Attachment B-Spec.pdf

Effective Date: January 23, 2020

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

WFIRST-ACS-SPEC-0074, Revision- Wide Field InfraRed Survey Telescope (WFIRST), Code 448

WFIRST Coarse Sun Sensor (CSS) Specification

GSFC WFIRST CMO

February 4, 2020

Released

WFIRST CSS Specification WFIRST-ACS-SPEC-0074, Revision -ii

WFIRST Component Standard Specification

Signature/Approval Page

Prepared By:

John Dixon

Approved By:

Jason Hylan

Electronic Signatures available online in CM tool iii

Preface

This document is a Wide Field InfraRed Survey Telescope (WFIRST) Configuration Management (CM)-controlled document. Changes to this document require prior approval of the applicable Configuration Control Board (CCB) Chairperson or designee.

Proposed changes shall be submitted to the WFIRST CM Office (CMO), along with supportive material justifying the proposed change.

In this document, a requirement is identified by “shall,” a good practice by “should,” permission by “may” or “can,” expectation by “will” and descriptive material by “is.”

Questions or comments concerning this document should be addressed to:

WFIRST Configuration Management Office Mail Stop 448 Goddard Space Flight Center Greenbelt, Maryland 20771 iv

Change History Log

Revision Effective Date Description of Changes (Reference the SCoRe & Approval Date)

Revision - January 23, 2020 Initial Release of document per WFIRST-CCR-0134 v

Table of TBDs/TBRs/TBSs

Action Item No.

