National Aeronautics and Space

Administration

GLIMR-06-002

PRELIMINARY

RELEASED DATE: 7/1/26

GEOSTATIONARY LITTORAL IMAGING AND

MONITORING RADIOMETER (GLIMR) MISSION

GLIMR Payload to Host Interface Definition Document

Revision: Preliminary Document No: GLIMR-06-002

Release Date: 7/1/2026 Page 2 of 76

Title: GLIMR Payload to Host Interface Definition Document

REVISION HISTORY PAGE

Revision No. Description Release Date

Preliminary Initial release, for use on GLIMR Mission Access to Space

Industry Concept Study

5/20/2024

Preliminary Approved release for request for proposal 7/1/2026

This document contains information that is as complete as possible. Where final numerical values or specification references are not available, best estimates are given and noted To Be Confirmed

(TBC). Items which are not yet defined are noted To Be Determined (TBD). The following table summarizes the TBD/TBC items in this revision of the document, and supplements the revision notice above.

TBD/TBC Section(s) Description of TBD/TBC items

TBC

Table 4-7

Tabulated MOI and Related Information for Azimuth Stage: Confirm as-built values.

TBC

Table 4-8

Tabulated MOI and Related Information for Elevation Stage:

Confirm as-built values.

TBC

Figure 5-1

Payload Power, Data and Telemetry Interfaces to the Host

Spacecraft: Confirm IRU RS-422 will be on Connector J2.

TBD

5.2.8

Remove Before Flight: Provide figure showing location(s) of

Remove Before Flight items.

TBC

Table 6-2

Operational Power Feeds: Confirm CBE power values.

TBC

Figure 6-1

Host I/F Connectors and PCU, DEB, and Sensor Assembly

Interconnects: Confirm connection for IRU RS-422 Inputs

TBC

Table 6-5

Host to Payload Interface Connectors: Confirm connection for IRU

RS-422 Inputs

TBC

9.2.1

Payload Purge Logbook: Confirm maximum accumulated time allowable outside of purge.

TBC

9.4

Host Spacecraft Cleanliness: Values for post-delivery to end-of-life particulate contamination impingement rate.

TBD

10.3

GMOC to GSDS Verbal Interfaces: Timing and format of verbal interfaces between the GMOC and GSDS.

Release Date: 7/1/2026 Page 3 of 76

TBD

10.7

GSDS Support Data Products Overview: Contents and formats of the

Support Data Products.

TBD

11.2

Payload Interface Connector J2: Determine pins to be used for the

IRU RS-422 connection to Host.

TBC Table C-1 Host Routed Cable Information: Confirm as-built cable lengths.

Release Date: 7/1/2026 Page 4 of 76

Table of Contents

1 INTRODUCTION

1.1 Scope and Purpose

1.2 Document Organization

1.3 Definitions

1.3.1 Payload

1.3.2 To Be Confirmed

1.3.3 To Be Determined

2 DOCUMENTS

2.1 Government Documents

2.2 Non-Government Documents

2.3 GLIMR Program Documents

2.4 Document Control

3 SYSTEM DESCRIPTION

3.1 Mission Description

3.2 Payload Description

3.2.1 Sensor Assembly

3.2.2 Decoder Electronics Box

3.2.3 Payload Control Unit

3.2.4 Dust Cover

4 KEY PAYLOAD PHYSICAL CHARACTERISTICS

4.1 Coordinate System Definition

4.2 Payload Dimensions

4.2.1 Sensor Assembly Stowed Volume

4.2.2 Telescope Swept Volume

4.2.3 PCU Volume

4.2.4 DEB Volume

4.2.5 Dust Cover Volume

4.3 Mass

4.4 Centers of Gravity and Moments of Inertia

4.4.1 Sensor Assembly Center of Gravity and Moments of Inertia

5 PAYLOAD TO HOST INTERFACE OVERVIEW

5.1 Mechanical Interface Overview

5.2 Payload Outlines and Mounting Interfaces

5.2.1 Host Interface Coordinate System Origin

5.2.2 Sensor Assembly Mounting Flange Outline

5.2.3 PCU Dimensions and Mounting

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5.2.4 DEB Dimensions and Mounting

