DOORS EXPORT Effective Date: TBD Expiration Date: TBD

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GSFC SOLAR Project

CMO

January 9, 2026

DRAFT

SOLAR-XRI-REQ-0029, Revision -

Space weather Observations at L1 to Advance Readiness (SOLAR), Code 493.0

Space weather Observations at L1 to

Advance Readiness (SOLAR)

X-Ray Irradiance Sensor (XRIS)

Requirements Specification (SPEC)

U.S. Department of Commerce (DOC)

National Oceanic and Atmospheric Administration (NOAA)

National Environmental Satellite, Data, and Information Service (NESDIS)

National Aeronautics and Space Administration (NASA) https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm

SOLAR Project XRIS SPEC SOLAR-XRI-REQ-0029, Revision - Effective Date: TBD ii Check https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm to verify that this is the correct version prior to use.

SOLAR X-Ray Irradiance Sensor (XRIS) Requirements

Specification (SPEC)

Review/Signature/Approval Page

Prepared by:

E. Matt Glanfield Date SOLAR XRIS Instrument Engineer SW Next SOLAR NASA Goddard Space Flight Center

Reviewed by:

C. Hudson DeLee Date SOLAR Instrument Systems Manager SW Next SOLAR NASA Goddard Space Flight Center

Approved by:

Jim Morrissey Date SOLAR Project Manager NASA/GSFC, Code 493

Electronic Approval available on-line at: IPD TDMS -Instruments Project Division TDMS (nasa.gov) iii Check https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm to verify that this is the correct version prior to use.

Preface

This document is under SWO Program configuration control. Once this document is approved, SWO approved changes are handled in accordance with Class I and Class II change control requirements as described in the SWO Configuration Management Procedure, and changes to this document shall be made by complete revision.

In this plan, all mandatory actions (i.e., requirements) are denoted by statements containing the term “shall.” The terms “may” or “can” denote discretionary privilege or permission; “should” denotes a good practice and is recommended but not required; “will” denotes expected outcome;

and “are/is” denotes descriptive material.

Any questions should be addressed to:

SOLAR Configuration Management Office

Mail Stop: 493

Goddard Space Flight Center

Greenbelt, MD 20771 iv Check https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm to verify that this is the correct version prior to use.

Change History Log

Change History Log will be provided by CM prior to release v Check https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm to verify that this is the correct version prior to use.

Deviations/Waivers Record

Section # /

Rqmt.

Deviation /

Waiver #

CCR # Date

Approved

Description

None. None. None. None. None.

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Table of TBCs/TBDs/TBRs/TBSs

Item

No.

Identifier Section Summary vii Check https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm to verify that this is the correct version prior to use.

Table of Contents

1 INTRODUCTION

1.1 General Information

1.2 Scope

2 APPLICABLE DOCUMENTS

3 CONTRACT DESCRIPTION

3.1 X-ray Irradiance Sensor Instrument

3.2 Ground Support Equipment Description

4 FUNCTIONAL/PERFORMANCE REQUIREMENTS

4.1 XRIS Performance Requirements

4.1.1 Solar X-Ray Irradiance Measurement Products

4.1.1.1 X-Ray Irradiance Band Ranges

4.1.1.2 X-Ray Irradiance Flux Ranges

4.1.1.3 X-ray Irradiance Accuracy

4.1.1.4 X-ray Flare Location

4.1.2 Refresh Rate and Latency

4.1.3 Field of View (FOV)

4.2 Resource Allocations

4.2.1 Mass Allocation

4.2.2 Physical Envelope

4.2.3 Operational Power Allocation

4.2.4 Peak Operational Power Allocation

4.2.5 Survival Heater Power Allocation

4.2.6 Telemetry

4.2.7 Transient Magnetic Field

4.2.8 Static Magnetic Field

4.3 XRIS Operating Modes

4.3.1 In-Flight Calibration

4.3.2 Safe Mode

4.3.3 Normal Operational Mode

4.3.4 On-Orbit Operations

4.3.5 Activation

4.4 Instrument Data Availability

4.5 Power

4.5.1 Voltage Range

4.5.2 Abnormal Voltages

4.5.3 Turn-on Current Transients

4.6 Electrical Grounding

4.6.1 Primary Power DC Isolation

4.6.2 Survival Power Isolation

4.6.3 Mechanical Contact Resistance

4.6.4 Grounding

4.6.5 Connector DC Resistance

4.7 Signal And Data Interfaces

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4.7.1 Passive Analog Telemetry

4.7.2 Data Signal Interface

4.7.2.1 Telemetry Source Packet Format

4.7.2.2 Command Source Packet Format

4.7.3 Clock Signal Interface

4.7.4 Command and Housekeeping telemetry

4.7.5 Commands for Autonomous Functions

4.8 Flight Software

4.8.1 Flight Load Non-volatile Memory

4.8.2 Software Updates

4.8.3 Software Table Updates

4.8.4 Flexibility and Ease of Software Modification

4.8.5 Version Identifiers

4.8.6 Warm Restart

4.8.7 Memory Tests

4.8.8 Memory Dump

4.8.9 Fault Detection and Correction Control

5 PHYSICAL REQUIREMENTS

5.1 Mass Properties Accuracy

5.1.1 Instrument Mass Accuracy

5.1.2 Center of Mass Accuracy

5.2 Mounting

5.3 Coordinate System and Alignment

6 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 (LV) Environmental Control System (ECS) Impingement

