PACE-GNC-SPEC-0064B_06-18-2018.pdf

PACE Star Sensor System SPEC PACE-GNC-SPEC-0064, Revision B

Effective Date: June 18, 2018 i Check https://ipdtdms.gsfc nasa.gov to verify that this is the correct version prior to use

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

PACE Star Sensor System Specification

Signature/Approval Page

Prepared By:

Linh Nguyen

Reviewed By:

Nikesha Davis

Jamie Eitnier

Craig Stevens

Dave Sohl

Jack Sanders

Marlon Enciso

Approved By:

Kathy McIntyre

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.

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

Change History Log

Revision Effective Date Description of Changes

(Reference the SCoRe & Approval Date)

Revision - 05/01/2018 Baseline release following approval of PACE-CCR-0293

Revision A 06/05/2018 Updated following the approval of PACE-CCR-0362

Revision B 06/18/2018 Updated following the approval of PACE-CCR-0379 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 Star Sensor System Description

3.2 Ground Support Equipment Description

4.0 Functional/Performance Requirements

4.1 Star Sensor System Flight Unit Functional/Performance Requirements

4.1.1 Attitude Products

4.1.1.1 Attitude Quaternion Solutions

4.1.1.2 Full Performance Attitude Accuracy

4.1.1.3 Attitude Data Output Rate

4.1.1.4 Attitude Data Time-Tag Accuracy

4.1.1.5 Attitude Data Latency

4.1.1.6 Angular Rate Output

4.1.1.7 Angular Rate Accuracy

4.1.1.8 Angular Rate Data Output Rate

4.1.2 Autonomous Attitude Acquisition Rate Capability

4.1.3 Rate and Acceleration Capability for Attitude Solution

4.1.4 Rate and Acceleration Capability for Rate Solution

4.1.5 1 Pulse Per Second Interface

4.1.6 Acquisition Time after Power on Time

4.1.7 Warm-Up/Cool-down Time

4.1.8 Bright Source Protection – Sun

4.1.9 Bright Source Protection – Earth or Moon

4.2 Resource Allocations

4.2.1 Mass Allocation

4.2.2 Nominal Power Allocation

4.2.3 Peak Power Allocation

4.3 Power

4.3.1 Primary (Unregulated) Power Input Requirements

4.3.1.1 Operating Voltage Range

4.3.1.2 Abnormal Voltages

4.3.1.3 Voltage Ripple

4.3.1.4 Voltage Transients

4.3.1.5 Sudden Removal of Power

4.3.1.6 Polarity Reversal Protection

4.3.1.7 Over-Current Protection

4.3.1.8 Primary Power Return Ground

4.3.1.9 Power Wiring Redundancy

4.3.2 Load Induced Noise Requirements

4.3.2.1 Turn-on Current Transients

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4.3.2.2 Operational Current Transients