5.2.5 Dust Cover Dimensions and Mounting

5.2.6 Stowed Position

5.2.7 Gimbal Range of Travel

5.2.8 Remove Before Flight

6 PAYLOAD TO HOST ELECTRICAL INTERFACES

6.1 Payload Power

6.1.1 Operational Power Feeds

6.1.2 Primary Power Voltage Range

6.1.3 Power Enable Signals

6.1.4 Maximum Payload Current

6.1.5 Payload Power Fusing

6.1.6 Survival Heater Power

6.1.7 Sensor Assembly Gimbal Launch Lock Release Power

6.2 Payload Power-On Sequence

6.3 Payload to Host Spacecraft Grounding

6.3.1 Chassis Grounding Diagram

6.3.2 PCU to Chassis Grounding

6.3.3 Multi-layer Insulation Grounding

6.4 Host Spacecraft to Payload Harnessing

6.4.1 Payload Interface Connectors

6.4.2 Host-Routed Cable Information

6.5 Analog Telemetry Interface (Optional)

6.5.1 Active Analog Telemetry

6.5.2 Passive Analog Telemetry

7 PAYLOAD TO HOST THERMAL INTERFACES

7.1.1 PCU Thermal Interface

7.1.2 DEB Thermal Interface

7.2 Orbital Thermal Environment

8 PAYLOAD TO HOST SPACECRAFT DATA INTERFACES

8.1 Host Spacecraft Data Interface

8.2 Command & Data Interfaces

8.2.1 1PPS Signal Interface

8.2.2 SpaceWire Interface

8.2.3 RS-422 Interface

8.2.4 Data Storage Requirements

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8.2.5 Data Error Correction

8.3 Payload Modes

8.3.1 Off/Survival Mode

8.3.2 Initialization Mode

8.3.3 Safe Mode

8.3.4 Operation Mode

8.3.5 Mode Transitions

8.4 Payload Internal Fault Detection and Correction

8.5 Host Spacecraft Commanding

8.5.1 Payload Emergency Power Off

8.5.2 Emergency Maneuvers

9 PAYLOAD CONTAMINATION CONTROL

9.1 Payload Protection

9.1.1 Payload Bagging

9.2 Payload Purge

9.2.1 Payload Purge Logbook

9.2.2 Purge Line Requirements

9.2.3 Purge Gas Requirements

9.2.4 Purge Gas Source

9.3 Host Integration & Test Facility

9.4 Host Spacecraft Cleanliness

9.5 Launch Site Integration & Test

9.6 Launch and On-Orbit Operations

10 GLIMR DATA FLOW OVERVIEW

10.1 GMOC Coordinate Systems and Units

10.2 GMOC Time Formats

10.3 GMOC to GSDS Verbal Interfaces

10.4 GSDS to GMOC Command Interface Overview

10.5 GMOC to GSDS Telemetry Interface Overview

10.6 Host Spacecraft Commands and Telemetry in Support of GLIMR

10.7 GSDS Support Data Products Overview

11 APPENDICES

11.1 Appendix A: Acronyms

11.2 Appendix B: Host Interface Connector Pinouts

11.3 Appendix C: Host-Routed Cable Information

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Table of Tables

Table 2-1: Government Guiding Documents Referenced Table 2-2: Additional Space Industry Documents Referenced Table 2-3: GLIMR Generated Documents Referenced Table 4-1: Sensor Assembly Stowed Volume (example only - see D4957031-500) Table 4-2: Sensor Assembly Swept Volume (example only - see D4957031-500) Table 4-3: PCU Volume (example only - see D4957031-500) Table 4-4: DEB Volume (example only - see D4957031-500) Table 4-5: Dust Cover Volume Table 4-6: Payload Mass by Assembly Table 4-7: Tabulated MOI and Related Information for Azimuth Stage [TBC] Table 4-8: Tabulated MOI and Related Information for Elevation Stage [TBC] Table 6-1: Payload NTE Power Allocations for Steady State and Peak Power Table 6-2: Payload CBE Power for each Host Spacecraft Power Feed [TBC] Table 6-3: Power Enable Signal Characteristics for PCU and FPE Power Feeds Table 6-4: Survival Heater Characteristics Table 6-5: Host to Payload Interface Connectors Table 6-6: Payload Active Analog Telemetry Table 6-7: Active Analog Telemetry Signal Characteristics Table 6-8: Passive Analog Telemetry List Table 7-1: Payload Assumed GEO Thermal Environment Table 8-1: 1-PPS Signal Characteristics Table 8-2: SpaceWire Message Types Table 8-3: Host Spacecraft Data Size and RMAP Message ID Table 8-4: Host Spacecraft Attitude and Ephemeris Table 8-5: Payload Telemetry Size, RMAP Message ID, and Rate Table 8-6: IRU Data Specifics Table 10-1: Host Spacecraft Commands and Telemetry in Support of GLIMR

Release Date: 7/1/2026 Page 8 of 76

Table of Figures

Figure 3-1: Planned Observations Figure 3-2: Payload Components Figure 3-3: Dust Cover Radius of Curvature Figure 4-1: Orbit Reference Frame (ORF) Figure 4-2: Graphical Representation of Stowed Volume (example only – see D4957031-500) Figure 4-3: Graphical Representation of Swept Volume (example only – see D4957031-500) Figure 4-4: Payload Control Unit Dimensions (example only – see D4957031-500) Figure 4-5: Decoder Electronics Box Dimensions (example only – see D4957031-500) Figure 4-6: Dust Cover Volume Figure 4-7: Center of Gravity for Sensor Assembly in Stowed Configuration (example only – see

D4957031-500) Figure 4-8: MOI & Related Information for Hardware Moved on Sensor Azimuth Stage [TBC] Figure 4-9: MOI & Related Information for Hardware Moved on Sensor Elevation Stage [TBC] ... 23 Figure 5-1: Payload to GLIMR Electronic Interfaces Figure 5-2: Payload View Showing Nadir-Facing Plane Figure 5-3: Cross Section of Sensor Assembly Gimbal Mounting Flange (example only – see

D4957031-500) Figure 5-4: Isometric View of the Payload Control Unit Figure 5-5: Payload Control Unit Dimensions (example only – see D4957031-500) Figure 5-6: DEB Dimension and Mounting Hole Locations (example only – see D4957031-500) ... 28 Figure 5-7: Dust Cover Mounting and Alignment Pin holes (View E-E in Figure 4 4) (example only

– see D4957031-500) Figure 5-8: Relative Location of Dust Cover to Payload Mounting Flange Figure 5-9: GLIMR Stowed Position - +23-deg. Elevation Figure 5-10: GLIMR Stowed Position – +44.8 deg. Azimuth Figure 6-1: Host I/F Connectors and PCU, DEB, and Sensor Assembly Interconnects Figure 6-2: Power Enable Electronics Interface Diagram Figure 6-3: Launch Lock Wiring Schematic Figure 6-4: Payload Grounding Diagram Figure 6-5: Payload Interface Connector Locations Figure 8-1: Payload Mode Transitions Figure 10-1: GLIMR Data Flow

Release Date: 7/1/2026 Page 9 of 76

1 INTRODUCTION

1.1 Scope and Purpose

GLIMR is a Category 3 (per NPR 7120.5E) Class D mission (per NPR 8705.4) being developed for the NASA EVI-5 program.

This document is the GLIMR Payload to Host Interface Definition Document (IDD), and defines the physical, mechanical, electrical, thermal, and data interfaces between the payload and Host

Spacecraft.

This document has precedence over any and all other written, verbal, or electronic statements addressing interface issues and is used in conjunction with the GLIMR Instrument Installation

Control Drawing (IICD) (D4957031-500).

1.2 Document Organization

Section 2 lists applicable and reference documents. Section 3 provides an overview of the

GLIMR mission and the payload description. Sections 4 through 12 contain the interface descriptions. Appendix A provides acronyms, Appendix B provides the Host Spacecraft interface connector pinouts, and Appendix C provides the Host Spacecraft routed cable information.

1.3 Definitions

1.3.1 Payload

When used in this document the term payload refers to the totality of the delivered GLIMR flight instrument.

1.3.2 To Be Confirmed

To Be Confirmed (TBC) denotation signifies that the requirement value is not yet firm.

1.3.3 To Be Determined

To Be Determined (TBD) denotation signifies that the referenced text is in development.

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2 DOCUMENTS

The documents listed in this section are for reference to the extent specified herein. In the event of a conflict between the documents referenced herein and the contents of this document, the contents of this document will be considered the superseding requirement. The documents’ version at the time of Request for Proposal release is considered applicable unless otherwise noted.

The GLIMR Host Spacecraft Statement of Work (SOW) references this IDD, and the SOW has precedence in the event of any discrepancies.

2.1 Government Documents

Table 2-1contains the guiding government documents referenced within this interface description document.

Table 2-1: Government Guiding Documents Referenced

Document Number Document Title

MIL-PRF-27401 Performance Specification: Propellant Pressurizing Agent, Nitrogen

MIL-STD-38999 Detail Specification: Connectors, Electrical, Circular, Miniature, High Density, Quick Disconnect (Bayonet, Threaded, and Breech

Coupling), Environment Resistant, Removable Crimp and

Hermetic Solder Contacts, General Specification For

MIL-STD-461 Electromagnetic Interference Characteristics of Subsystems and

Equipment Requirements

NPD 7120.4E NASA Engineering and Program/Project Management Policy

S-311-P-18 Thermistor, (Thermally Sensitive Resistor), Insulated and

Uninsulated, Negative Temperature Coefficient, Specification for

2.2 Non-Government Documents

The documents listed in Table 2-2 are non-government industry generated documents for reference to the extent specified herein. The latest revision should be utilized.