6.9 On-Orbit Dynamic Environment

6.9.1 Dynamic Linear Acceleration

6.9.2 Dynamic Angular Acceleration

6.10 Thermal Requirements

6.10.1 Flight Interface Design Temperature Limits

6.11 Charged Particle Radiation Requirements

6.11.1 Definitions

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6.11.2 Radiation Hardness Assurance

6.11.3 Charging Environment

6.11.3.1 Charging Environments During Magnetic Substorm

6.11.3.1.1 Radiation Belt Geomagnetic Sub-Storm Surface Charging

6.11.3.1.2 Radiation Belt Geomagnetic Sub-Storm Bulk Charging (Energetic

Particle) Environment

6.11.3.2 L1 Surface Charging Environment

6.11.3.3 Solar Photon Environment

6.11.4 Total Ionizing Dose

6.11.5 Displacement Damage Dose Levels

6.11.6 Single Event Effects

6.11.7 Solar Cycle Performance

6.11.8 Solar Particle Event Operations

6.11.9 Solar Flare Survival

7 CONTAMINATION CONTROL

7.1 Surface Cleanliness Requirements

7.2 Contamination Generation

7.2.1 Particulate Generation

7.2.2 Molecular Contamination Generation

7.2.2.1 Molecular Material Restrictions

7.3 Vacuum Bakeouts

7.4 Venting Requirements

7.5 Reserved

7.6 Cleanability and Protection

8 DESIGN AND CONSTRUCTION REQUIREMENTS

8.1 Electrical

8.1.1 Connector Specifications

8.1.2 Test Interfaces

8.1.2.1 Spacecraft Charging Limits

8.2 Safety

8.3 Electromagnetic Compatibility

8.3.1 Conducted Emissions

8.3.1.1 Applicability of Conducted Emissions

8.3.1.2 CE101 – Differential Mode Conducted Emissions Limits

8.3.1.3 CE03 – Differential Mode Conducted Emissions Limits

8.3.1.4 CMBCE – Common Mode Bulk Conducted Emissions Limits

8.3.2 Conducted Susceptibility

8.3.2.1 Applicability of Conducted Susceptibility

8.3.2.2 CS101 – Conducted Susceptibility Limit

8.3.2.3 CS114 – Conducted Susceptibility Differential Mode Limit

8.3.2.4 CS114 – Conducted Susceptibility, Common Mode, Interconnected Power and

Signal Cables

8.3.2.5 CS06 – Conducted Susceptibility, Transients Limits

8.3.2.6 CS115 - Conducted Susceptibility, Bulk Cable Injection, Impulse Excitation ... 72

8.3.3 Radiated Emissions

8.3.3.1 RE102 - Electric Field Emissions Limits

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8.3.4 Radiated Susceptibility

8.3.4.1 RS103 - Radiated Susceptibility, Electric Field, Launch Limit

8.3.4.2 RS103 - Radiated Susceptibility, Electric Field, On-Orbit Limit

8.4 Identification and Marking

8.5 Workmanship

8.5.1 Connectors

8.6 Reliability and Mission Lifetime

8.6.1 Mission Life

8.6.2 Operating Time

8.6.3 Trouble-Free Time

8.7 Ground Handling

8.7.1 GSE Cleanliness

8.7.2 GSE Bakeout

8.7.3 Test Harness

9 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 Factors of Safety

9.3 Mechanisms

10 LOGISTICS

10.1 Ground Support Equipment

10.1.1 Electrical System Test Equipment

10.1.2 Instrument Emulator

10.2 Transportation Equipment

10.2.1 Shipping Container

11 VERIFICATION REQUIREMENTS

11.1 Test Requirements

11.1.1 Definitions

11.1.2 Test Factors

11.1.3 Test Tolerances

11.1.4 Test Restrictions

11.1.4.1 Failure During Tests

11.1.4.2 Modification of Hardware

11.1.4.3 External Adjustment

11.1.4.4 Re-Test Requirements

11.1.4.5 Configuration

11.2 Required Tests

11.2.1 Performance Tests

11.2.1.1 Comprehensive Performance Test

11.2.1.2 Limited Performance Test

11.2.1.3 Abbreviated Functional Test

11.2.2 Mass Properties Measurement

11.2.3 Static Loads/Strength Test

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11.2.3.1 Sine Burst

11.2.3.2 Static Pull

11.2.4 Sine Sweep Survey

11.2.5 Sine Vibration

11.2.6 Random Vibration

11.2.7 Acoustic Test

11.2.8 Shock

11.2.9 Thermal Vacuum Bake-out

11.2.10 Thermal Vacuum Test

11.2.10.1 Thermal Vacuum Test Parameters

11.2.10.2 Thermal Vacuum Test Profile

11.2.11 Magnetics Test

APPENDIX A ABBREVIATIONS AND ACRONYMS

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List of Figures Figure 1. Telemetry Source Packet Definition Figure 2. Time Code Format Figure 3. Command Source Packet Figure 4. Spacecraft Time Message Packet Figure 5. Spacecraft Reference System Figure 6. Sine Vibration Environment Figure 7. Random Vibration Environment Figure 8. Candidate LV Shock Environment Figure 9. SOLAR Acoustic Design Envelope Figure 10. Allowable Analytical, Flight Operation, Flight Acceptance Test, and

Qualification (ProtoFlight or Prototype) Test Temperature Ranges Figure 11. Integral Electron Fluence vs Energy Figure 12. Direction Axes Used in L1 Plasma Flux Environments Figure 13. Total Ionizing Dose-Depth Curve Figure 14. Displacement Damage Dose in Silicon vs. Shielding Thickness Dose-Depth

