Attachment C-418-XO-PSPEC-0051_V_1_3_final.pdf

Effective Date: February 18, 2022 418-XO-PSPEC-0051 Responsible Organization: GeoXO Flight Project/Code 418 Baseline 1.0

To verify the correct version of this document, please contact the GeoXO Series Configuration Management Office.

Geostationary Extended Observations (GeoXO)

Spacecraft (SC) Functional and Performance Specification (F&PS)

Signature/Approval Page

Prepared by:

Email approved by:

02/22/2022

Thomas M. Kenney GeoXO Flight Project, Deputy Mission Systems Engineer NASA GSFC, Code 418

Date

Reviewed by:

Electronically approved by:

02/16/2022

Steven W. Bidwell Date GeoXO Flight Project, Mission Systems Engineer

Concurred by:

Electronically approved by:

02/16/2022

Michelle P. Rizzo Date GeoXO Flight Project, Observatory Manager

Approved by:

Electronically approved by Christopher Wheeler for:

02/18/2022

Jason H. Hair Date GeoXO Flight Project Manager

/GeoXO Flight Project Spacecraft

SCFPS

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification (F&PS)

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Contents

1 1Introduction

1.1 1Scope

1.2 1GeoXO Mission Objectives

1.3 1Spacecraft Description (Common)

1.4 1East /West Spacecraft Description

1.5 2Central Spacecraft Description

1.6 2Mission Phases

1.7 3Definitions

2 8Documents

2.1 8Applicable Documents

2.2 9Reference Documents

3 10Requirements

3.1 10Spacecraft Payload Definition

3.2 10Satellite System Requirements

3.2.1 10Mission Orbit and Location

3.2.2 11Reliability

3.2.3 11Compatibility Requirements

3.2.4 11Launch Vehicle

3.2.5 12Fault Management

3.2.6 12Autonomous Operations

3.2.7 12Yaw Flip

3.2.8 13Orbital Debris

3.2.9 13Natural and Induced Environment

3.2.10 13Contamination Requirements

3.2.11 13Measurement Units

3.2.12 13Spacecraft Access

3.2.13 13Reserved (CCR-X00148A)

3.2.14 13Spacecraft Command & Telemetry During Launch and Orbit Raising (CCR- X00148A)

3.2.15 14Data Latency

3.2.16 14Spacecraft Resource Margins

3.2.17 15Electrostatic Arc-Discharge Susceptibility

3.2.18 15External Surface-to-Surface Direct Discharge

3.2.19 15Deep Dielectric Charging

Project: GeoXO Flight Project Spacecraft Module: SCFPS Baseline Version: 1.3

Contents ii

3.3 15Interfaces

3.3.1 15Instrument Interfaces

3.4 16Spacecraft Requirements

3.4.1 16Electrical Power

3.4.1.1 16Instrument Payload with Power Margin

3.4.1.2 16Energy

3.4.1.3 16Power Conditioning and Distribution

3.4.1.4 17Solar Array

3.4.1.5 17Lithium-Ion Batteries

3.4.1.5.1 17Battery Functional and Performance Requirements

3.4.1.5.2 18Battery Design and Construction Requirements

3.4.1.5.3 19Battery Launch Vehicle Payload Safety Requirements

3.4.2 19Command and Data Handling

3.4.2.1 19Command Verification

3.4.2.2 19Spacecraft Command Status Reporting

3.4.2.3 19Command Security

3.4.2.4 20Watchdog Timer

3.4.2.5 20Command Storage

3.4.2.6 20Command Order

3.4.2.7 20Packetized Telemetry

3.4.2.8 20Non-Packetized Telemetry

3.4.2.9 20Diagnostic Telemetry Mode

3.4.2.10 20Telemetry and Sensor Data Formatting

3.4.2.11 21Transfer Frame Routing

3.4.2.12 21Telemetry Monitor

3.4.2.13 21On-Board Computer (OBC)

3.4.2.14 22Time and Timing

3.4.3 22Communications

3.4.3.1 22Spacecraft Telemetry and Command Data Communications

3.4.3.1.1 22Orbit Raising, Tracking, Telemetry, and Commanding (ORTT&C) Communications