Table 2-2: Additional Space Industry Documents Referenced

Document # Document Name

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

Interface Circuits

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

Volatile Condensable Materials from Outgassing in a Vacuum

Environment

ECSS-E-ST-50-12C Rev.1 SpaceWire – Links, Nodes, Routers and Networks

ECSS-E-ST-50-51C SpaceWire Protocol Identification

ECSS-E-ST-50-52C SpaceWire – Remote Memory Access Protocol

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Document # Document Name

IEST-STD-CC1246 Product Cleanliness Levels – Applications, Requirements, and

Determination

ISO-8601 Date and Time Format

2.3 GLIMR Program Documents

The documents listed in Table 2-3 are GLIMR program generated documents for reference to the extent specified herein. The latest revision should be utilized.

Table 2-3: GLIMR Generated Documents Referenced

Document # Document Name

D4957031-500 Instrument Mechanical Installation Control Drawing (Raytheon)

D4982947 Cable Assy, DEB, Op Htr/Ts, W303A (Raytheon)

D4983021 Cable Assy, Op Htr/Ts, PCU, W310A (Raytheon)

D4983024 Cable Assy, Op Htr/Ts, DEB, PCU, Az Motor Drv, W320A

(Raytheon)

D4983059 Cable Assy, LL, Surv Htr/Ts, W323A/W324A (Raytheon)

D4983064 Cable Assy, SpaceWire, W330A (Raytheon)

D4983067 Cable Assy, Power, Host, W331A (Raytheon)

D4983070 Cable Assy, SPW Cmd & Data, W333 – Host (Raytheon)

D4983071 Cable Assy, Op Htr/Ts, W335A (Raytheon)

D4982950-100 Payload Control Unit (PCU) Assembly (Raytheon)

D4954207 Encoder Assembly (Raytheon)

GLIMR_05.07.001 GLIMR Software User Manual (Raytheon)

GLIMR_05.09.001 Command Database (Raytheon)

GLIMR_05.09.002 Telemetry Database (Raytheon)

GLIMR_02.11.001 GLIMR Fault Handling (Raytheon)

2.4 Document Control

This document is managed by the Earth System Science Pathfinder (ESSP) Program Office.

After initial approval, it will be placed under formal configuration control. This IDD can be updated continuously throughout the life of the mission upon agreement between the Host

Spacecraft and payload team. NASA is responsible for the composition of the IDD.

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3 SYSTEM DESCRIPTION

This section describes the system architecture for GLIMR mission and provides context for descriptions in later sections. Refer to the GLIMR Concept of Operations Document (UNH-

GLIMR-CONOPS-01) for a more thorough description of the mission functions and operations.

3.1 Mission Description

GLIMR is a hyperspectral ocean color sensor which targets the Gulf of America and other coastal and ocean waters of North and South America. With its vantage point from geostationary orbit, GLIMR will be the first hyperspectral ocean color sensor in the Western Hemisphere to study ocean processes at the diurnal timescales required to observe the dynamic ecological, biogeochemical, and physical processes typical of coastal and ocean waters. GLIMR is uniquely designed to capture the spatiotemporal evolution of phytoplankton blooms, which grow and dissipate on scales of hours to days, monitor phytoplankton physiology, growth rates and shifts in community composition, resolve coastal features, fronts, eddies, and upwelling features, and monitor biogeochemical fluxes and land-ocean exchanges at sub-diurnal to multi-day scales.

GLIMR will improve quantification of primary productivity, biomass, and carbon fluxes within coastal aquatic ecosystems and across the land-sea interface and contribute observations that will aid development and evaluation of models that project future changes in carbon cycling and marine ecosystems for use in ecological forecasting. The two main science goals of GLIMR are

(1) to understand the processes contributing to rapid changes in phytoplankton growth rate and community composition, and (2) quantify how high frequency fluxes of sediments, organic matter, and other materials between and within coastal ecosystems regulate the productivity and health of coastal ecosystems. The science application objectives aim to study the formation, magnitude, and trajectories of harmful algal blooms, oil spills, and Sargassum accumulation events.

At the heart of the GLIMR payload is a hyperspectral imaging (HSI) spectrometer with a spectral range between 340-1040nm, a spectral resolution of about 10 nm, and the capability to scan the

Gulf of America in about 70 minutes. A higher spatial resolution landmark imager (LMI) co-aligned with the spectrometer enables image navigation and registration of the ocean color pixels collected by the spectrometer. A 28 cm diameter telescope provides high throughput and magnification for high signal to noise and spatial resolution.

During the daylight hours, the instrument will scan various regions of interest in the western hemisphere. These regions include opportunities for calibration, including simultaneous measurements with other instruments and star looks. During the nighttime, the instrument will be commanded to the stowed position.

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Figure 3-1: Planned Observations

3.2 Payload Description

The complete GLIMR payload is made up of four separately mounted systems: the sensor assembly (encompassing all moving parts), the Decoder Electronics Box (DEB), the Payload

Control Unit (PCU), and the dust cover. The payload components are shown in Figure 3-2.

Payload Components (Not To Scale)

Payload Control Unit

Decoder Electronics Box

Sensor Assembly

Dust Cover

Figure 3-2: Payload Components

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3.2.1 Sensor Assembly

The fully integrated sensor assembly consists of the gimbal assembly, the telescope assembly, and portions of the heat exchange subsystem.

3.2.1.1 Gimbal Assembly Subsystem

The Gimbal Assembly Subsystem is the part of the payload that is commanded in azimuth and elevation across the various scan regions. Each of the two axes contains bearings, a motor, and an optical encoder for position monitoring.

3.2.1.2 Telescope Assembly

The Telescope Assembly Subsystem collects the optical signal from the scene and contains a barrel-like sunshade for mitigating stray light onto the detectors. Out of the back of the telescope assembly, the incoming light is directed through the telecentric relay towards the Landmark

Imager (LMI) used for geolocation and the Hyperspectral Imager (HSI) used to collect science data.

3.2.1.3 Heat Exchange Subsystem

The payload Heat Exchange Subsystem contains components both physically on and off the sensor assembly. It manages temperatures for the payload with a combination of heaters, radiators, thermal straps, multilayer insulation and custom thermal isolation components.

Operational heaters maintain operational temperatures of the gimbal assembly, optical assemblies, payload focal plane arrays, and the spectrometer focal plane array. Survival heaters using bang-bang thermostat control are used to maintain temperatures of all sensor assembly components within their survival limits.

3.2.2 Decoder Electronics Box

The DEB contains electronics to integrate sensor assembly encoder read heads and report rotational positions via a serial interface for each axis to the PCU. The DEB is functionally part of the sensor assembly but is mechanically mounted separately from the sensor assembly on the

GLIMR Spacecraft structure.