Location Summary Individual/ Organization

Actionee vi

Table of Contents

1.0 INTRODUCTION

1.1 General Information

1.2 Scope

1.3 Related Documentation

1.3.1 Applicable Documents [and Forms]

1.3.2 Reference Documents

2.0 CONTRACT DESCRIPTION

2.1 Coarse sun sensor Description

2.2 Ground Support Equipment Description

3.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

3.1 CSS Flight Unit Functional/Performance Requirements

3.1.1 Performance

3.1.1.1 Current Output

3.1.1.2 Maximum Current

3.1.1.3 Calibrated Accuracy

3.1.1.4 Output Variability

3.2 Resource Allocations

3.2.1 Reserved

3.2.2 Nominal Power Allocation

3.3 Reserved

3.4 Electrical Grounding

3.4.1 Reserved

3.4.2 Reserved

3.4.3 Reserved

3.4.4 Mechanical Contact Resistance

3.4.5 Mating Method

3.4.6 Reserved

3.4.7 Reserved

3.4.8 Connector and Back-Shell DC Resistance

3.5 Reserved

3.6 Reserved

3.7 Reserved

3.8 Reserved

4.0 PHYSICAL REQUIREMENTS

4.1 Interface Documentation

4.1.1 Mechanical Interface

4.1.2 Electrical Interface

4.2 Mass Properties

4.2.1 Component Masses

4.2.2 Center of Mass Location

4.2.3 Reserved

4.2.4 Reserved

vii

4.3 Physical Envelope

4.4 Mounting

4.5 Field of View

4.6 Alignment

4.6.1 Alignment Location/Orientation Accuracy

4.6.2 Alignment Knowledge Accuracy

4.6.3 Alignment Stability

4.6.4 Alignment Method

4.6.5 Reserved

5.0 ENVIRONMENTAL REQUIREMENTS

5.1 Quasi-Static Acceleration

5.2 Frequency Requirement

5.2.1 Stowed Fundamental Launch Frequencies

5.3 Vibration

5.3.1 Sinusoidal Vibration

5.3.2 Random Vibration

5.4 Shock

5.5 Reserved

5.6 Pressure

5.6.1 Operating Pressure Range

5.6.2 Maximum Depressurization Rate

5.6.3 Launch Vehicle Environmental Control System Impingement Velocity

5.7 On-Orbit Dynamic Environment

5.8 Ground Environments

5.9 Thermal Requirements

5.9.1 Flight Interface Design Temperature Limits

5.9.2 Reserved

5.10 Charged Particle Radiation Requirements

5.10.1 Definitions

5.10.2 Total Ionizing Dose

5.10.2.1 Minimum TID Tolerance for EEE Parts and Materials

5.10.3 Displacement Damage Dose

5.10.4 Single Event Effects

5.10.5 Charging Environment

6.0 CLEANLINESS

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 Particulate Contamination

6.1.1.2 Molecular Contamination – Exposed Surfaces

6.1.1.3 Molecular Contamination – Covered Surfaces

6.1.2 Surface Contamination Generation

6.1.2.1 Particulate Generation

6.1.2.2 Molecular Generation

6.1.2.3 Intentional and Unintentional Vents

6.1.2.4 Intentional and Unintentional Vents – Filters

6.1.2.5 Intentional Vents – Pressure Buildup Prevention

viii

6.1.3 Cleanability

6.1.3.1 Cleanability – Sensitive Surfaces

6.1.3.2 Cleanability – Sensitive Surface Cleaning Methods

6.2 Electrostatic Cleanliness

6.2.1 Conductive Surface Ground Path

6.2.2 Conductive Surface Resistivity

6.2.3 Closeout of Gaps and Apertures

6.2.3.1 Conductive Tape Surface Resistivity

6.2.3.2 Conductive Tape, Grounding

6.2.4 Exposed Harness Specific Requirements

6.2.5 Reserved

7.0 DESIGN & CONSTRUCTION REQUIREMENTS

7.1 Parts, Materials & Processes (PMP)

7.1.1 EEE Parts

7.1.2 Materials

7.1.2.1 Material Conductivity

7.1.3 Reserved

7.2 Electrical

7.2.1 Test Sensors

7.2.2 Interface Requirements

7.2.2.1 Connector Selection

7.2.2.2 Signal Segregation

7.2.2.3 Reserved

7.2.2.4 Reserved

7.2.3 Mitigation of Internal Charging

7.2.3.1 Mitigation Strategies for Internal Charging

7.2.3.2 Floating Conductors

7.2.3.3 Dielectric Structures

7.3 Reserved

7.4 Reserved

7.5 Identification and Marking

7.6 Workmanship

7.6.1 Workmanship Standards

7.6.2 Connector

7.6.2.1 GSE Cable Connectors

7.6.2.2 Prevention of Connector Mismating

7.6.2.3 Reserved

7.6.2.4 Connector Identification

7.6.2.5 Protection of Unused Test Connectors

7.7 Reliability and Mission Lifetime

7.7.1 Mission Life

7.7.2 Operating Time

7.7.2.1 Reserved

7.7.2.2 Failure Free Time

7.8 Ground Handling

7.8.1 Ground Support Equipment (GSE) Design

ix

7.8.2 Reserved

7.8.3 Reserved

7.8.4 Reserved

7.8.5 Reserved

7.8.6 Reserved

7.8.7 GSE Cleanliness

7.8.8 GSE Bakeout

7.8.9 Test Harness

8.0 MECHANICAL DESIGN REQUIREMENTS

8.1 Mechanical Factors of Safety

8.2 Fracture Control Requirements

8.3 Reserved

8.4 Reserved

8.5 Materials

8.5.1 Dissimilar Metals

8.5.2 Material Allowables and Stiffness

8.5.3 CTE Data

8.5.4 Stress Corrosion Cracking

8.6 Joints

8.6.1 Joint Edge Distance

8.6.2 Bonded Joint Allowable

8.6.3 Reserved

8.6.4 Fastened Joints

8.6.4.1 Fastener Locking

8.6.4.2 Fastened Joint Margin of Safety

8.6.4.3 Installation Torque Documentation

8.6.4.4 Seal Analysis

8.6.4.5 Critical Fasteners

8.7 Reserved

8.8 Reserved

9.0 LOGISTICS

9.1 Reserved

9.2 Ground Support Equipment

9.3 Transportation Equipment

10.0 VERIFICATION REQUIREMENTS

10.1 Verification Methods

10.1.1 Inspection

10.1.2 Analysis

10.1.3 Test

10.2 Inspection Requirements

10.2.1 Visual Inspection

10.2.2 Physical Measurement

10.2.3 Documentation Search

10.3 Analysis Requirements

10.4 Test Requirements

x

10.4.1 Definitions

10.4.2 Test Factors

10.4.3 Thermal Cycle Prior to Structural Testing

