Attachment_B_-_SPEC.pdf

Effective Date: October 16, 2018

Expiration Date: October 16, 2023

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

PACE-SC-SPEC-0084, Revision B

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

PACE Spacecraft Inertial Reference Unit

(IRU) Specification

Goddard Space Flight Center

Greenbelt, Maryland

National Aeronautics and

Space Administration

GSFC PACE CMO

10/16/2018

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

PACE IRU Spec PACE-GNC-SPEC-0084, Revision B ii

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

PACE Spacecraft Inertial Reference Unit (IRU)

Signature/Approval Page

Prepared By:

Agbontaen Imasuen

Reviewed By:

John Blackwood

Gary Davis

Beth Weinstein

Paul Mason

Craig Stevens

Eric Rogstad

Approved By:

Andre’ Dress

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

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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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Change History Log

Revision Effective

Date

Description of Changes

(Reference the SCoRe & Approval Date)

Revision - 05/16/2018 Baseline Release following the approval of PACE-CCR-0321

Revision A 06/15/2018 Updated following the approval of PACE-CCR-0374

Revision B 10/16/2018 Updated following the approval of PACE-CCR-0442 v

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Table of TBDs/TBRs/TBSs

Action Item

No.

Location Summary Individual/

Organization

Actionee vi

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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 IRU Description

4.0 Functional/Performance Requirements

4.1 IRU Flight Unit Functional/Performance Requirements

4.1.1 Description

4.1.2 Survival Rates

4.1.3 Performance Rates

4.1.4 Polarity

4.1.5 Cross Axis Coupling

4.1.6 IRU Performance

4.1.6.1 Scale Factor Resolution

4.1.6.2 Scale Factor Stability

4.1.6.3 Bandwidth

4.1.6.4 Rate Bias

4.1.6.5 Sensor Noise

4.2 Resource Allocations

4.2.1 Mass Allocation

4.2.2 Nominal Power Allocation

4.2.3 Peak Power Allocation

4.2.4 Telemetry

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

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

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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 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

5.0 Physical Requirements

5.1 Interface Documentation

5.1.1 Mechanical Interface

5.1.2 Electrical Interface

5.2 Mass Properties

5.2.1 Component Masses

5.2.2 Center of Mass Location

5.2.3 Center of Mass Accuracy

5.2.4 Determination of Moments and Products of Inertia

5.3 Physical Envelope

5.4 Mounting

5.5 NA

5.6 Alignment

5.6.1 Alignment Reference Surfaces

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 NA

6.7 Transportation

6.7.1 Transportation Cleanliness

6.8 Pressure

6.8.1 Operating Pressure Range

6.8.2 Maximum Depressurization Rate

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6.9 On-Orbit Dynamic Environment

6.9.1.1 Dynamic Linear Acceleration

6.9.1.2 Dynamic Angular Acceleration

6.10 Ground Environments

6.11 Thermal Requirements

6.11.1 Flight Interface Design Temperature Limits

6.12 Charged Particle Radiation Requirements

6.12.1 Definitions

6.12.2 Total Ionizing Dose

6.12.2.1 Minimum TID Tolerance for EEE Parts and Materials

6.12.2.2 ELDR Evaluation

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)

6.12.5 Charging Environment

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

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

8.2 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 – Differential Mode Current Emission Limits

8.4.1.3 CE102 – Differential Mode Current Emission Limits

8.4.1.4 Common Mode Current Emissions 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 NA

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

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

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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 NA

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

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

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11.5.2 Mass Properties Measurement

11.5.3 Static Loads/Strength Test

11.5.3.1 Sine Burst

11.5.3.2 Static Pull

11.5.4 Sine Sweep Survey

11.5.5 Sine Vibration

11.5.6 Random Vibration

11.5.7 NA

11.5.8 NA

11.5.9 Shock

11.5.10 Thermal Vacuum Bake-out

11.5.11 Thermal Vacuum Test

11.5.11.1 Thermal Vacuum Test Parameters

11.5.11.2 Thermal Vacuum Test Profile

11.5.12 NA

11.5.13 NA

11.5.14 NA

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 Differential Mode Current Emission Limits