3.2.3 Payload Control Unit

The PCU is the main interface electronics box mounted separately from the sensor assembly and controls the gimbal, distributes power, controls the thermal environment, and interfaces to the

GLIMR Spacecraft.

3.2.4 Dust Cover

The payload dust cover dish is mechanically mounted separate from the sensor assembly in a fixed location on the GLIMR Spacecraft structure. When in the stowed position, the telescope is oriented in a position that aligns the telescope sunshade with the dust cover to provide contamination protection during launch and thermal protection during non-observing periods on-orbit.

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Both the outer edge of the sunshade and the surface of the dust cover facing the telescope sunshade have a radius of curvature as shown in Figure 3-3 that prevents interference when the telescope moves through its full range of motion.

Figure 3-3: Dust Cover Radius of Curvature

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4 KEY PAYLOAD PHYSICAL CHARACTERISTICS

4.1 Coordinate System Definition

The GLIMR design team utilizes the community standard body reference frame coordinate system where:

+Z is in the direction of nadir;

+X is in the direction of orbital velocity; and

+Y completes the right-handed system.

Figure 4-1: Orbit Reference Frame (ORF)

The origin of the coordinate system is indicated on the GLIMR IICD (D4957031-500).

4.2 Payload Dimensions

4.2.1 Sensor Assembly Stowed Volume

The volume of the stowed sensor assembly is specified in the GLIMR IICD (D4957031-500).

Note: In the stowed position, the telescope is pointed at the dust cover, not at nadir.

Figure 4-2: Graphical Representation of Stowed Volume (example only – see

D4957031-500)

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Table 4-1: Sensor Assembly Stowed Volume (example only - see D4957031-500)

Parameter Maximum Expected Value x-extent (parallel to orbit) 0.94 m y-extent (normal to orbit) 1.55 m z-extent (nadir) 0.95 m

Total Volume 1.4 m3

4.2.2 Telescope Swept Volume

The volume of the swept sensor assembly is specified in the GLIMR IICD (D4957031-500).

The swept volume represents the keep-out volumes for other payloads or structures on the Host

Spacecraft bus.

Figure 4-3: Graphical Representation of Swept Volume (example only – see D4957031-500)

Table 4-2: Sensor Assembly Swept Volume (example only - see D4957031-500)

Parameter Maximum Expected Value x-extent (parallel to orbit) 1.13 m y-extent (normal to orbit) 1.26 m z-extent (nadir) 1.13 m

Total Volume 1.60 m3

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4.2.3 PCU Volume

The volume of the PCU is specified in the GLIMR IICD (D4957031-500).

Figure 4-4: Payload Control Unit Dimensions (example only – see D4957031-500)

Table 4-3: PCU Volume (example only - see D4957031-500) x-extent 0.445 m y-extent 0.489 m z-extent 0.218 m

Total Volume 0.047 m3

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4.2.4 DEB Volume

The volume of the DEB is specified in the GLIMR IICD (D4957031-500).

Figure 4-5: Decoder Electronics Box Dimensions (example only – see D4957031-500)

Table 4-4: DEB Volume (example only - see D4957031-500) x-extent 0.107 m y-extent 0.156 m z-extent 0.199 m

Total Volume 0.003 m3

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4.2.5 Dust Cover Volume

The volume of the dust cover is specified in the GLIMR IICD (D4957031-500)

Figure 4-6: Dust Cover Volume

Table 4-5: Dust Cover Volume

Parameter Maximum Expected Value x-extent 0.35 m y-extent 0.08 m z-extent 0.36 m

Total Volume 0.01 m3

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4.3 Mass

The payload will have a maximum total mass allocation Not to Exceed (NTE) of 175 kg.

Table 4-6: Payload Mass by Assembly

Subsystem Not to Exceed Value

Sensor Assembly 136 kg

PCU + cables 34 kg

DEB 3 kg

Dust Cover 2 kg

Total Payload 175 kg

The mass of the heat exchange system has been included in the sensor assembly. The mass of the total payload includes all hardware including internal payload connectors and harnesses, thermal blankets, and all additional hardware delivered with the payload and utilized in the mounting and/or final closeout of the unit on the Host Spacecraft bus.

The payload NTE value does not include any sensor assembly interface mating structure or dust cover mounting structure provided by the Host Spacecraft which would mount directly to the

Host Spacecraft.

4.4 Centers of Gravity and Moments of Inertia

4.4.1 Sensor Assembly Center of Gravity and Moments of Inertia

The GLIMR IICD (D4957031-500) specifies the center of gravity of the sensor assembly. he

Moments of Inertia (MOI) values are shown in Figure 4-8 and Table 4-7 for the azimuth-mounted hardware and in Figure 4-9 and Table 4-8 for the elevation-mounted hardware.

Figure 4-7: Center of Gravity for Sensor Assembly in Stowed Configuration (example only

– see D4957031-500)

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Figure 4-8: MOI & Related Information for Hardware Moved on Sensor Azimuth Stage

[TBC]

Table 4-7: Tabulated MOI and Related Information for Azimuth Stage [TBC]

CG (m) with respect to GIMBAL-MNT-INTFC coordinate frame

X Y Z 1.0619337E-02 4.9844002E-01 -1.4960601E-02

INERTIA with respect to GIMBAL-MNT-INTFC coordinate frame

Inertia Tensor (kg * m^2) Ixx Ixy Ixz 3.2047396E+01 -5.8600539E-01 4.2166219E-01

Iyx Iyy Iyz -5.8600539E-01 6.0590752E+00 -2.4773043E-01

Izx Izy Izz 4.2166219E-01 -2.4773043E-01 3.3548200E+01

INERTIA AT CG with respect to GIMBAL-MNT-INTFC coordinate frame

Inertia Tensor (kg * m^2) Ixx Ixy Ixz 8.9158892E+00 -9.3628880E-02 4.0688358E-01

Iyx Iyy Iyz -9.3628880E-02 6.0277648E+00 -9.4139416E-01

Izx Izy Izz 4.0688358E-01 -9.4139416E-01 1.0427024E+01

PRINCIPAL MOMENTS OF INERTIA: (kg * m^2)

I1 I2 I3 5.8347493E+00 8.8191782E+00 1.0716751E+01

ROTATION MATRIX from GIMBAL-MNT-INTFC orientation to PRINCIPAL AXES

0.00329 -0.97417 0.22579

0.97968 -0.04214 -0.19609

0.20054 0.22185 0.95424

ROTATION ANGLES from GIMBAL-MNT-INTFC orientation to PRINCIPAL AXES (deg) about x about y about z

11.61 13.05 89.81

RADII OF GYRATION with respect to PRINCIPAL AXES (m):