4.3.2.3 Repetitive Transients

4.4 Electrical Grounding

4.4.1 Primary Power DC Isolation

4.4.2 Internally Generated Secondary Returns

4.4.3 Internally Generated Secondary to Primary DC Isolation

4.4.4 Mechanical Contact Resistance

4.4.5 Mating Method

4.4.6 Ground Strap, Non-Conductive Surface

4.4.6.1 Ground Lug Contact Area

4.4.7 NA

4.4.8 Connector DC Resistance

4.5 Signal And Data Interfaces

4.5.1 NA

4.5.2 NA

4.5.3 NA

4.5.4 Data and Clock Signals

4.5.4.1 Command and Telemetry Interface

4.5.4.2 1PPS Clock Signal Interface

4.5.4.3 RS-422 Interface

4.5.4.4 1553 Bus Interface

4.5.4.5 SpaceWire Interface

4.6 Operating Modes

4.6.1 Minimum Operating Modes

4.7 Command and Data Services

4.7.1 Commands

4.7.2 Data

4.8 Flight Software

4.8.1 Flight Software and Star Catalog

4.8.2 Software Version Identification

4.8.3 Software Updates

4.8.4 Software Table Updates

4.8.5 Non-Volatile Memory Storage of Flight Code

4.8.6 Boot Code

4.8.7 Star Catalog Version

4.8.8 Software Table Dump

4.8.9 Packet Checksum

4.8.10 Memory Load

4.8.11 Memory Dump

4.8.12 Memory Checksum

4.8.13 Software Instructions

5.0 Physical Requirements

5.1 Interface Documentation

5.1.1 Mechanical Interface

5.1.2 Electrical Interface

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5.1.3 Data Interface

5.2 Mass Properties

5.2.1 Component Masses

5.2.2 Center of Mass Location

5.2.3 Center of Mass Accuracy

5.2.4 Determination of Moments and Products of Inertia

5.3 Physical Envelope

5.4 Mounting

5.5 Fields of View

5.5.1 Sun Exclusion Angle

5.5.2 Moon and Earth Exclusion Angle

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

5.6.5 Alignment Measurement Tooling/Features

6.0 Environmental Requirements

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 Acoustics

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 Component Temperature Limits

6.11.2 NA

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 ELDR Evaluation

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6.12.3 Displacement Damage Dose

6.12.4 Single Event Effects

6.12.4.1 Destructive Events (SELs)

6.12.4.2 Destructive Events (SEBs and SEGRs)

6.12.4.3 Non-Destructive Events (SEUs, SETs, SEFIs, SHEs, and MBUs) ... 35

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

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 Material Limitations for Debris Casualty Area

8.1.3 Software Assurance

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

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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 Electromagnetic Compatibility

8.4.1 Conducted Emissions

8.4.1.1 Applicability of Conducted Emissions

8.4.1.2 CE101 – Conducted Emissions, Differential Mode Limits

8.4.1.3 CE102 – Conducted Emissions, Differential Mode Limits

8.4.1.4 Conducted Emissions, Common Mode Limits

8.4.2 Conducted Susceptibility

8.4.2.1 Applicability of Conducted Susceptibility

8.4.2.2 CS101 – Power Leads, 30 Hz to 150 kHz Limit

8.4.2.3 CS114 – Power Leads, 150 kHz to 50 MHz Limit

8.4.2.4 CS106 - Transients

8.4.2.5 CS114 – Power and Signal Cables, Common Mode

8.4.3 Radiated Emissions

8.4.3.1 RE102 - Electric Field Emissions

8.4.4 Radiated Susceptibility

8.4.4.1 RS103 - Electric Field

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

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

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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

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

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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

11.5.1.4 NA

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 Acoustic Test

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

11.5.12 Thermal Balance Requirement

11.5.13 Thermal Cycling Testing

11.5.14 Harness Tests

11.5.14.1 Continuity/Hi-Pot Tests

11.5.14.2 Harness Bake-Out

11.5.15 EMI/EMC Tests

11.5.15.1 Conducted Emissions Tests

11.5.15.2 Conducted Susceptibility Tests

Appendix A Abbreviations and Acronyms xii

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List of Figures

Figure Page

Figure 4-1 Inrush Current Limits

Figure 6-1 Sine Vibration Environment

Figure 6-3 Shock Envelope

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

Figure 6-5 Integral LET Spectra for Galactic Cosmic Ray Ions

Figure 6-6 Solar Heavy Ion Linear Energy Transfer Spectrum

Figure 8-1 CE101/CE102 Conducted Emissions Differential Mode Limits

Figure 8-2 Conducted Emissions, Common Mode Limits

Figure 8-3 Power Lead Conducted Susceptibility (CS101) Voltage Limit

Figure 8-4 CS106 Transient Waveform

Figure 8-5 Bulk Cable Injection (CS114) Calibrated Test Levels

Figure 8-6 Component Level RE102 Radiated Electric Field Emission Limits

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 Star Sensor System Design Limit Loads

Table 6-3 Star Sensor System Sine Vibration Environment

Table 6-4 Star Sensor System Random Vibration Environment, (22.7 kg, or less)

Table 6-6 Qualification Level Shock Response Spectrum

Table 6-7 Flight/Acceptance Level Acoustic Environments

Table 6-8 Protoflight/Qualification Level Acoustic Environments

Table 6-9 Transportation Loads

Table 6-10 Component Temperature Limits

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

Table 6-12 Integral LET Spectra for Galactic Cosmic Ray Ions

Table 6-13 Solar Heavy Ion Linear Energy Transfer Spectrum

Table 8-1 Limited materials for debris casualty area

Table 8-2 Emission and Susceptibility Requirements

Table 8-3 Receiver Notch RE102 Test Limit Levels (TBR)

Table 8-4 RS103 Radiated Susceptibility Levels

Table 11-1 Test Factors and Durations

Table 11-2 Test Tolerances

Table 11-3 Thermal Vacuum Test Parameters

PACE Star Sensor System Spec PACE-GNC-SPEC-0064, Revision B

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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 Star Sensor System 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

Star Sensor System 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 Star Sensor System Statement of Work (PACE-GNC-SOW-0034), in which case the

Statement of Work takes precedence.

