Attachment B Requirements Specification.pdf

Effective Date : Oct. 21, 2019 422-L1-MAGSPEC-0008

Expiration Date : TBD Version 0.1

Responsible Organization: SWFO-L1/Code 422

422-L1-MAGSPEC-0008

Version 0.1

Space Weather Follow On - Lagrange 1 (SWFO – L1)

Code 422

SWFO – L1

Magnetometer (MAG)

Requirements Specification (SPEC)

Goddard Space Flight Center

Greenbelt, Maryland

National Aeronautics and Space Administration

SWFO Magnetometer SPEC 422-L1-MAGSPEC-0008 ii

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SWFO – L1 Magnetometer Specification Requirements (SPEC)

Signature/Approval Page

Prepared by:

Michael Honaker Date

SWFO-L1 Magnetometer Instrument Engineer

NASA/GSFC, Code 422

Reviewed by:

Ronald J. Hooker, Sc.D Date

SWFO-L1 Instrument Systems Manager

Approved by:

Gene Martin Date

SWFO-L1 Project Manager iii

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CM FOREWORD

This document is a Space Weather Follow On - Lagrange 1 (SWFO-L1) Project Configuration

Management (CM)-controlled document. Changes to this document require prior approval of the applicable Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the SWFO-L1 CM Office (CMO), along with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

Questions or comments concerning this document should be addressed to:

NASA/Goddard Space Flight Center

SWFO-L1 Project Office, Code 422

Attention: Configuration Management Office

Greenbelt, Maryland 20771 https://goessp.ndc.nasa.gov/ iv

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Change History Log

REV

LEVEL

DESCRIPTION OF CHANGE APPROVED

BY

DATE

APPROVED

0.0 Initial release

0.1

In section 3.1, clarify that the magnetometer electronics can consist of a single unit per sensor if needed (2 maximum).

Rewrote section 8.9.2 to clarify the 1000-hour operating time requirement.

v

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Table of TBDs/TBRs/TBPs

Action Item

No.

Location Summary Individual/

Organization

Actionee vi

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Table of Contents

1.0 Introduction

1.1 General Information

1.2 Scope

2.0 Applicable Documents

3.0 Contract Description

3.1 Magnetometer Description

3.2 Ground Support Equipment Description

4.0 Functional/Performance Requirements

4.1 Magnetometer Performance Requirements

Operational Measurement Range Measurement Accuracy

Orientation Accuracy Ground Test Measurement Range Refresh Rate and Latency Noise

Quantization Error

4.2 Resource Allocations

Mass Allocation Nominal Power Allocation Peak Power Allocation

Survival Heater Power Allocation Telemetry

Instrument-to-Spacecraft Volume Transient Magnetic Field

Static Magnetic Field

4.3 Magnetometer Operating Modes

Deterministic Power-on Configuration Safe Mode

4.3.2.1 Provision of Safe Mode

4.3.2.2 Entry into Safe Mode

Fail-safe Recovery Mode Normal Operational Mode On-Orbit Operations

Station Keeping Selective Performance/Contingency Operating Mode

4.4 Power

Voltage Range Abnormal Voltages Power Transients Sudden Removal of Power

Over-Current Protection Primary Power Return Ground Turn-on Current Transients

Operational Current Transients vii

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4.5 Electrical Grounding

Primary Power DC Isolation Survival Power Isolation

Internally Generated Secondary to Primary DC Isolation Internally Generated Secondary Return Mechanical Contact Resistance Grounding Connector DC Resistance

4.6 Signal And Data Interfaces

Passive Analog Telemetry Data Signal Interface

4.6.2.1 Telemetry Source Packet Format

4.6.2.2 Command Source Packet Format

Clock Signal Interface Command and Housekeeping telemetry

Commands for Autonomous Functions Limits and Triggers

On-Board Processor Reset Microswitches

4.7 Flight Software

Flight Load Non-volatile Memory Software Updates

Software Table Updates Flexibility and Ease of Software Modification

Version Identifiers Warm Restart

Memory Tests Memory Dump

Fault Detection and Correction Control Health and Safety Monitoring

5.0 Physical Requirements

5.1 Interface and Design Units and Exculsions

5.2 Mass Properties Accuracy

Magnetometer Instrument Mass Accuracy Center of Mass Location Center of Mass Accuracy Determination of Moments and Products of Inertia