10.4.4 Powered-on Conditions for Test

10.4.5 Proof Testing Factors

10.4.6 Test Tolerances

10.4.7 Test Restrictions

10.4.7.1 Failure During Tests

10.4.7.2 Modification of Hardware

10.4.7.3 External Adjustment

10.4.7.4 Re-Test Requirements

10.5 Required Tests

10.5.1 Performance Tests

10.5.1.1 Comprehensive Performance Test

10.5.1.2 Limited Performance Test

10.5.1.3 Abbreviated Functional Test

10.5.1.4 Reserved

10.5.2 Mass Properties Measurement

10.5.3 Static Loads/Strength Test

10.5.3.1 Sine Burst

10.5.3.2 Static Pull

10.5.4 Frequency Signature Survey

10.5.5 Reserved

10.5.6 Sine Vibration

10.5.7 Random Vibration

10.5.8 Reserved

10.5.9 Reserved

10.5.10 Shock

10.5.11 Thermal Vacuum Bake-out

10.5.12 Thermal Vacuum Test

10.5.12.1 Thermal Vacuum Test Parameters

10.5.12.2 Thermal Vacuum Test Profile

10.5.13 Reserved

10.5.14 Thermal Cycling Testing - Ambient

10.5.15 Reserved

10.5.16 Reserved

xi

List of Figures

Figure Page

Figure 10-1. Thermal Vacuum Profile

List of Tables

Table Page

Table 5-1. CSS Design Limit Loads Table 5-2 Item Name Generic Sine Vibration Environment Table 5-3 Random Vibration Limit Levels for Components Weighing Less Than 22.7 kg Table 5-4 ASD Limit Load Reduction Table 5-5 Shock Limit Levels Table 5-6. Temperature Limits at CSS Mounting Interface Table 8-1 Design Factors of Safety Table 8-2 Minimum Edge Distance Table 10-1. Test Factors and Durations Table 10-2 Proof Test Factors Table 10-3. Test Tolerances Table 10-4. Thermal Vacuum Test Parameters

1.0 INTRODUCTION

1.1 General Information

The Wide Field InfraRed Survey Telescope (WFIRST) is a mission responding to the 2010 National Research Council New Worlds, New Horizons (NWNH) Astronomy and Astrophysics Decadal Survey top priority recommendation in the large space mission category. The science program includes two dedicated investigations to tackle outstanding questions in dark energy research and exoplanet exploration, and includes a substantial General Observer program to enable targeted investigations of astrophysical phenomena to advance other goals from the Decadal Survey. A coronagraph instrument is included in the payload for purposes of advancing the present state of the art of coronagraph technology. This document defines the requirements for the Coarse Sun Sensor.

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) WFIRST Mission.

1.3 Related Documentation

WFIRST documents can be obtained in the WFIRST CM Tool.

1.3.1 Applicable Documents [and Forms]

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 CSS 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 WFIRST CSS Statement of Work (WFIRST-ACS-SOW-0022), in which case the Statement of Work takes precedence.

Document Number Title ANSI/TIA/EIA-422-B Electrical Characteristics of Balanced Voltage Digital Interface

Circuits

ANSI/TIA/EIA-644-A-

Electrical Characteristics of Low Voltage Differential Signaling (LVDS) Interface Circuits

ASTM E-595-07 Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment

ECSS-E-ST-50-12A SpaceWire - Links, Nodes, Routers and Networks EEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, and Derating IEST-STD-CC1246E Product Cleanliness Levels And Contamination Control Program MIL-STD-1553B Department of Defense Interface Standard for Digital Time

Division Command/Response Multiplex Data Bus

MIL-STD-461F Military Standard, Electromagnetic Emission And Susceptibility Requirements For The Control Of Electromagnetic Interference

(EMI)

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

MSFC-STD-3029A Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments

NFPA 70 National Fire Protection Association National Electric Code NASA-STD-5017A Design and Development Requirements for Mechanisms NASA-STD-5019A Fracture Control Requirements for Spaceflight Hardware NASA-STD-5020 Requirements for Threaded Fastening Systems in Spaceflight

Hardware NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety

Requirements GSFC-STD-7000A General Environmental Verification Standard (GEVS) 541-WI-5330.1.41 Fastener Locking Using Arathane 5753

WFIRST-SYS-PLAN-

WFIRST Contamination Control Plan

WFIRST-SYS-SPEC-0033 WFIRST Cleanliness Specification WFIRST-ACS-LIST-0033 WFIRST Coarse Sun Sensor Deliverable Items List and

Schedule (DILS) WFIRST-ACS-SOW-0022 WFIRST Coarse Sun Sensor Statement of Work

WFIRST-SYS-ANYS-

WFIRST-SYS-ANYS-0154, WFIRST Radiation Environment, For External

1.3.2 Reference Documents

The following documents are referenced herein and amplify or clarify the information presented in this document. These documents are not binding on the content of this document.

Document Number Title

2.0 CONTRACT DESCRIPTION

2.1 Coarse sun sensor Description

Twenty Six (26) Coarse Sun Sensors (CSS) with 24 baffles will be used by the WFIRST Attitude Control System (ACS) subsystem. These Coarse Sun Sensors will be mounted upon the exterior of the WFIRST observatory.

The Coarse Sun Sensors will be used to provide sun position information as a key part of the WFIRST attitude control system.

