Attachment B - Requirements Specification (SPEC).pdf

L1 Series CSPEC L1SERIES-COR-REQ-0018, Revision - Effective Date: February 16th, 2024 ii Check https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm to verify that this is the correct version prior to use.

SW L1 Coronagraph Requirement Specification (CSPEC) Review/Signature/Approval Page

Prepared by:

Originally Approved on TDMS By: 02-14-2024 Steve Wasserzug Date SWO L1 Series NASA GSFC, Code 493

Reviewed by:

Originally Approved on TDMS By: 02-12-2024 Charles “Hudson” DeLee Date SWO L1 Series NASA GSFC, Code 493

Approved by:

Originally Approved on TDMS By: 02-16-2024 Tim Vansant Date L1 Series 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:

SWO Configuration Management Office

NASA/GSFC

Code 493.0 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

Revision Effective Date Description of Changes (Reference the CCR & CCB/ERB Approval Date)

Revision - February 16, 2024 This document was reviewed for baseline in CCR L1SERIES-CCR-0064 and was approved on February 16, 2024.

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Table of Contents

1 INTRODUCTION

1.1 General Information

1.2 Scope

2 APPLICABLE DOCUMENTS

3 CONTRACT DESCRIPTION

3.1 Coronagraph Description

3.2 Ground Support Equipment Description

4 FUNCTIONAL/PERFORMANCE REQUIRMENTS

4.1 Coronagraph Performance Requirements

4.1.1 Scene Coverage

4.1.2 Spatial Resolution

4.1.3 Measurement Range

4.1.4 Accuracy

4.1.5 Field of View (FOV)

4.1.6 Refresh Rate and Latency

4.1.7 Detector Well Depth

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 Coronagraph Operating Modes

4.3.1 In-Flight Calibration

4.3.2 Mode Transitions

4.3.3 Deterministic Power-on Configuration

4.3.4 Safe Mode

4.3.4.1 Entry Into Safe Mode

4.3.5 Fail-safe Recovery Mode

4.3.6 Normal Operational Mode

4.3.7 On-Orbit Operations

4.3.8 Station Keeping

4.3.9 Activation

4.4 Instrument Data Availability

4.5 Power

4.5.1 Voltage Range

4.5.2 Abnormal Voltages

4.5.3 Sudden Removal of Power

4.5.4 Over-Current Protection

4.5.5 Primary Power Return Ground

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4.5.6 Turn-on Current Transients

4.6 Electrical Grounding

4.6.1 Primary Power DC Isolation

4.6.2 Survival Power Isolation

4.6.3 Internally Generated Secondary to Primary DC Isolation

4.6.4 Internally Generated Secondary Return

4.6.5 Mechanical Contact Resistance

4.6.6 Grounding

4.6.7 Connector DC Resistance

4.7 Signal and Data Interfaces

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

4.7.6 Limits and Triggers

4.7.7 On-Board Processor Reset

4.7.8 Micro-switches

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

4.8.10 Health and Safety Monitoring

5 PHYSICAL REQUIREMENTS

5.1 Interface and Design Units and Exclusions

5.2 Mass Properties Accuracy

5.2.1 Instrument Mass Accuracy

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 Mounting

5.3.1 Surface Flatness

5.3.2 Method

5.4 Coordinate System and Alignment

6 ENVIRONMENTAL REQUIREMENTS

6.1 Mechanical Factors of Safety

6.2 Quasi-Static Acceleration

6.3 Frequency Requirement

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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