R1 R2 R3 2.5044801E-01 3.0790769E-01 3.3942048E-01

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Figure 4-9: MOI & Related Information for Hardware Moved on Sensor Elevation Stage

[TBC]

Table 4-8: Tabulated MOI and Related Information for Elevation Stage [TBC]

CG (m) with respect to AZ-EL_AXIS_INTERSECT coordinate frame

X Y Z -1.3513660E-02 3.8890183E-02 -1.9131286E-02

INERTIA with respect to AZ-EL_AXIS_INTERSECT coordinate frame

Inertia Tensor (kg * m^2) Ixx Ixy Ixz 4.2787102E+00 1.8039408E-01 -4.1401512E-01

Iyx Iyy Iyz 1.8039408E-01 4.6817370E+00 1.1239268E+00

Izx Izy Izz -4.1401512E-01 1.1239268E+00 4.6979268E+00

INERTIA AT CG with respect to AZ-EL_AXIS_INTERSECT coordinate frame

Inertia Tensor (kg * m^2) Ixx Ixy Ixz 4.1583439E+00 1.4671829E-01 -3.9744895E-01

Iyx Iyy Iyz 1.4671829E-01 4.6465825E+00 1.0762520E+00

Izx Izy Izz -3.9744895E-01 1.0762520E+00 4.5893114E+00

PRINCIPAL MOMENTS OF INERTIA: (kg * m^2)

I1 I2 I3 3.3524109E+00 4.3273177E+00 5.7145091E+00

ROTATION MATRIX from AZ-EL_AXIS_INTERSECT orientation to PRINCIPAL AXES

0.43833 0.89143 -0.11497 -0.60356 0.38670 0.69726

0.66602 -0.23624 0.70754

ROTATION ANGLES from AZ-EL_AXIS_INTERSECT orientation to PRINCIPAL AXES (deg) about x about y about z -44.581 -6.602 -63.816

RADII OF GYRATION with respect to PRINCIPAL AXES (m):

R1 R2 R3 2.2873161E-01 2.5987048E-01 2.9863247E-01

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5 PAYLOAD TO HOST INTERFACE OVERVIEW

Figure 5-1 shows the high-level categories of electronic interfaces between the GLIMR Payload and the GLIMR spacecraft. These interfaces are described in the following sections and in key reference documentation. There are mechanical mounting and thermal interfaces not shown in diagram that are also described in the following sections and detailed in the GLIMR IICD

(D4957031-500). As shown, the Host Spacecraft provides the power for the thermostatically controlled survival heaters and temperature sensors. Analog telemetry monitors the secondary voltages prior to the SpaceWire communications being enabled.

Figure 5-1: Payload to GLIMR Electronic Interfaces

5.1 Mechanical Interface Overview

The payload mechanical interfaces are defined and controlled in the GLIMR IICD (D4957031-

500). This drawing depicts the payload isometric views and general exterior dimensions, Host

Spacecraft mounting interfaces and dimensions, stowed, and swept volumes. The figures in this section are examples only.

The payload mechanical interfaces to the Host Spacecraft consists of the sensor assembly mounting flange, the PCU baseplate and thermal strapping, the DEB baseplate, and the dust cover mounting flange. The sensor assembly mounting interface is the azimuthal axis base of the gimbal assembly.

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5.2 Payload Outlines and Mounting Interfaces

5.2.1 Host Interface Coordinate System Origin

Requested origin location and orientation relative to the NASA standard Host Spacecraft interface coordinate system is depicted on the GLIMR IICD (D4957031-500).

Nadir-facing Plane

Figure 5-2: Payload View Showing Nadir-Facing Plane

5.2.2 Sensor Assembly Mounting Flange Outline

The sensor assembly is mounted to the Host Spacecraft in such a manner as to support nadir pointing of the telescope with the azimuth gimbal base mounted in the X-Z plane.

Figure 5-3: Cross Section of Sensor Assembly Gimbal Mounting Flange (example only – see

D4957031-500)

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5.2.3 PCU Dimensions and Mounting

The PCU is mounted as close to the gimbal azimuth base as feasible while considering the payload harness lengths and PCU thermal requirements. The primary limitation on PCU placement is the W330A cable length of 1.27 m (50 inches) which carries the SpaceWire and

One Pulse Per Second (1PPS) signals from the PCU to the Azimuth cable wrap. There are other cables from the PCU to the Azimuth cable wrap and DEB, however their length is not as critical for achieving functionality.

(Note: this maximum distance value represents the baseline configuration and potentially can be increased if desired by Host Spacecraft vendor pending a signal integrity analysis).

The PCU is shown in an isometric view in Figure 5-4 with items labelled as follows:

1-Chassis

2-Input Power Filter

3-Top Cover

The PCU dimensions are shown on the GLIMR IICD (D4957031-500).

Figure 5-4: Isometric View of the Payload Control Unit

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A thermal heat load will primarily be conducted from two thermal interfaces on the side of the

PCU chassis. These thermal interfaces are identical and are located on opposite sides of the unit as shown in Figure 5-5.

Figure 5-5: Payload Control Unit Dimensions (example only – see D4957031-

500)

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5.2.4 DEB Dimensions and Mounting

The DEB is mounted as close to the gimbal azimuth base as feasible while considering the payload harness lengths. The primary limitations on the DEB location are the W303A cable between the PCU and DEB (DEB command and data), and the W320A cable which contains the read heads. The distance between W303A P12 (PCU J12) and W303A P1 (DEB J1) is a maximum of 0.60 m (23.7 inches). The distance between W320A P4 and P5 (DEB J4 and J5) and W320A J29 (Azimuth Cable Wrap, W389 P29) is a maximum of 1.25 m (49.4 inches).

(Note: these maximum distance values represent the baseline configuration and potentially can be increased if desired by Host Spacecraft vendor pending a signal integrity analysis).

The DEB dimensions and its mounting plate interface bolt pattern are shown on GLIMR IICD

(D4957031-500).

5.2.5 Dust Cover Dimensions and Mounting

The dust cover relative location with respect to the sensor assembly is shown in GLIMR IICD

(D4957031-500).

Figure 5-7: Dust Cover Mounting and Alignment Pin holes (View E-E in Figure 4 4)

(example only – see D4957031-500)

Figure 5-6: DEB Dimension and Mounting Hole Locations (example only – see D4957031-500)

Release Date: 7/1/2026 Page 29 of 76

Figure 5-8: Relative Location of Dust Cover to Payload Mounting Flange

5.2.6 Stowed Position

The gimbal stowed position is Az = 30.8 deg., El = 23 deg.

Nadir is defined as Az = -14 deg., El = 0 deg.

(0 deg. Z_ORF = -14 deg. Az_gimbal)

Figure 5-9: GLIMR Stowed Position - +23-deg. Elevation

Release Date: 7/1/2026 Page 30 of 76

Figure 5-10: GLIMR Stowed Position – +44.8 deg. Azimuth

5.2.7 Gimbal Range of Travel

The gimbal range of travel is:

• Az: +34.5 deg. to -34.5 deg.