5.3 Mounting

Surface Flatness Method

5.4 Coordinate System and Alignment

6.0 Environmental Requirements

6.1 Mechanical Factors of Safety

6.2 Quasi-Static Acceleration

6.3 Frequency Requirement

Fundamental Launch Frequencies viii

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6.4 Vibration

Sinusoidal Vibration Random Vibration

6.5 Shock

6.6 Acoustics

6.7 Transportation

Transportation Cleanliness

6.8 Pressure

Operating Pressure Range Maximum Depressurization Rate Launch Vehicle (LV) Environmental Control System (ECS) Impingement

6.9 On-Orbit Dynamic Environment

6.10 Thermal Requirements

Flight Interface Design Temperature Limits

6.11 Charged Particle Radiation Requirements

Definitions Non-Destructive Events (SEUs, SETs, SEFIs, SHEs, and MBUs)

6.11.2.1 Single-Event Effect Rate Calculations

Destructive Events

6.11.3.1 Single Event Latchups (SELs)

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

Charging Environment

Total Ionizing Dose Displacement Damage Dose

7.0 Contamination control

7.1 Magnetometer Surface Cleanliness Requirements

7.2 Contamination Generation

Particulate Generation

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

Conductive Surface Ground Path Conductive Surface Resistivity Closeout of Gaps and Apertures

Exposed Harness Specific Requirements Thermal Blankets

8.0 Design and Construction Requirements

8.4 Electrical

Test Sensors MGSE Grounding Connector Specifications

8.4.3.1 Contact Derating

ix

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8.4.3.2 Redundant Contact Derating

8.4.3.3 Signal Segregation

8.4.3.4 Test and Flight Signal Isolation

Test Interfaces

8.4.4.1 Facility-Induced Noise

8.4.4.2 Facility-Induced ESD GSE Malfunction

8.4.4.3 Facility-Induced GSE Malfunction

Mitigation of Internal Charging

8.4.5.1 Mitigation Strategies for Internal Charging

8.4.5.2 Floating Conductors

8.4.5.3 Dielectric Structures

8.5 Safety

8.6 Electromagnetic Compatibility

Conducted Emissions

8.6.1.1 Applicability of Conducted Emissions

8.6.1.2 CE101 – Differential Mode Current Emission Limits

8.6.1.3 CE102 – Differential Mode Current Emission Limits

Common Mode Current Emissions Limits Conducted Susceptibility

8.6.3.1 Applicability of Conducted Susceptibility

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

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

8.6.3.4 CS114 – Power and Signal Cables, Common Mode

Radiated Emissions

8.6.4.1 RE102 - Electric Field Emissions

Radiated Susceptibility

8.6.5.1 RS103 - Electric Field

8.7 Identification and Marking

8.8 Workmanship

Connectors

8.9 Reliability and Mission Lifetime

Mission Life Operating Time

Trouble-Free Time

8.10 Ground Handling

Ground Support Equipment (GSE) Design Lifting Hardware Manual Lifting Hardware

GSE Cleanliness GSE Bakeout

Test Harness

8.11 Interface Documentation

Mechanical Interface Electrical Interface Data Interface

9.0 Mechanical Design Requirements

x

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9.1 Structural Requirements

Component Fatigue Fracture Control Requirements

9.2 Fastening Systems

Factors of Safety Supplemental Factor Ultimate Design Loads Yield Design Loads

Design Separation Load 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

Fastened Joints Criteria

10.0 Logistics

10.4 I&T Deliverables

10.5 Ground Support Equipment

Electrical System Test Equipment Magnetometer Emulator (STISE)

10.6 Transportation Equipment

Shipping Container

11.0 Verification Requirements

11.4 Verification Methods

Inspection Analysis

Test

11.5 Inspection Requirements

Visual Inspection

Physical Measurement Documentation Search

11.6 Analysis Requirements

11.7 Test Requirements

Definitions Test Factors Test Tolerances

Test Restrictions

11.8 Required Tests

Performance Tests Mass Properties Measurement Static Loads/Strength Test Sine Sweep Survey Sine Vibration Random Vibration xi

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Acoustic Test

Shock Thermal Vacuum Bake-out

Thermal Vacuum Test Magnetics Test Harness Tests EMI/EMC Tests

Appendix A ABBREVIATIONS and Acronyms xii

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List of Figures

Figure Page

Figure 6-1 Sine Vibration Environment [TBR]

Figure 6-2 Mag Acceptance [TBR] Random Vibration Environment Figure 6-3 SWFO-L1 S/C Acceptance [TBR] Shock Envelope Figure 6-4 SWFO-L1 S/C Protoflight [TBR] Acoustic Envelope Figure 6-5 Total ionizing dose-depth curve for a 5-year mission at L1 Figure 6-6 Silicon displacement damage dose-depth curve for a 5-year mission at L1

Figure 6-7 Gallium Arsenide displacement damage dose-depth curve for a 5-year mission at L1 ..

