as required.

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6.0 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements for the spacecraft components are specified in this section.

The CSS shall meet its performance requirements in section 4.0 during and after exposure to the environments specified in this section.

6.1 MECHANICAL FACTORS OF SAFETY

a. The CSS, as well as Mechanical Ground Support Equipment (MGSE), shall demonstrate positive Margins of Safety under limit loads for all yield and ultimate failures using the

Factors of Safety (FS) defined in Table 6-1. Margin of Safety (MS) is defined as follows:

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

Table 6-1. Factors of Safety

b. Primary and secondary structure comprised of composite materials, Beryllium, bonded joints and/or bonded inserts shall be proof tested to 1.25 x Limit Load; qualification by analysis only is not acceptable. Actual flight component testing is preferred, but testing of representative sets of hardware with a similarity qualification argument can be used if approved by the NASA/GSFC COR.

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6.2 QUASI-STATIC ACCELERATION

Quasi-static acceleration represents the combination of steady-state accelerations and the low frequency mechanically transmitted dynamic accelerations that occur during launch.

The CSS shall be designed to withstand the quasi-static design limit loads defined in the mass-acceleration curve (MAC) shown in Table 6-2 without damage or degradation of performance.

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 CSS 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.

Table 6-2. CSS Design Limit Loads

6.3 FREQUENCY REQUIREMENT

6.3.1 Fundamental Launch Frequencies

The CSS shall have a fundamental frequency greater than 100 Hz when hard mounted at its spacecraft interface. Any component, which fails to meet the specified fundamental frequency, must supply a finite element model, correlated to modal survey test results up to 50 Hz, to be

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

used in coupled loads analyses. Requirements for the submitted finite element model are in the

SOW and discussed in Section 11.5.4.

6.4 VIBRATION

6.4.1 Sinusoidal Vibration

The CSS 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 6-3. Stiff components showing a first mode greater than 150Hz can be exempted from sine vibration testing upon approval by the NASA/GSFC COR. A generic sine vibration specification is provided for Protoflight (PFT), Qualification (QT) and Acceptance (AT) levels applied at the

PACE to CSS interface. See Section 11.4.1 for definitions of Protoflight, Qualification, and

Acceptance.

Table 6-3. CSS Sine Vibration Environment

Frequency Protoflight, Qual Level Flight, Acceptance Level

5-20 Hz 0.63 in (double amplitude) 0.50 in (double amplitude)

20-100 Hz 12.5 G 10.0 G

Figure 6-1. Sine Vibration Environment

Levels may be notched to not exceed 1.25 times the design limit load outlined in section 6.2.

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.

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6.4.2 Random Vibration

a. The CSS shall demonstrate its ability to meet its performance requirements after being subjected to the random vibration environment in Table 6-4, for components weighing

22.7 kg (50 lb) or less, applied at the Spacecraft to CSS interface.

Table 6-4. CSS Random Vibration Environment, (22.7 kg, or less)

Frequency ASD Level (g2/Hz)

(Hz) Qualification Acceptance

20-50

50-800

800-2000

0.026

+6 dB/oct

0.16

-6 dB/oct

0.026

0.013

+6 dB/oct

0.08

-6 dB/oct

0.013

Overall 14.1 Grms 10.0 Grms

Table 6-5. NA

Figure 6-2. NA

This environment will be updated with random vibration analysis. Note for CSS, the highest design loads may be from this random vibration environment.

b. The contractor shall provide random vibration analysis along with static loads analysis.

Please see NASA-HDBK-7005 and NASA-STD-7001 for more information.

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

c. Notching shall be limited to -12 dB of the original input and to a bandwidth of less than

100 Hz to limit the random vibration responses to 3dB above design limit levels.

Notching beyond these limits will require NASA/GSFC COR approval.

6.5 SHOCK

a. The CSS shall be designed to meet its performance requirements after being subjected to the shock environment in Table 6-6, applied at the CSS interface to the PACE spacecraft structure.

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Table 6-6. Qualification Level Shock Response Spectrum

Freq (Hz) SRS (G)

100 100

800 1700

8000 1700

10000 2200

b. A shock susceptibility and attenuation assessment shall be performed on all Spacecraft components. If the flight shock environment as shown on a Shock Response Spectra

(SRS) plot (Q=10) is enveloped by the curve shown below (Figure 6-3), then the shock environment can be considered benign and there is low risk in deferring the shock test to the Observatory level.