Curve Figure 15. Displacement Damage Dose in Gallium Arsenide Dose-Depth Curve Figure 16. LET Spectrum of GCR Background Figure 17. Solar LET Spectrum Background Figure 18. Solar Proton Background Spectra Figure 19. Peak Solar Particle Event Flux Figure 20. Peak Solar Proton Flux Figure 21. CE101/CE03 Differential Mode Conducted Emissions Limits Figure 22. Common Mode Bulk Conducted Emissions Limit Figure 23. Power Lead Conducted Susceptibility (CS101) Voltage Limit Figure 24. Conducted Susceptibility Power Limit (30 Hz to 150 kHz) Figure 25. Conducted Susceptibility, Power Leads, 150 kHz to 50 MHz Limit Figure 26. Conducted Susceptibility, Power Leads, 150 kHz to 50 MHz Current Limits .. 70 Figure 27. Common Mode Conducted Susceptibility Calibration Limit Figure 28. Conducted Susceptibility Transient Waveform Figure 29. CS115 Impulse Waveform Figure 30. Unit Level RE102 Radiated Electric Field Emission Limits

List of Tables

Table 1. Applicable Documents Table 2. Factors of Safety1,2 Table 3. Design Limit Loads

Table 4. Candidate LV Shock Environment Table 5. SOLAR Acoustic Design Envelope Table 6. Transportation Loads Table 7. Orbit Thermal Parameters Table 8. Geo-magnetic Sub-storm Region Surface Charging Environment Model Parameters . 35 Table 9. Geomagnetic Sub-storm Bulk Charging (Energetic Electron) Environment xiii Check https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm to verify that this is the correct version prior to use.

Table 10. Plasma Moments for L1 Solar Wind Environment Surface Charging Table 11. Directionality for L1 Surface Charging Analysis

Table 12. Solar Photon Flux Density Table 13. Total Ionizing Dose vs. Shielding Thickness Table 14. Displacement Damage Dose in Silicon vs. Shielding Thickness Table 15. Displacement Damage Dose in Gallium Arsenide vs. Shielding Thickness Table 16. LET Spectrum of GCR Background

Table 17. Solar Background LET Spectrum Table 18. Solar Proton Background Flux Table 19. Abbreviate Peak Solar Particle Event Flux Table 20. Peak Solar Proton Flux Table 21. Lifetime Contamination Requirements

Table 22. Emission and Susceptibility Requirements Table 23. RS103 Radiated Susceptibility Levels Table 24. Test Factors and Durations

Table 25. Thermal Vacuum Test Parameters

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1 INTRODUCTION

1.1 General Information

1.2 Scope

This specification describes the performance and electrical, mechanical, environmental, and verification testing requirements for a space-qualified solar X-ray irradiance sensor for the Space

Weather Observations at L1 to Advance Readiness (SOLAR) Mission.

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2 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

XRIS to the extent specified herein. Where requirements that point to these standards provide alternatives, they will be accepted. In the event of conflict between this specification and any referenced document, this specification will govern, with the exception of the SOLAR XRIS

Statement of Work (SOLAR-XRIS-SOW-0007), in which case the Statement of Work takes precedence.

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

Table 1. Applicable Documents

Document Number Title

SOLAR-XRIS-SOW-0007 SOLAR X-ray Irradiance Sensor (XRIS) Instrument

Statement of Work (SOW)

SOLAR-XRIS-REQ-0031 SOLAR X-ray Irradiance Sensor (XRIS) Instrument Contract

Deliverables Requirements List (CDRL)

SOLAR-SYS-REQ-0016 Radiation Hardness Assurance Requirements Document

(RHARD)

ANSI/TIA/EIA-422 Revision -

B

Electrical Characteristics of Balanced Voltage Digital

Interface Circuits

ANSI/TIA/EIA-644 Revision -

A-2001

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-6016C Standard Materials and Processes Requirements for

Spacecraft

NASA-HDBK-7005 Dynamic Environment Criteria

NASA-STD-7001 Payload Vibroacoustic Test Criteria

IEST-STD-CC-1246E Product Cleanliness Levels and Contamination Control

Program

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

Volatile Condensable Materials from Outgassing in a

Vacuum Environment

MIL-DTL-5541 Chemical Conversion Coatings on Aluminum and Aluminum

Alloys

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

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

MIL-STD-461G Requirements for the Control of Electromagnetic Interference

Characteristics of Subsystems and Equipment

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Document Number Title

GSFC-STD-7000B General Environmental Verification Standard (GEVS) for

GSFC Flight Programs and Projects

NASA-STD-5019A Fracture Control Requirements for Spaceflight Hardware

NASA-STD-5020B Requirements for Threaded Fastening Systems in Spaceflight

Hardware

NASA-STD-5017B Design and Development Requirements for Mechanisms

NASA-HDBK-4002A Mitigating In-Space Charging Effects—A Guideline

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

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

Corrosion Cracking Resistance in Sodium Chloride

Environments

MIL-STD-462, Notice 4 Electromagnetic Interference Characteristics, Measurement of, 1 May 1970

ISO 14644-1:2015 Part 1: Classification of air cleanliness by particle concentration

ISO 14644-3: 2005 Cleanrooms and associated controlled environments -- Part 3:

Test methods

CCSDS 133.0-B-2 Space Packet Protocol, Blue Book, Issue 2, June 2020

CCSDS 301.0-B-4 Time Code Formats. Blue Book. Issue 3, January 2002

NA-GSFC-2005-05 Latent Damage Reliability Issues Due to Single-Event

Latchup and High-Current Error Modes

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3 CONTRACT DESCRIPTION

3.1 X-ray Irradiance Sensor Instrument

The X-ray Irradiance Sensor (XRIS) instrument is used to measure the energy power from solar radiation per unit area. This can be used to detect changes in solar activity like solar flares and help predict solar energetic particle events that can penetrate the Earth’s magnetic field. It is also important to understand the X-ray input into the Earth’s upper atmosphere. The instrument is essential in monitoring X-ray flares that can cause changes in the Earth’s ionosphere like strong radio bursts that may interfere with radio frequencies and cause problems for satellite communication and navigation systems. Coronal Mass Ejections (CMEs) are sometimes associated with solar flares which can lead to geomagnetic storms.

