L1 Series STIS SPEC v2 (Draft).pdf

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L1 Series, Code L1SERIES-STIS-REQ-0009, Revision -

L1 Series Suprathermal Ion Sensor (STIS) Requirements Specification

(SPEC)

SWO CMO

November 21, 2023

DRAFT

U.S. Department of Commerce (DOC) National Oceanic and Atmospheric Administration (NOAA) NOAA Satellite and Information Service (NESDIS) National Aeronautics and Space Administration (NASA)

Effective Date:

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L1 Series Suprathermal Ion Sensor (STIS) Requirements Specification (SPEC)

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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

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Change History Log

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Deviations/Waivers Record

Section # / Rqmt.

Deviation / Waiver #

CCR # Date Approved

Description

None. None. None. None. None.

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Table of TBR/TBD/TBS

Item No.

Identifier Section Summary

1 STISPEC8 Applicable Documents (TBR)

2 STISPEC51 4.2.5

The STIS shall have a survival heater power consumption of ≤ 4W [TBR].

3 STISPEC74

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

4 STISPEC75

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 STIS contractor and command the STIS into safe mode when any of those key values have been exceeded.

5 STISPEC253 6.5

The STIS shall be designed to meet its performance requirements after being subjected to the shock environment in Figure 8 (TBR), applied at the STIS interface to the L1 Series spacecraft structure.

6 STISPEC261 6.6

The STIS shall be designed to withstand, without any damage or degradation of performance, when acoustic tested to the equivalent Protoflight (TBR) levels, which are +3db over those shown in Figure 9.

7 STISPEC277 6.8.2

The STIS shall be designed to meet all performance requirements after exposure to a maximum depressurization rate of -50 mbar/sec (-0.72 psi/sec) (TBR) experienced during launch and ascent.

8 STISPEC288 6.10

The STIS electronics box heat transferred to the spacecraft shall be limited to 4 Watts (TBR).

9 STISPEC951

The spacecraft interface temperature ranges (operational and survival) will be defined in the spacecraft to STIS ICD. (TBR) vii Check https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm to verify that this is the correct version prior to use.

10 STISPEC907 6.11.5

Shielded EEE parts and materials shall be selected according to the 95% confidence level dose-depth curve shown in Figure 11 (TBR) and Table 5.

11 STISPEC909 6.11.7

Silicon shielded EEE parts and materials shall be selected according to the 95% confidence level displacement damage dose-depth curve shown in Figure 11 (TBR) and Table 5 (TBR).

12 STISPEC910 6.11.7

Gallium Arsenide shielded EEE parts and materials shall be selected according to the 95% confidence level displacement damage dose-depth curve shown in Figure 12 (TBR) and Table 6 (TBR.)

13 STISPEC594 8.6.1

The STIS shall be designed for a 5-year continuous in-orbit, plus 6 months of commissioning, and 2-years (TBR) of ground storage period.

14 STISPEC718 10.2.1

The ESTE shall generate and maintain command logs

[TBR].

15 STISPEC726 10.2.2

The STIS emulator shall communicate with a spacecraft emulator for instrument command and telemetry using a standard interface (TBD) that matches the STIS instrument design.

16 STISPEC734 10.2.2

The STIS emulator shall be certified fully functional and operates correctly as intended on system using the United States Government Configuration Baseline (USGCB, formerly the Federal Desktop Core Configuration, FDCC). Reference https://csrc.nist.gov/Projects/United-States- Government-Configuration-Baseline/FAQs#yJUv.

(TBR)

17 STISPEC870 11.5.10

(2) The component shall be powered and critical parameters monitored using an LPT as defined in 7.2.3 (TBR) during chamber evacuation.

For RF components that are expected to be powered during launch, the absence of corona or multipaction effects shall be verified during instrument component thermal vac testing.