6.11.2 Radiation Hardness Assurance

6.11.2.1 Single-Event Effect Rate Calculations

6.11.3 Charging 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 Contamination Analyses

7.5 Venting Requirements

7.6 Instrument Purging

7.7 Cleanability and Protection

7.8 Electrostatic Cleanliness

7.8.1 Conductive Surface Ground Path

7.8.2 Conductive Surface Resistivity

7.8.3 Closeout of Gaps and Apertures

7.8.4 Exposed Harness Specific Requirements

7.8.5 Thermal Blankets

8 DESIGN AND CONSTRUCTION REQUIREMENTS

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8.1 Electrical

8.1.1 Test Sensors

8.1.2 MGSE Grounding

8.1.3 Connector Specifications

8.1.3.1 Contact Derating

8.1.3.2 Redundant Contact Derating

8.1.3.3 Signal Segregation

8.1.3.4 Test and Flight Signal Isolation

8.1.4 Test Interfaces

8.1.4.1 Facility-Induced Noise

8.1.4.2 Facility-Induced ESD GSE Malfunction

8.1.4.3 Facility-Induced GSE Malfunction

8.1.5 Mitigation of Internal Charging

8.1.5.1 Mitigation Strategies for Internal Charging

8.1.5.2 Floating Conductors

8.1.5.3 Dielectric Structures

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, Power Cables

8.3.2.5 CS06 – Conducted Susceptibility, Transients Limits

8.3.2.6 CS115 - Conducted Susceptibility, Bulk Cable Injection, Impulse Excitation ... 71

8.3.3 Radiated Emissions

8.3.3.1 RE102 - Electric Field Emissions Limits

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 Ground Support Equipment (GSE) Design

8.7.2 Lifting Hardware

8.7.3 Manual Lifting Hardware

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8.7.4 GSE Cleanliness

8.7.5 GSE Bakeout

8.7.6 Test Harness

8.8 Interface Documentation

8.8.1 Mechanical Interface

8.8.2 Electrical Interface

8.8.3 Data Interface

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

9.2.2 Supplemental Factor

9.2.3 Ultimate Design Loads

9.2.4 Yield Design Loads

9.2.5 Design Separation Load

9.2.6 Fastener Locking and Retention

9.2.6.1 Thread Engagement

9.2.6.2 Locking Feature Verification

9.2.6.3 Locking Feature Installation

9.2.6.4 Snap Ring and Cotter Pin Use Limitation

9.2.6.5 Snap Ring and Cotter Pin Use

9.2.6.6 Liquid Locking Compounds

9.2.6.6.1 Verification

9.2.6.7 Locking Features

9.2.6.7.1 Specification

9.2.6.7.2 Validation

9.2.7 Installation Torque Specification and Control

9.2.7.1 Installation Torque Range

9.2.7.2 Running Torque Identification

9.2.8 Fastened Joints Criteria

9.2.8.1 Minimum and Maximum Preload

9.2.8.2 Analysis Addressing Potential Rupture

9.2.8.3 Ultimate Strength Analysis

9.2.8.3.1 Applied Shear Loading

9.2.8.3.2 Shear Loading

9.2.8.4 Simultaneous Applied Tensile and Shear Loads

9.2.8.5 Allowable Yield Tensile Load

9.2.8.6 Separation Analysis

9.2.8.7 Seal Analysis

10 LOGISTICS

10.1 I&T Deliverables

10.2 Ground Support Equipment

10.2.1 Electrical System Test Equipment

10.2.2 Instrument Emulator

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10.3 Transportation Equipment