2.2 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 24 CSS stimulators and a control panel / panels, which shall meet the requirements as defined in Section 9.2.

3.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

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

3.1 CSS Flight Unit Functional/Performance Requirements

3.1.1 Performance

3.1.1.1 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.

3.1.1.2 Maximum Current

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

3.1.1.3 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.

3.1.1.4 Output Variability

Each CSS shall have a Beginning of Life (BOL) nominal output within ±5 % of the average of all 26 CSS corresponding to full solar illumination on the boresight axis.

3.2 Resource Allocations

3.2.1 Reserved

3.2.2 Nominal Power Allocation

a. The CSS shall be a current source.

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

3.3 Reserved

3.4 Electrical Grounding

3.4.1 Reserved

3.4.2 Reserved

3.4.3 Reserved

3.4.4 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.

3.4.5 Mating Method

The primary mating method for a WFIRST 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.

3.4.6 Reserved

3.4.7 Reserved

3.4.8 Connector and Back-Shell DC Resistance

Component connectors including the EMI back-shells shall be electrically connected to chassis ground of the electronics box with a DC resistance less than or equal to 100 mΩ.

3.5 Reserved

3.6 Reserved

3.7 Reserved

3.8 Reserved

4.0 PHYSICAL REQUIREMENTS

4.1 Interface Documentation

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

• SI units with parenthetical English units are permitted for engineering and manufacturing drawings

• Angular measurement may be expressed in degree of arc or in an appropriate subdivision of degree of arc such as minute or second of arc (arc-min or arc-sec).

• Hardware that has been previously built, or new hardware of similar design heritage, may be specified in English units where use of metric equivalents would lead to additional risk and cost to the program, but interfaces with that hardware must use metric with parenthetical English units.

4.1.1 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.

4.1.2 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.

4.2 Mass Properties

4.2.1 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.

4.2.2 Center of Mass Location

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

4.2.3 Reserved

4.2.4 Reserved

4.3 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.

4.4 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.

4.5 Field of View

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

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

4.6 Alignment

4.6.1 Alignment Location/Orientation Accuracy

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

4.6.2 Alignment Knowledge Accuracy

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

4.6.3 Alignment Stability

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

4.6.4 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.

4.6.5 Reserved

5.0 ENVIRONMENTAL REQUIREMENTS

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

The CSS shall meet its performance requirements in section 3.0 during and after exposure to the environments specified in this section.

5.1 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 5-1 without damage or degradation of performance.

Note:

1. For masses within the breakpoints, linear interpolation is used to determine load.

2. Loads are considered to act in any direction, one axis at a time.

Table 5-1. CSS Design Limit Loads

Component Mass (Kg) Limit Load (g)

0.5 35.9 1 35

5.2 Frequency Requirement

5.2.1 Stowed Fundamental Launch Frequencies

The CSS shall have a fundamental frequency greater than 100 Hz when hard mounted at its spacecraft interface.

Note:

1. Requirements are met assuming rigid stiffness to restrained Degree of Freedom (DOF).

2. Components with first fundamental frequency greater than 75 Hz may verify through low level signature tests. Components with a frequency less than 75 Hz are required to perform modal testing and provide a test correlated FEM for dynamic analysis. Requirements for the FEM are in the SOW and discussed in Section 10.5.5.

3. Frequency requirements apply to all modes with modal effective mass fraction (MEMF) >5% of FEM mass. MEMF is compared to the total mass of the element/subsystem/component, not against higher levels of assembly. e.g. OBS modes MEMF is compared to OBS mass

4. For modes with MEMF between 5% and 10% that do not meet the frequency requirement(s), additional assessments can be made by the observatory systems group to determine the effects of the mode on observatory dynamics and subsystem/element design limit loads.

5.3 Vibration

5.3.1 Sinusoidal Vibration

The Coarse Sun Sensor shall withstand the sinusoidal vibration levels defined in Error!

Reference source not found.Table 5-2.

Note:

1. Levels defined are acceptance levels. Analysis is conducted against acceptance with the safety factors defined in Table 8-1. Testing factors are defined in Table 10-1.

2. Peak levels at the low end of the frequency range (5 - 20 Hz) may be ramped up as needed to accommodate table limitations.

3. The sine sweep vibration levels shown in Table 5-2 are defined at the interface of the element/subsystem.

4. Input levels may be notched to limit the test specimens CG response to 1.25 times its Design Limit Load outlined in Table 5-1.

5. Test verification is required over the range 5 to 50 Hz. Analytical verification is required from 50 to 100 Hz.

6. Elements/Subsystems/components with a first fundamental frequency equal to or greater than 75 Hz can forgo sine vibration testing from 5 to 50 Hz, while verification through analysis of the 50 to 100 Hz is still required.