3.4.3.1.2 22Command and Data Acquisition Station (CDAS) Communications

3.4.3.2 23Spacecraft Raw Data Transmission

3.4.3.3 23Data Collection System (DCS) Communications (CCR-X00148A)

3.4.3.4 23Spacecraft Tracking

3.4.4 23Mechanical and Structural

3.4.4.1 23Structural Integrity

3.4.4.2 23Instrument Payload Mass

3.4.4.3 23Structural Analysis

Contents iii

3.4.4.4 23Design Limit Loads

3.4.4.5 24Structural Nonlinear Loads

3.4.4.6 24Structural Strength Design Factors

3.4.4.7 24Structural Deflections

3.4.4.8 24Material Properties

3.4.4.9 24Critical Member Design Values

3.4.4.10 24Redundant Member Design Values

3.4.4.11 24Selective Design Values

3.4.4.12 24Structural Reliability

3.4.4.13 25Mechanical Attachment

3.4.4.14 25Fluid Filter Performance Under Contamination

3.4.4.15 25Pressurized System

3.4.4.16 25Mechanisms

3.4.5 27Thermal

3.4.5.1 27Thermal Control

3.4.5.2 27Thermal Design Requirements

3.4.5.2.1 27Heaters

3.4.5.2.2 27Radiators

3.4.5.2.3 27Heat Pipes

3.4.6 28Guidance, Navigation, and Control (GN&C)

3.4.6.1 28Orbit Determination

3.4.6.2 28Attitude Determination

3.4.6.3 28Attitude Control

3.4.6.4 29Spacecraft to Instrument Disturbances

3.4.6.5 30Instrument to Spacecraft Disturbances

3.4.6.6 30Instrument Interface Outages

3.4.6.7 31Momentum Management

3.4.6.8 31Spacecraft Contingency Modes

3.4.6.8.1 32Coarse Earth-Pointing Mode

3.4.6.8.2 32Sun-Pointing Safe-Hold Mode

3.4.6.8.3 33Robotic Servicing (CCR-X00220)

3.4.6.8.3.1 33Robotic Servicing ACS (CCR-X00220)

3.4.6.8.3.2 34Robotic Servicing Fiducials (CCR-X00220)

3.4.6.8.3.3 34Robotic Servicing Mechanical Interface (CCR-X00220)

3.4.6.8.3.4 34Robotic Servicing Electrical Interfaces (CCR-X00220)

3.4.6.9 34GN&C Commands and Telemetry

3.4.6.9.1 34End-to-End GN&C Phasing

3.4.6.10 34GN&C Design Requirements

Contents iv

3.4.6.10.1 34Rigid-Body Stability Margins

3.4.6.10.2 35Time Delay

3.4.6.10.3 35Flexible-Body Modes

3.4.6.10.4 35Reaction Wheel Momentum Margin

3.4.6.11 35Propulsion System

3.4.6.11.1 35Purging of Residual Test Fluids

3.4.6.11.2 35Safety Electrical Disconnects

3.4.6.11.3 35Overtemperature Protection

3.4.6.11.4 36Unintended Propellant Vapor Ignition

3.4.6.11.5 36Propellant Leakage Dual Fault Tolerant

3.4.6.11.6 36Thruster and Venting Impingement

3.4.6.11.7 36Pressure Surges (CCR-X00148A)

3.4.7 36Flight Software

3.4.7.1 36Language and Methodology

3.4.7.2 36Flight Software Upload

3.4.7.3 37Flexibility and Ease of Software Modification

3.4.7.4 37Version Identifiers

3.4.7.5 37Flight Processor Resource Sizing

3.4.7.6 37Software Event Logging

3.4.7.7 38Soft Reset

3.4.7.8 38Memory Integrity

3.4.7.9 38Memory Dump

3.4.7.10 38Telemetry

3.4.7.11 39Stored Command Processing

3.4.7.11.1 39Command Execution

3.4.7.11.2 39Stored Command Buffer Management

3.4.7.12 39Bootstrap Software (CCR-X00185)

3.5 39Spacecraft Ground Support Equipment (GSE)

3.5.1 40GSE General Requirements

3.5.2 40Electrical Ground Support Equipment (EGSE)