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Table of Contents

1 INTRODUCTION

1.1 General Information

1.2 Scope

2 APPLICABLE DOCUMENTS

3 CONTRACT DESCRIPTION

3.1 STIS Description

3.2 Ground Support Equipment Description

4 FUNCTIONAL/PERFORMANCE REQUIREMENTS

4.1 STIS Performance Requirements

4.1.1 Maximum and Minimum Flux

4.1.2 Flux Measurement Range

4.1.3 Accuracy

4.1.4 Field of View (FOV)

4.1.5 Refresh Rate and Latency

4.1.6 Response Stability

4.2 Resource Allocations

4.2.1 Mass

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 STIS 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 Normal Operational Mode

4.3.6 On-Orbit Operations

4.3.7 Station Keeping

4.3.8 Activation

4.4 Instrument Data Availability

4.5 Power

4.5.1 Voltage Range

4.5.2 Abnormal Voltages

4.5.3 Power Transients

4.5.4 Sudden Removal of Power

4.5.5 Over-Current Protection

4.5.6 Primary Power Return Ground

4.5.7 Turn-on Current Transients

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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 Microswitches

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 STIS 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

6.3.1 Fundamental Launch Frequencies

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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

6.9.1.1 Dynamic Linear Acceleration

6.9.1.2 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 Destructive Events

6.11.3.1 Single Event Gate Rupture (SEGRs), Single Event Burnout (SEB)

6.11.4 Charging Environment

6.11.5 Total Ionizing Dose

6.11.6 Displacement Damage Dose Levels

6.11.7 Single Event Effects

6.11.8 Solar Cycle Performance

6.11.9 Solar Particle Event Operations

6.11.10 Solar Flare Survival

7 CONTAMINATION CONTROL

7.1 STIS 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 Cleanability and Protection

7.7 Electrostatic Cleanliness

7.7.1 Conductive Surface Ground Path

7.7.2 Conductive Surface Resistivity

7.7.3 Closeout of Gaps and Apertures

7.7.4 Exposed Harness Specific Requirements

7.7.5 Thermal Blankets

8 DESIGN & 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 Current Emission Limits

8.3.1.3 CE03 – Differential Mode Current Emission Limits

8.3.1.4 CMBCE - Common Mode Bulk Conducted Emissions Limits

8.3.1.5 CEDTE - Conducted Emissions, Time Domain, Transients 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 and Signal Cables . 59

8.3.2.5 CS06 - Conducted Susceptibility, Transients, Power Leads

8.3.2.6 CS115 - Conducted Susceptibility, Bulk Cable Injection, Impulse Excitation ... 61

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

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8.7.3 Manual Lifting Hardware

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 & Cotter Pin Use Limitation

9.2.6.5 Snap Ring & 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

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10.2.2 STIS Emulator (STISE)

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

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 STIS Acceptance Random Vibration Environment Figure 8 Candidate LV Shock Environment Figure 9 L1 Series S/C Protoflight [TBR] Acoustic Envelope Figure 10 Allowable Analytical, Flight Operation, Flight Acceptance Test, and Qualification

(ProtoFlight or Prototype) Test Temperature Ranges Figure 11 Total Ionizing Dose -Depth Curves Figure 12 DDD in Si Dose-Depth Curve Figure 13 DDD in GaAs Dose-Depth Curve Figure 14 LET Spectrum of GCR Background Figure 15 Solar LET Spectrum Background Figure 16 Solar Proton Background Spectra Figure 17 Peak Solar Particle Event Flux Figure 18 Peak Solar Proton Flux Figure 19 CE101/CE102 (CE03) Differential Mode Current Emission Limits Figure 20 Common Mode Bulk Conducted Emissions Limit Figure 21 Inrush Current Transient Default Limit Figure 22 Power Lead Conducted Susceptibility (CS101) Voltage Limit Figure 23 Conducted Susceptibility Power Limit (30 Hz to 150 kHz) Figure 24 Conducted Susceptibility, Power Leads, 150 kHz to 50 MHz Limit Figure 25 Common Mode Conducted Susceptibility Calibration Limit Figure 26 Conducted Susceptibility Transient Waveform Figure 27 CS115 Impulse Waveform Figure 28 Unit Level RE102 Radiated Electric Field Emission Limits Figure 29 Thermal Vacuum Profile

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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 Suprathermal Ion Sensor (STIS) for the Space Weather Next L1 Series Mission.