7. Subsystems/components with a fundamental frequency equal to or greater than 150 Hz can forgo the sine vibration environment until higher levels of assembly upon approval by the

NASA/GSFC COR.

Table 5-2 Item Name Generic Sine Vibration Environment

Frequency (Hz) Limit Level

5 to 100 10.0 g

Levels may be notched to not exceed 1.25 times the design limit load outlined in section 5.1.

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.

5.3.2 Random Vibration

a. The Coarse Sun Sensor shall withstand the random vibration environment in Table 5-3 applied at the interface to the Coarse Sun Sensor.

The random vibration inputs are based on GEVS and may be notched to limit the test article’s cg response to 1.25 times DLL. Notched inputs must envelope minimum workmanship levels in order to screen for design or manufacturing flaws. A delicate optic, detector, sensor, etc., which is shown analytically to have negative margin at minimum workmanship levels may notch further at those frequency bands with project concurrence.

Table 5-3 shows the limit level random vibration environment for components with mass less than

22.7 kg. For components weighing more than 22.7 kg, relief can be provided in accordance with Table 5-4. Levels defined are acceptance levels. Analysis is conducted against acceptance with the safety factors defined in Table 8-1. Testing factors are defined in Table 10-1. Random vibration specs may be updated based on acoustic analysis

During testing, force limiting control is recommended and the control method must be approved by the NASA/GSFC COR.

Table 5-3 Random Vibration Limit Levels for Components Weighing Less Than 22.7 kg

Frequency (Hz) ASD Limit Level (g2/Hz)

20 0.013 20 - 50 +6 dB/oct

50 - 800 0.08 800 - 2000 -6 dB/oct

2000 0.013 Overall 10.0 grms

Table 5-4 ASD Limit Load Reduction

ASD Limit Load Reduction for Flight Hardware Weighing More Than 22.7 kg(1, 2, 3) dB reduction =10 log(W/22.7) ASD (50 - 800 Hz) = 0.08 x (22.7/W)

(1) W represents weight.

(2) Slopes are maintained at plus and minus 6dB/oct for components weighing up to 59kg.

Above 59 kg, slopes are adjusted to maintain an ASD level of 0.01 g2/Hz at 20 and 2000Hz.

(3) For components weighing more than 182 kg, levels are maintained at the 182 kg level.

5.4 Shock

The Coarse Sun Sensor shall be designed to meet its performance requirements after being subjected to the shock environment in Table 5-5, applied at the Coarse Sun Sensor interface to the WFIRST spacecraft structure. Levels defined are acceptance levels.

Testing factors are defined in 10.4.2.

Table 5-5 Shock Limit Levels

Frequency (Hz) Limit Load (g)

100 100 2000 2500 10000 2500

Note: A shock susceptibility and attenuation assessment is to be performed on CSS. This assessment can be based on past shock tests or other relevant information.

5.5 Reserved

5.6 Pressure

5.6.1 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).

5.6.2 Maximum Depressurization Rate

The CSS shall be designed to meet all performance requirements after exposure to a depressurization rate of 2.48 kPa/sec (0.36 psi/sec) or less with a single brief allowable peak of up to 5.03 kPa/sec (0.73 psi/sec).

Note: Electronic boxes which meet the requirements of Electronic Box Minimum Vent Area of Section 8.4 may waive maximum depressurization rate analysis.

5.6.3 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.

5.7 On-Orbit Dynamic Environment

The CSS shall meet all requirements while subjected to loads resulting from the following rate and accelerations:

Angular acceleration = 4.5 E-3 rad/sec^2, per axis Angular velocity = 3.5 E-3 rad/sec, per axis Lateral acceleration = 0.0035 m/sec^2, per axis Axial acceleration = 0.045 m/sec^2, thrust direction (+X)

Note:

1. Angular rates and accelerations are about the Observatory CG

2. These loads are to be combined with other loads (thermal, etc.).

3. Angular velocity is applied so that centripetal force adds to the lateral or axial load applied.

5.8 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.

5.9 Thermal Requirements

5.9.1 Flight Interface Design Temperature Limits

a. The CSS shall be capable of surviving indefinitely when its temperatures are within the survival limits shown in Table 5-6 without damage or permanent performance degradation.

b. The CSS shall meet all performance requirements when operational anywhere within the Operational and Protoflight/Qualification limits shown in Table 5-6.

c. The CSS shall operate at the Minimum and Maximum Protoflight/Qualification (in spec) limits shown in Table 5-6.