Figure 6-3. Shock Envelope

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

d. Any component determined to be susceptible to the shock environment (e.g., shock levels are above the curve) shall have shock testing performed at the component level

(preferably on a qualification unit).

e. Component self-induced shock testing shall be accomplished by two actuations at the component level for each self-induced shock source (in order to account for the scatter associated with the actuation of the device) for the first flight unit, and a single actuation on subsequent units.

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6.6 NA

Table 6-7. NA

Table 6-8. NA

6.7 TRANSPORTATION

In addition to the launch loads shown above, the CSS shall also be designed to withstand the maximum transportation loads shown in Table 6-9 without damage or degradation of performance.

Table 6-9. Transportation Loads

6.7.1 Transportation Cleanliness

Materials and enclosure used for transportation and storage shall not generate molecular or particle contaminants or degrade the surface cleanliness of the item or adjacent items.

6.8 PRESSURE

6.8.1 Operating Pressure Range

The CSS shall be designed to 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).

6.8.2 Maximum Depressurization Rate

The CSS shall be designed to meet all performance requirements after exposure to a maximum depressurization rate of -50 mbar/sec (-0.72 psi/sec) experienced during launch and ascent.

6.8.3 Launch Vehicle Environmental Control System Impingement Velocity

The CSS exterior surfaces shall not suffer damage or degradation when exposed to the Launch

Vehicle Environmental Control System airflow velocity of 10 m/sec.

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6.9 ON-ORBIT DYNAMIC ENVIRONMENT

The CSS shall be designed to handle all linear and angular acceleration requirements at the same time, i.e. the linear and angular acceleration are not mutually exclusive.

6.9.1.1 Dynamic Linear Acceleration

The CSS shall survive on-orbit when subjected to the maximum linear acceleration due to nominal thruster firing of 0.060 m/sec2.

6.9.1.2 Dynamic Angular Acceleration

The CSS shall survive the maximum angular acceleration of 2.5E-5 rad/sec2.

6.10 GROUND ENVIRONMENTS

a. The CSS shall meet all of their performance requirements during exposure to air temperature between +5 and +30 degrees C and relative humidity between 30% and 70%.

b. After being powered “OFF” and exposed to air temperatures of +5 to +30 degrees C and relative humidity of 0 to 70%, the CSS shall meet all of its performance requirements.

6.11 THERMAL REQUIREMENTS

6.11.1 Flight Interface Design Temperature Limits

a. When powered “OFF”, the CSS shall be capable of surviving indefinitely when its temperatures are within the survival limits shown in Table 6-10 without damage or permanent performance degradation.

b. The CSS shall meet all performance requirements when powered “ON” anywhere within the Operational and Protoflight/Qualification limits shown in Table 6-10.

c. The CSS shall demonstrate turn on at the Minimum and Maximum

Protoflight/Qualification (in spec) limits shown in Table 6-10.

Table 6-10. Temperature Limits at Box Mounting Interface

Minimum Temperature (ºC) Maximum Temperature (ºC) Operational (In Spec) -90 +80

Protoflight/Qualification (In Spec) -100 +90

Survival (Unpowered) -135 +90

6.12 CHARGED PARTICLE RADIATION REQUIREMENTS

Components containing electronic parts will be exposed to a natural space radiation environment that consists of: (1) trapped particles which include electrons, protons, and heavier ions; (2)

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

particles from solar events (coronal mass ejections and flares); and (3) galactic cosmic ray particles.

6.12.1 Definitions

Total Ionizing Dose (TID) - the mean energy deposited by ionizing radiation in a device region divided by the mass of the region. This is often given in units of rad(Si), where 1 rad(Si) = 100 erg deposited per gram of silicon.

Radiation Design Margin (RDM) - the ratio of the derated hardness capability of the part to the estimated dose at the part location.

Enhanced Low Dose Rate Sensitivity (ELDRS) - used to refer to a part that shows enhanced radiation-induced damage at dose rates below 50 rad(Si)/s. The enhancement is the result of true dose rate effects.

Displacement Damage Dose (DDD) - the mean energy deposited by ionizing radiation in a device region that goes into atomic displacements divided by the mass of the region. There is no official unit for DDD. One such unit is MeV/g.