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 STIS to the extent specified herein. In the event of conflict between this specification and any referenced document, this specification will govern, with the exception of the L1 Series STIS Statement of Work (L1SERIES-STIS-SOW-0004), 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-STIS-SOW-

L1 Series Suprathermal Ion Sensor (STIS) Statement of Work

(SOW)

L1SERIES-STIS-REQ-

L1 Series Suprathermal Ion Sensor (STIS) Contract Deliverables Requirements List (CDRL)

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-6016 Standard Materials and Processes Requirements for Spacecraft NASA-HDBK-7005 Dynamic Environment Criteria NASA-STD-7001 Payload Vibroacoustic Test Criteria IEST-STD-CC-1246E Product Cleanliness Levels and Contamination Control Program ASTM E-595-07 Standard Test Method for Total Mass Loss and Collected Volatile

Condensable Materials from Outgassing in a Vacuum Environment MIL-DTL-5541 Chemical Conversion Coatings on Aluminum and Aluminum

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

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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 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, June 2020 CCSDS 301.0-B-4 Time Code Formats. Blue Book. November 2010

3 CONTRACT DESCRIPTION

3.1 STIS Description

The Suprathermal Ion Sensor (STIS) is an ion spectrometer device that measures ions across a broad range of energies to characterize solar ejecta including Coronal Mass Ejections (CMEs), co-rotating interaction regions (CIRs) and interplanetary shocks. In particular, STIS is a low energy charged particle detector that is capable of measuring the ion flux population as a function of energy. This ion population with energies higher than that of the bulk plasma are produced by local solar acceleration as well as acceleration from a CME shock front. Analysis of these spectra can aid in estimating the arrival time and strength of CMEs shocks.

The STIS contract includes the delivery of two (2) Flight Models (FM) instrument and one (1) Engineering Development Unit (EDU), flight harnesses between instrument boxes, if applicable, and enough spares and built-up sub-assemblies for one additional FM, an electronics unit, and harnesses.

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3.2 Ground Support Equipment Description

The STIS contract includes the delivery of (2) sets of the Electrical System Test Equipment (ESTE), (4) STIS Emulators (STISEs), (1) Flight Software Development Environments (FSDEs), and (1) One 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 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).

4 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performances requirements for the L1 Series STIS.

The requirements in this Specification pertain to the STIS ‘system’, which may include all instrument hardware, software, and ground processing algorithms. The STIS 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.

4.1 STIS Performance Requirements

4.1.1 Maximum and Minimum Flux

STISPEC22: The STIS shall provide in situ solar wind low energy proton flux observations with flux observations in the following range ([cm^2 sec sr KeV]^-1):

Minimum Flux: 2.48x102 * E(KeV)-^1.6

Maximum Flux: 1.01x107 * E(KeV)-^1.6

4.1.2 Flux Measurement Range

STISPEC27: The STIS shall provide in situ solar wind low energy proton flux observation with observational energy range of 10 to 2,000 keV.

STISPEC28: The STIS shall determine the differential ion fluxes in sufficient evenly spaced logarithmic energy bands to meet the accuracy requirement defined in section 4.1.3.

4.1.3 Accuracy

STISPEC30: The STIS shall provide suprathermal proton flux observations with a flux accuracy of 20% over the required energy range.

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STISPEC928: The STIS instrument shall provide >3 bands in a logarithmic bin per energy decade.

4.1.4 Field of View (FOV)

STISPEC32: The spacecraft shall provide the STIS a FOV that is 80 degrees in the ecliptic by 60 degrees centered 50 degrees in the ecliptic off of the sun-earth line in the “ahead” direction (the ahead direction is the direction the earth travels in the ecliptic). The unit vector for the centroid of the FOV of the instrument should be (cos 50 degrees, 0, sin 50 degrees) or (0.643, 0,0 .766).