Table 5-6. Temperature Limits at CSS Mounting Interface

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

-100 +90

Survival (Unpowered) -135 +90

5.9.2 Reserved

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

5.10.1 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).

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.

5.10.2 Total Ionizing Dose

5.10.2.1 Minimum TID Tolerance for EEE Parts and Materials

All EEE parts shall tolerate a minimum Total Ionizing Dose (TID) defined in WFIRST-SYS-

ANYS-0154.

5.10.3 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.

The silicon and gallium arsenide EEE parts susceptible to displacement damage degradation shall be able to tolerate a minimum DDD as defined in WFIRST-SYS-ANYS-0154.

5.10.4 Single Event Effects

a. The Coarse Sun Sensor shall be designed to avoid or tolerate errors due to non-destructive Single Event Effects (SEE).

b. All EEE parts shall have LET thresholds for SEL greater than 75 (MeV·cm2)/mg for the destructive events. Parts with LET thresholds between 37 MeV-cm2/mg and 75 MeV-cm2/mg could be used if the evaluation of the probability and outcome of the destructive event shows that there would be no mission impact. Reference the WFIRST Radiation Specification (WFIRST-SYS-SPEC-0083) for details on the environments.

5.10.5 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 WFIRST mission orbit.

6.0 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.

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 Particulate Contamination

The CSS shall meet IEST-STD-CC1246E VC-0.5-1000 + UV, or equivalent, when inspected with both UV and white light in a darkened room prior to be integrated to WFIRST.

6.1.1.2 Molecular Contamination – Exposed Surfaces

The CSS shall meet IEST-STD-CC1246E NVR Cleanliness Level R1 and shall also be visually inspected and meet Level VC-0.5-1000 + UV, or equivalent, when inspected with both UV and white light in a darkened room prior to be integrated to WFIRST.

6.1.1.3 Molecular Contamination – Covered Surfaces

a. All CSS surfaces not exposed to the space environment shall be visually inspected and meet

IEST-STD-CC1246E Level VC-0.5-1000 + UV, or equivalent, when inspected in a darkened room with white and UV light prior to be integrated to WFIRST.

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

6.1.2 Surface Contamination Generation

6.1.2.1 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)).

• 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.

6.1.2.2 Molecular Generation

6.1.2.2.1 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.

6.1.2.2.2 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.

6.1.2.2.3 Assembly Outgassing

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

b. The CSS outgassing measurements shall 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. The CSS outgassing shall meet an outgassing rate of 5E-9 g/s with the QCM at -20 degrees C. It is recommended that certification be conducted after thermal cycling to take advantage of heat exposure times.

d. 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.

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

6.1.2.3 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.).

6.1.2.4 Intentional and Unintentional Vents – Filters

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

6.1.2.5 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.

6.1.3 Cleanability

6.1.3.1 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.

6.1.3.2 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.

6.2 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 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.

6.2.1 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.

6.2.2 Conductive Surface Resistivity

The CSS external conductive surfaces shall have a resistivity less than 10^8 ohm-cm or 10^9 ohms/square.

6.2.3 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.

6.2.3.1 Conductive Tape Surface Resistivity

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

6.2.3.2 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.

6.2.4 Exposed Harness Specific Requirements

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

6.2.5 Reserved

7.0 DESIGN & CONSTRUCTION REQUIREMENTS

7.1 Parts, Materials & Processes (PMP)

7.1.1 EEE Parts

The CSS contractor’s Quality Assurance system for EEE parts will be in accordance with the requirements in the SOW, WFIRST-ACS-SOW-0022.

7.1.2 Materials

The CSS will be comprised of materials and processes in accordance with the requirements in the SOW, WFIRST-ACS-SOW-0022.

7.1.2.1 Material Conductivity

All parts should be passivated and mounting surfaces on CSS shall be conductive as defined in Section 3.4.

7.1.3 Reserved

7.2 Electrical

7.2.1 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.

7.2.2 Interface Requirements

7.2.2.1 Connector Selection

7.2.2.1.1 Connector Specifications

a. Selected connector types shall meet the Connector and Contact Requirements defined in Section C2 of EEE-INST-002.

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. Flying leads, greater than 45.7 cm (18 in) in length, shall be attached to the included terminals to facilitate a splice.

7.2.2.1.2 External Box Power Connectors

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

7.2.2.1.3 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.

7.2.2.1.4 Redundant Contact Derating

When redundant contacts are used, each contact shall meet the required derating criteria.

7.2.2.2 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.