Non-Ionizing Energy Loss (NIEL) - a measure of the rate of energy loss due to atomic displacements as a particle traverses a material.

Single Event Effect (SEE) - any measurable effect to a circuit due to an ion strike. This includes, but is not limited to, single event upsets (SEUs), single event transients (SETs), single hard errors (SHEs), single event latchups (SELs), single event functional interrupts (SEFIs), single event burnouts (SEBs), single event gate ruptures (SEGRs), and single event dielectric ruptures

(SEDRs).

Single Event Upset (SEU) - a change of state or transient induced by an energetic particle such as a cosmic ray or proton in a device. This may occur in digital or analog, circuits and may have effects in surrounding interface circuitry (a subset known as SETs). These are “soft” errors in that a reset or rewriting of the device will usually return the device to normal behavior thereafter.

The general goal for non-destructive events such as SEUs or SETs is not to avoid them completely, but to manage their impact through robust circuit design, automatic correction, and/or operational activities based on knowledge from ground radiation tests and circuit/system analysis.

Single Hard Error (SHE) - a SEU that causes a permanent change to the operation of a device.

An example is a stuck bit in a memory device.

Multiple Bit Upset (MBU) - an event induced by a single energetic particle such as a cosmic ray or proton that causes multiple upsets or transients during its path through a device or system in a single logical structure (ex., 2 bits affected in a single 16-bit word).

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Single Event Functional Interrupt (SEFI) - a condition that causes loss of device functionality due to a single event in a control portion of a device. It generally requires a device reset or a re-initialization to resume normal device operations, but, for some devices, a power cycle is necessary to resume normal device operations. The general goal for non-destructive events such as SEFI is to avoid them, however, managing their impact through robust circuit design, automatic correction, and/or operational activities may be considered.

Single Event Latchup (SEL) - a condition that may cause device failure due to a single event induced high current state. A SEL may or may not cause permanent device damage, but requires power cycling of the device to resume normal device operations. In addition, susceptible devices have the concern for latent damage (device does not fail from the immediate single particle event, but reliability is degraded and premature failure may occur).

Single Event Burnout (SEB) - a condition that can cause device destruction due to a high current state in a power transistor.

Single Event Gate Rupture (SEGR) - a single ion induced condition in power MOSFETs that may result in the formation of a conducting path in the gate oxide.

Linear Energy Transfer (LET) - a measure of the energy deposited per unit length as an energetic particle travels through a material. The common LET unit is MeV*cm2/milligram (mg) of material.

Threshold LET (LETth) - the maximum LET at which no SEE is observed at a particle fluence of

107 ions/cm2.

6.12.2 Total Ionizing Dose

6.12.2.1 Minimum TID Tolerance for EEE Parts and Materials

The top-level total ionizing dose requirement is shown in Figure 6-4 Error! Reference source not found.and Table 6-11. The dose values are calculated as a function of aluminum shield thickness in units of krad in silicon. For a nominal 2.54 mm (100 mils) of equivalent aluminum shielding and a 3-year mission life, the expected dose is 6.67 krad-Si. This includes a factor of 2 margin. EEE parts and materials shall be selected according to the level of shielding shown in

Figure 6-4 Error! Reference source not found.and Table 6-11.

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Figure 6-4. Total Ionizing Dose-Depth Curve (includes x2 margin)

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Table 6-11. Dose (including x2 margin) as a Function of Shielding

6.12.2.2 NA

6.12.3 Displacement Damage Dose

The displacement damage dose or non-ionizing dose is the mean energy deposited in a material that goes into atomic displacements divided by the mass of the material. It is analogous to the ionizing dose except that the energy considered produces displacements in a semiconductor lattice.

Parts sensitive to displacement damage degradation (e.g., optical detectors, precision bipolar integrated circuits, crystal oscillators, and power devices) shall be able to tolerate a minimum

DDD of 1.67 x 108 MeV/g (or a 10 MeV equivalent proton fluence of 2.12×1010 cm-2) for silicon devices or 1.53 x 108 MeV/g (or a 10 MeV equivalent proton fluence of 2.32×1010 cm-2) for gallium arsenide devices, respectively.