The XRIS contract includes the delivery of one Flight Models (FM) instrument, one spare Flight

Model instrument and one (1) Engineering Development Unit (EDU), flight harnesses between instrument boxes, if applicable, and spares, an electronics unit, and harnesses.

3.2 Ground Support Equipment Description

The XRIS contract includes the delivery of two (2) sets of the Electrical System Test Equipment

(ESTE), four (4) X-ray Irradiance Sensor Emulators (XRISE), and one (1) Ground Processing

Development System (GPDS). Additionally, Mechanical Ground Support Equipment (MGSE) such as lifting fixtures/handles, shipping containers, purge carts, drill templates, test fixtures, non-flight protective covers, etc. are included in the delivery. Electrical Ground Support

Equipment (EGSE) includes, but is not limited to, test cables and break-out boxes as required.

Ground Support Equipment is necessary to operate the instrument during spacecraft testing.

Supplying this equipment is considered part of the delivery for the instrument. More information is provided in the Contract Deliverable Requirement List (CDRL).

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4 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the SOLAR XRIS.

The requirements in this Specification pertain to the XRIS "system", which may include all instrument hardware, software, and ground processing algorithms. The XRIS contractor is not responsible for the operational ground system, but certain specifications will require ground processing after collection but before data distribution for which the contractor is responsible for defining and developing.

4.1 XRIS Performance Requirements

4.1.1 Solar X-Ray Irradiance Measurement Products

4.1.1.1 X-Ray Irradiance Band Ranges

XRIS-116: The XRIS shall measure solar X-ray irradiance through the band ranges of 0.05 – 0.4 nm and 0.1 – 0.8 nm.

4.1.1.2 X-Ray Irradiance Flux Ranges

XRIS-118: The XRIS shall measure solar X-ray irradiance with a flux range of 1x10-9–1x10-3

Wm-2 in the 0.05 – 0.4 nm band range and 1x10-8–1x10-3 Wm-2 in the 0.1 – 0.8 nm band range.

4.1.1.3 X-ray Irradiance Accuracy

XRIS-2642: The XRIS shall measure solar X-ray irradiance with a measurement accuracy of +/-

20%.

4.1.1.4 X-ray Flare Location

XRIS-2662: The XRIS shall provide the flare location for flares classification M1 and above

(above 10-5 W/m2) within a five (5) arcmin accuracy.

4.1.2 Refresh Rate and Latency

The SOLAR mission will make solar X-ray irradiance data available to SWPC forecasters within

30 seconds of the completion of each data acquisition.

XRIS-134: The XRIS shall have a refresh rate of the solar X-ray irradiance data acquisition and output of ≤ 3 sec.

XRIS-2663: The XRIS shall provide flare location measurement and output with a refresh rate of 60 seconds or less.

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XRIS-2664: The XRIS shall provide flare location measurements with a latency of 120 seconds or less.

XRIS-135: The processing of a completed on-orbit data collection by the instrument and the transmission to the spacecraft shall contribute ≤1 second to the overall latency of the X-ray irradiance and flare location data products.

XRIS-136: The XRIS ground processing algorithms contribution to data latency of the XRIS algorithm outputs shall be ≤ 3 seconds.

Note: This applies to both X-ray irradiance and flare location products.

4.1.3 Field of View (FOV)

XRIS-138: The XRIS FOV shall not exceed a 15-degree half-width cone centered on the +X

SRF axis (XRIS- Coordinate System is defined in Section 5.3).

Rationale: Allows for glint or other interference of spacecraft structures at angles larger than this.

XRIS-139: The width of the FOV shall be sufficient to ensure measurement requirements can be met 99.9% of the time under normal operating conditions. See Section 5.3 for spacecraft pointing performance under normal operating conditions.

XRIS-140: The XRIS solar X-ray irradiance observation extent shall be from 0 – 1. 3 Rsun.

The spacecraft will provide the XRIS +Z URF face a 2pi steradian FOV to space, excepting a

FOV blockage with less than 80 W/m2 back loading.

The XRIS Unit Reference Frame (URF) is defined as the following:

The URF is orthogonal and right-handed

The URF is fixed to the body of the XRIS instrument

The location of the XRIS origin will be specified by the XRIS provider

The +YURF axis is approximately aligned with the instrument line of sight

The +ZURF axis is approximately perpendicular with the instrument mounting plane

The +XURF axis completes the triad

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4.2 Resource Allocations

4.2.1 Mass Allocation

XRIS-144: The XRIS (including mounting hardware, thermal blankets, any separate electronics units and cabling) shall have a mass of less than or equal to 6.0 kg. Note: This mass includes mass margin.

XRIS-145: The XRIS instrument mass margin shall comply with GSFC-STD-1000H Rule 1.06.

4.2.2 Physical Envelope

XRIS-147: The XRIS volume (length*width*height), including mounts, thermal blankets and connectors, for both stowed and operational configurations shall have dimensions that do not exceed a total volume of 11424 cm3. These dimensions pertain to both static and dynamic envelopes of the instrument.