There are many applicable documents (standards, procedures, etc.) throughout this SPEC. 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-0034 PACE Star Sensor System Statement of Work

PACE-GNC-LIST-0017 PACE Star Sensor System Deliverable Items List and Schedule (DILS)

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

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

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Document Number Title

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)

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

MIL-STD-1553B Department of Defense Interface Standard for

Digital Time Division Command/Response

Multiplex Data Bus

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

ECSS-E-ST-50-12A SpaceWire - Links, Nodes, Routers and Networks

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3.0 CONTRACT DESCRIPTION

3.1 STAR SENSOR SYSTEM DESCRIPTION

The main science instrument for PACE resides upon a tilting platform that is attached to the spacecraft body. This tilt platform rotates +20 to -20 degrees from the nominal attitude about the pitch axis twice per orbit. During the scan, the rotation rate is 1.33 deg/sec. The PACE architecture allows for:

a) 1 to 2 star sensors to be mounted on the instrument in order to provide attitude orientation knowledge of the instrument for the post-processing of the science data, and

b) 1 to 2 star sensors upon the spacecraft body in order to provide attitude pointing and optional attitude rate information for the purposes of spacecraft attitude control.

As a result, the orientation of the Star Sensor(s) on the instrument will change relative to the Star

Sensor(s) on the spacecraft body. The instrument star sensors are only required to meet knowledge requirements at the two final tilt orientations (not during the slew).

The Star Sensor System shall consist of the processing electronics, 2-to-4 optical heads with star sensors, optical head baffles for bright object protection, required operational power and signal cabling between processing electronics and optical heads, and test optical Electronic GSE for ground stimulation of star sensors.

3.2 GROUND SUPPORT EQUIPMENT DESCRIPTION

a) In accordance with the SOW, the Star Sensor System GSE and all unique software required to operate the Star Sensor System shall be provided for integration and testing of the star sensor systems at NASA/GSFC.

b) The Star Sensor System shall be delivered with vacuum compatible optical stimulators for limited functional and performance testing at the spacecraft level, including thermal vacuum testing.

c) Mounting interface and alignment hardware shall be provided for all necessary GSE optics.

d) The contractor shall provide non-flight baffle covers.

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4.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

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

4.1 STAR SENSOR SYSTEM FLIGHT UNIT

FUNCTIONAL/PERFORMANCE REQUIREMENTS

4.1.1 Attitude Products

4.1.1.1 Attitude Quaternion Solutions

The Star Sensor System shall provide three-axis time-tagged inertial attitude quaternion solutions corrected for systematic errors relative to the sensor axes and the star magnitude (from the ECI J2000 frame to the star sensor frame).

4.1.1.2 Full Performance Attitude Accuracy

The Star Sensor System attitude quaternion solutions shall meet or exceed the following accuracies, except during periods of occultation by the Earth, Moon, or Sun:

a) 17 arcsec (3 sigma) per axis with a 7 arcsec (3 sigma) Noise Equivalent Angle for the star sensors mounted on the instrument.

b) 17 arcsec (3 sigma) per axis with a 7 arcsec (3 sigma) Noise Equivalent Angle for the star sensors mounted on the spacecraft body.

4.1.1.3 Attitude Data Output Rate

The Star Sensor System shall provide attitude quaternion solutions at a 5 Hz rate.

4.1.1.4 Attitude Data Time-Tag Accuracy

The contractor shall define and provide the time and relative location within the integration period less than or equal to 0.1 milliseconds relative to a 1 pulse per second (PPS) signal.

4.1.1.5 Attitude Data Latency

Attitude data shall have a latency of no greater than 300 msec. Latency is defined as the time delay from the time data is acquired by the Star Sensor System to the time the data is output to the spacecraft.

4.1.1.6 Angular Rate Output

The Star Sensor System shall provide a three-axis time-tagged angular rate vector solution.