The spacecraft will provide an unobstructed field of view.

4.1.5 Refresh Rate and Latency

STISPEC35: The STIS shall provide in situ solar wind low energy proton flux observation at a refresh rate of 300 sec.

STISPEC36: The STIS shall provide in situ solar wind low energy proton flux observations that meet an overall latency requirement of 5 minutes.

STISPEC37: The STIS ground processing algorithms contribution to data latency of the STIS Level 1b algorithm output shall be ≤ 116 seconds.

4.1.6 Response Stability

STISPEC39: The STIS instrument shall provide monotonically increasing fluxes up to three times the maximum flux for a given energy range.

4.2 Resource Allocations

4.2.1 Mass

STISPEC42: The STIS, including the sensor unit, electronics, box, and sensor standoff bracket (if required), and the intra-instrument harness, shall have a combined mass of less than or equal to 5.2 kg.

STISPEC1940: The STIS instrument mass margin shall comply with GSFC-STD-1000H Rule 1.06.

4.2.2 Physical Envelope

STISPEC44: The STIS volume (length*width*height), including the sensor unit, electronics box, mounts, sensor standoff bracket (if required), intra-instrument harness, thermal blankets, and connectors, for both stowed and operational configurations shall have dimensions that do not

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exceed a total volume of 27,000 cm3. These dimensions pertain to both static and dynamic envelopes of the instrument.

STISPEC45: The STIS maximum dimension shall be less than 30 cm.

4.2.3 Operational Power Allocation

STISPEC47: The STIS shall have a nominal operational power of ≤4W.

4.2.4 Peak Operational Power Allocation

STISPEC49: The STIS shall have a peak (instantaneous) operational power of ≤ 8 W for a duration of 20 msec at 30 V.

4.2.5 Survival Heater Power Allocation

STISPEC51: The STIS shall have a survival heater power consumption of ≤ 4W [TBR].

4.2.6 Telemetry

STISPEC53: The STIS 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 1 kilobits per second (kbps) averaged over any 4.5 second period.

4.2.7 Transient Magnetic Field

STISPEC55: The STIS 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.

4.2.8 Static Magnetic Field

STISPEC57: The STIS units shall limit the static magnetic field to less than 1 nT in any axis 3.5 meters from any face of the unit for any operating mode.

4.3 STIS Operating Modes

STISPEC59: The STIS current operating mode shall be identified by a flag in its telemetry stream.

4.3.1 In-Flight Calibration

STISPEC61: The STIS shall provide an in-flight calibration mode that provides a test input to support calibration and testing both on the ground and in space.

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STISPEC63: The STIS in-flight calibration shall be both self-terminating and able to be terminated by a ground command.

4.3.2 Mode Transitions

STISPEC65: The STIS instrument shall transition from the current mode to any other mode without causing damage to itself.

4.3.3 Deterministic Power-on Configuration

STISPEC67: The STIS shall initialize upon power-up into a predetermined configuration.

STISPEC929: The STIS initialization configuration shall not enable any high voltage power supplies.

STISPEC930: High voltage power supplies shall be enabled by ground command only.

4.3.4 Safe Mode

STISPEC69: The STIS 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.

STISPEC70: In safe mode any STIS high-voltage power supplies shall be disabled. High voltage power supplies shall be re-enabled by ground command only.

STISPEC71: The STIS instrument shall be capable of remaining in a safe configuration for at least 120 hours without ground intervention.

4.3.4.1 Entry into Safe Mode

STISPEC73: The STIS instrument shall enter Safe Mode upon detection of internal faults that are capable of causing damage to the instrument.

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

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4.3.5 Normal Operational Mode

STISPEC79: In normal operational mode, the STIS instrument 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.

4.3.6 On-Orbit Operations

STISPEC81: The STIS instrument shall operate normally, within specification, while flying aboard a 3-axis stabilized spacecraft with orbital limit constraints as stated in this specification.

4.3.7 Station Keeping

STISPEC83: The STIS instrument shall continuously operate during all spacecraft maneuvers.