XRIS-148: The XRIS maximum dimension in the sun-pointing vector shall be less than 32 cm.

XRIS-2666: The XRIS maximum height from spacecraft interface plate shall be less than 21 cm.

4.2.3 Operational Power Allocation

XRIS-150: The XRIS operational power, including operational heater power, shall be ≤ 12.1W.

4.2.4 Peak Operational Power Allocation

XRIS-152: The XRIS shall have a peak (instantaneous) operational power of ≤ 16 W for a duration of <= 20 msec at 34 V.

4.2.5 Survival Heater Power Allocation

XRIS-154: The XRIS shall have a survival heater power consumption of ≤ 5.5 W.

4.2.6 Telemetry

XRIS-156: The XRIS instrument science and engineering data rate, including all overhead associated with Consultative Committee for Space Data Systems (CCSDS) packetization by the instrument at the spacecraft interface, shall not exceed 30. 0 kilobits per second (kbps).

The Spacecraft-to-instrument Interface Control Document (ICD) will document peak data rate and average data rate for each instrument mode.

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4.2.7 Transient Magnetic Field

XRIS-158: The XRIS instrument shall limit varying magnetic field less than 5 Hz to less than 4 nT zero-to-peak in any axis 1 meter from any face of the unit for any operating mode and changing configurations.

Rationale: Derived from a magnetic cleanliness budget and based on magnetic characteristics of similar instruments.

4.2.8 Static Magnetic Field

XRIS-160: The XRIS units shall limit the static magnetic field to less than 40 nT in any axis 1 meter from any face of the unit for any operating mode.

Rationale: Derived from a magnetic cleanliness budget and based on magnetic characteristics of similar instruments.

4.3 XRIS Operating Modes

XRIS-162: The XRIS current operating mode shall be identified in its telemetry stream.

Rationale: Common best practices deem that being able to quickly identify an instrument’s operating mode is critical for safe on-orbit operations.

4.3.1 In-Flight Calibration

XRIS-164: The XRIS shall provide an in-flight calibration mode that provides a test input to support calibration and testing both on the ground and in space.

Rationale: Common best practices for trending/identifying any functional/performance issues with the electronics of the instrument.

XRIS-166: The XRIS in-flight calibration shall be both self-terminating and able to be terminated by a ground command.

Rationale: To ensure health and safety of the instrument; if one cannot terminate in-flight calibration sequence, the instrument cannot return to the normal operations mode.

4.3.2 Safe Mode

XRIS-174: The instrument shall provide a Safe Mode. Safe mode is an operating mode of the instrument during which all non-essential subsystems are shut down and only essential functions are active. Safe mode represents a configuration in which the instrument is thermally, mechanically, and optically safe without receiving commands from the Spacecraft.

Rationale: To provide a mode where the instrument is in thermally, mechanically, and optically “safe” configuration without receiving commands from the Spacecraft. This mode supports the Observatory Safe Mode strategy.

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XRIS-177: The instrument shall be capable of remaining in a safe configuration for at least 168 hours without ground intervention.

Rationale: To ensure survivability of the instrument should an extended safe mode configuration be required or due to any ground outages.

4.3.3 Normal Operational Mode

XRIS-185: In normal operational mode, the XRIS shall be in a fully functional configuration in which designed measurements are made in accordance with the performance requirements listed in section 4.1 of this document and are available to be sent to the spacecraft for downlink.

Rationale: This is the mode in which the instrument will be operated most of the time.

4.3.4 On-Orbit Operations

XRIS-189: The XRIS shall continuously operate during all spacecraft maneuvers.

Rationale: Simplifies on-orbit operations if no special configurations are required for spacecraft maneuvers. Minimizes any impact to instrument performance due to thermal variations if changing configurations.

XRIS-2667: XRIS shall survive any spacecraft orientation.

XRIS-190: XRIS shall meet all performance requirements within 300 seconds after the spacecraft interface has returned to being within specification following spacecraft maneuvers.

4.3.5 Activation

XRIS-192: The instrument shall require no active commanding prior to fourteen (14) days after launch.

Rationale: The 14 days allows time for the Spacecraft to be fully checked out.

Rationale: To support ground on-orbit anomaly troubleshooting.

4.4 Instrument Data Availability

XRIS-197: The XRIS shall limit on-orbit calibrations to no more than four (4) calibrations per year with each calibration lasting no more than 30 minutes.

XRIS-198: The instrument shall be capable of maintaining nominal observation mode without ground interaction for a minimum of 4 days.

XRIS-199: The instrument shall be capable of re-scheduling planned instrument events that occur during NWS-declared Critical Space Weather Days.

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Rationale: Once a Critical Space Weather Day is declared by NWS it is expected that any pre-planned instrument activities (e.g., calibration maneuvers) will have the flexibility to be re-scheduled without impacting instrument performance. L2RD-239 requires that the sum of daily data gaps for in situ Solar Wind HAP be less than 5 min and any instrument’s pre-planned event has the potential to impact Solar Wind HAP acquisition.

The expected cadence and duration of NWS-declared Critical Space Weather Days could last between 3 and 21 consecutive days.

XRIS-200: The XRIS shall provide a 99% or greater probability of providing all of the solar X-ray Irradiance data during S4 class solar radiation storms or solar flare conditions including class

X50, verifying through analysis using the “Worst Week” October 1989 model in CREME96 or equivalent.