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4.1.1.7 Angular Rate Accuracy

The Star Sensor System angular rate data shall be accurate to within 100 arcsec/sec (3-sigma) per axes.

4.1.1.8 Angular Rate Data Output Rate

The Star Sensor System shall provide angular rate vector solutions at a 5 Hz rate.

4.1.2 Autonomous Attitude Acquisition Rate Capability

The Star Sensor System shall have an autonomous attitude acquisition capability that operates up to a maximum spacecraft body rate of 4 deg/sec.

4.1.3 Rate and Acceleration Capability for Attitude Solution

The Star Sensor System shall maintain its attitude solution performance accuracy up to:

a) a maximum rate of 2 deg/sec, and

b) an acceleration of 1 deg/sec2.

4.1.4 Rate and Acceleration Capability for Rate Solution

The Star Sensor System shall maintain its rate solution performance accuracy up to:

a) a maximum rate of 2 deg/sec, and

b) an acceleration of 1 deg/sec².

4.1.5 1 Pulse Per Second Interface

a) The Star Sensor System shall accept a 1 Pulse Per Second (1PPS) signal from the spacecraft, an active high, 100 ms (+/- 5ms) wide pulse used to synchronize the Star Sensor System to

Spacecraft time.

b) The Star Sensor System shall synchronize its internal time to within 0.1 millisecond of the 1PPS pulse.

4.1.6 Acquisition Time after Power on Time

The Star Sensor System, with no prior attitude knowledge, shall survey the star field and output a correct quaternion output within 1 minute after warm up/cool down is complete.

4.1.7 Warm-Up/Cool-down Time

The Star Sensor System shall meet performance requirements within 5 minutes of the application of power.

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4.1.8 Bright Source Protection – Sun

The Star Sensor System shall not be damaged when subjected to exposure of the sun or directly reflected sun in its FOV for 2 minutes, powered or unpowered.

4.1.9 Bright Source Protection – Earth or Moon

The Star Sensor System shall not be damaged when subjected to exposure of the earth or moon in its FOV indefinitely, powered or unpowered.

4.2 RESOURCE ALLOCATIONS

4.2.1 Mass Allocation

The Star Sensor System shall have a total mass of less than or equal to 10.0 kg (not including a harness).

4.2.2 Nominal Power Allocation

The Star Sensor System shall have a nominal (continuous) power consumption of less than or equal to 21.0 W for a duration of 100 minutes.

4.2.3 Peak Power Allocation

The Star Sensor System shall have a peak (instantaneous) power consumption of less than or equal to 24.0 W for a duration of 60 seconds.

4.3 POWER

4.3.1 Primary (Unregulated) Power Input Requirements

4.3.1.1 Operating Voltage Range

The Star Sensor System shall be designed to operate over the bus voltage range of +27 to +35

VDC at their primary power inputs during all normal mission phases and for all expected load conditions (except when turned off).

4.3.1.2 Abnormal Voltages

The Star Sensor System shall survive without performance degradation after indefinite exposure to an anomalous voltage range of 0 to +40 VDC. This requirement is to be verified by analysis or test of a non-flight unit.

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4.3.1.3 Voltage Ripple

The Star Sensor System shall be designed to meet its performance requirements in the presence of ripple voltages specified in the CS101 and CS114 requirements in paragraphs 8.4.2.2 and

8.4.2.3, respectively.

4.3.1.4 Voltage Transients

The Star Sensor System shall be designed to meet its performance requirements in the presence of transient voltages specified in the CS106 requirements in paragraph 8.4.2.4.

4.3.1.5 Sudden Removal of Power

The Star Sensor System shall meet its performance requirements without degradation after exposure to an abrupt, unannounced removal of power.

4.3.1.6 Polarity Reversal Protection

The Star Sensor System should have built-in protection to prevent damage due to polarity reversal at the power inputs (up to -35V), where feasible.

4.3.1.7 Over-Current Protection

The Star Sensor System should not use non-resetting over-current protection (i.e., fuses) internal to the unit, but if they are required by the contractor to isolate an internal fault, prior approval shall be obtained from the NASA/GSFC Contracting Officer’s Representative (COR).

4.3.1.8 Primary Power Return Ground

Star Sensor System shall provide a dedicated Primary Power return in the same connector as the primary power.