STISPEC84: The STIS instrument may not meet performance requirements during these maneuvers but shall meet them again after each maneuver is completed within 300 secs.

4.3.8 Activation

STISPEC86: The STIS instrument shall require no active commanding prior to fourteen (14) days after launch.

4.4 Instrument Data Availability

STISPEC1936: The STIS Instrument shall limit on-orbit calibrations to no more than four (4) calibrations per year with each calibration lasting no more than 30 minutes.

4.5 Power

4.5.1 Voltage Range

STISPEC89: The STIS 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).

STISPEC965: The STIS survival heaters shall operate over the voltage range of +26 to +34 VDC at the power inputs for all expected load conditions.

4.5.2 Abnormal Voltages

STISPEC91: The STIS shall survive without performance degradation after indefinite exposure to an anomalous voltage range of 0 to +40 VDC.

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4.5.3 Power Transients

STISPEC93: The STIS shall meet its performance requirements in the presence of transients specified in MIL-STD-461F, Figure CS115-1 and Figure CS116-2.

4.5.4 Sudden Removal of Power

STISPEC95: The STIS shall meet its performance requirements without degradation after exposure to an abrupt, unannounced removal of power.

4.5.5 Over-Current Protection

STISPEC97: The STIS shall not use non-resetting over-current protection (i.e., fuses) internal to the unit.

4.5.6 Primary Power Return Ground

STISPEC99: STIS shall provide a dedicated Primary Power return in the same connector as the primary power.

4.5.7 Turn-on Current Transients

STISPEC101: The STIS shall be capable of receiving limit turn-on input voltage (0 – 34V) with a rise time of 500us.

STISPEC103: The STIS instrument shall limit the instrument operational power turn-on current in-rush transient(s) to 1.0A for a period less than 20 milliseconds.

STISPEC104: Following initial in-rush, the STIS instrument shall limit the instrument operational power current transient(s) to 0.4A peak to peak.

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

4.6 Electrical Grounding

4.6.1 Primary Power DC Isolation

STISPEC111: The STIS 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

STISPEC113: The STIS survival heater power interfaces shall be isolated from the unit chassis by a DC resistance of greater than or equal to 10 Megaohms.

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4.6.3 Internally Generated Secondary to Primary DC Isolation

STISPEC115: Secondary power inputs shall be isolated from primary power by a DC resistance of greater than 10 Megaohms.

4.6.4 Internally Generated Secondary Return

STISPEC117: The STIS instrument shall reference its secondary returns (power and signal grounds) to the unit chassis ground by connecting them at one or more places.

4.6.5 Mechanical Contact Resistance

STISPEC119: 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 mΩ DC resistance.

4.6.6 Grounding

STISPEC122: The STIS shall provide a ground lug for a grounding strap to be attached from the STIS chassis for connection to the spacecraft conductive structure.

STISPEC123: The grounding lug location on the STIS instrument chassis or the tie points in contact with the ground strap shall be 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

STISPEC125: STIS connectors backshells shall be electrically connected to Electronic Unit chassis with a DC resistance ≤ 2.5 mΩ.

4.7 Signal And Data Interfaces

4.7.1 Passive Analog Telemetry

STISPEC128: The STIS shall provide no more than 6 analog signals to monitor critical temperature points when the STIS is powered off.

4.7.2 Data Signal Interface

STISPEC130: The STIS Instrument data transfer interface to or form 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.

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4.7.2.1 Telemetry Source Packet Format

STISPEC132: The STIS shall transfer all data to the spacecraft using the CCSDS 133.0-B-1 Section 4.1 Protocol Data Unit definition shown in the Telemetry Source Packet Definition Figure 1.

Figure 1 Telemetry Source Packet Definition

STISPEC134: The STIS telemetry Source packets shall be variable length with a maximum data zone of 8192 octets including Secondary Header.

STISPEC135: The STIS shall set the telemetry source packet Secondary Header Flag to the value 1.

STISPEC136: The STIS shall set the Telemetry Source Packet Sequence Flags to the value of 11.