Rationale: The CREME96 model includes the October 1989 solar particle event worst week model that conservatively bounds the single event effects environment in S4 and

X50 solar events. It is expected that the probability of occurrence of outages as well as

Bit Error Rates (BER) due to SEE will be used in the analysis that verifies this probability. Further definition of the worst week model can be found here:

https://creme.isde.vanderbilt.edu/CREME-MC/help/solar-energetic-particle-environment

4.5 Power

The spacecraft will supply one operational power bus for normal instrument operation.

The spacecraft will supply one survival power bus for the instrument survival heaters.

4.5.1 Voltage Range

XRIS-205: The instrument shall operate over the operational bus voltage range of +28 to +34

VDC at the primary power inputs for all expected load conditions (except when turned off).

XRIS-206: The instrument survival heaters shall operate over the voltage range of +26 to +34

VDC at the power inputs for all expected load conditions.

Rationale: The spacecraft power services required to be between +28 to +34 Volts DC.

This is based on an unregulated lithium-ion battery.

4.5.2 Abnormal Voltages

XRIS-208: The instrument shall survive without performance degradation after indefinite exposure to an anomalous voltage range of 0 to +40 VDC.

Rationale: During a spacecraft power circuit failure, the instrument operational power bus could experience an overvoltage transient until the spacecraft turns off the faulty circuit. The overvoltage transient to +40 V peak assumes the high end of the dc voltage range on the spacecraft bus.

https://creme.isde.vanderbilt.edu/CREME-MC/help/solar-energetic-particle-environment

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4.5.3 Turn-on Current Transients

XRIS-216: The instrument shall meet its performance requirements without degradation after exposure to a turn-on input voltage (0 – 34 VDC) with a rise time of 500µs.

Rationale: This is required to be compatible with the power service requirements of the spacecraft.

Note: The spacecraft bus will not drop below 26 V during instrument turn on.

4.6 Electrical Grounding

4.6.1 Primary Power DC Isolation

XRIS-220: The instrument primary power interfaces, primary power and primary power returns, shall be isolated from the unit chassis by a DC resistance of greater than or equal to 1

Megaohms.

Rationale: Separating the instrument power ground from the instrument chassis will prevent the instrument power supply noise current from flowing into the instrument secondary grounds.

4.6.2 Survival Power Isolation

XRIS-222: The instrument survival heater power interfaces shall be isolated from the unit chassis by a DC resistance of greater than or equal to 1 Megaohms.

Rationale: Returning the instrument survival heater power returns to the spacecraft single point ground and isolating the survival heater power returns from the instrument chassis helps control common mode noise current and minimize magnetic field effects due to ground loops.

4.6.3 Mechanical Contact Resistance

XRIS-228: The DC resistance of the mechanical contact between two conductive mating surfaces (internal to the unit) shall be less than or equal to 2.5 milliohm DC resistance.

Rationale: Low resistance contacts are essential for maintaining a zero potential ground.

4.6.4 Grounding

XRIS-230: The DC resistance between the XRIS chassis and the observatory chassis shall be ≤

2.5 milliohms.

Rationale: Low resistance contact between instrument chassis to spacecraft structure is essential for maintaining a zero potential ground.

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XRIS-231: The XRIS shall provide a ground lug for a grounding strap to be attached from the

XRIS chassis for connection to the spacecraft conductive structure.

Rationale: This provides a method/location for tying the instrument ground to the spacecraft conductive structure.

XRIS-232: The grounding lug location on the XRIS instrument chassis or the tie points in contact with the ground strap will be as defined in the Mechanical Interface Control Drawing

(MICD). The ground lug contact area must remain free of any material finish that may affect the reliability of the ground connection and will be shown in the MICD.

Rationale: Minimum contact area is a best practice for maintaining integrity of the ground.

4.6.5 Connector DC Resistance

XRIS-234: XRIS connector backshells shall be electrically connected to Electronic Unit chassis with a DC resistance ≤ 2.5 milliohms.

Rationale: Low resistance contacts are essential for maintaining a zero potential ground.

4.7 Signal And Data Interfaces

4.7.1 Passive Analog Telemetry

XRIS-237: The XRIS shall require no more than two analog signals to monitor critical temperature points when the instrument is powered off.

Rationale: The Spacecraft will be able to monitor health and safety of the instrument if the instrument is in any configuration other than the normal operational configuration, e.g. safe or survival mode. Two signals is also deemed reasonable given the complexity of the instrument.

4.7.2 Data Signal Interface

The spacecraft will provide an interface for sending commands to the instrument and will provide a receiving interface for instrument telemetry. The spacecraft will also provide an output line for a Pulse Per Second (PPS) time signal.

XRIS-240: The instrument data transfer interface to or from the spacecraft shall either use a serial Universal Asynchronous Receiver Transmitter (UART) using ANSI/TIA/EIA-422-B

Electrical Characteristics of Balanced Voltage Differential Interface Circuits, or ANSI/TIA/EIA-

644-A-2001 Electrical Characteristics of Low Voltage Differential Signaling (LVDS) Interface

Circuits.

4.7.2.1 Telemetry Source Packet Format

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XRIS-242: The instrument shall transfer all data to the spacecraft using the CCSDS 133. 0-B-2

Section 4.1 Protocol Data Unit definition shown in the Telemetry Source Packet Definition

Figure 1 Telemetry Source Packet Definition.

Rationale: CCSDS is a data format interface standard developed by NASA to create commonality to simplify data processing and routing.

Figure 1. Telemetry Source Packet Definition

XRIS-245: The instrument telemetry Source packets shall be variable length with a maximum data zone of 8192 octets including Secondary Header.

Rationale: CCSDS command pack definition and SWSRD 636

XRIS-246: The instrument shall set the telemetry source packet Secondary Header Flag to the value 1.