4.3.1.9 Power Wiring Redundancy

The Star Sensor System should utilize redundant power wiring, if possible.

4.3.2 Load Induced Noise Requirements

4.3.2.1 Turn-on Current Transients

With an input voltage (0 – 35V) rise time no faster than 500us, the inrush current shall not exceed the no-trip limit in Figure 4-1 below for durations greater than 100 microseconds.

4.3.2.1.1 Turn-on Current Transients (less than 100 microseconds)

For durations less than 100 microseconds, inrush current shall not exceed 600% of switch current rating of 1 amp.

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Figure 4-1 Inrush Current Limits

The compliance of this requirement can be demonstrated using a solid-state power switch with the similar turn on characteristics as the flight power control switch.

4.3.2.2 Operational Current Transients

a) The Star Sensor System operational current transients shall be less than or equal to 5A above the steady state current for a period of less than 1 millisecond.

b) The rate of change of the current shall be less than or equal to 20 milliamperes/microsecond.

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4.3.2.3 Repetitive Transients

Any repetitive current transients, in the frequency range between 1 Hz to 200 Hz, shall not exceed 0.2A peak-to-peak.

This measurement will be performed with an oscilloscope current probe in time domain.

4.4 ELECTRICAL GROUNDING

4.4.1 Primary Power DC Isolation

At the Star Sensor System primary power interfaces, primary power and primary power returns shall be isolated from signal grounds and from the component chassis by a DC resistance of greater than or equal to 1 MΩ.

4.4.2 Internally Generated Secondary Returns

Secondary returns (power and signal grounds) shall be referenced to the component chassis ground by connecting them at one or more places.

4.4.3 Internally Generated Secondary to Primary DC Isolation

Secondary power (or signal) inputs shall be isolated from primary power by a DC resistance of greater than 1 MΩ.

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 Ground Strap, Non-Conductive Surface

The Star Sensor System shall contain provisions for a grounding strap to be attached from the component chassis for connection to the spacecraft conductive structure.

4.4.6.1 Ground Lug Contact Area

The grounding lug location on the component chassis or the tie points in contact with the ground strap shall have a minimum contact area of 80 mm². The ground lug contact area must remain

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free of any material finish that may affect the reliability of the ground connection and be shown in the MICD.

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.

4.5 SIGNAL AND DATA INTERFACES

4.5.1 NA

4.5.2 NA

4.5.3 NA

4.5.4 Data and Clock Signals

4.5.4.1 Command and Telemetry Interface

The command and telemetry interface to the Star Sensor System shall be either digital differential RS-422 (ANSI/TIA/EIA-422-B), MIL-STD-1553B, or Spacewire interface that accommodates the Star Sensor System Transmit/Receive interfaces.

4.5.4.2 1PPS Clock Signal Interface

The 1PPS signal interface to the Star Sensor System shall be a digital differential RS-422

(ANSI/TIA/EIA-422-B) interface.

4.5.4.3 RS-422 Interface

The RS-422 interface circuit and line termination shall be documented in the Star Sensor System

Electrical ICD, using ANSI/TIA/EIA-422-B, Electrical Characteristics of Balanced Voltage

Digital Interface Circuits as a guide.

Signals that use RS-422 interface should be designed to minimize the chassis ground current loop. The chassis current due to the line imbalance of the 3.3V CMOS interface should be less than 30 micro-amps.

4.5.4.4 1553 Bus Interface

Any 1553 Bus interface shall be designed per MIL-STD-1553B, Digital Time Division

Command/Response Multiplex Data Bus.

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4.5.4.5 SpaceWire Interface

Signals that use the SpaceWire interface shall adhere to the electrical characteristics and termination requirements found in ECSS-E-ST-50-12A, SpaceWire - Links, Nodes, Routers and

Networks.

4.6 OPERATING MODES

4.6.1 Minimum Operating Modes

The Star Sensor System shall have stand-by and attitude operating modes at a minimum. The

Star Sensor(s) will be off during launch and ascent.