• El: +24.0 deg. to -11.4 deg.

5.2.8 Remove Before Flight

Locations of the Remove Before Flight items are identified in Figure (TBD).

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6 PAYLOAD TO HOST ELECTRICAL INTERFACES

The interconnect diagram presents a high-level overview of the connections between various components of the payload and the Host Spacecraft interface connectors. Payload electrical interface connections required for operation will be per MIL-STD-38999 and pinouts for the interface connectors are listed in Appendix B: Host Interface Connector Pinouts.

Figure 6-1: Host I/F Connectors and PCU, DEB, and Sensor Assembly Interconnects

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6.1 Payload Power

Note: All payload power numbers listed are based on Program Office allocations with incorporated contingency, margin, and reserve, and not based on payload MEV values. Payload

Critical Design Review (CDR) maturity Current Best Estimate (CBE) values when listed are noted as such.

6.1.1 Operational Power Feeds

Payload allocations and minimum accommodations required for operational steady-state average power and operational peak power.

Steady-state average power is defined as being over any 180-minute period.

Peak power is defined as the maximum power consumed averaged over any 10 ms period.

Table 6-1: Payload NTE Power Allocations for Steady State and Peak Power

Power Feed NTE Operational

Power Allocation (W)

NTE Peak Power Allocation

(W)

Power Feed #1 – FPE 120 130

Power Feed #2 – PCU 150 175

Power Feed #3 – Operational Heaters 130 195

Total Payload 400 500

Table 6-2: Payload CBE Power for each Host Spacecraft Power Feed [TBC]

Power Feed CBE Operational

Power (W)

CBE Peak Power

(W)

Power Feed #1 – FPE 73 80

Power Feed #2 – PCU 48 60

Power Feed #3 – Operational Heaters 86 131

Total Payload 207 271

In addition to the three operational power feeds, the payload requires survival heater power

(Section 6.1.6) and Launch Release Mechanism (LRM) power (Section 6.1.7).

6.1.2 Primary Power Voltage Range

The Host Spacecraft provides a steady state voltage range of 32 ± 2.0 VDC on each of the three operational power feeds to the payload.

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6.1.3 Power Enable Signals

The Host Spacecraft provides two discrete power enable signals, one for Power Feed #1 – FPE and one for Power Feed #2 – PCU. Power enable signal electrical characteristics are described

Table 6-3 and Figure 6-2 illustrates the Power Enable Interface Circuit. These signals must be sent continuously during payload operations.

Table 6-3: Power Enable Signal Characteristics for PCU and FPE Power Feeds

Power Enable Characteristic Limit

On Circuit Voltage Range +30 VDC to +34 VDC

Maximum Leakage Current (Off) 20 microAmp

Maximum Load Current 10 microAmp

Figure 6-2: Power Enable Electronics Interface Diagram

6.1.4 Maximum Payload Current

The maximum current draw is calculated as the maximum peak power accommodation values on a given feed divided by the minimum allowable bus voltage. The maximum current draw for each feed is:

• Power Feed #1 – FPE: 5.0A (130W / 26V),

• Power Feed #2 – PCU: 6.7A (175W / 26V), and

• Power Feed #3 – Operational Heaters: 9.6A (250W / 26V).

6.1.5 Payload Power Fusing

The payload does not contain internal fusing, and Host Spacecraft-side resettable overcurrent protection is expected.

6.1.6 Survival Heater Power

In addition to the three operational power feeds, the sensor assembly requires survival heater power for its payload-supplied thermostatically controlled survival heaters.

The Host Spacecraft provides a 250W power feed for sensor assembly survival heater thermostats at all times. The Host Spacecraft will provide for the measurement of 18 payload temperature sensors for use when the payload is powered off, so that the mission may verify thermostats are operational and that no temperature limits were exceeded in OFF mode.

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The Host Spacecraft survival power feed operates an array of 12 thermostatically controlled survival heaters on the sensor assembly, with no Host Spacecraft-side temperature control loop functionality. The payload will only draw power if the payload temperatures drop below the thermostat setpoints.

The Host Spacecraft enables or disables the thermostat array by applying or removing feed power.

If the DEB and PCU require thermostatically controlled survival heaters to meet the survival temperature ranges, these would be the responsibility of the Host Spacecraft.

6.1.6.1 Sensor Assembly Survival Heater Characteristics

Relevant electrical characteristics for the survival heaters in the sensor assembly.

Table 6-4: Survival Heater Characteristics

Ref

Des Heater Location

Set

Point

(K)

Dead

Band

(K)

Nominal

Resistance

(Ω)

Tolerance

Power (W)

Max

Curren t (A)

Allowable

Current

(A)

Power

Flux

Density

(W/in^2) Min

Heater

Max

Circuit

Max

HS01 Secondary Mirror 244 8 65.4 5 11.0 15.4 15.43 0.498 0.625 4.7

HS04 FPE 233 8 20.3 5 34.5 49.1 49.42 1.594 2.000 4

HS05 Gimbal Ring 244 8 65.4 5 11.0 15.4 15.43 0.498 0.625 3

HS06 Yoke Encoder 244 8 65.4 5 11.0 15.4 15.43 0.498 0.625 1.1

HS07 Yoke Motor 244 8 65.4 5 11.0 15.4 15.43 0.498 0.625 0.7

HS08 LMI Housing 244 8 207.3 5 3.5 4.9 4.88 0.157 0.625 9.6

HS09 HSI Radiator 244 8 207.3 5 3.5 4.87 4.88 0.157 0.625 8.3

HS11 LMI FPA 244 8 207.3 5 3.5 4.9 4.88 0.157 0.625 5.7

HS12 Yoke Az Bearing 244 8 65.4 5 11.0 15.4 15.43 0.498 0.625 2.2

HS13 HSI Housing 244 8 65.4 5 11.0 15.4 15.43 0.498 0.625 3.6

HS18 Aft Metering Structure 244 8 65.4 5 11.0 15.4 15.43 0.498 0.625 3.7

HS19 Gimbal Base Az Motor 244 8 65.4 5 11.0 15.4 15.43 0.498 0.625 1.4 Max Draw 133.0 186.9 187.59 6.000

6.1.7 Sensor Assembly Gimbal Launch Lock Release Power

The Host Spacecraft provides power to release sensor assembly launch locks on the azimuth and elevation gimbals.

The launch lock actuators are Ensign-Bickford Aerospace & Defense (EBAD) P100 type Pin

Pullers (vendor P/N: N81724-2). In fully reset mode, the pin puller internal circuits are closed and do not require Host Spacecraft power to stay engaged. When sufficient power is applied, the current raises the temperature of the Shape Memory Alloy (SMA) wire and subsequently contracts at its transition temperature allowing internal ball locks to release and internal drive springs to retract the pin into the body of the device. Once retracted, the circuit is open allowing no power to continue to heat the SMA wire and provides positive feedback that the device has actuated.