3.5.2.1 40EGSE Command and Data Handling

3.5.2.2 40EGSE Command Interface to Mission Management Center

3.5.2.3 40EGSE Telemetry Interface to Mission Management Center

3.5.2.3.1 40EGSE Telemetry to Instrument EGSE

3.5.2.3.2 40EGSE Failure

3.6 40Flight Software Development Environment

3.7 41Spacecraft Hardware Simulator (SHS)

3.7.1 41SHS General Requirements

Contents v

3.7.2 42Unique SHS Instantiations

3.7.3 42SHS Control

3.7.4 42SHS Status

3.7.5 42SHS Data Logging

3.7.6 43SHS Modeling Requirements

3.7.7 43SHS Performance

3.7.8 44SHS Interface

3.8 44Satellite Software Simulator (S3)

3.8.1 44S3 General Requirements

3.8.2 45S3 Control

3.8.3 45S3 Status

3.8.4 45S3 Data Logging

3.8.5 46S3 Modeling Requirements

3.8.6 47S3 Performance

3.8.7 47S3 Interface

3.9 47Spacecraft to Instrument Interface Simulator

3.10 47Spacecraft Command and Telemetry Simulator (SCTS) (CCR-X00199)

4 48Design Verification Requirements

4.1 48Electrical Functional and Performance Verification Requirements

4.1.1 48Electrical Interface Tests

4.1.2 48RESERVED (CCR-X00220)

4.1.3 48RESERVED (CCR-X00220)

4.1.4 48Performance Operating Time and Trouble-Free Performance

4.2 48Structural and Mechanical Verification Requirements

4.2.1 49Mechanical Test Factors and Duration

4.2.2 51Minimum Workmanship

4.2.3 52Testing in Flight Configuration

4.2.4 52Structural Proof Testing

4.2.5 53Model Survey Characterization

4.2.6 53Structural Qualification

4.2.7 53Deployment and Articulation Verification

4.2.8 53Life Test

4.2.9 54Mechanical Clearance Verification

4.3 54RF Air-link Test

4.4 54Thermal Verification / Thermal-Vacuum Requirements

4.4.1 54Pressure (Unit/Component Level and Satellite System Level)

4.4.2 54Temperature Transition Rates (Unit/Component Level and Satellite System Level)

4.4.3 54Corona Operation (Unit/Component Level and Satellite System Level)

Contents vi

4.4.4 54Heater Verification (Satellite System Level)

4.4.5 55Hot and Cold Start Demonstrations (Unit/Component Level and Satellite System Level)

4.4.6 55Test Temperatures (Unit/Component Level and Satellite System Level)

4.4.6.1 56Qualification, Protoflight, and Acceptance Temperatures (Unit/Component Level and Satellite System Level)

4.4.6.2 56Test Temperature Control (Unit/Component Level and Satellite System Level)

4.4.7 56Cumulative Cycles (Unit/Component Level and Satellite System Level)

4.4.8 56Unit/Component Level Thermal-Vacuum Testing

4.4.9 57Satellite System Level Thermal-Vacuum Testing

4.4.10 57Satellite System Level Thermal-Balance (TB)

4.4.10.1 57Thermal Control Objectives

4.4.10.2 57Thermal Analytical Model Correlation Objectives

4.4.10.3 58Test Sequence

4.4.10.4 58Balance Points

4.4.10.5 58Satellite Thermal-Balance Configuration

4.4.10.6 58TB Steady State Criteria

4.5 58Cell and Battery Verification Requirements

4.5.1 59Battery Life Test

4.5.1.1 59Battery Cell Qualification/Life Test (CCR-X00200)

4.5.1.2 59Battery Accelerated Geostationary-Earth-Orbit (GEO) Life Testing

4.6 59Solar Array Verification Requirements

4.6.1 59Solar Array Component Qualification

4.6.2 60Solar Array Panel Qualification

4.6.3 60Flight Solar Array Panel Testing

4.6.4 61Solar Array Spacecraft Level Tests

4.7 62Test Condition Tolerances

5 63Appendix A: Battery System Definitions

6 65Appendix B: Acronyms

Contents vii

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1 Introduction

1.1 Scope

This document defines the high level functional, and performance requirements for the core Spacecraft, located within the East, West, and Central longitudinal slots, for the next generation Geostationary Operational Environmental – Extended Observation series (GeoXO) of Satellites. NASA manages the design and acquisition of the Spacecraft for the NOAA customer in a similar paradigm as preceding GOES series of Spacecraft. Throughout this document, the term ‘Spacecraft’ refers to all Spacecraft.