Dose

(mm): (mils): (g/cm2): (krad-Si):

0.002935 0.115532 0.000792 7.98E+03

0.004039 0.159003 0.00109 5.56E+03

0.005521 0.217353 0.00149 3.96E+03

0.007596 0.299042 0.00205 2.94E+03

0.010412 0.409907 0.00281 2.28E+03

0.014302 0.563075 0.00386 1.74E+03

0.019601 0.771675 0.00529 1.34E+03

0.0269 1.059047 0.00726 1.05E+03

0.036941 1.454366 0.00997 8.06E+02

0.050761 1.998477 0.0137 6.12E+02

0.069658 2.742436 0.0188 4.54E+02

0.095595 3.763556 0.0258 3.38E+02

0.131165 5.163948 0.0354 2.58E+02

0.180073 7.089488 0.0486 1.91E+02

0.246767 9.715225 0.0666 1.38E+02

0.339027 13.34749 0.0915 9.40E+01

0.466857 18.38015 0.126 6.32E+01

0.637297 25.09037 0.172 4.16E+01

0.87443 34.42632 0.236 2.66E+01

1.204194 47.40913 0.325 1.70E+01

1.64882 64.91404 0.445 1.10E+01

2.263885 89.12916 0.611 7.28E+00

3.108674 122.3885 0.839 5.40E+00

4.260995 167.7554 1.15 4.14E+00

5.854237 230.4813 1.58 3.22E+00

8.040313 316.5471 2.17 2.54E+00

11.04154 434.7053 2.98 2.06E+00

15.15432 596.6257 4.09 1.65E+00

20.78625 818.3545 5.61 1.32E+00

Aluminum Shield Thickness

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This meets the DDD exposure for a 3-year mission with nominal 2.54 mm (100 mils) of aluminum shielding and includes a 2x margin.

6.12.4 NA

Figure 6-5. NA

Table 6-12. NA

Figure 6-6. NA

Table 6-13. NA

6.12.5 Charging Environment

The CSS shall be designed to withstand the degradation of surface materials and associated surface charging effects due to the radiation environment for the PACE mission orbit.

6.13 ATOMIC OXYGEN FLUENCE

When a component has surfaces that will be exposed to the external space environment, all external materials shall survive an atomic oxygen fluence (Observatory orbit velocity direction) of 2.0E+20 atoms/cm2 without loss of structural integrity or loss of critical performance criteria.

6.13.1 Atomic Oxygen Analysis

An analysis shall be performed for all external materials and finishes to verify the compatibility with the AO environment. The analysis should show that any AO degradation does not pose a contamination hazard for other components on the spacecraft (source of particles, molecular films, debris, etc.).

6.13.2 Atomic Oxygen Testing

If no data exists for the proposed material and finishes, a test shall be performed, exposing a representative sample to AO and verifying no loss of structural integrity or loss of critical performance criteria.

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7.0 CLEANLINESS

The requirements in this section ensure the cleanliness of the CSS at delivery, so as not to adversely affect its own performance, as well as not be a source of contamination to other items, including not generating contaminants following delivery in excess of that permitted below by virtue of its design, materials of construction, or operation.

7.1 SURFACE CONTAMINATION

7.1.1 Surface Contamination Levels at Delivery

7.1.1.1 Particulate Contamination

The CSS shall meet IEST-STD-CC1246E VC-0.5-1000 + UV, or equivalent, with NASA/GSFC

COR approval, when inspected with both UV and white light in a darkened room.

7.1.1.2 Molecular Contamination – Exposed Surfaces

If the CSS has surfaces exposed to the space environment, all such surfaces shall meet a molecular surface cleanliness level of IEST-STD-CC1246E R3.3E-1 on all external and critical surfaces and IEST-STD-CC1246E VC-0.5-1000 + UV, or equivalent, with NASA/GSFC COR approval.

7.1.1.3 Molecular Contamination – Covered Surfaces

a. All CSS surfaces not exposed to the space environment shall meet IEST-STD-CC1246E

VC-0.5-1000 + UV, or equivalent, with NASA/GSFC COR approval when inspected in a darkened room with white and UV light.

b. The surfaces shall be free of molecular contamination – for example films, spots, or other.

7.1.2 Surface Contamination Generation

7.1.2.1 Particulate Generation

The CSS contractor shall not employ any of the following particle generating materials or processes into the CSS design or construction without prior approval by the NASA/GSFC COR:

Paints prone to shedding due to large paint pigment molecules, overspray, poor adhesion, etc.