10.3.1 Shipping Container

11 VERIFICATION REQUIREMENTS

11.1 Verification Methods

11.1.1 Inspection

11.1.2 Analysis

11.1.3 Test

11.1.4 Demonstration

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

11.4.4.5 Configuration

11.5 Required Tests

11.5.1 Performance Tests

11.5.1.1 Comprehensive Performance Test

11.5.1.2 Limited Performance Test

11.5.1.3 Abbreviated Functional Test

11.5.2 Mass Properties Measurement

11.5.3 Static Loads/Strength Test

11.5.3.1 Sine Burst

11.5.3.2 Static Pull

11.5.4 Sine Sweep Survey

11.5.5 Sine Vibration

11.5.6 Random Vibration

11.5.7 Acoustic Test

11.5.8 Shock

11.5.9 Thermal Vacuum Bake-out

11.5.10 Thermal Vacuum Test

11.5.10.1 Thermal Vacuum Test Parameters

11.5.10.2 Thermal Vacuum Test Profile

11.5.11 Magnetics Test

11.5.12 Harness Tests

11.5.13 EMI/EMC Tests

12 APPENDIX A ABBREVIATIONS AND ACRONYMS

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List of Figures

Figure 1 Allowable volumes: Telescope Module (left), Electronics Unit(s) (right) Figure 2 Coronagraph allowable door-deployment volume Figure 3 Telemetry Source Packet Definition Figure 4 Time Code Format Figure 5 Command Source Packet Figure 6 Spacecraft Time Message Packet Figure 7 Spacecraft Reference System Figure 8 Sine Vibration Environment Figure 9 Random Vibration Environment Figure 10 Candidate LV Shock Environment Figure 11 L1 Series Acoustic Design Envelope (TBR) Figure 12 Allowable Analytical, Flight Operation, Flight Acceptance Test, and Qualification

(ProtoFlight or Prototype) Test Temperature Ranges Figure 13 Total Ionizing Dose-Depth Curve Figure 14 Displacement Damage Dose in Silicon vs. Shielding Thickness Dose-Depth Curve 41 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 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 Figure 31 Thermal Vacuum Profile

List of Tables

Table 1 Applicable Documents (TBR) Table 2 Factors of Safety1,2 Table 3 Design Limit Loads Table 4 Candidate LV Shock Environment Table 5 L1 Series Acoustic Design Envelope (TBR) Table 6 Transportation Loads (TBR) Table 7 Orbit Thermal Parameters Table 8 Total Ionizing Dose vs. Shielding Thickness xv Check https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm to verify that this is the correct version prior to use.

Table 9 Displacement Damage Dose in Silicon vs. Shielding Thickness Table 10 Displacement Damage Dose in Gallium Arsenide vs. Shielding Thickness Table 11 LET Spectrum of GCR Background Table 12 Solar Background LET Spectrum Table 13 Solar Proton Background Flux Table 14 Peak Solar Particle Event Flux Table 15 Peak Solar Proton Flux Table 16 Lifetime Contamination Requirements Table 17 Thermal Blanket Area vs. Grounding Tabs Table 18 Emission and Susceptibility Requirements Table 19 RS103 Radiated Susceptibility Levels Table 20 Test Factors and Durations Table 21 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 Coronagraph Instrument for the Space Weather Next L1 Series 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 Coronagraph to the extent specified herein. In the event of conflict between this specification and any referenced document, this specification will govern, except for the L1 Series Coronagraph Statement of Work (L1SERIES-COR-SOW-0006), in which case the Statement of Work takes precedence.

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

Table 1 Applicable Documents (TBR)

Document Number Title

L1SERIES-COR-SOW-

L1 Series Coronagraph (COR) Statement of Work (SOW)

L1SERIES-COR-REQ-

L1 Series Coronagraph (COR) Contract Deliverables Requirements List (CDRL)

L1SERIES-COR-REQ-

Mission Assurance Requirements (MAR)

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 Military Standard, Electromagnetic Emission And Susceptibility

Requirements For The Control of Electromagnetic Interference

(EMI)

GSFC-STD-7000B General Environmental Verification Standard (GEVS) NASA-STD-5019A Fracture Control Requirements for Spaceflight Hardware NASA-STD-5020B Requirements for Threaded Fastening Systems in Spaceflight

Hardware

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Document Number Title NASA-STD-5017B Design and Development Requirements for Mechanisms NASA-HDBK-4002B 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 4, November 2010

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3 CONTRACT DESCRIPTION

3.1 Coronagraph Description

The Coronagraph instrument will monitor the solar corona and detect and characterize coronal mass ejections (CMEs) that are directed towards Earth. CMEs can be remotely detected with white light imagery of the upper solar corona and the Coronagraph will capture this white light imagery. Sequences of CME images can be used to determine size, velocity, and density of CMEs. The Coronagraph instrument consists of a Telescope Module, an Electronics Unit, and the associated harnesses.