7.2.2.3 Reserved

7.2.2.4 Reserved

7.2.3 Mitigation of Internal Charging

Internal charging refers to the physical effect where high energy electrons deposit charge in a dielectric, if the charging rate is higher than the leakage rate eventually a point is reached where the dielectric discharges to the nearby structure.

7.2.3.1 Mitigation Strategies for Internal Charging

There are a number of mitigation strategies possible to prevent charge build-up and subsequent arcing within components. Designers shall implement one, or more, of the following strategies:

(1) Provide sufficient (greater than 3 mils of Al equivalent) shielding for exposed cables or components. If possible, use the shielded cable to provide the additional shielding.

(2) Use low pass filters at either end of the cable that absorb the energy from the discharge without creating dangerous voltages.

(3) Provide circuits at either end that have sufficiently low impedance that they are not harmed by the discharges

Note, that if the contractor uses strategies (2) or (3), all affected circuit designs shall be analyzed to demonstrate compliance with the requirements, using a transient circuit analysis tool such as PSpice.

7.2.3.2 Floating Conductors

Floating conductors, if present, shall have a bleed path of less than 10 MΩ to the component structure. This requirement is not applicable to small floating conductors (645mm2 (1 inch2) or less) or short (127 mm (5 inch) or less) unterminated traces or wires that are inside of the components.

7.2.3.3 Dielectric Structures

7.2.3.3.1 Bulk Resistivity

Dielectric structures shall have a bulk resistivity less than 1012 ohm-cm.

7.2.3.3.2 Charge Bleed-Off

All dielectric structures shall have a charge bleed path to the spacecraft interface, designed to route the discharge into the spacecraft structure in a controlled fashion.

7.3 Reserved

7.4 Reserved

7.5 Identification and Marking

a. Each unit shall be permanently marked with the part number and a unique sequential serial number in the area designated on the interface control drawing in a manner to be approved by the NASA/GSFC COR.

b. All markings shall use alcohol proof ink, engraving, or laser etching.

7.6 Workmanship

7.6.1 Workmanship Standards

The workmanship standards and processes outlined in the SOW will be used.

7.6.2 Connector

7.6.2.1 GSE Cable Connectors

GSE cable connectors that mate with flight test connectors shall be flight-approved connectors.

7.6.2.2 Prevention of Connector Mismating

Connector mismating prevention requirements are identified in this section.

7.6.2.2.1 Connector Uniqueness

Physically adjacent connectors shall be of different sizes or of different sexes or uniquely keyed to facilitate proper mating.

7.6.2.2.2 Connector Facing

Physically adjacent Rectangular “D” connectors should not face the same way.

7.6.2.2.3 Connector Keying

Circular connectors shall be positively keyed.

7.6.2.2.4 Accessibility

The CSS Spacecraft interface connectors shall be spaced far enough apart to allow the mate and demate operations to be performed without a special tool.

7.6.2.2.5 Reserved

7.6.2.2.6 “Scoop-Proof” Connectors shall be “scoop-proof” to prevent the connector shell from contacting pins during mating.

7.6.2.3 Reserved

7.6.2.4 Connector Identification

Each connector shall be labeled and clearly visible to facilitate proper mating.

7.6.2.5 Protection of Unused Test Connectors

Test connectors shall be capped with flight-approved RF and static control covers when not in use.

7.7 Reliability and Mission Lifetime

7.7.1 Mission Life

The CSS shall meet all performance specifications through four (4) years of ground testing and five (5) years of operation in space, following a three (3) month commissioning period.

7.7.2 Operating Time

7.7.2.1 Reserved

7.7.2.2 Failure Free Time

a. The CSS shall receive at least 32 hours of failure free operating time in vacuum at the hot dwell temperature and 32 hours of failure free operating time at the cold dwell temperature prior to delivery to NASA GSFC.

b. The CSS shall demonstrate failure-free performance for the last 40 hours of operation prior to delivery to NASA GSFC.

7.8 Ground Handling

7.8.1 Ground Support Equipment (GSE) Design

All electrical EGSE or support equipment shall be in compliance with the National Electric Code (NFPA 70) or equivalent standard.

7.8.2 Reserved

7.8.3 Reserved

7.8.4 Reserved

7.8.5 Reserved

7.8.6 Reserved

7.8.7 GSE Cleanliness

All Ground and Test support equipment shall be compatible with the flight component and the environment where the flight component or test component will reside (cleanroom, thermal vacuum chamber, vibration cell, etc.)

7.8.8 GSE Bakeout

Thermal Vacuum GSE shall be baked out and the outgassing rate certified prior to the test.

7.8.9 Test Harness

Test harnesses that will be used in vacuum during ground operations shall be vacuum compatible.