The sensor assembly launch lock release power forms have the following characteristics:

• Four power lines are required, a primary and redundant for each gimbal axis.

• Each power line will be provided with 4A of current for 200-300ms to activate the device. The operating voltage range for the initiators is 18V to 39V.

Release Date: 7/1/2026 Page 35 of 76

• The circuit resistance is 7 Ohms, +/- 1 Ohm.

• The listed voltage range and resistance do not account for any system cabling resistance.

This must be considered when designing the driving interface.

• The lines will interface with a D38999/24GG35PA connector with interface pinout as listed in Appendix B: Host Interface Connector Pinouts - J4. The payload lines will use

26 AWG wires.

6.1.7.1 Launch Lock Monitoring

The Host Spacecraft monitors the launch lock power forms and reports them in the real-time telemetry.

There is no independent release monitor built onto the payload. The only indication the LRM pin pullers have successfully activated is that both primary and secondary circuits switch open once the pin is retracted (i.e., deployed).

6.2 Payload Power-On Sequence

The following steps represent a baseline power-on sequence for the payload during initial power-on during commissioning.

1. Upon receiving command from the ground, the Host Spacecraft provides primary power to the payload power feeds.

2. Upon receiving command from the ground, the Host Spacecraft provides the power enable signal to the PCU. This command will be sent after the ground has confirmed telemetry that the PCU is within its operational temperature range.

Upon application of Power Feed #2 – PCU prime power and receipt of its discrete power enable signal, the PCU will initiate the boot-up sequence. At the completion of the boot-up sequence, the PCU will respond to commands and provide telemetry including confirmation of successful power enable. This step may take several minutes to complete.

The timing for when the Operational Heater feed is energized is not critical, as this power goes to the HDM within the PCU which doesn’t distribute power to heaters until after PCU start-up.

3. Upon receiving command from the ground, the Host Spacecraft provides the power enable signal to the FPE. This command will be sent after the ground has

Figure 6-3: Launch Lock Wiring Schematic

Release Date: 7/1/2026 Page 36 of 76 confirmed telemetry that the PCU has completed initialization and the FPE is within its operational temperature range.

At the conclusion of FPE power-on the payload will be in Standby state awaiting commands.

At this point, the commanding for launch lock release can be performed. This incorporates a motor flinch test to energize the motors prior to release and a gimbal functional test after release.

6.3 Payload to Host Spacecraft Grounding

6.3.1 Chassis Grounding Diagram

Figure 6-4: Payload Grounding Diagram

The mechanical interfaces between the payload and the Host Spacecraft are electrically conductive and EMI-grounded to the bus structure. The Direct Current (DC) electrical resistance between the payload and the Host Spacecraft is to be less than or equal to 0.0025 Ohms.

The Host Spacecraft-to-payload wire harness shields are terminated to the bus structure.

The resistance between any shield and the closest reference point on the payload structure is less than 10 mΩ for a current of 10 mA.

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6.3.2 PCU to Chassis Grounding

The payload PCU secondary ground is tied to PCU chassis on the PCU Backplane.

6.3.3 Multi-layer Insulation Grounding

All Multi-layer Insulation (MLI) on or within 24 inches of the sensor assembly is grounded and has a DC resistance of less than or equal to 100 milliohms between the MLI and ground plane.

6.4 Host Spacecraft to Payload Harnessing

6.4.1 Payload Interface Connectors

Interface connectors from the Host Spacecraft to the payload are in the GLIMR Mechanical

Installation Control Drawing (D4957031-500). Connector pinouts are contained in Appendix B:

Host Interface Connector Pinouts.

Table 6-5: Host to Payload Interface Connectors

Reference

Designator Connector Type Description Destination

J1 D38999 / 24GE35PN SpaceWire & 1 PPS PCU

J2 D38999 / 24GD35PN

Analog Telemetry

(optional) & IRU RS-

422 [TBC]

PCU

J3 D38999 / 24GG35PN Power PCU

J4 D38999 / 24GG35PA Launch Locks Az Cable Wrap

J5 D38999 / 24GH35PN Survival Heaters Az Cable Wrap

J6 D38999 / 24GH35PA Survival Heaters Az Cable Wrap

Figure 6-5: Payload Interface Connector Locations

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6.4.2 Host-Routed Cable Information

Information for the cable harnesses that will need to be routed by the Host Spacecraft is contained in Appendix C: Host-Routed Cable Information.

6.5 Analog Telemetry Interface (Optional)

6.5.1 Active Analog Telemetry

The active analog signals originate from the secondary side of the power supplies in the FPE and the PCU and indicate the voltage of the secondary power forms generated within the units.

The payload provides at least one pin for the return of analog monitors for each electronics unit.

This interface should be sampled as a differential interface to avoid chassis grounding of the analog telemetry return, thus preserving the payload’s single point ground architecture.

Table 6-6: Payload Active Analog Telemetry

C o u n t

Signal Name Signal Type /

Description Scale Factor

Scaled Range in

Volts with offsets

1 PCU_HST_P30VTLM_H

PCU P30V Power

Telemetry

0.167 +/- 6% 4.0 to 6.0V

2 PCU_HST_P5VTLM_H

PCU P5V Power

Telemetry

0.98 +/- 6% 4.0 to 6.0V

3 PCU_HST_P15VTLM_H

PCU P15V Power

Telemetry

0.382 +/- 6% 4.0 to 6.0V

4 PCU_HST_P5P6VTLM_H

PCU P5.6V Power

Telemetry

0.846 +/- 6% 4.0 to 6.0V

5 P15VDC_TLM

FPE P15V Power

Telemetry 0.33+/-6% 4.0 to 6.0V

6 P5VDC_TLM

FPE P5V Power

Telemetry

0.98 +/- 6% 4.0 to 6.0V

Measured telemetry voltage divided by the scale factor provides the voltage of the secondary power forms.

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Table 6-7: Active Analog Telemetry Signal Characteristics

Parameter Limit

User source signal 0.0 to 10.0 Vdc

User fault voltage (not to exceed) (1) -16.0 to +16.0 V

Current (during sampling) <=10 µA

User source resistance <=10 kOhm

(1) Fault propagation will occur if user applies an Active Analog voltage signal outside of this range to the interface

6.5.2 Passive Analog Telemetry

The Host Spacecraft monitors and reports 18 passive analog payload survival temperature sensors for use when the payload is powered off.

The passive analog telemetry thermistors (survival temperature sensors) are of type S-311-P-18 per GSFC standard (PN: 311P18-07A101).