The Coronagraph contract includes the delivery of one (1) Engineering Development Unit (EDU) with representative spacecraft mechanical and electrical interfaces, three (3) Flight Units plus parts for a robust sparing philosophy, with enough flight spares to mitigate risk of issues that may occur late in the launch integration flow.

3.2 Ground Support Equipment Description

The Coronagraph contract includes the delivery of four (4) sets of the Electrical System Test Equipment (ESTE), four (4) Coronagraph Emulators, one (1) Flight Software Development Environment (FSDE), 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 includes, but is not limited to, test cables and break-out boxes.

The Coronagraph delivery includes the ground support equipment required to power and operate the Coronagraph at the vendor’s and spacecraft contractor’s facilities. The delivery also includes standalone emulators that will interface to a spacecraft simulator for the purposes of interface testing, operations product development, and training.

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 CDRL.

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4 FUNCTIONAL/PERFORMANCE REQUIRMENTS

This section defines the functional and performance requirements for the L1 Series Coronagraph.

The requirements in this Specification pertain to the Coronagraph ‘system’, which may include all instrument hardware, software, and ground processing algorithms. The Coronagraph contractor is not responsible for the operational ground system, but certain data products may require ground processing between data collection and data distribution; the contractor is responsible for defining, developing, and distributing the associated ground processing algorithms.

4.1 Coronagraph Performance Requirements

4.1.1 Scene Coverage

CSPEC-22: The Coronagraph shall capture radial coronal scenes from ≤ 3 Rsun to ≥ 22 Rsun.

CSPEC-23: The Coronagraph shall capture the radial coronal scene with a ≥ 290° coverage satisfying all spatial and photometric requirements.

Rationale: The inner field of view represents a geometric cutoff, not a photometric cutoff, and does not account for spacecraft motions. The spatial range for the field of view ensures that a CME travelling at ~3000 km/s can have sufficient observations (with the cadence given in Section 4.1.6) to fit model parameters. The specification ensures at least three images for each CME observed, giving sufficient margin for determining the CME properties necessary for prediction. Note that the external occulter (EO) will vignette the coronal scene (reduce the A1 aperture that can view the coronal scene) and add to the image noise from the diffracted straylight. The angular scene coverage of 290° allows use of pylon(s) to support the occulter. The EO pylon(s) will vignette the coronal scene (reduce the A1 aperture that can view the coronal scene) and add to the image noise from the diffracted straylight off of the EO pylon edge and reflected straylight off of the aft face of the pylon(s). The impact of the EO pylon(s) on the vignetting and straylight will decrease with the polar angle from the center of any EO pylon(s). Therefore, it is recognized the spatial and photometric requirements will have localized effects near these instrument features.

4.1.2 Spatial Resolution

CSPEC-26: The Coronagraph shall capture coronal images with spatial resolution ≤ 70.0 arcsec over the un-occulted field of view.

Rationale: The image spatial resolution is defined as the average of the optical spatial resolution in the sagittal and tangential planes of the optical system. The instrument spatial resolution applies to an image acquired under nominal pointing conditions (with no boresight jitter or pointed windowed stability during the averaged image acquisition time). The spatial resolution is defined as the distance, in arcsec, between two point-sources of equal brightness located at infinity. The relative intensity between the peaks of

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the imaged sources (100% intensity) and the intensity at the mid-point between the peaks is 90% or less.

4.1.3 Measurement Range

CSPEC-29: The Coronagraph shall capture coronal white light imagery with intensity of 1E-11 BSun to 1E-8 BSun.

Rationale: The Coronagraph’s dynamic range shall be such that the expected signal of the F + K corona, the residual straylight, and the CME all as modulated by the vignetting will be covered over the entire field of view. Bsun (or B0) is the surface brightness of the solar disk, which is an extended source. It is equivalent to the observed flux density per solid angle.

4.1.4 Accuracy

CSPEC-31: The Coronagraph shall have photometric accuracy for the corona image of ≤ 10% in mean solar brightness.