4.7 COMMAND AND DATA SERVICES

4.7.1 Commands

The Star Sensor System shall accommodate the following set of commands or their equivalents, at a minimum.

a) Reset: The Star Sensor System shall accept a reset command that causes the data processor to start over.

b) Software Table Dump: The Star Sensor System shall accept a command to dump a software table.

c) Memory Dump: The Star Sensor System shall accept a command to dump memory.

d) Software Table Update: The Star Sensor System shall accept a command to update a software table.

e) Software Update: The Star Sensor System shall accept a command to update software.

f) Operating Mode: The Star Sensor System shall accept a command to change the unit’s operating mode.

g) Memory Checksum: The Star Sensor System shall accept a command to generate a memory checksum.

h) Image Acquisition: The Star Sensor System shall accept a command to download an optical head image.

i) Time Message: The Star Sensor System shall accept a time at tone command message

400-900 (TBR) milliseconds prior to the 1PPS pulse.

4.7.2 Data

The Star Sensor System shall provide the following data, at a minimum, in telemetry packets that include packet identification, packet length and a time stamp.

a) Attitude Quaternions: The Star Sensor System shall provide three-axis inertial attitude quaternion solutions for each optical head corrected for systematic errors relative to the sensor axes and the star magnitude (from the ECI J2000 frame to the star sensor frame).

b) Attitude Validity: The Star Sensor System shall provide telemetry that indicates the validity of the attitude solution.

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c) Attitude Quality: The Star Sensor System shall provide telemetry that indicates the quality of the attitude solution.

d) Angular Rate Vector: The Star Sensor System shall provide a three-axis time-tagged angular rate vector solution.

e) Number of Objects Imaged: The Star Sensor System shall provide telemetry that indicates the number of objects detected in its current image.

f) Housekeeping Data: The Star Sensor System shall provide telemetry that provides data regarding the system’s housekeeping status, such as temperatures, voltages and currents.

4.8 FLIGHT SOFTWARE

4.8.1 Flight Software and Star Catalog

The Star Sensor System shall host its flight software and star catalog.

4.8.2 Software Version Identification

The Star Sensor System flight software shall be uniquely identifiable (i.e., be able to verify SW version).

4.8.3 Software Updates

The Star Sensor System shall have the capability to update flight software during integration and test phases and in-flight operations.

4.8.4 Software Table Updates

The Star Sensor System shall have the capability to update flight software tables during integration and test phases as well as during flight operations.

4.8.5 Non-Volatile Memory Storage of Flight Code

Flight software executable code shall reside in non-volatile storage devices that can be reprogrammed during flight operations.

4.8.6 Boot Code

Boot (Start-up or Kernel) code shall reside in non-volatile storage devices that are not alterable in flight.

4.8.7 Star Catalog Version

The Star Sensor System flight star catalog version shall be uniquely identifiable.

4.8.8 Software Table Dump

The Star Sensor System shall provide the capability to dump a software table in a telemetry dataset.

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4.8.9 Packet Checksum

The Star Sensor System shall provide the capability to compute the checksum for a packet.

4.8.10 Memory Load

The Star Sensor System shall provide the capability to load memory.

4.8.11 Memory Dump

The Star Sensor System shall provide the capability to dump memory in a telemetry dataset.

4.8.12 Memory Checksum

The Star Sensor System shall provide the capability to compute the checksum for a specified memory region.

4.8.13 Software Instructions

Instructions for software loads shall be provided to the PACE COR for software loads at I&T and on-orbit operation.

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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: Component 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.

5.1.3 Data Interface

The data interface shall be defined in the Data Interface Control Document, which will be developed between the contractor and NASA GSFC.

5.2 MASS PROPERTIES

5.2.1 Component Masses

a) The combined mass of the Star Sensor optical heads and baffles on the instrument shall be less than or equal to 3.0 kg (not including harnessing).

b) The combined mass of the Star Sensor optical heads, baffles, and processing electronics on the spacecraft body shall be less than or equal to 7.0 kg (not including harnessing).

c) The total mass of the Star Sensor System shall be measured to within 25 g.

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5.2.2 Center of Mass Location

The contractor shall define the center of mass of each component of the Star Sensor System in the Mechanical ICD.

5.2.3 Center of Mass Accuracy

The center of mass of each component of the Star Sensor System shall be determined to within

±3.0 mm relative to an external reference.

5.2.4 Determination of Moments and Products of Inertia

The final moments and products of inertia of the Star Sensor System shall be calculated to within

10%.