Rationale: CCSDS 133.0-B-1 Space Packet Protocol, Blue Book, Issue 1, September

XRIS-249: The instrument shall set the Telemetry Source Packet Time Code per CCSDS 301.0-

B-4 Time Code Formats, Day Segmented format in the Time Code Format Figure 2.

Rationale: CCSDS is a data format interface standard developed by NASA to create commonality to simplify data processing and routing.

Figure 2. Time Code Format

Note: The P-Field is implied and not included in the actual time message.

The spacecraft on-board reference time will be Coordinated Universal Time (UTC).

4.7.2.2 Command Source Packet Format

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XRIS-255: The instrument shall receive all data from the spacecraft formatted per CCSDS

133.0-B-2 Section 4.1 Protocol Data Unit definition shown in the Command Source Packet

Definition Figure 3.

Figure 3. Command Source Packet

XRIS-258: The instrument Command Source Packets shall be variable length with a maximum data zone of 1013 octets.

Rationale: CCSDS 133.0-B-1 Space Packet Protocol, Blue Book, Issue 1, September

4.7.3 Clock Signal Interface

XRIS-263: The instrument will receive a Pulse Per Second (PPS) time pulse from spacecraft clock line by the interface defined in section 4.7.2 of this document.

Rationale: RS-422 and LVDS are two of the common physical interfaces used on spacecraft. The NOAA RFI responses indicate most spacecraft vendors can support one or the other or both.

The Spacecraft will maintain Spacecraft time correlation to within +/- 0.95 sec of UTC and will provide the instrument a 1 PPS with a time code message.

XRIS-265: The instrument will receive from the spacecraft a time code message on the data line as defined in Spacecraft Time Message Packet Figure 4 Spacecraft Time Message Packet. The time code message is the time applicable to receipt of the 1 PPS.

Rationale: CCSDS 133.0-B-1 Space Packet Protocol, Blue Book, Issue 1, September

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Figure 4. Spacecraft Time Message Packet

4.7.4 Command and Housekeeping telemetry

XRIS-269: The instrument shall provide command and housekeeping telemetry functions in all powered modes.

Rationale: Housekeeping telemetry are data required to monitor instrument operation, health, and safety. Command functions may be required to maintain the health and safety of the instrument.

4.7.5 Commands for Autonomous Functions

XRIS-271: The instrument shall execute commands to individually enable and disable each autonomous function.

Rationale: Allows flexibility to maintain the health and safety of the instrument during on-orbit operations.

4.8 Flight Software

4.8.1 Flight Load Non-volatile Memory

XRIS-282: The instrument flight software image shall be contained in its entirety in non-volatile memory at launch.

Rationale: Allows for power-cycling without loss of the stored flight software image.

4.8.2 Software Updates

XRIS-284: The flight software modules shall be reprogrammable.

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Rationale: Provides flexibility should operational changes be required following an anomaly, as the instrument ages, etc.

4.8.3 Software Table Updates

XRIS-286: Instrument configuration data (e. g. , Table Loads or Configuration Parameters) shall be reprogrammable during integration and test phases and on-orbit without computer restart.

Rationale: Common best practices dictate this capability to avoid unnecessary power cycles of the instrument.

XRIS-287: Modified instrument configuration data (e. g. , Table Loads or Configuration

Parameters) shall be committed to operational use by ground command.

Rationale: Common best practice to ensure “a person in the loop” before making a configuration change.

4.8.4 Flexibility and Ease of Software Modification

XRIS-290: Instrument configuration data (e.g., Table Loads or Configuration Parameters) shall be referenced such that data can be loaded and dumped by the ground without reference to memory address.

Rationale: Common best practices dictate this capability for ease of operation and to reduce risk of loading to the wrong memory address.

XRIS-291: The definition of instrument commands within the ground database (excluding memory diagnostic/load commands) shall not be dependent on physical memory addresses within the flight software.

Rationale: To avoid the need to rebuild the command and telemetry database every time the flight software is recompiled; modifying the code results in address changes leading to command changes and then re-testing commands.

4.8.5 Version Identifiers

XRIS-293: The instrument software and firmware versions shall each have an internal identifier

(embedded in the executive program) that can be included in the instrument engineering data.

Rationale: Common best practice to ensure that knowledge of version is always accessible.

4.8.6 Warm Restart

XRIS-296: The instrument flight software shall provide a restart by command with preservation of instrument configuration data and memory tables.

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4.8.7 Memory Tests

XRIS-298: The instrument flight software shall provide a mechanism to verify the contents of all memory areas.

4.8.8 Memory Dump

XRIS-300: The instrument flight software, and associated on-board computer hardware, shall provide the capability to dump any memory location.

XRIS-301: The flight software memory dump capability shall not disturb normal operations and instrument data processing.

4.8.9 Fault Detection and Correction Control

XRIS-303: If applicable, the instrument shall provide the capability to enable and disable any internal Fault Detection and Correction (FDC) features.

Rationale: Allows flexibility in the on-orbit operations.

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5 PHYSICAL REQUIREMENTS

5.1 Mass Properties Accuracy

5.1.1 Instrument Mass Accuracy

XRIS-314: The mass of the instrument shall be measured to within ±0.1 kg.

Rationale: Defines an allowable tolerance on instrument unit weight measurements.

5.1.2 Center of Mass Accuracy

XRIS-318: The instrument center of mass shall be determined to within ±5 mm relative to a reference that will be defined in the MICD.

Rationale: Defines an allowable tolerance on instrument center of mass measurements.