Figure 8-1 CE101/CE102 Differential Mode Current Emission Limits Figure 8-2 Common Mode Current Emission Limits Figure 8-3 Power Lead Conducted Susceptibility (CS101) Voltage Limit Figure 8-4 Unit Level RE102 Radiated Electric Field Emission Limits [TBR]

Figure 11-1 Thermal Vacuum Profile

List of Tables

Table Page

Table 2-1 Applicable Documents

Table 6-1 Factors of Safety Table 6-2 Mag Design Limit Loads [TBR]

Table 6-3 Qualification [TBR] Level Shock Response Spectrum Table 6-4 Transportation Loads Table 6-5 Temperature Limits at Instrument Mounting Interface Table 6-6 Dose requirements as a function of shielding for TID in silicon, DDD in silicon and

DDD in gallium arsenide Table 7-1 Magnetometer Lifetime Contamination Requirements Table 7-2 Thermal Blanket Area vs. Grounding Tabs Table 8-1 Emission and Susceptibility Requirements Table 8-2 RS103 Radiated Susceptibility Levels

Table 11-1 Test Factors and Durations

Table 11-2 Test Tolerances

Table 11-3 Thermal Vacuum Test Parameters

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1.0 INTRODUCTION

1.1 GENERAL INFORMATION

1.2 SCOPE

This specification describes the performance and electrical, mechanical, environmental, and verification testing requirements for a space-qualified Magnetometer Insturment for the Space

Weather Follow-On at Lagrange 1 (L1) Mission (SWFO-L1).

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2.0 APPLICABLE DOCUMENTS

The following documents and drawings in effect on the day this specification was signed shall apply to the fabrication and to the electrical, mechanical, and environmental requirements of the

Magnetometer 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 SWFO-L1

Magnetometer Statement of Work (422-L1-MAGSOW-0007), in which case the Statement of

Work takes precedence.

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

Table 2-1 Applicable Documents

Document Number Title

422-L1-MAGSOW-0007 SWFO-L1 Magnetometer Statement of Work

422-L1-MAGCDRL-0009 SWFO-L1 Magnetometer Contract Deliverables Requirements List (CDRL)

ANSI/TIA/EIA-422-B Electrical Characteristics of Balanced Voltage Digital Interface Circuits

ANSI/TIA/EIA-644-A-

Electrical Characteristics of Low Voltage Differential Signaling (LVDS)

Interface Circuits

NFPA 70 National Fire Protection Association National Electric Code

NASA-STD-5001B Structural Design And Test Factors Of Safety For Spaceflight Hardware

NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety Requirements

NASA-STD-6016 Standard Materials and Processes Requirements for Spacecraft

NASA-HDBK-7005 Dynamic Environment Criteria

NASA-STD-7001 Payload Vibroacoustic Test Criteria

IEST-STD-CC-1246E Product Cleanliness Levels And Contamination Control Program

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

Condensable Materials from Outgassing in a Vacuum Environment

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

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

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-7000A General Environmental Verification Standard (GEVS)

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

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

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

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

FAA AC 20-71 Federal Aviation Administration Advisory Circular (AC) 20-71, “Dual

Locking Devices on Fasteners".

NASM 33540 Safety Wiring, Safety Cabling, Cotter Pinning, General Practices for

SAE AS567 Safety Cable, Safety Wire, Key Washers, and Cotter Pins for Propulsion

Systems, General Practices for Use of

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

Cracking Resistance in Sodium Chloride Environments

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

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Document Number Title

CCSDS 133.0-B-1 Space Packet Protocol, Blue Book, Issue 1, September 2003

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

3.0 CONTRACT DESCRIPTION

3.1 MAGNETOMETER DESCRIPTION

The Magnetometer instrument (Mag) or suite consists of two magnetometer sensor units (each containing a three-axis magnetic field sensor), up to two electronic unit(s) (one per sensor if needed) and the associated interconnecting cables. It is expected that the magnetometer sensors will be mounted on a deployable boom provided by the SWFO-L1 spacecraft contractor. The exact location, mounting configuration and distances between the sensors and the electronics will be negotiated between the magnetometer contractor and the spacecraft contractor in order to develop the required interface control documents for the magnetometer.