Rationale: Setting absolute photometric accuracy, i.e. calibration, on the observed scene allows for determination of CME mass. Much of the image noise in the K corona image is shot noise with a Poisson distribution for the electrons collected at each pixel based on the total brightness of the instrumental straylight, the coronal scene, vignetting, and the detector dark current. The estimation error over a block of pixels dominated by Poisson statistical noise can be reduced by averaging over this block of pixels if the source brightness is not varying greatly over this pixel block.

4.1.5 Field of View (FOV)

CSPEC-33: The Coronagraph shall allow for an unobstructed field of view of no more than 50 deg half cone angle.

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

4.1.6 Refresh Rate and Latency

CSPEC-35: The Coronagraph shall produce an image during nominal operations at a refresh rate of every 15 minutes or less.

CSPEC-36: The total time for the instrument portion of data product generation—from end of image acquisition to delivery of Level-1b products—shall be 1300 sec or less.

Rationale: The limitation ensures at least three images for each CME observed, giving sufficient margin for determining the CME properties necessary for prediction. Note that the Coronagraph instrument latency is only one component of the end-to-end latency budget. The 1300 seconds allocated for the instrument includes all instrument data

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processing time (on-board and on the ground). The latency requirement is driven by the time needed for forecasters to analyze the images and by the cadence of real-time operational models.

4.1.7 Detector Well Depth

CSPEC-38: The Coronagraph instrument exposure time shall be such that total brightness measurement in electrons does not exceed 95% full well of detector.

Rationale: Provides margin against saturating the detector.

4.2 Resource Allocations

4.2.1 Mass Allocation

CSPEC-41: The Coronagraph instrument, including the imaging module, power supply box, intra-instrument harness, and thermal blankets, shall have a combined mass of ≤ 29.0 kg (includes margin).

CSPEC-4066: The Coronagraph instrument mass margin shall comply with GSFC-STD-1000H Rule 1.06.

4.2.2 Physical Envelope

CSPEC-43: The Telescope Module—including mounts, sensor standoff bracket (if required), intra-instrument harness, thermal blankets, and connectors—in both stowed and operational configurations shall have dimensions that do not exceed 100 cm (XCRF) x 40 cm (ZCRF) x 50 cm (YCRF) (TBR—all 3 dimensions). These dimensions pertain to both static and dynamic envelopes of the instrument. See Figure 1.

CSPEC-4271: If a separate Electronics Unit is required, it—including mounts, brackets, thermal blankets, and connectors—shall have dimensions that do not exceed the dimensions shown in Figure 1 (TBR).

The spacecraft will locate the Coronagraph Telescope Module and Electronics Unit such that the length of the cable between them is ≤ 1.0 meter. Relative orientation of the two separate volumes will be negotiated during ICD development between spacecraft and instrument providers.

CSPEC-44: The Coronagraph maximum instrument volume during deployment of mechanisms shall have dimensions that do not exceed the dimensions shown in Figure 2 (TBR).

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Figure 1 Allowable volumes: Telescope Module (left), Electronics Unit(s) (right)

Figure 2 Coronagraph allowable door-deployment volume

4.2.3 Operational Power Allocation

CSPEC-46: The Coronagraph average operational EOL power, including operational heater power, shall not exceed 29 Watts when averaged over any 15-minute period.

4.2.4 Peak Operational Power Allocation

CSPEC-48: The Coronagraph shall have a peak (instantaneous) power consumption of less than or equal to 47 Watts (TBR) for a duration of 20msec @ 34Vdc.

4.2.5 Survival Heater Power Allocation

CSPEC-50: The Coronagraph survival heater power shall not exceed 22.5 (TBR) Watts when averaged over any 60-minute period.

4.2.6 Telemetry

CSPEC-52: The Coronagraph total 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 140 kbps averaged over 5 seconds.

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4.2.7 Transient Magnetic Field

CSPEC-54: The Coronagraph 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.

4.2.8 Static Magnetic Field

CSPEC-56: The Coronagraph 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.

4.3 Coronagraph Operating Modes

CSPEC-4071: The Coronagraph current operating mode shall be identified by a flag in its telemetry stream.