Dry lubricants (e.g. molybdenum disulfide).

Surfaces prone to corrosion or oxides because of a lack of corrosion protection or dissimilar metals in close contact.

Fabrics with brittle constituents (e.g., composites, graphite or glass).

Perforated materials when material is highly susceptible to tear propagation (e.g., multi-layer insulation (MLI)).

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Metal oxides (bare [untreated] aluminum and magnesium, iron, non-corrosion resistant steel, etc.).

Braided metallic or synthetic wires, ropes, slings, etc. unless measures have been taken to contain any broken filaments or fibers (sheathing, sealing with polymers, covering, etc.).

Woven materials especially cut or unfinished ends (metal braid, EMI shielding, lacing cord, expando sleeving), unless measures have been taken to prevent fraying or generation of particles (cut with a hot knife, seal with polymer, bag, etc.).

Materials with thin films known to erode or crack or flake when subjected to normal handling (e.g., indium tin oxide [ITO] or other rigid or brittle semiconductor or ceramic coating on flexible substrates, Teflon, MLI, etc.).

Foams, highly textured materials.

Trapped debris in holes.

7.1.2.2 Molecular Generation

7.1.2.2.1 Material Selection

The CSS materials shall have a total mass loss (TML) less than 1.00% and a collected volatile condensable mass (CVCM) less than 0.10%, when measured in accordance with ASTM E-595 unless a materials usage agreement has been generated and approved by the NASA/GSFC COR.

7.1.2.2.2 Material Selection - Silicones

Silicones on external (to the spacecraft) surfaces shall not be exposed to the space environment unless approved by the NASA/GSFC COR. Silicones should be avoided or minimized. It is highly recommended that silicones be baked out at a high temperature prior to integration into the system to prevent extended bakeouts of the entire assembly.

7.1.2.2.3 Assembly Outgassing

a. The CSS outgassing shall be measured in a vacuum of 1E-5 torr at the unit under test’s maximum hot survival temperature based on Table 6-10. The hot operating temperature plus 5 degrees may be used with the NASA/GSFC COR’s approval.

b. The CSS outgassing shall not exceed 2E-8 g/sec per kg of unit under test’s mass that is condensable on a Quartz Crystal Monitor (QCM) that is operated at -20 degrees C. The measurement will be made in a chamber that has been certified clean (back ground outgassing rate and free of silicones and other high molecular weight contaminants) and has been modeled by the GSFC Contamination Analyst to account for mass sinks (cold fingers, pumps, cold surfaces, etc.) that could influence the source outgassing rate.

c. External components shall meet an outgassing rate of 1E-13 g/cm2-s with the QCM at -

20 degrees C.

d. All other external components shall meet an outgassing rate of 1E-11 g/cm2-s with the

QCM at -20 degrees C. It is recommended that certification be conducted after thermal

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cycling to take advantage of heat exposure times. Items with significant quantities of organics are baked out at the hottest temperature that will not degrade the performance of the item. This is especially true of harnesses.

e. A cold finger and/or a scavenger plate shall be used in tests for components that will be mounted externally unless approved otherwise by the NASA/GSFC COR.

f. Any motors inside a contamination critical enclosure shall have its outgassing rate certified while it is operated in its extreme hot operating environment.

7.1.2.3 Intentional and Unintentional Vents

When located external to the spacecraft, intentional vents (dedicated ascent vents or thermal blanket vents for example) and unintentional vents (through holes, gaps at seams and faying surfaces, thermal blanket openings and other openings) shall not impinge on the surface of instruments or other contamination sensitive surfaces (star trackers, calibration standards, contamination sensitive thermal surfaces like radiators, etc.).

7.1.2.4 Intentional and Unintentional Vents – Filters

Any vents shall have filters or other means of preventing the egress of particles larger than 35 microns.

7.1.2.5 Intentional Vents – Pressure Buildup Prevention

a. Intentional vents shall be sized to prevent a buildup of pressure during ascent that may result in mechanical damage.

b. Intentional vents shall be sized to prevent a buildup of pressure during high vacuum due to outgassing that could result in discharges in higher voltage circuits.