The term “Magnetometer Instrument” or suite includes the magnetometer sensors, electronics, cabling between sensors and electronics, up to the spacecraft provided interface. The

“Magnetometer Electronics” includes the electronic components and may consist of a single box for both sensor units or two separate boxes, one for each sensor unit. Each “Magnetometer

Sensor” includes a tri-axial sensor. The “Magnetometer Harness” includes the Magnetometer

Instrument supplier provided harness and cabling necessary to connect each electronics/sensor pair.

The Magnetometer contract includes the delivery of a flight magnetometer suite (including the two (2) flight sensors, flight electronics unit(s) and flight harness(es) between the instrument boxes), one (1) Engineering Development Unit (EDU) magnetometer suite, and enough spares and built-up sub-assemblies for two additional FM sensors, electronics unit and harnesses.

3.2 GROUND SUPPORT EQUIPMENT DESCRIPTION

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

(ESTE), (2) Magnetometer Emulators (MAGEs), and (1) One Ground Processing Development

System (GPDS). Additionally, Mechanical Ground Support Equipment (MGSE) such as lifting fixtures/handles, shipping containers, 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 Magnetometer suite delivery includes ground support equipment required to power and operate the magnetometers at the vendor’s and spacecraft contractor’s facilities. The delivery also includes standalone emulator 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 Contract Deliverables Requirements List (CDRL).

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4.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the SWFO MAG.

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

4.1 MAGNETOMETER PERFORMANCE REQUIREMENTS

Operational Measurement Range

The Mag shall have an in space measurement range of at least ±250nT in each of the three orthogonal axes.

Measurement Accuracy

The Mag shall measure each Cartesian component of the magnetic field to an accuracy of ≤ ±0.5 nT [TBR] when measuring a magnetic field with magnitude ≤ 100 nT.

The Mag shall measure each component of the magnitude of magnetic field to an accuracy (after in-flight calibration and temperature correction for gains and offsets) of ≤ ± 0.5% [TBR] when measuring a magnetic field with magnitude >100 nT.

This accuracy requirement must be met with all Mag error sources included for the duration of the mission. One calibration maneuver per month (s/c roll about the sun line) can be assumed.

It is expected that a calibration period for the Mag will occur following launch (during the cruise phase) where inflight calibration will be performed.

Orientation Accuracy

The relative orientation Mag axes shall be known and stable to better than +/- 0.1 degrees in the local instrument frame. Absolute orientation of the Mag axes shall be calibrated relative to the instrument mounting surface to better than +-/0.1 degrees.

Ground Test Measurement Range

The Mag shall have for ground based testing an ambient measurement range of at least ±65 𝜇𝑇 in each of the three orthogonal axes.

Refresh Rate and Latency

The SWFO-L1 mission will make Solar Wind data (including data from the Mag) available to

SWPC forecasters within 5 minutes of the completion of each data acquisition.

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The mag instrument shall sample and send to the spacecraft one magnetic field vector per second.

The Mag ground processing algorithm (ref Sec 4.0) shall produce (output) an ambient magnetic field vector measurement ≤ 60 seconds derived from the 1 Hz magnetic field measurements.

The Mag instrument measurements shall be timed-tagged to an accuracy ≤ 100 msec relative to the 1 PPS spacecraft reference signal.

The processing of a completed on-orbit data collection by the Mag and the transmission to the spacecraft shall contribute ≤ 2 [TBR] seconds to the overall latency of the solar wind data product.

The Mag ground processing algorithms contribution to data latency of the Mag Level 1b algorithm output shall be ≤ 236 [TBR] seconds.

Noise

The Mag noise shall be less than 0.137 nT rms when measured as the square root of the integral of the power spectral density between 0.05 and 0.5 Hz.

Quantization Error

The Mag shall have quantization error of ≤0.05 nT while measuring a field in the ±250 nT range

[TBR].

The Mag shall have quantization error of ≤13 nT while measuring a field in the ±65000 nT range

4.2 RESOURCE ALLOCATIONS

Mass Allocation

The Magnetometer Sensor Unit shall have a mass of ≤0.5 kg each [TBR].