The GeoXO East and GeoXO West Spacecraft will accommodate identical instrument payloads which will differ from the instrument payload on the GeoXO Central Spacecraft. The Spacecraft and Satellites may be referred to as “East,” ‘West,” and “Central” when distinction is necessary and “East/West” for both “East” and “West” without distinction.

1.2 GeoXO Mission Objectives

The GeoXO Primary Mission Objectives are:

a) To maintain GOES mission continuity and quality in environmental observations in the 2030 – 2050 timeframe.

b) To provide enhanced environmental data products.

c) To improve services and data provided to the customer.

d) To be responsive to technology infusion and to meet evolving customer needs.

1.3 Spacecraft Description (Common)

The GeoXO Satellite will be launched from the Kennedy Space Center (KSC), Florida.

The government is responsible for all ground networks, and communications required for performing the launch and orbit raising activities. All launch and orbit raising activities will be performed from the NOAA Satellite Operations Facility (NSOF) in Suitland, MD. Once geostationary orbit is achieved, all activities will be performed using the NOAA Command and Data Acquisition Station (CDAS) at Wallops Island, Virginia as the primary ground station, and a Consolidated Backup (CBU) Facility at Fairmont, West Virginia.

During the transition to the GeoXO Satellite series, the operational constellation may contain GOES R-U Satellites and incumbent Satellites that have not yet been de-orbited.

The Spacecraft will be commanded throughout their mission lifetime from the NOAA Satellite Operational Control Center (SOCC) located at NSOF with the ground station radio frequency (RF) interface located at the Wallops CDAS and the CBU. The Satellite health and safety telemetry streams are received by the CDAS and/or the CDU and ground relayed to the SOCC for processing and monitoring. Similarly, the payload instrument science data will be received by the CDAS and/or the CBU.

1.4 East /West Spacecraft Description

This section addresses the functional and performance specifications for the East/West Spacecraft. The primary function of the GeoXO East/West Spacecraft is to accommodate and support its science payload comprised of the following instruments:

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⦁ GeoXO Imager (GXI)

⦁ GeoXO Ocean Color (OCX)

⦁ GeoXO Lightning Mapper (LMX)

An auxiliary function of the East/West Spacecraft is to accommodate and support the:

⦁ Data Collection Systems (DCS)

Note: In other documents the East/West Spacecraft and Satellites may be referred to as “GeoI” where the “I” references the Imager Instrument.

(CCR-X00148A)

1.5 Central Spacecraft Description

This section addresses the functional and performance specifications for the Central Spacecraft. The primary function of the GeoXO Central Spacecraft is to accommodate and support its science payload comprised of the following instruments:

⦁ GeoXO Sounder (GXS)

⦁ GeoXO Atmospheric Composition (ACX)

⦁ GeoXO Partner Payload

Note in other documents the Central Spacecraft and Satellites may be referred to as “GeoS” with “S” referencing the Sounder instrument.

1.6 Mission Phases

The GeoXO mission is divided into four phases as follows:

a) Pre-launch: The pre-launch phase provides for the design, fabrication, integration, and testing of each Spacecraft, Satellite integration and test, shipment to the launch site, and launch site activities. This phase also includes the development, installation, and testing of any associated ground equipment, software, and facilities required during all subsequent mission phases. This phase begins at Spacecraft contract award and continues until one day before launch (L-1 day).

b) Launch and Orbit Raising: The launch and orbit raising (if needed) phase consists of the launch, transfer orbit operations, orbital maneuvers to the checkout station, appendage deployments, and Spacecraft functional checkout when Spacecraft communications and other subsystems are validated on-orbit.

This phase begins at day L-1 and continues until engineering handover to NASA.

Note: launch vehicle and launch services will be provided by government.