Note: Segmentation services are not permitted.

STISPEC138: The STIS shall set the Telemetry Source Packet Time Code per CCSDS 301.B-4 Time Code Formats, Day Segmented format in the Time Code Format Figure 2.

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

STISPEC143: The STIS shall receive all data from the spacecraft formatted per CCSDS 133.0- B-1 Section 4.1 Protocol Data Unit definition shown in the Command Source Packet Definition Figure 3.

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Figure 3 Command Source Packet

STISPEC145: The STIS command Source packets shall be variable length with a maximum data zone of 1013 octets.

STISPEC146: The STIS shall receive Command Source Packet with Secondary Header Flag set to the value 0.

STISPEC147: The STIS shall receive Command Source Packet with the Sequence Flags set to the value of 11.

Note: Segmentation services are not permitted.

4.7.3 Clock Signal Interface

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

STISPEC150: The STIS shall receive a Pulse Per Second (PPS) time pulse from spacecraft clock line by the interface defined in STISPEC130

STISPEC152: The STIS shall receive from the spacecraft a time code message on the data line as defined in Spacecraft Time Message Packet Figure 4. The time code message is the time applicable to receipt of the 1 PPS.

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Figure 4 Spacecraft Time Message Packet

4.7.4 Command and Housekeeping telemetry

STISPEC155: The STIS instrument shall provide command and housekeeping telemetry functions in all powered modes.

4.7.5 Commands for Autonomous Functions

STISPEC157: The STIS shall execute commands to individually enable and disable each autonomous function.

4.7.6 Limits and Triggers

STISPEC159: The STIS autonomous limits and triggers shall be changeable by command.

4.7.7 On-Board Processor Reset

STISPEC161: The STIS On-Board Processor shall be reset by command.

STISPEC162: An autonomous power-on reset occurrence shall be unambiguously identifiable via telemetry. Note: This does not imply real-time telemetry as the reset is occurring.

4.7.8 Microswitches

STISPEC164: Micro-switches shall only be used for information and not use to initiate on-board autonomous activity or as an on-board interlock.

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4.8 Flight Software

4.8.1 Flight Load Non-volatile Memory

STISPEC167: The STIS flight software image shall be contained in its entirety in non-volatile memory at launch.

4.8.2 Software Updates

STISPEC169: The flight software modules shall be reprogrammable.

STISPEC170: Activation of uploaded modified software shall not require an upload of the entire flight software image.

4.8.3 Software Table Updates

STISPEC172: Instrument configuration data (e.g., Table Loads or Configuration Parameters) shall be reprogrammable during integration and test phases and on-orbit without computer restart.

STISPEC173: Modified instrument configuration data (e.g., Table Loads or Configuration Parameters) shall be committed to operational use by ground command.

4.8.4 Flexibility and Ease of Software Modification

STISPEC175: The STIS flight software shall be deterministic in terms of scheduling and prioritization of critical processing tasks to ensure their timely completion.

STISPEC176: 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.

STISPEC177: 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.

4.8.5 Version Identifiers

STISPEC179: The STIS software and firmware versions shall each have an internal identifier (embedded in the executive program) that can be included in the instrument engineering data.

STISPEC180: This software identifier shall be configuration management controlled.

4.8.6 Warm Restart

STISPEC182: The STIS 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

STISPEC184: The STIS flight software shall provide a mechanism to verify the contents of all memory areas.

4.8.8 Memory Dump

STISPEC186: The STIS flight software, and associated on-board computer hardware, shall provide the capability to dump any memory location.

STISPEC187: The flight software memory dump capability shall not disturb normal operations and instrument data processing.

4.8.9 Fault Detection and Correction Control

STISPEC189: If applicable, STIS shall provide the capability to enable and disable any internal Fault Detection and Correction (FDC) features.

4.8.10 Health and Safety Monitoring

STISPEC191: The STIS flight software shall provide health and safety monitoring, including memory checksum and watchdog timer, during integration and test phases and on-orbit.