The Magnetometer Electronics Unit(s) shall have a mass of ≤1.0 kg [TBR].

The Magnetometer Electronics Unit to Sensor Units Cable shall have a mass of ≤ 1.0 kg [TBR].

Nominal Power Allocation

The Mag shall have a nominal (continuous) operational power, including operational heater power, of ≤ 3.4 W.

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Peak Power Allocation

The Mag shall have a peak (instantaneous) power consumption of less than or equal to 5 W

[TBR] for a duration of 60 seconds.

Survival Heater Power Allocation

The Mag shall includes survival heaters with power consumption of less than or equal to 3.4 W

[TBR].

Telemetry

The Mag 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 400 [TBR] bits per second (bps) averaged over five [TBR] seconds.

Instrument-to-Spacecraft Volume

Each Magnetometer Sensor Units volume (length*width*height), including mounts, thermal blankets and connectors, for both stowed and operational configurations shall have dimensions that do not exceed a total volume of 1000 cm3 [TBR] each.

The Magnetometer Electronics Unit(s) total volume (length*width*height), including mounts, thermal blankets and connectors, for both stowed and operational configurations shall have dimensions that do not exceed a total volume of 9000 cm3 [TBR].

Transient Magnetic Field

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

Static Magnetic Field

The Mag 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 [TBR].

4.3 MAGNETOMETER OPERATING MODES

The Mag shall be able to be commanded to any one of its dynamic ranges, if it has more than one. If the magnetometer has any diagnostic modes (e.g. a mode to cycle through available dynamic ranges), such a mode shall be both self-terminating and able to be terminated by a ground command.

The Mag shall provide an in-flight calibration that allows scale factor to be determined.

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

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The Mag’s current operating mode shall be identified by a flag in its telemetry stream.

Deterministic Power-on Configuration

The Mag On-Board Processor shall initialize upon power-up into a predetermined configuration.

Safe Mode

4.3.2.1 Provision of Safe Mode

The Mag 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.

The Mag shall be capable of remaining in a safe configuration for at least 120 hours without ground intervention.

4.3.2.2 Entry into Safe Mode

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

The SWFO-L1 spacecraft will monitor up to four analog instrument health and safety parameters defined by the Mag contractor and command the Mag into safe mode when any of those key values have been exceeded.

The SWFO-L1 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 Mag contractor and command the Mag into safe mode when any of those key values have been exceeded.

Fail-safe Recovery Mode

The Mag 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.

Normal Operational Mode

In normal operational mode, the Mag 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 and in which both magnetometer sensors are operating.

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On-Orbit Operations

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

Station Keeping

The Mag shall continuously operate during all spacecraft maneuvers.

Selective Performance/Contingency Operating Mode

The Mag shall have a mode in which either of the 3-axis sensors can be operated independently.

4.4 POWER

The spacecraft will provide a total of up to two power feeds to the Magnetometer Electronics

Units(s) if needed.

Voltage Range

The Mag shall operate over the bus voltage range of +27 to +35 VDC at the primary power inputs during all normal mission phases and for all expected load conditions (except when turned off).

Abnormal Voltages

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

Power Transients

The Mag shall meet its performance requirements in the presence of transients specified in MIL-

STD-461G, Figure CS115-1 and Figure CS116-2.

Sudden Removal of Power

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

Over-Current Protection

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

Primary Power Return Ground

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

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Turn-on Current Transients

The Mag shall be capable of receiving limit turn-on input voltage (0 – 35V) with a rise time of

500us or longer.

The Mag shall limit its instrument operational power turn-on current ramp rate to less than

2A/microsecond.

The Mag shall limit the instrument operational power turn-on current in-rush transient(s) to 1.0A

[TBR] for a period less than 20 milliseconds [TBR].

Following initial in-rush, the Mag shall limit the instrument operational power current transient(s) to 0.4A peak to peak [TBR].

Note: The spacecraft bus will not drop below 27V during the Magnetometer instrument turn on.

Operational Current Transients

The Mag shall limit any change in operational power current at any time (including initial power turn-on) to less than 0.2A/µs [TBR].

The Mag shall limit the maximum delta change in operational current at any time to 0.1 A

4.5 ELECTRICAL GROUNDING

Primary Power DC Isolation

The Mag 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.

Survival Power Isolation

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

Internally Generated Secondary to Primary DC Isolation

Secondary power (or signal) inputs shall be isolated from primary power by a DC resistance of greater than 10 Megaohms.