5.2 Mounting

XRIS-2058: Instrument mounting shall not require access from inside the spacecraft.

Rationale: To simplify installation and reduce risk for damage.

5.3 Coordinate System and Alignment

XRIS-327: The instrument shall use the spacecraft on-orbit coordinate system which uses a right-hand orthogonal, body-fixed XYZ coordinate system, Spacecraft Reference Frame (SRF), as follows: the +X-axis is aligned with the nominal direction of the spacecraft to Sun vector pointing towards the Sun, the +Y-axis nominal to normal of the ecliptic plane and pointing to celestial North, and the Z-axis is obtained from the right-hand rule. The roll, pitch and yaw rotations are defined about the X, Z, and Y axes, respectively. The origin of the coordinate system is at the center of the Launch Vehicle separation ring attachment plane. See Figure 5

Spacecraft Reference System.

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Figure 5. Spacecraft Reference System

The spacecraft will physically align the XRIS axes to within ±0.25 degrees of the direction specified in section 4.1.3 and provide alignment knowledge of the sensor axes to within ±0.1 degrees of the direction specified in section 4.1.3.

XRIS-331: The instrument shall have clear fiducial marks to enable alignment during integration with spacecraft.

Rationale: Required for accurate mechanical alignment during integration onto the spacecraft.

The instrument unit mounting frame is an orthogonal reference frame defined by the locations of the spacecraft side of the instrument unit mounting points. A rigorous definition of this frame will be documented in the ICD.

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6 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements for the instrument are specified in this section.

XRIS-335: The instrument shall meet its performance requirements in section 4 after exposure to the environments specified in this section.

Rationale: Required to demonstrate that the instrument will survive launch, transportation, storage, and on-orbit environments.

6.1 Mechanical Factors of Safety

XRIS-337: The instrument flight hardware shall demonstrate positive Margins of Safety under limit loads for all yield and ultimate failures using the Factors of Safety (FS) defined in Table 2.

Margin of Safety (MS) is defined as follows:

Rationale: The yield factor of safety ensures that a prototype or protoflight test can be conducted with low risk of the hardware experiencing detrimental yielding. The ultimate factor of safety provides adequate separation between yield and ultimate failure modes and ensures that the hardware will not experience an ultimate failure under expected loading conditions.

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

Table 2. Factors of Safety1,2

Type Static Sine Random/Acoustic4,5

Metallic Yield 1.253 1.25 1.6

Metallic Ultimate 1.43 1.4 1.8

Stability Ultimate 1.4 1.4 1.8

Beryllium Yield 1.4 1.4 1.8

Beryllium Ultimate 1.6 1.6 2.0

Composite Ultimate 1.5 1.5 1.9

Bonded Inserts/Joints Ultimate 1.5 1.5 1.9

1 Factors of safety for pressurized systems to be compliant with AFSPCMAN 91-710

(Range Safety).

2 Factors of safety for glass and structural glass bonds specified in NASA-STD-5001

3 If qualified by analysis only, positive margin must be shown for factors of safety of

2.0 on yield and 2.6 on ultimate. See section 2.4.1.1.1 GSFC-STD-7000B

4 Factors shown should be applied to statistically derived peak response based on RMS level. As a minimum, the peak response shall be calculated as a 3-sigma value.

5 Factors shown assume that qualification/protoflight testing is performed at acceptance level plus 3dB. If difference between acceptance and qualification levels is less than

3dB, then above factors may be applied to qualification level minus 3dB instead of analyzing to acceptance level.

XRIS-383: 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

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analysis only is not acceptable. Actual flight hardware testing is preferred, but testing of representative sets of hardware with a similar qualification argument can be used if approved by the NASA/GSFC COR.

Rationale: In order to account for uncertainties in workmanship, material properties and local stress levels special screening/criteria is required per NASA-STD-5001.

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.

XRIS-386: The instrument shall be designed to withstand the quasi-static design limit loads defined in the mass-acceleration curve (MAC) shown in Table 3 without damage or degradation of performance and are to be applied one axis at a time. 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 instrument 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 3. Design Limit Loads

Mass (kg) Accel (g)

1 or less 51.3

5 42.4

10 36.6

20 30.3

40 24.2

60 20.9

80 18.7

100 or greater 18.1

6.3 Frequency Requirement

6.3.1 Fundamental Launch Frequencies

XRIS-419: The instrument shall have a fundamental frequency greater than 100 Hz when hard mounted at its spacecraft interface. Any unit which fails to meet the specified fundamental frequency must supply a finite element model, correlated to modal survey test results or measured responses from vibration testing, so that it accurately represents the frequency and response amplitudes from the test. Requirements for the submitted finite element model are shown in the CDRL.

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6.4 Vibration

6.4.1 Sinusoidal Vibration

XRIS-422: The instrument shall undergo qualification, protoflight or acceptance (level depends on qualification status or qualification approach for the unit) sine vibration testing in all three axes at the levels shown in Figure 6. Instruments with a first mode greater than 150 Hz can be exempted from sine vibration testing. upon approval by the NASA/GSFC COR. See Section

11.1.2 for definitions of Protoflight, Qualification, and Acceptance

Figure 6. 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.

6.4.2 Random Vibration

XRIS-427: The instrument shall demonstrate its ability to meet its performance requirements after being subjected to the random vibration environment in Figure 7 for units weighing 22.7 kg

(50 lb.) or less, applied at the Spacecraft to instrument interface.

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Rationale: Test as you fly. GEVS standard to verify performance after (or during) environmental tests if component/ subassembly or assembly is on during launch.

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Figure 7. Random Vibration Environment

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The random…