7.1.3 Cleanability

7.1.3.1 Cleanability – Sensitive Surfaces

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

7.1.3.2 Cleanability – Sensitive Surface Cleaning Methods

Alternate cleaning methods shall be identified and appropriate documentation provided for any surfaces that are not cleanable with Isopropyl alcohol.

7.2 ELECTROSTATIC CLEANLINESS

The following paragraphs provide requirements and guidelines for minimizing the magnitude and variations in the radiated electric field from the external surfaces of the CSS when exposed

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to the space plasma. All external Observatory surfaces that are exposed to the space plasma will be sufficiently conductive and be connected to spacecraft ground through low impedance paths.

7.2.1 Conductive Surface Ground Path

All CSS external conductive surfaces shall be connected to the spacecraft interface with a resistance less than 5 ohms, either through the use of ground wire(s) or through metal-to-metal mounting contact.

7.2.2 Conductive Surface Resistivity

The CSS external conductive surfaces shall have a resistivity less than 10^9 ohms/square.

7.2.3 Closeout of Gaps and Apertures

All gaps and apertures not required for component fields of view, deployments, or observatory venting shall be closed out with conductive tape or thermal blankets.

7.2.3.1 Conductive Tape Surface Resistivity

Conductive tapes used to closeout gaps or apertures shall be designed to meet the conductive surface resistivity requirements.

7.2.3.2 Conductive Tape, Grounding

The external surface of conductive tapes shall be connected to the spacecraft interface through the use of folded under tabs, or conductive adhesives.

7.2.4 Exposed Harness Specific Requirements

Harnesses that are exposed to sunlight or the ambient plasma shall be bundle shielded from source to destination.

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8.0 DESIGN & CONSTRUCTION REQUIREMENTS

8.1 PARTS, MATERIALS & PROCESSES (PMP)

8.1.1 EEE Parts

The CSS contractor’s Quality Assurance system for EEE parts will be in accordance with the requirements in the SOW, PACE-GNC-SOW-0037.

8.1.2 Materials

The CSS will be comprised of materials and processes in accordance with the requirements in the SOW, PACE-GNC-SOW-0037.

8.1.2.1 Material Conductivity

All parts should be passivated and mounting surfaces on CSS shall be conductive as defined in

Section 4.4.

8.1.2.2 NA

Table 8-1. NA

8.2 ELECTRICAL

8.2.1 Test Sensors

a. Test sensors shall be designed for flight.

b. Unless specified to be removed before flight, test sensors shall not be removed prior to flight.

8.2.2 Interface Requirements

8.2.2.1 Connector Selection

8.2.2.1.1 Connector Specifications

a. Selected connector types shall meet the Connector and Contact Requirements defined in

Section C2 of EEE-INST-002, or equivalent, with NASA/GSFC COR approval GSFC.

b. Environmental seals shall not be used in connectors especially if made from silicone without the explicit approval by the NASA/GSFC COR.

c. If a connector is not provided for the CSS, solder terminals shall be included on the CSS to facilitate electrical connection.

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8.2.2.1.2 External Box Power Connectors

The use of micro-miniature D connectors should be avoided for power interfaces.

8.2.2.1.3 Contact Derating

The current carrying capacity of the contacts shall be derated for continuous operation at the required current levels in a vacuum, as defined in Section C2 of EEE-INST-002, or equivalent, with NASA/GSFC COR approval.

8.2.2.1.4 Redundant Contact Derating

When redundant contacts are used for a single power source, each contact shall meet the required derating criteria.

8.2.2.2 Signal Segregation

a. Wherever possible, different classes of signals (power, digital, analog, etc.) shall be separated by using separate connectors.

b. If separate connectors are not feasible, classes of signals within a common connector shall be isolated from one another. Connector pin assignments should be such that sensitive circuits are separated from potential interference sources.

8.2.2.3 Test and Flight Signal Isolation

Test signals and flight signals shall not be located in the same connector.

8.2.2.4 Test Interfaces

Component test signals that require access during observatory-level testing will be handled as follows:

8.2.2.4.1 Facility-Induced Noise

All test signals should be protected or isolated from facility-induced noise.

8.2.2.4.2 Facility-Induced ESD GSE Malfunction

All test signals shall be protected or isolated from facility-induced ESD GSE malfunction.

8.2.2.4.3…