Figure 8-2. Common Mode Current Emission 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. IRU Design Limit Loads

Table 6-3. IRU Sine Vibration Environment

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

Table 6-6. Qualification Level Shock Response Spectrum

Table 6-9. Transportation Loads

Table 6-10. Temperature Limits at Box Mounting Interface

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

Table 6-12. 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

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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 IRU for the NASA Goddard Space Flight Center (GSFC)

PACE Mission.

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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

IRU 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 IRU

Statement of Work (PACE-GNC-SOW-0049), in which case the Statement of Work takes precedence.

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

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

Table 2-1. Applicable Documents

Document Number Title

PACE-GNC-SOW-0049 PACE IRU Statement of Work

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

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

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Document Number Title

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

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3.0 CONTRACT DESCRIPTION

3.1 IRU DESCRIPTION

The IRU provides three axis rate information to the PACE Guidance, Navigation and Control

(GNC) System. PACE is an orbiting spacecraft in a circular sun-synchronous polar orbit at 676.5 km altitude with an inclination of 98 degrees and a 1 pm local ascending node crossing time.

PACE is currently targeted to launch in the 2022-2023 timeframe.

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4.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the IRU as defined in

Section 3.1.

4.1 IRU FLIGHT UNIT FUNCTIONAL/PERFORMANCE REQUIREMENTS

4.1.1 Description

The IRU shall provide three axis rate output information based on the use of flight qualified high-grade inertial sensors. The IRU is intended for use as a high accuracy, fine resolution attitude reference.

4.1.2 Survival Rates

The IRU shall survive, and meet all of this specification's requirements, after exposure to angular rates up to, and including, 10 degrees per second.

4.1.3 Performance Rates

After the removal of the Survival input rates specified above, and rates less than or equal to 3 degrees per second, the IRU shall operate within full performance within 3 minutes.

4.1.4 Polarity

The IRU shall output the proper polarity for rates to at least 10 degrees per second.

Dual rate range solutions are acceptable.

4.1.5 Cross Axis Coupling

a. The cross-axis coupling for rates within 3 degrees per second shall be less than 0.25% including gyro intra-axis misalignment.

b. The bandwidth for this requirement may be limited to 1 Hz. Over the total bandwidth of

10 Hz, the cross-axis coupling shall not exceed 2%.

c. The cross axis coupling shall be stable within 0.05% of the input rate during continuous operation up to 30 days.

4.1.6 IRU Performance

The following requirements apply to the outputs when operating within +/- 3 degrees per second input rates.

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4.1.6.1 Scale Factor Resolution

a. The IRU data shall have a resolution less than or equal to 0.25 arcseconds per second.

b. The scale factor shall be measured prior to delivery to an accuracy of 0.1%.

4.1.6.2 Scale Factor Stability

a. The scale factors of the outputs shall be stable within 0.05% (3 sigma) per 30 days of the nominal scale factor.

b. This stability requirement is based on continuous operation, and shall apply over the

Survival and Performance Rates specified above.

4.1.6.3 Bandwidth

Signal bandwidth in all modes shall be greater than or equal to 10 Hz to the -3 dB point.

4.1.6.4 Rate Bias

The Rate Bias shall comply with the performance requirements in this section (4.1.6.4).

4.1.6.4.1 Absolute Value

The absolute value of the Rate Bias shall be less than or equal to 7.0 arcseconds per second using a 10-minute sampling period.

4.1.6.4.2 Short Term Rate Bias Stability

The Short Term Rate Bias shall be stable to within 0.03 arcseconds per second (3 sigma) for 8 hours of continuous operation.

4.1.6.4.3 Long Term Rate Bias Stability

The Long Term Rate Bias shall be stable to within 5.0 arcseconds per second (3 sigma) for 30 days of start/stop operation. Compliance may be shown by comparison of pre and post environmental data.