5 PHYSICAL REQUIREMENTS

5.1 Interface and Design Units and Exclusions

STISPEC194: 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.

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5.2 Mass Properties Accuracy

5.2.1 STIS Instrument Mass Accuracy

STISPEC200: The mass of the STIS shall be measured to within ±0.1 kg.

5.2.2 Center of Mass Location

STISPEC202: The contractor shall define the center of mass in the MICD.

5.2.3 Center of Mass Accuracy

STISPEC204: The center of mass of the STIS shall be determined to within ±5 mm relative to a reference that will be defined in the MICD.

5.2.4 Determination of Moments and Products of Inertia

STISPEC206: The Contractor shall determine by analysis the launch and on-orbit moments and products of inertia to an accuracy of ±5.0 percent of the maximum principal moment of inertia, referenced to the coordinate axes with an origin at the center of gravity.

5.3 Mounting

5.3.1 Surface Flatness

STISPEC209: Mounting interface flatness, and co-planarity requirements for the STIS side of the interface (including brackets, if any, and shims) shall be as defined in the MICD.

5.3.2 Method

The method by which the STIS is mounted to the spacecraft will be defined in the Spacecraft to STIS Interface Control Document (ICD).

5.4 Coordinate System and Alignment

STISPEC213: The L1 Series observatories shall use the spacecraft on-orbit coordinate system which uses a right-hand orthogonal, body-fixed XYZ coordinate system as follows: the +X-axis is aligned with 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, Y, and Y axes, respectively. The origin of 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 align the STIS to within 0.5 deg in each axis.

STISPEC215: The STIS shall have clear fiducial marks to enable alignment during integration with 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.

6 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements for the STIS Instrument are specified in this section.

STISPEC219: The STIS shall meet its performance requirements in section 4.0 after exposure to the environments specified in this section.

6.1 Mechanical Factors of Safety

STISPEC221: The STIS 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:

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

Table 2 Factors of Safety

Type Static Sine Random/Acoustic 4,5 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

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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.

STISPEC225: Primary and secondary structure comprised of composite materials, Beryllium, bonded joints and/or bonded inserts shall be proof tested to 1.25 x Limit Load; qualification by analysis only is not acceptable. Actual flight 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.

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.

STISPEC228: The STIS 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 STIS 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 STIS 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

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6.3 Frequency Requirement

6.3.1 Fundamental Launch Frequencies

STISPEC237: The STIS 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 up to 100 Hz, to be used in coupled loads analyses. Requirements for the submitted finite element model are shown in the STIS SOW and discussed in Section 11.8.4 of this document.

6.4 Vibration

6.4.1 Sinusoidal Vibration

STISPEC240: The STIS 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

11.4.2 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.

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6.4.2 Random Vibration

STISPEC245: The STIS 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 STIS interface.

Figure 7 STIS Acceptance Random Vibration Environment

The random environment will be updated by NASA once more information is available at the mission level. Note for lightweight STIS, the highest design loads may be from this random vibration environment.

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STISPEC249: The contractor shall provide random vibration analysis along with static loads analysis. Please see NASA-HDBK-7005 and NASA-STD-7001 for more information.

During the random vibration test, the test input level will be reduced (notched) at critical frequencies, if required, to limit the random vibration loads and/or acceleration responses to 3 dB above design limit levels.

STISPEC251: Notching shall be limited to -12 dB of the original input and to a bandwidth of less than 100 Hz to limit the random vibration responses to 3dB above design limit levels.

Notching beyond these limits will require NASA/GSFC COR approval.

6.5 Shock

STISPEC253: The STIS shall be designed to meet its performance requirements after being subjected to the shock environment shown in Table 4 and Figure 8, applied at the STIS interface to the L1 Series spacecraft structure. The shock environment will not exceed the values shown in Table 4 and Figure 8.

The spacecraft contractor will develop an attenuated shock environment for the STIS to spacecraft interface and document that environment in the spacecraft to STIS ICD.

Note: Figure 8 is the acceptance shock input at the L1 Series S/C to Launch Vehicle…