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c) Post-Launch Test (PLT): The post-launch test phase consists of all activities required to confirm compliance with the Spacecraft requirements (this document), Instrument requirements, and Ground System requirements. The PLT addresses all Spacecraft subsystems, instruments, instrument calibration, ground system characterization, and product assurance. It is during PLT, that all Image Navigation and Registration (INR) requirements are confirmed. This phase begins at the engineering handover and continues until the Spacecraft is declared operational.

d) Operations: The operational phase consists of routine Satellite and data handling activities performed throughout the life of the mission. This phase also includes any period when the Spacecraft is in on-orbit storage. The ground system processes instrument data and generates and distributes data products to the user community. During the Operations phase, NOAA is responsible for mission management including planning and scheduling, building, verifying, and sending command loads, routing of mission data, and monitoring and control of the Spacecraft. Mission life includes PLT, on-orbit storage, and Operations durations and begins once the Spacecraft has achieved geosynchronous orbit.

e) End of Life (EOL): At the end of its operational life, operators raise the Spacecraft orbit to nominally 300 kilometers above geosynchronous altitude to vacate its orbital slot.

1.7 Definitions

Throughout this document, the following definitions apply. Additional definitions applicable to the battery system are found in Appendix A: Battery System Definitions.

Accuracy: Refers to the error in a measurement that is the difference between the measured and true value. It includes both systematic and random errors. Systematic errors must be estimated from an analysis of the experimental conditions and techniques. Random errors can be determined, and reduced, through repeated measurements under identical conditions and a Standard Deviation calculated. The magnitude of a random error is taken as three standard deviations (3).

Abnormal Operation: Encompasses unforeseen circumstances that are not handled via established contingency plans and operational states such as anomalous conditions or failures.

Absolute Time Commands: Stored Commands that have a time tag containing an absolute Spacecraft time.

Absolute Time Command Buffer: An allocated memory area used to store the absolute time commands.

Absolute Time Sequence Command: Sequences of commands that are executed at the absolute time associated with each command in the sequence.

Allowable Flight Temperature: See Mission Allowable Temperature

Attitude Knowledge: Attitude knowledge is the difference between the true attitude and estimated attitude.

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Battery System: The battery system is the Satellite battery which can be a parallel configuration of batteries or a single battery comprised of a parallel/series combination of battery cells. Battery system interfaces can include telemetry, command, cell balancing, and test which can reside outside of the physical configuration of the battery.

Component: A functional subdivision of a subsystem and generally a self-contained combination of items performing a function necessary for the subsystem’s operation and synonymous with Unit. Examples include electronic box, transmitter, gyro package, actuator, motor, battery, instrument electronics unit, sensor unit.

Computer Software Configuration Item (CSCI): A group of software treated as a single entity by a configuration management (CM) system.

Critical Command: A command issued by the ground to the Satellite producing changes in operational parameters that have the potential to adversely affect the health and safety of the Satellite (Spacecraft and/or instruments) or resulting in irreversible changes to the operational state of the Satellite.

Critical Structural Members: Structural members are classified as critical when their failure would result in loss of structural integrity of the flight units.

Credible Failure: A condition that has a reasonable likelihood of occurrence over the mission life. With consideration of reliability and engineering experience, the government technical authority will determine credibility of failure. For the purposes of this document, failures of structure, pressure vessels, and pressurized lines and fittings are not considered credible failure modes if those elements comply with the applicable requirements of this document.

Data Latency: The time interval between the end of a data collection sequence and the time that the data are available on the Ground.

Design Life: The minimum duration during which the Spacecraft must be capable of performing all mission operational requirements.

Eclipse: Defined as when the solar disk is completely or partially occulted by the Earth or Moon as viewed from the Spacecraft.

Essential Loads: Power loads that are essential for safety and health of the space vehicle.

(CCR-X00199)

Fully Functional Configuration: Being able to collect the full complement of science data; determine instrument response changes; acquisition of sensor health and status data; generation of sensor, calibration, monitoring, health and status data streams; and reception and execution of command and control data.

Fault Management: Process of detecting and reacting to the occurrence of a fault or anomaly, whether in hardware or software.

Hazardous Command: A command whose execution (including inadvertent, out-of-sequence, or incorrectly executed) could lead to an identified critical or catastrophic

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hazard, or a command whose execution can lead to a reduction in the control of a hazard (including reduction in failure tolerance against a hazard or the elimination of an inhibit against a hazard).

Launch: The period between lift off and the separation of the GeoXO Satellite from the launch vehicle.