4.1.6.4.4 Environment

a. The Rate Bias shall be stable to within 2.0 arcseconds per second (3 sigma) through each environment.

b. This change shall be determined from a comparison of the pre and post environmental test data including data taken at hot, cold, and nominal temperatures.

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4.1.6.5 Sensor Noise

The requirements below (4.1.6.5.1, 4.1.6.5.2, 4.6.5.3 and 4.1.6.5.4) shall be verified via an 8 hour performance test.

4.1.6.5.1 Quantization

Quantization shall be less than or equal to 0.25 arcseconds per count.

4.1.6.5.2 Angle Random Walk

Angle Random Walk (ARW) shall be less than or equal to 0.005 degrees per root-hour (1 sigma)

4.1.6.5.3 Rate Random Walk

Rate Random Walk (RRW) shall be less than or equal to 0.005 degrees per hour per root-hour (1 sigma).

4.1.6.5.4 Angle White Noise

Angle White Noise (AWN) shall be less than or equal to 0.2 arc-second (1 sigma).

4.2 RESOURCE ALLOCATIONS

4.2.1 Mass Allocation

The IRU shall have a mass of less than or equal to 5 kg.

4.2.2 Nominal Power Allocation

The IRU shall have a nominal (continuous) power consumption of less than or equal to 33 W for a duration of 100 minutes.

4.2.3 Peak Power Allocation

The IRU shall have a peak (instantaneous) power consumption of less than or equal to 45W for a duration of 60 seconds.

4.2.4 Telemetry

The housekeeping telemetry rate of the IRU shall be defined in the Electrical ICD.

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4.3 POWER

4.3.1 Primary (Unregulated) Power Input Requirements

4.3.1.1 Operating Voltage Range

The IRU 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 IRU 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.

4.3.1.3 Voltage Ripple

The IRU 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 IRU 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 IRU shall survive without performance degradation after an abrupt, unannounced removal of power.

4.3.1.6 Polarity Reversal Protection

The IRU 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 IRU 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

IRU shall provide a dedicated Primary Power return in the same connector as the primary power.

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4.3.1.9 Power Wiring Redundancy

The IRU 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 2 amps.

Figure 4-1. Inrush Current Limits

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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 IRU 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.

4.3.2.3 Repetitive Transients

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

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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 IRU 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 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 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 IRU shall be a MIL-STD-1553B interface that accommodates the IRU command and telemetry signals.

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

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

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

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

5.1.1 Mechanical Interface

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

5.1.2 Electrical Interface

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

5.2 MASS PROPERTIES

5.2.1 Component Masses

The mass of the IRU shall be less than or equal to 5 kg.

5.2.2 Center of Mass Location

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

5.2.3 Center of Mass Accuracy

The center of mass of the IRU shall be determined to within ±2.5 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 IRU shall be calculated to within 10%.

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5.3 PHYSICAL ENVELOPE

The IRU outside dimensions, including mounting flanges and connectors, shall not exceed the volume envelope of 50 cm length x 25 cm width x 28 cm height as defined in the Mechanical

ICD.

5.4 MOUNTING

a. The package will be hard-mounted on a mechanical surface of the spacecraft structure.

Proper fit/alignment of the IRU 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. 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.

c. The IRU 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.

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5.6 ALIGNMENT

5.6.1 Alignment Reference Surfaces

Permanent optical reference surfaces that permit alignment on the spacecraft to the mechanical coordinate frame shall be provided on the IRU.

5.6.2 Alignment Knowledge Accuracy

Knowledge of the sensor axes alignment to the optical or mechanical reference shall be to 10 arc seconds or better in each axis.

5.6.3 Alignment Stability

Throughout environments, the sensor axes shall not shift with respect to the optical or mechanical reference by more than 30 arc seconds in each axis.

5.6.4 Alignment Method

a. Proper fit/alignment of the IRU 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.

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

b. The contractor shall provide a drill template with alignment cube, as required, to meet any alignment requirements.