4.3.1 In-Flight Calibration

CSPEC-4074: Coronagraph in-flight calibration shall be both self-terminating and able to be terminated by a ground command.

4.3.2 Mode Transitions

CSPEC-59: The instrument shall transition from the current mode to any other mode without causing damage to itself.

4.3.3 Deterministic Power-on Configuration

CSPEC-61: The instrument shall initialize upon power-up into a predetermined configuration.

4.3.4 Safe Mode

CSPEC-64: 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.

CSPEC-65: The instrument shall be capable of remaining in a safe configuration for at least 168 hours without ground intervention.

4.3.4.1 Entry Into Safe Mode

CSPEC-67: The instrument shall enter Safe Mode upon detection of internal faults that are capable of causing damage to the instrument.

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

The L1 Series spacecraft will monitor up to six (TBR) analog instrument health and safety parameters defined by the Coronagraph contractor and command the Coronagraph into safe mode when any of those key values have been exceeded.

The L1 Series spacecraft will monitor up to 12 (TBR) digital telemetry points that are part of the normal instrument generated data stream for instrument health and safety parameters defined by the Coronagraph contractor and command the Coronagraph into safe mode when any of those key values have been exceeded.

4.3.5 Fail-safe Recovery Mode

CSPEC-4077: The instrument shall provide a failsafe recovery mode dependent on a minimal hardware configuration capable of accepting and processing a minimal command subset sufficient to load and dump memory.

4.3.6 Normal Operational Mode

CSPEC-70: In normal operational mode, the Coronagraph 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.

4.3.7 On-Orbit Operations

CSPEC-72: The Coronagraph shall operate normally, within specification, while flying aboard a 3-axis stabilized spacecraft with orbital limit constraints as stated in this specification.

4.3.8 Station Keeping

CSPEC-74: The Coronagraph shall continuously operate during all spacecraft maneuvers.

CSPEC-75: The instrument shall meet all performance requirements within 300 seconds (TBR) after the spacecraft interface has returned to being within specification following spacecraft maneuvers. Note: this requirement assumes that no direct sunlight illuminates the Coronagraph radiator during the maneuver.

4.3.9 Activation

CSPEC-4079: The instrument shall require no active commanding prior to fourteen (14) days after launch.

4.4 Instrument Data Availability

CSPEC-4081: The Coronagraph shall limit on-orbit calibrations to no more than four (4) calibrations per year with each calibration lasting no more than 30 minutes.

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

Rationale: These limits provide margin against data availability requirements, and they apply only after commissioning.

CSPEC-4082: The instrument shall be capable of maintaining nominal observation mode without ground interaction for a minimum of 4 days.

CSPEC-4083: The instrument shall be capable of re-scheduling planned instrument events that occur during NWS-declared Critical Space Weather Days.

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-39 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.

CSPEC-4084: The Coronagraph shall provide a 99% or greater probability of providing all of the instrument 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

CSPEC-80: The instrument shall operate over the operational bus voltage range of +26 to +34 VDC at the primary power inputs for all expected load conditions (except when turned off).

CSPEC-4085: The instrument survival heaters shall operate over the voltage range of +24 to +34 VDC at the power inputs for all expected load conditions.

4.5.2 Abnormal Voltages

CSPEC-82: The instrument shall survive without performance degradation after indefinite exposure to an anomalous voltage range of 0 to +40 VDC.

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

4.5.3 Sudden Removal of Power

CSPEC-86: The instrument shall meet its performance requirements without degradation after exposure to an abrupt, unannounced removal of power.

4.5.4 Over-Current Protection

CSPEC-88: The instrument shall not use non-resetting over-current protection (i.e., fuses) internal to the unit.

4.5.5 Primary Power Return Ground

CSPEC-90: The instrument shall provide a dedicated Primary Power return in the same connector as the primary power.

4.5.6 Turn-on Current Transients

CSPEC-92: 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.

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

CSPEC-95: 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 10 Megaohms.