Limit Loads: Limit loads are defined as all worst-case load conditions including temperature effects from the environments expected during all phases of the structure's service life including manufacturing, ground handling, transportation, environmental testing, integration, pre-launch, launch and on-orbit operations and storage.

Margin (%) is represented by the equation: Margin (%) = [(Available Resource - Worst-case Estimate) / (Available Resource)] x 100.

Measurement Resolution: Resolution of the A/D converter.

Mission Critical Events: Critical events are the events in a mission that must occur in sequence to achieve overall mission success. If the event is not executed properly and without anomaly, it could result in failure to accomplish the mission. Examples of mission-critical events include: launch and ascent, separation from the launch vehicle, power-up of major components or subsystems, deployment of mechanisms and/or mission-critical appendages, initial thruster firings and all planned propulsive maneuvers required to establish mission orbit and/or achieve safe attitude.

Mission Allowable Temperature: Mission Allowable Temperature (MAT) is defined as the worst case thermal analysis predictions with a minimum of 5°C uncertainty for all mission phases launch, orbit raising, ascent and on-orbit operation from the beginning of life to the end of life for operational mode. The uncertainty value during the design phase is expected to be larger, narrowing as the design matures, and at the conclusion of satellite-level thermal testing the MAT still contains 5°C uncertainty.

Mission Management Functionality: The Mission Management functions of the GeoXO Ground System which is required to safely and reliably control and communicate with the Satellite, monitor Satellite systems performance, and capture all raw science data necessary for production of mission data.

Non-Essential Loads: Power loads that can be powered off without adversely affecting the minimum controllability and commandability of the Spacecraft.

Non-Operational mission allowable Temperature: Non-Operational mission allowable Temperature (NOT) is defined as the worst case thermal analysis predictions with a minimum of 5°C uncertainty for all mission phases launch, orbit raising, ascent and on-orbit operation from the beginning of life to the end of life for dormant, non-operational and not powered components and subsystems. The uncertainty value during the design phase is expected to be larger, narrowing as the design matures, and at the conclusion of satellite-level thermal testing the non-operating MAT still contains the 5°C uncertainty.

Normal Operation: Nadir pointing mode to support science data collection.

Operational Loads: Operational load is defined as the expected on-orbit structural loads.

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Payload: Self-Contained instrument, sensor, or device that fulfills some mission objective.

Random Errors: Unrepeatable statistical fluctuations in the measured data due to the precision limitations of the measurement.

Red Limit: A red limit is the value of a telemetry point requiring TMON action when the point is either greater than or less than, depending on the point monitored, the stored limit value.

Relative Time Commands: Stored commands that have a time tag containing a time that is relative to when the previous command was executed.

Relative Time Command Sequence: A named sequence of Relative Timed Commands, which perform a specific function.

Relative Time Sequence Buffer: An allocated memory area used to store the named Relative Command Sequences.

Relative Time Sequences Command: Sequences of commands that can be sent from the on-board processor following a pre-defined sequence with execution times relative when the previous command in the sequence was issued.

Redundant Structural Members: Structural members are classified as redundant when their failure would result in the redistribution of applied loads to other structural members without loss of structural integrity.

Resolution: Ability to distinguish two adjacent features in the spectral, spatial, or temporal domain.

Satellite: The GeoXO Satellite consists of the Spacecraft and GFE science instrument payload.

Single-Point Failure: Single component, wiring, connector failure, software glitch, or computer failure that results in the permanent loss of the Spacecraft's ability to perform its primary mission for the intended design lifespan.

Space Segment: The Space Segment (SS) consists of a constellation of geostationary Satellites and required pre-launch ground support equipment.

Spacecraft: The Spacecraft consists of the flight hardware and software required to accommodate the science instrument payload. The Spacecraft includes the DCS auxiliary communication.

Spacecraft Bus: Satellite without instruments and DCS auxiliary communication.

Station Keeping: On-orbit Spacecraft maneuver that corrects for orbital drifts.

Stored Commands: Commands that are stored in the on-board computers RAM that are executed at a specified time. Each stored command contains a either an absolute or relative time tag that indicates when the command will be executed.

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Systematic Errors: Consistently reproducible inaccuracies in the measured data due to the precision limitations of the measurement and bias.