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 IRU.

b. Optical reference surfaces shall be at least 0.6 x 0.6 inches.

c. Optical reference surfaces shall be capable of producing an auto collimation indication.

d. ABSOLUTE ALIGNMENT: The electrical axes shall be aligned to the mechanical axes within 10 arcminutes.

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

6.0 ENVIRONMENTAL REQUIREMENTS

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

The IRU 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 IRU, 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 IRU 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 IRU 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. IRU Design Limit Loads

6.3 FREQUENCY REQUIREMENT

6.3.1 Fundamental Launch Frequencies

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

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

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

SOW and discussed in Section 11.5.4.

6.4 VIBRATION

6.4.1 Sinusoidal Vibration

The IRU 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 IRU interface. See Section 11.4.1 for definitions of Protoflight, Qualification, and

Acceptance.

Table 6-3. IRU 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

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

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.

6.4.2 Random Vibration

a. The IRU 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 IRU interface.

Table 6-4. IRU 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 IRU, 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 IRU shall be designed to meet its performance requirements after being subjected to the shock environment in

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b. Table 6-5, applied at the IRU interface to the PACE spacecraft structure.

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

Table 6-5. Qualification Level Shock Response Spectrum

Freq (Hz) SRS (G)

100 100

800 1700

8000 1700

10000 2200

c. 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

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

e. 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).

f. Component self-induced shock testing shall be accomplished by two actuations at the component level for each self-induced shock source (in order to account for the scatter

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

associated with the actuation of the device) for the first flight unit, and a single actuation on subsequent units.

6.6 NA

6.7 TRANSPORTATION

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

Table 6-6. 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 IRU 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 IRU 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.

6.9 ON-ORBIT DYNAMIC ENVIRONMENT

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

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

6.9.1.1 Dynamic Linear Acceleration

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

6.9.1.2 Dynamic Angular Acceleration

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

6.10 GROUND ENVIRONMENTS

a. The IRU 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 IRU shall meet all of its performance requirements.

6.11 THERMAL REQUIREMENTS

6.11.1 Flight Interface Design Temperature Limits

a. When powered “OFF”, the IRU shall be capable of surviving indefinitely when its temperatures are within the survival limits shown in Table 6-7 without damage or permanent performance degradation.

b. The IRU shall meet all performance requirements when powered “ON” anywhere within the Operational and Protoflight/Qualification limits shown in Table 6-7.

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

Protoflight/Qualification (in spec) limits shown in Table 6-7.

Table 6-7. Temperature Limits at Box Mounting Interface

Minimum Temperature (ºC) Maximum Temperature (ºC) Operational (In Spec) -10 +40

Protoflight/Qualification (In Spec) -20 +50

Survival (Unpowered) -25 +55

6.12 CHARGED PARTICLE RADIATION REQUIREMENTS

Components containing electronic parts will be exposed to a natural space radiation environment that consists of: (1) trapped particles which include electrons, protons, and heavier ions; (2) particles from solar events (coronal mass ejections and flares); and (3) galactic cosmic ray particles.

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

6.12.1 Definitions

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

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

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

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

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

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

(SEDRs).

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

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

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

An example is a stuck bit in a memory device.

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

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-

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

initialization to resume normal device operations, but, for some devices, a power cycle is necessary to resume normal device operations. The general goal for non-destructive events such as SEFI is to avoid them, however, managing their impact through robust circuit design, automatic correction, and/or operational activities may be considered.

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

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

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

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

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

107 ions/cm2.

6.12.2 Total Ionizing Dose

6.12.2.1 Minimum TID Tolerance for EEE Parts and Materials

The top-level total ionizing dose requirement is shown in Figure 6-3 Error! Reference source not found.and

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

Table 6-8. The dose values are calculated as a function of aluminum shield thickness in units of krad in silicon. For a nominal 2.54 mm (100 mils) of equivalent aluminum shielding and a 3-year mission life, the expected dose is 6.67 krad-Si. This includes a factor of 2 margin. EEE parts and materials shall be selected according to the level of shielding shown in Figure 6-3 Error!