4.6.2 Survival Power Isolation

CSPEC-97: The instrument survival heater power interfaces shall be isolated from the unit chassis by a DC resistance of greater than or equal to 10 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 Internally Generated Secondary to Primary DC Isolation

CSPEC-99: Secondary power inputs shall be isolated from primary power by a DC resistance of greater than 10 Megaohms.

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

Rationale: Isolating the instrument secondary power returns from the instrument primary power returns helps control common mode noise currents and minimizes magnetic field effects due to ground loops.

4.6.4 Internally Generated Secondary Return

CSPEC-101: The instrument shall reference its secondary returns (power and signal grounds) to the unit chassis ground by connecting them at one or more places.

Rationale: NASA-HDBK-4001 ELECTRICAL GROUNDING ARCHITECTURE FOR

UNMANNED SPACECRAFT

4.6.5 Mechanical Contact Resistance

CSPEC-103: 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.

4.6.6 Grounding

CSPEC-105: The DC resistance between the Coronagraph 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.

CSPEC-106: The Coronagraph shall provide a ground lug for a grounding strap to be attached from the Coronagraph Electronics Unit chassis for connection to the spacecraft conductive structure.

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

CSPEC-107: The grounding lug location on the Coronagraph Electronics Unit chassis or the tie points in contact with the ground strap shall 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.

4.6.7 Connector DC Resistance

CSPEC-109: Coronagraph connector backshells shall be electrically connected to Electronics Unit chassis with a DC resistance ≤ 2.5 milliohms.

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

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

4.7 Signal and Data Interfaces

4.7.1 Passive Analog Telemetry

CSPEC-112: The Coronagraph shall use no more than six (TBR) analog signals to monitor critical temperature points when the instrument is powered off.

Rationale: Common best practices for on-orbit operations deems that 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. Six signals are also deemed reasonable given the complexity of the instrument.

4.7.2 Data Signal Interface

The spacecraft will provide two data signal interface feeds to the instrument for command and telemetry. The spacecraft will also provide an interface line for a Pulse Per Second (PPS) time pulse.

CSPEC-114: The instrument science data transfer interface to the spacecraft shall be SpaceWire (ECSS-E-ST-50-12C-rev.1) data bus. For command and telemetry, the instrument may use either SpaceWire or a serial Universal Asynchronous Receiver Transmitter (UART) using ANSI/TIA/EIA-422-B Electrical Characteristics of Balanced Voltage Differential Interface Circuits.

Rationale: SpaceWire LVDS and RS-422 are the preferred standard high-rate data interfaces.

4.7.2.1 Telemetry Source Packet Format

CSPEC-117: 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 3.

Rationale: Use of CCSDS standards simplifies command and telemetry data handling system and interface design.

Figure 3 Telemetry Source Packet Definition

CSPEC-120: The instrument telemetry Source packets shall be variable length with a maximum data zone of 8192 octets including Secondary Header.

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

Rationale: CCSDS 133.0-B-2 Space Packet Protocol, Blue Book, Issue 1, September

CSPEC-121: The instrument shall set the telemetry source packet Secondary Header Flag to the value 1.

Rationale: CCSDS 133.0-B-2 Space Packet Protocol, Blue Book, Issue 1, September

CSPEC-122: The instrument shall set the Telemetry Source Packet Sequence Flags to the value of 11.

Rationale: CCSDS 133.0-B-2 Space Packet Protocol, Blue Book, Issue 1, September

Note: Segmentation services are not permitted.

CSPEC-123: 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 4.

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

Figure 4 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

CSPEC-130: 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 5.

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

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

Figure 5 Command Source Packet

CSPEC-133: The instrument Command Source Packets shall be variable length with a maximum data zone of 1013 octets.

Rationale: CCSDS 133.0-B-2 Space Packet Protocol, Blue Book, Issue 1, September

CSPEC-134: The instrument shall receive Command Source Packet with Secondary Header Flag set to the value 0.

Rationale: CCSDS 133.0-B-2 Space Packet Protocol, Blue Book, Issue 1, September

CSPEC-135: The instrument shall receive Command Source Packet with the Sequence Flags set to the value of 11.