Internally Generated Secondary Return

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

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Mechanical Contact Resistance

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.

Grounding

The DC resistance between the Mag chassis and the observatory chassis shall be ≤ 2.5 milliohms.

The Mag shall provide a ground lug for a grounding strap to be attached from the Mag chassis for connection to the spacecraft conductive structure [TBR].

The grounding lug location on the Mag chassis or the tie points in contact with the ground strap shall have a minimum contact area of 80 mm². The ground lug contact area must remain free of any material finish that may affect the reliability of the ground connection and will be shown in the MICD.

Connector DC Resistance

Mag connectors shall be electrically connected to chassis with a DC resistance ≤ 2.5 mΩ.

4.6 SIGNAL AND DATA INTERFACES

Passive Analog Telemetry

The Mag shall utilize no more than 6 [TBR] analog signals to monitor critical temperature points when the Magnetometer is powered off.

Data Signal Interface

The Mag shall interface for data transfer to or from the spacecraft by either RS-422 or Low

Voltage Differential Signal ECS-E50-12A (LVDS) data bus complying with (TBP), or

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, as applicable.

The spacecraft will provide a total of up to two data signal interface feeds to the Magnetometer

Electronics Unit(s).

4.6.2.1 Telemetry Source Packet Format

The Mag 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.

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TELEMETRY SOURCE PACKET DEFINITION

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

The Mag shall set the telemetry source packet Secondary Header Flag to the value 1.

The Mag shall set the Telemetry Source Packet Sequence Flags to the value of 11.

Note: Segmentation services are not permitted.

The Mag 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.

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

The spacecraft on-board reference time will be International Atomic Time (TAI) [TBR].

4.6.2.2 Command Source Packet Format

The Mag 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.

COMMAND SOURCE PACKET FIGURE

SEC.

HDR

FLAG

13 - 8K octets

3 bits 1 bit 1 bit 11 bits 2 bits 14 bits 16 bits 72 bits

DATA

VARIABLE

TIME CODE PACKET

LENGTH

SECONDARY

HEADER

PACKET

SEQUENCE

COUNT

SEQUENCE

FLAGS

PRIMARY HEADER

PACKET IDENTIFICATION PACKET SEQUENCE CONTROL

VERSION

NUMBER

TYPE APPLICATION

PROCESS ID

SEC.

HDR

FLAG

13 - 8K octets

3 bits 1 bit 1 bit 11 bits 2 bits 14 bits 16 bits

DATA

VARIABLE

PACKET

LENGTH

PACKET

SEQUENCE

COUNT

SEQUENCE

FLAGS

PRIMARY HEADER

PACKET IDENTIFICATION PACKET SEQUENCE CONTROL

VERSION

NUMBER

TYPE APPLICATION

PROCESS ID

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The Mag command Source packets shall be variable length with a maximum data zone of 8192 octets.

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

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

Note: Segmentation services are not permitted.

Clock Signal Interface

The Mag shall receive a Pulse per Second time code sequence from spacecraft clock line by either RS-422 or Low Voltage Differential Signal (LVDS) data bus complying with

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, as applicable.

The spacecraft will provide the Mag a 1 PPS time code sequence accurate to +/- 1 msec [TBR] relative to TAI.

The spacecraft will provide a total of up to two clock interface feeds to the Magnetometer

Electronics Unit(s).

The Mag shall receive from the spacecraft a time code message on the data line as defined in

Spacecraft Time Message Packet Figure. The time code message is the time applicable to receipt of the 1 PPS.

Command and Housekeeping telemetry

The Mag instrument shall provide command and housekeeping telemetry functions in all powered modes.

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Commands for Autonomous Functions

The Mag shall execute commands to individually enable and disable each autonomous function.

Limits and Triggers

The Mag autonomous limits and triggers shall be changeable by command.

On-Board Processor Reset

The Mag On-Board Processor shall be reset by command.

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.

Microswitches

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

4.7 FLIGHT SOFTWARE

Flight Load Non-volatile Memory

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

Software Updates

The flight software modules shall be reprogrammable.

Activation of uploaded modified software shall not require an upload of the entire flight software image [TBR].

Software Table Updates

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

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

Flexibility and Ease of Software Modification

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

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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.

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 [TBR].

Version Identifiers

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

This software identifier shall be configuration management controlled.