Release Date: 7/1/2026 Page 40 of 76

Table 6-8: Passive Analog Telemetry List

No Location

Low

Resolution

Temperature

Range

Temperature Sensor

Description

Temp

Resolution

[K]

TSTAT*

(Y/N)

Low

Temp

(K)

High

Temp

(K) Type Config

Resistance

@ 273K

[Ω ]

Secondary

Mirror Cover 230 340 Thermistor 2-wire 10k 1 Y

2 DEB 230 340 Thermistor 2-wire 10k 1 Y**

3 PCU 230 340 Thermistor 2-wire 10k 1 Y**

4 FPE 230 340 Thermistor 2-wire 10k 1 Y

5 Gimbal Ring 230 340 Thermistor 2-wire 10k 1 Y

Gimbal

Yoke

Encoder

230 340 Thermistor 2-wire 10k 1 Y

Gimbal

Yoke Motor 230 340 Thermistor 2-wire 10k 1 Y

LMI

Housing 230 340 Thermistor 2-wire 10k 1 Y

Relay

Housing 230 340 Thermistor 2-wire 10k 1 Y

10 N/A N/A N/A N/A N/A N/A N/A N/A

11 LMI FPA 230 340 Thermistor 2-wire 10k 1 Y

Yoke

Azimuth

Bearing

230 340 Thermistor 2-wire 10k 1 Y

13 HSI FPA 230 340 Thermistor 2-wire 10k 1 Y

Azimuth

Encoder 230 340 Thermistor 2-wire 10k 1 N

Elevation

Encoder 230 340 Thermistor 2-wire 10k 1 N

FPE

Radiator 230 340 Thermistor 2-wire 10k 1 N

17 HSI Radiator 230 340 Thermistor 2-wire 10k 1 N

Aft Metering

Structure 230 340 Thermistor 2-wire 10k 1 Y

Gimbal Base

Azimuth

Motor

230 340 Thermistor 2-wire 10k 1 Y

* TSTAT (Thermostat) indicates that this location is also monitored by a thermostatically controlled survival heater (see Section 6.1.6).

** If the DEB and PCU require thermostatically controlled survival heaters to meet the survival temperature ranges, these would be the responsibility of the Host Spacecraft.

Conversion of temperature to resistance can be achieved by applying a piecewise-linear fit of the temperature vs resistance curves included in GSFC standard S-311-P-18 Rev G, Table V.

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7 PAYLOAD TO HOST THERMAL INTERFACES

7.1.1 PCU Thermal Interface

The Host Spacecraft provides all necessary thermal control for the PCU. The PCU will dissipate less than 65 W into the Host Spacecraft thermal interface hardware (i.e., heat pipes or straps mounted to PCU side walls) averaged over any 180-minute period, and less than 80 W peak as averaged over any 10-millisecond period. For analysis purposes, the payload recommends using a predicted average power dissipation value of 48 W (consistent with CDR timeframe analysis), distributed as 24 W to each PCU side wall interface.

7.1.2 DEB Thermal Interface

The Host Spacecraft provides all necessary thermal control for the DEB. The DEB will dissipate less than 6.0 W into the Host Spacecraft thermal interface hardware (i.e., to the DEB baseplate) averaged over any 180-minute period, and less than 10 W peak as averaged over any 10-millisecond period. For analysis purposes, the payload recommends using a predicted average power dissipation value of 5.4 W (consistent with CDR timeframe analysis).

7.2 Orbital Thermal Environment

The payload orbital thermal environment is assumed to be that of a typical geosynchronous

(GEO) earth orbit.

Table 7-1: Payload Assumed GEO Thermal Environment

Season Beta Angle

Solar Flux

(W/m2)

Albedo

(W/m2)

IR

(W/m2)

Summer -23.4 1280 0.29 227

Winter 23.4 1420 0.31 241

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8 PAYLOAD TO HOST SPACECRAFT DATA INTERFACES

8.1 Host Spacecraft Data Interface

The Host Spacecraft provides the Host Spacecraft Attitude, and Ephemeris data over the

SpaceWire interface to support the payload LOS disturbance rejection. The angular rate data from the IRU comes over an RS-422 interface.

8.2 Command & Data Interfaces

For the payload-to-Host Spacecraft command and data interface, the payload supports a primary

SpaceWire physical connection to the Host Spacecraft. The command and data interface contains

Host Spacecraft commands to the payload, periodic payload telemetry, and payload mission science data.

The command and data telemetry are specified in GLIMR_05.09.001 GLIMR Command

Database and GLIMR_05.09.002 GLIMR Telemetry Database.

8.2.1 1PPS Signal Interface

The Host Spacecraft provides to the payload a One Pulse per Second (1PPS) RS-422 discrete signal for the SpaceWire timestamp synchronization.

Table 8-1: 1-PPS Signal Characteristics

Parameter Characteristic

Signal Type RS-422 (3.3V)

Rise Time <= 100 ns

Fall Time <= 100 ns

Pulse Width 10 µs ± 1 µs

Pulse Period 1.0 s ± 200 ns

Source termination resistance 51 Ohm

8.2.2 SpaceWire Interface

The Host Spacecraft to payload SpaceWire link has a signaling rate of 100 MHz +/-1%. This rate supports the full-duplex throughput of up to 72 Mbps per direction.

The payload supports logical addressing for SpaceWire communication.

For SpaceWire messages sent from the Host Spacecraft/ground to the payload, the Target

Logical Address (TLA) assignments are:

• Receiver TLA = 0xE0

• Sender TLA = 0xF0

For SpaceWire messages sent from the payload to the Host Spacecraft, the TLA assignments are:

• Receiver TLA = 0xF0

• Sender TLA = 0xE0

Release Date: 7/1/2026 Page 43 of 76

These TLA addresses are programmable and can be changed if required.

The Host Spacecraft encodes and decodes all payload SpaceWire interface messages per RMAP in accordance with ECSS-E-ST-50-52C. The detailed message level protocol is described in the

GLIMR_05.07.001 GLIMR Software User Manual.

Table 8-2: SpaceWire Message Types

Message

Type Direction Frequency Size Throughput

Normal

Telemetry From PL 1 Hz ≤ 4096 bytes < 4 kbps

Mission

Data (HSI) From PL 500-1000 Hz ≤ 4096 bytes 18 Mbps

Mission

Data (LMI) From PL 1000-3000 Hz ≤ 4096 bytes 10-50 Mbps

Memory

Download From PL <= 1000 Hz ≤ 4096 bytes 4 Mbps

Payload

Commands To PL <= 50 Hz ≤ 4096 bytes < 4 kbps

Host

Spacecraft

Attitude &

Ephemeris

To PL 1 Hz ≤ 4096 bytes < 4 kbps

Memory

Upload To PL <= 1000 Hz ≤ 4096 bytes 4 Mbps

8.2.2.1 SpaceWire Message Formats

All SpaceWire packets for commands and Host Spacecraft data from the Host Spacecraft, or telemetry to the Host Spacecraft.