TMON Rule: A telemetry monitor (TMON) Rule defines the action to be taken when a red limit violation is detected.

Unit: A functional subdivision of a subsystem and generally a self-contained combination of items performing a function necessary for the subsystem’s operation and synonymous with Component. Examples include electronic box, transmitter, gyro package, actuator, motor, battery, instrument electronics unit and instrument sensor unit.

Yaw Flip: An on-orbit maneuver that rotates the Spacecraft 180° about the Spacecraft z axis (yaw). The net effect reverses the directions of the roll and pitch axes relative to the Earth, while maintaining the yaw axis pointing at nadir.

(CCR-X00148A) (CCR-X00199)

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2 Documents

2.1 Applicable Documents

The following documents form a part of this specification to the extent specified herein.

1. GeoXO General Interface Requirements Document (GIRD), 418-XO-GIRD-0041

2. GXI Unique Instrument Interface Document (UIID), 418-XO-GXIUIID-0023

3. OCX Unique Instrument Interface Document (UIID), 418-XO-OCXUIID-0062

4. LMX Unique Instrument Interface Document (UIID), 418-XO-LMXUIID-0067

5. GXS Unique Instrument Interface Document (UIID), 418-XO-GXSUIID-0030

6. ACX Unique Instrument Interface Document (UIID), 418-XO-GXSUIID-0072

7. Partner Payload Unique Instrument Interface Document (PPXUIID), 418-XO-

PPXUIID-0141

8. SS to Ground Located Command, Control, and Communications (SS-C3S) Interface Requirements Document (IRD), 418-XO-IRD-0043

9. SS to Data Collection System (SS-DCS) Interface Requirements Document (IRD), 418-XO-IRD-0044

10. GeoXO Payload Resources Allocation Document (PRAD), 418-XO-RAD-0037

11. Consultative Committee for Space Data Systems, Recommendation for Space Data System Standards, AOS Space Data Link Protocol, Blue Book, September 2015, Specification, CCSDS 732.0-B-3

12. Consultative Committee for Space Data Systems, Recommendation for Space Data System Standards, Communications Operation Procedure-1, Blue Book, September 2010, CCSDS 232.1-B-2

13. NASA-STD-8719.24, NASA Expendable Launch Vehicle Payload Safety Requirements (Base and Annex).

14. MMPDS-15, Metallic Materials Properties Development and Standardization (MMPDS) Handbook,1 July 2020

15. NASA-STD-5019A-C3, Fracture Control Requirements for Spaceflight Hardware, 14 August 2020

16. MSFC-STD-3029, REV. A, Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments, NASA MSFC, 25 February 2005

17. The Radiation Environment for Electronic Devices on GeoXO Satellites, 418-XO- RPT-0042, August 14, 2021

18. AIAA S-111A-2014, Qualification and Quality Requirements for Space Solar Cells

19. AIAA S-112A-2013, Qualification and Quality Requirements for Electrical Components on Space Solar Panels

20. RESERVED.

21. GSFC-STD-7000B, General Environmental Verification Standard (GEVS) For GSFC Flight Programs and Projects, April 28, 2021

22. Standard for the Design and Fabrication of Ground Support Equipment, NASA- STD-5005D, Change 1, 14 June 2013

23. NASA-STD-8719.14B, Process for Limiting Orbital Debris

24. NPR 8715.6B, NASA Procedural Requirements for Limiting Orbital Debris and Evaluating the Meteoroid and Orbital Debris Environments, February 16, 2017, Page 9 of 66 Printed Tuesday, August 8, 2023

ID

SCFPS28

SCFPS29

SCFPS30

Object Number

2.1.0-1

2.2

2.2.0-1

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification

(F&PS)

NASA-STD-8719.14B, Process for Limiting Orbital Debris

25. IEEE/ASTM SI 10-2016, American National Standard for Metric Practice

26. NIST Special Publication 800-53, Rev. 5, Security and Privacy Controls for Information Systems and Organizations, September 2020

(CCR-X00185)

2.2 Reference Documents

The following documents provide reference material for part of this document.