Rationale: CCSDS 133.0-B-2 Space Packet Protocol, Blue Book, Issue 1, September

Note: Segmentation services are not permitted.

4.7.3 Clock Signal Interface

CSPEC-137: The instrument shall receive a Pulse Per Second (PPS) time pulse from spacecraft clock line by the interface defined in section 4.7.2 of this document.

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.

CSPEC-140: The instrument shall receive from the spacecraft a time code message on the data line as defined in Spacecraft Time Message Packet Figure 6. The time code message is the time applicable to receipt of the 1 PPS.

Rationale: CCSDS 133.0-B-2 Space Packet Protocol, Blue Book, Issue 1, September

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

Figure 6 Spacecraft Time Message Packet

4.7.4 Command and Housekeeping telemetry

CSPEC-144: 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

CSPEC-146: 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. An autonomous function is any task or series of tasks executed by the instrument without any commands, instructions, or other intervention from the ground or the spacecraft.

4.7.6 Limits and Triggers

CSPEC-148: The instrument autonomous limits and triggers shall be changeable by command.

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

4.7.7 On-Board Processor Reset

CSPEC-150: The instrument On-Board Processor shall be reset by command.

Rationale: The ability to command a software reset is a common best practice that dictates this capability for on-orbit anomaly troubleshooting, e.g. to terminate hung processes.

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

CSPEC-151: An autonomous power-on reset occurrence shall be unambiguously identifiable via telemetry.

Rationale: Common best practices dictate this capability for on-orbit anomaly troubleshooting

Note: This does not imply real-time telemetry as the reset is occurring.

4.7.8 Micro-switches

CSPEC-153: Micro-switches shall be used for information only and not used to initiate on-board autonomous activity or as an on-board interlock.

Rationale: Common best practices dictate this capability for on-orbit anomaly troubleshooting. Micro-switches are not reliable enough for autonomous actions.

4.8 Flight Software

4.8.1 Flight Load Non-volatile Memory

CSPEC-156: 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

CSPEC-158: The flight software modules shall be reprogrammable.

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

4.8.3 Software Table Updates

CSPEC-160: 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.

CSPEC-161: 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.

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

4.8.4 Flexibility and Ease of Software Modification

CSPEC-163: The instrument flight software shall be deterministic in terms of scheduling and prioritization of critical processing tasks to ensure their timely completion.

Rationale: To ensure that a critical task gets executed in the intended timeframe.

CSPEC-164: 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.

CSPEC-165: 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

CSPEC-167: 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.

CSPEC-168: This software identifier shall be configuration management controlled.

Rationale: Common best practice.

4.8.6 Warm Restart

CSPEC-170: The instrument flight software shall provide a restart by command with preservation of instrument configuration data and memory tables.

4.8.7 Memory Tests

CSPEC-172: The instrument flight software shall provide a mechanism to verify the contents of all memory areas.

Rationale: Common best practice.

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

4.8.8 Memory Dump

CSPEC-174: The instrument flight software, and associated on-board computer hardware, shall provide the capability to dump any memory location.

Rationale: Common best practice.

CSPEC-175: The flight software memory dump capability shall not disturb normal operations and instrument data processing.

Rationale: Common best practice.

4.8.9 Fault Detection and Correction Control

CSPEC-177: 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.

4.8.10 Health and Safety Monitoring

CSPEC-179: The instrument flight software shall provide health and safety monitoring, including memory checksum and watchdog timer.

Rationale: Common best practice for maintaining health and safety of the instrument.

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

5 PHYSICAL REQUIREMENTS

5.1 Interface and Design Units and Exclusions

CSPEC-182: The contractor shall use metric units to design the instrument and for developing interfaces with the spacecraft including any drawings, documents, models, except for the following cases:

Heritage Component or unit: Components or units 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.

5.2 Mass Properties Accuracy

5.2.1 Instrument Mass Accuracy

CSPEC-185: The mass of the instrument shall be measured to within ±0.1 kg.

5.2.2 Center of Mass Location

CSPEC-187: The contractor shall define the instrument center of mass in…