Warm Restart

The Mag flight software shall provide a restart by command with preservation of instrument configuration data and memory tables.

Memory Tests

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

Memory Dump

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

The flight software memory dump capability shall not disturb normal operations and instrument data processing [TBR].

Fault Detection and Correction Control

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

Health and Safety Monitoring

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

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5.0 PHYSICAL REQUIREMENTS

5.1 INTERFACE AND DESIGN UNITS AND EXCULSIONS

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

Magnetometer Instrument Mass Accuracy

The mass of the Mag shall be measured to within ±0.1 kg.

Center of Mass Location

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

Center of Mass Accuracy

The center of mass of the Mag shall be determined to within ±5 mm [TBR] relative to a reference that will be defined in the MCID.

Determination of Moments and Products of Inertia

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

5.3 MOUNTING

Surface Flatness

Mounting interface flatness, and co-planarity requirements for the Mag side of the interface

(including brackets, if any, and shims) shall be as defined in the MICD.

Method

The method by which the Mag is mounted to the spacecraft will be defined in the Spacecraft to

Magnetometer ICD.

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5.4 COORDINATE SYSTEM AND ALIGNMENT

The SWFO spacecraft on-orbit coordinate system uses a right-hand orthogonal, body fixed XYZ coordinate system (Spacecraft Reference Frame - SRF) as follows: the positive XSRF-axis points toward the Sun, the positive YSRF-axis points towards the north ecliptic pole and the ZSRF axis is obtained from the right hand rule. The roll, pitch and yaw rotations are defined about the X, Z, and Y axes, respectively. The origin of coordinate system is at the center of the ESPA ring attachment plane.

The spacecraft will align the Mag sensor axes to within +/-2 [TBR] degrees of the nominal spacecraft reference frame axes.

The Mag 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.

The instrument unit mounting frame shall be parallel to the spacecraft reference frame [TBR].

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6.0 ENVIRONMENTAL REQUIREMENTS

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

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

6.1 MECHANICAL FACTORS OF SAFETY

The Mag 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 6-1. Margin of

Safety (MS) is defined as follows:

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

Table 6-1 Factors of Safety

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.

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6.2 QUASI-STATIC ACCELERATION

Quasi-static acceleration represents the combination of steady-state accelerations and the low frequency mechanically transmitted dynamic accelerations that occur during launch.

The Mag shall be designed to withstand the quasi-static design limit loads defined in the mass-acceleration curve (MAC) shown in Table 6-2 without damage or degradation of performance 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 Magnetometer weight. Note that these design limit loads are intended to cover only the low frequency launch environment and must be used in conjunction with the random vibration environments to assess structural margins.

Table 6-2 Mag Design Limit Loads [TBR]

MAC Breakpoints

Mass (kg) GSFC (g) Moog* RSS (g)

1.0 51.3 125.7

5.0 42.4 60.8

10.0 36.6 44.5

20.0 30.3 32.5

40.0 24.2 23.8

60.0 20.9 19.8

80.0 18.7 17.4

100.0 17.2 15.7

120.0 15.9 14.5

140.0 15.0 13.5

160.0 14.2 12.7

181.0 13.5 12.0

200 12.9 11.5

300 10.8 9.6

400 9.5 8.4

Notes:

1. Moog MAC values are taken from “ESPA User’s Guide, The EELV Secondary Payload

Adapter, November 2018

2. Numbers in red are extrapolated from the Moog MAC

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6.3 FREQUENCY REQUIREMENT

Fundamental Launch Frequencies

The Mag shall have a fundamental frequency greater than 100 Hz [TBR] 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 50 Hz, to be used in coupled loads analyses. Requirements for the submitted finite element model are shown in the

Magnetometer SOW and discussed in Section 11.8.4 of this document.

6.4 VIBRATION

Sinusoidal Vibration

The Mag 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 6-1. Instruments with a first mode greater than 150 Hz can be exempted from sine vibration testing upon approval by the NASA/GSFC COR. A generic sine vibration specification is provided for acceptance levels applied at the SWFO-L1 to Mag interface is shown (Note: in Figure 6-1, axial refers to the launch vehicle thrust axis and lateral to either of the other two orthogonal directions). See Section 11.7.1 for definitions of Protoflight, Qualification, and Acceptance.

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Figure 6-1 Sine Vibration Environment [TBR]

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.