8.0 MECHANICAL DESIGN REQUIREMENTS

For additional guidance on the design and analysis of threaded fastening systems in NASA spaceflight hardware, consult NASA-STD-5020, “Requirements for Threaded Fastening Systems In Spaceflight Hardware”.

8.1 Mechanical Factors of Safety

Positive Margin of Safety (MS) shall be demonstrated analytically for Coarse Suns Sensor, including interfaces to MGSE, using the MS formula below with the appropriate Factors of Safety (FS) defined in Table 8-1 applied to flight limit loads.

Note:

1. MS is defined as follows:

𝑀𝑀𝑆𝑆 =

𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴 𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿 𝐿𝐿𝐴𝐴𝐴𝐴𝐿𝐿 𝐴𝐴𝑜𝑜 𝑆𝑆𝐿𝐿𝑜𝑜𝐴𝐴𝑆𝑆𝑆𝑆

𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐿𝐿𝐴𝐴𝐿𝐿 𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿 𝐿𝐿𝐴𝐴𝐴𝐴𝐿𝐿 𝐴𝐴𝑜𝑜 𝑆𝑆𝐿𝐿𝑜𝑜𝐴𝐴𝑆𝑆𝑆𝑆 ∗ 𝐹𝐹𝑆𝑆

− 1

2. Positive MS in joints that are both bonded and fastened is demonstrated assuming only the bond carries the load, only the bolt carries the load, or if the capability of the joint requires both the bond and fasteners, a detailed analysis is performed.

a. The adhesive portion of the joint is usually the stiffest load path and thus carries most of the load. In this scenario, the full strength of the bolted interface is not realized until after the failure of the adhesive.

3. Positive MS is demonstrated for adhesively bonded joints for all potential failure modes under all loading conditions and temperatures.

a. Failure modes may include cohesive failure in the adhesive, failure at the adhesive to adhered interface, and adherent failure. Composite adherent failure modes may include in-plane tension/compression, in-plane shear, peel, interlaminar shear, and interactions between failure modes.

b. Fiber orientation should be considered in the computation of strength for a bonded joint where one or more of the adherends are made of composite material.

c. Interaction between adhesively bonded joints in close proximity should be included when analyzing bonded joints.

d. Temperature stresses in bonded joints and sandwich panels should consider bond and consolidation cure temperatures.

e. Adhesive bond fillets and bondline thickness should be considered in bonded joint design and analysis under temperature loading.

Table 8-1 Design Factors of Safety

Type of Hardware Static /Sine Random/Acoustic 3,4

Tested Metallic Yield 1.252 1.6

Tested Metallic Ultimate 1.42 1.8

Stability/Buckling Ultimate 1.4 1.8 Beryllium Yield 1.4 1.8

Type of Hardware Static /Sine Random/Acoustic 3,4

Beryllium Ultimate 1.6 2.0 Composite Ultimate1 1.5 1.9 Bonded inserts/Joints Ultimate 1.5 1.9 Glass/Ceramic 3.05 3.8 Bond in Glass/Ceramic 1.5 1.9

(1) All composite structures must be tested to 1.25 x limit loads.

(2) For qualification by analysis only, positive margin must be shown with FS of 2.0 and 2.6 for metallic yield and ultimate, respectively.

(3) Factors shown are applied to statistically derived peak response based on limit/acceptance RMS level.

As a minimum, the peak response must be calculated as a 3-sigma value.

(4) Factors shown assume that qualification/prototype testing is performed at 3dB above limit/acceptance limits. If the difference between acceptance and qualification is less than 3dB then the above factors may be applied to qualification minus 3dB.

(5) For qualification by analysis only, a no-test safety factor of 5.0 may be used for glass (not ceramic), per

NASA-STD-5001B Table 3.

8.2 Fracture Control Requirements

The vendor shall evaluate all hardware to determine fracture criticality. A four-tier system is used to evaluate fracture criticality of hardware; each tier has associated mitigation requirements that shall be followed, as specified below.

Tier 1 fracture classification: Light-weighted optics, ceramic structural elements used to transfer load from one component to another, and any glass/ceramic hardware designated high risk by WFIRST shall mitigate risk of catastrophic failure due to cracks/flaws through the following requirements:

a. Proving design is fail-safe or,

a. Industry standard fracture analysis using NASGRO, Flaw Growth Analysis per NASA-STD- 5018, or other fracture analysis software demonstrating safe-life for 4 times number of mission cycles.

and,

b. Positive MS using B-basis Weibull accounting for Weibull scale effect per ASTM C1683 and,

c. Slow crack growth analysis of glass/ceramics made of materials…