5.3 PHYSICAL ENVELOPE

The Star Sensor System shall not exceed the internal and external envelopes defined below:

a) Processing Electronics Volume: 180 x 180 x 110 mm high

b) Spacecraft Body Optical Head with Baffle Volume: 250 x 250 x 500 mm high

c) Instrument Optical Head with Baffle Volume: 155 x 155 x 310 mm high

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 Star Sensor System shall have a flatness of 0.25 mm from any 2 points on the mounting interface for a component <0.3 m2 and 0.50 mm from any 2 points on the mounting interface for a component >0.3 m2 and <0.6 m2.

5.5 FIELDS OF VIEW

All Star Sensor System fields of view shall be as shown in the MICD.

5.5.1 Sun Exclusion Angle

Each star sensor shall maintain full performance when the Sun is greater than or equal to 39 degrees from the optical head boresight.

5.5.2 Moon and Earth Exclusion Angle

Each star sensor shall maintain full performance when the Moon or the Earth is greater than or equal to 23 degrees from the optical head boresight.

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5.6 ALIGNMENT

5.6.1 Alignment Location/Orientation Accuracy

The Star Sensor System alignment error between the alignment reference and the boresight of each optical head shall be less than 30 arc-seconds.

5.6.2 Alignment Knowledge Accuracy

The Star Sensor System angular alignment knowledge of the optical head boresight axis with respect its mounting surface shall be known to an accuracy of 20 arc-seconds (3 sigma) or better about each axis. This includes measurement accuracy and shifts due to the launch environment, but not long term temperature changes which will be calibrated out.

5.6.3 Alignment Stability

a) Throughout environmental testing, the optical axes shall not shift with respect to the alignment reference by more than 30 arc-seconds (3 sigma) in each axis.

b) The alignment reference shall be measured before and after environmental testing.

5.6.4 Alignment Method

Proper fit/alignment of the Star Sensor System 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.

5.6.5 Alignment Measurement Tooling/Features

a) Two adjacent surfaces of the optical reference shall be available for alignment measurements without affecting the basic integrity of the Star Sensor System.

b) Optical reference surfaces shall be at least 15.24 x 15.24 mm

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6.0 ENVIRONMENTAL REQUIREMENTS

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

The Star Sensor System 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 Star Sensor System, 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 Star Sensor System 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 Star Sensor System 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 Star Sensor System Design Limit Loads

6.3 FREQUENCY REQUIREMENT

6.3.1 Fundamental Launch Frequencies

The Star Sensor System shall have a fundamental frequency greater than 50 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

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results up to 50 Hz, to be 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 Star Sensor System 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 Star Sensor System interface. See Section 11.4.1 for definitions of

Protoflight, Qualification, and Acceptance.

Table 6-3 Star Sensor System 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 Star Sensor System 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 Star Sensor

System interface.

Table 6-4 Star Sensor System 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

This environment will be updated with random vibration analysis. Note for lightweight Star

Sensor System, 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 Star Sensor System shall be designed to meet its performance requirements after being subjected to the shock environment in b)

c) Table 6-5, applied at the Star Sensor System interface to the PACE spacecraft structure.

Table 6-5 Qualification Level Shock Response Spectrum

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Freq (Hz) SRS (G)

100 100

800 1700

8000 1700

10000 2200

d) 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 6-2), then the shock environment can be considered benign and there is low risk in deferring the shock test to the Observatory level.

Figure 6-2 Shock Envelope

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

f) 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).

6.6 ACOUSTICS

a) The Star Sensor System shall be designed to withstand, without any damage or degradation of performance, the equivalent Protoflight levels shown in Table 6-7.

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Table 6-6 Flight/Acceptance Level Acoustic Environments

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Table 6-7 Protoflight/Qualification Level Acoustic Environments

b) An acoustic test shall be performed unless an assessment of the component indicates that is not susceptible to the expected acoustic environment, or responses are enveloped by random vibration testing, or that testing at higher levels of assembly provides sufficient exposure at an acceptable level of risk to the program as determined by the NASA/GSFC

COR.

c) Component acoustic tests shall be to Protoflight levels (Acceptance + 3 dB) with a duration of 1 minute.

6.7 TRANSPORTATION

In addition to the launch loads shown above, the Star Sensor System shall also be designed to withstand the maximum transportation loads shown in Table 6-8 without damage or degradation of performance.

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Table 6-8 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 Star Sensor System 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 Star Sensor System 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…