Random Vibration

The Mag shall demonstrate its ability to meet its performance requirements after being subjected to the random vibration environment in Figure 6-2 for units weighing 22.7 kg (50 lb.) or less, applied at the Spacecraft to Mag interface.

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Figure 6-2 Mag Acceptance [TBR] 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.

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.

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

The Mag shall be designed to meet its performance requirements after being subjected to the shock environment in Table 6-3, applied at the Magnetometer interface to the SWFO-L1 spacecraft structure.

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Table 6-3 Qualification [TBR] Level Shock Response Spectrum

Freq (Hz) SRS (G)

100 100

800 1700

8000 1700

10000 2200

A shock susceptibility and attenuation assessment (analysis) shall be performed on all Spacecraft units or instruments. If the flight shock environment as shown on a Shock Response Spectra

(SRS) plot (Q=10) is enveloped by the curve shown below, then the shock environment can be considered benign and there is low risk in deferring the shock test to the Observatory level and the Magnetometer contractor can request a waiver to not perform the shock test at the instrument level. Analysis supporting this conclusion (i.e. deferral of shock testing) shall be provided to the

NASA/GSFC COR for review and approval with the waiver request.

Figure 6-3 SWFO-L1 S/C Acceptance [TBR] Shock Envelope

Any unit determined to be susceptible to the shock environment (e.g., where shock levels are expected to be above the shown curve) shall have shock testing performed at the unit level

(preferably on a qualification unit).

Unit self-induced shock testing shall be accomplished by two actuations at the unit level for each self-induced shock source (in order to account for the scatter associated with the actuation of the device) for the first flight unit, and a single actuation on subsequent units.

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6.6 ACOUSTICS

The Mag shall be designed to withstand, without any damage or degradation of performance, the equivalent Protoflight [TBR] levels shown in Figure 6-4. The Mag contractor shall use envelope option 1 for the SWFO-L1 design, analysis and testing if required. Note: Envelop option 1 and 2 in Figure 6-4 are the same above 800 Hz.

Figure 6-4 SWFO-L1 S/C Protoflight [TBR] Acoustic Envelope

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

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

6.7 TRANSPORTATION

In addition to the launch loads shown above, the Mag shall also be designed to withstand the maximum transportation loads shown in Table 6-4 without damage or degradation of performance.

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Table 6-4 Transportation Loads

Transportation Cleanliness

Materials and enclosure used for transportation and storage shall not generate molecular or particle contaminants or degrade the surface cleanliness of the item or adjacent items to level below those required in section 7.

6.8 PRESSURE

Operating Pressure Range

The Mag shall be designed to be capable of aliveness testing over a pressure range of 1.08 x 105

N/m2 (813 Torr) to 1.3 x 10-12 N/m2 (1 x 10-14 Torr).

The Mag shall meet all performance requirements while under vacuum 1.3 x 10-12 N/m2 (1 x 10-

14 Torr).

Maximum Depressurization Rate

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

Launch Vehicle (LV) Environmental Control System (ECS) Impingement

The Mag exterior surfaces shall not suffer damage or degradation when exposed to the LV ECS airflow velocity of 10 m/sec.

6.9 ON-ORBIT DYNAMIC ENVIRONMENT

The Mag shall be designed to handle all permutation of the following linear and angular acceleration requirements simultaneously.

Dynamic Linear Acceleration

The Mag shall survive on-orbit when subjected to the maximum linear acceleration due to nominal thruster firing of ±0.060 m/sec2 in all axes.

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Dynamic Angular Acceleration

The Mag shall survive the maximum angular acceleration of ±2.5E-5 rad/sec2 about all axes.

6.10 THERMAL REQUIREMENTS

The Mag shall be thermally conductively coupled to the spacecraft for thermal control (TBP).

The Mag heat transferred to the spacecraft shall be limited to 4 Watts [TBR].

The Mag shall meet all of its performance requirements after exposure to air temperature between +5 and +30 degrees C and relative humidity between 30% and 70%.

The Mag thermal design shall employ thermal coatings properties validated to be accurate for materials and mission flight parameters over the lifecycle of the mission.

Flight Interface Design Temperature Limits

When powered “OFF”, the Mag shall be capable of surviving indefinitely when the spacecraft interface temperatures are within the survival limits shown in Table 6-5 without damage or permanent performance degradation.

The Mag shall meet all performance requirements when within the Operational and

Protoflight/Qualification limits shown in Table 6-5.

The Mag shall demonstrate turn on at the Minimum Survival and…