Reference source not found.and

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Table 6-8.

Figure 6-3. Total Ionizing Dose-Depth Curve (includes x2 margin)

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Table 6-8. Dose (including x2 margin) as a Function of Shielding

6.12.2.2 ELDR Evaluation

Linear bipolar and BiCMOS device types that are known to be, or potentially susceptible to

Enhanced Low Dose Rate Sensitivity (ELDRS) for TID, shall be evaluated using test data at a low dose rate not to exceed 0.01 rad(Si)/sec in accordance with the latest version of MIL-STD-883, Test Method 1019 for all bias cases to be used in the Spacecraft Observatory equipment (i.e., biased and unbiased), unless existing data shows that a lower dose rate is required to manifest the

ELDRS effect as determined by the NASA/GSFC COR.

6.12.3 Displacement Damage Dose

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

Dose

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

0.002935 0.115532 0.000792 7.98E+03

0.004039 0.159003 0.00109 5.56E+03

0.005521 0.217353 0.00149 3.96E+03

0.007596 0.299042 0.00205 2.94E+03

0.010412 0.409907 0.00281 2.28E+03

0.014302 0.563075 0.00386 1.74E+03

0.019601 0.771675 0.00529 1.34E+03

0.0269 1.059047 0.00726 1.05E+03

0.036941 1.454366 0.00997 8.06E+02

0.050761 1.998477 0.0137 6.12E+02

0.069658 2.742436 0.0188 4.54E+02

0.095595 3.763556 0.0258 3.38E+02

0.131165 5.163948 0.0354 2.58E+02

0.180073 7.089488 0.0486 1.91E+02

0.246767 9.715225 0.0666 1.38E+02

0.339027 13.34749 0.0915 9.40E+01

0.466857 18.38015 0.126 6.32E+01

0.637297 25.09037 0.172 4.16E+01

0.87443 34.42632 0.236 2.66E+01

1.204194 47.40913 0.325 1.70E+01

1.64882 64.91404 0.445 1.10E+01

2.263885 89.12916 0.611 7.28E+00

3.108674 122.3885 0.839 5.40E+00

4.260995 167.7554 1.15 4.14E+00

5.854237 230.4813 1.58 3.22E+00

8.040313 316.5471 2.17 2.54E+00

11.04154 434.7053 2.98 2.06E+00

15.15432 596.6257 4.09 1.65E+00

20.78625 818.3545 5.61 1.32E+00

Aluminum Shield Thickness

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

ionizing dose except that the energy considered produces displacements in a semiconductor lattice.

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

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

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

6.12.4 Single Event Effects

a. The IRU shall be designed to avoid or tolerate errors due to non-destructive Single Event

Effects (SEE).

b. The integral galactic cosmic ray and solar heavy ion linear energy transfer spectrum in

Figure 6-4 (Table 6-9) and Figure 6-5 (Table 6-10), respectively, shall be used for prediction of heavy ion-induced single-event effects.

Note: CREME96 may be used to generate the environments, provided that the

CREME96 calculations are shown to match the referenced figures and tables.

c. The IRU shall meet the PACE Observatory performance requirements for the background environment.

d. The IRU shall survive the peak proton flux of a worst-case solar flare.

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Figure 6-4. Integral LET Spectra for Galactic Cosmic Ray Ions

Table 6-9. Integral LET Spectra for Galactic Cosmic Ray Ions

LET Integral Flux LET Integral Flux

MeV-cm2/mg particles/cm2/day MeV-cm2/mg particles/cm2/day

0.101035 770.3394291 3.59474 2.345403496

0.131919 479.0803133 4.48083 1.425135001

0.164436 401.7760225 5.58533 0.857893053

0.204969 320.9865242 6.96209 0.511771777

0.255493 252.4658249 8.67822 0.305525666

0.31847 202.0551758 10.8174 0.180199342

0.396972…