1. General Specification for Assemblies, Moving Mechanical, for Space and Launch Vehicles, Document Number MIL-A-83577B, February 1, 1988

2. Space Mechanisms Handbook, Document Number NASA TP-1999-206988

Page 10 of 66 Printed Tuesday, August 8, 2023

ID

SCFPS31

SCFPS32

SCFPS33

SCFPS34

SCFPS35

SCFPS36

SCFPS37

SCFPS38

SCFPS39

Object Number

3.0-1

3.1

3.1.0-1

3.1.0-2

3.2

3.2.1

3.2.1.0-1

3.2.1.0-2

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification

(F&PS)

3 Requirements

The following requirements terminology is used throughout this document:

The use of “shall” designates a requirement that must be met.

The use of “will” designates a statement of fact or intention of the government.

The use of “may” designates that permission has been granted by the government.

The term “(TBD)” means, “to be determined” and is used when no value is available with subsequent study needed to obtain it.

The term “(TBR)” means “to be refined/reviewed” for a value that is subject to review for appropriateness and is subject to revision. The vendor is liable for compliance with the information marked “TBR” as if the “TBR” notation did not exist.

3.1 Spacecraft Payload Definition

Each East/West Spacecraft in the GeoXO constellation shall accommodate the following science payload instrumentation and communication services.

a) GeoXO Imager (GXI)

b) GeoXO Ocean Color (OCX)

c) GeoXO Lightning Mapper (LMX)

d) Data Collection System (DCS)

(CCR-X00148A)

Each Central Spacecraft in the GeoXO constellation shall accommodate the following science payload instrumentation and communication services.

a) GeoXO Sounder (GXS)

b) GeoXO Atmospheric Composition (ACX)

c) Partner Payload

3.2 Satellite System Requirements

3.2.1 Mission Orbit and Location

The GeoXO East/West Spacecraft will provide coverage from two locations, one (East) at 75 west longitude and another (West) at 137 west longitude. Each of these locations may be adjusted within a ±10 window during operations. An additional one or two East/West Spacecraft will be maintained as spares and periodically checked while in on-orbit storage positioned at 94 (primary storage) or 92 (secondary storage) west longitude. GeoXO East/West Spacecraft post-launch test and checkout activities will be conducted at 88 west longitude.

The GeoXO Central Spacecraft will provide coverage from one location at 105 west longitude. The Central Spacecraft location may be adjusted within a ±15 window during operations. An additional Central Spacecraft will be maintained as a spare and periodically checked in on-orbit storage positioned at 94 primary location) or 92 (secondary location) west longitude. GeoXO Central Spacecraft post-launch test and checkout activities will be conducted at 88west longitude.

Page 11 of 66 Printed Tuesday, August 8, 2023

ID

SCFPS40

SCFPS41

SCFPS43

SCFPS44

SCFPS45

SCFPS46

SCFPS47

SCFPS48

SCFPS49

SCFPS50

SCFPS51

SCFPS52

SCFPS53

SCFPS54

SCFPS55

Object Number

3.2.1.0-3

3.2.1.0-4

3.2.1.0-5

3.2.1.0-6

3.2.1.0-7

3.2.2

3.2.2.0-1

3.2.2.0-2

3.2.3

3.2.3.0-1

3.2.3.0-2

3.2.4

3.2.4.0-1

3.2.4.0-2

3.2.4.0-3

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification

(F&PS)

The Spacecraft shall provide delta-V to execute the following station location changes during its on-orbit life:

a) Four on-orbit relocations at a drift rate of 1°/day including 4 start and 4 stop delta-V’s.

b) Two emergency relocations at a drift rate of 3°/day including 2 start and 2 stop delta-V’s.

c) Boost from geostationary orbit at end-of-life longitude to end-of-life super-synchronous orbit with a perigee of no less than 300 km above geostationary altitude.

For station location changes, the time to achieve a drift rate up to 3°/day shall not exceed 13 hours.

The Spacecraft shall provide station-keeping to maintain a north/south position of ± 0.05° about the equator and an east/west position of ±0.05° of the on-station longitude during normal operational phases of the mission.

The Spacecraft shall achieve the mission orbit after separation from the launch vehicle, with up to 3-sigma launch dispersions.

The Spacecraft shall maintain the mission orbit at station locations for a minimum of 15 years.

3.2.2 Reliability

The Spacecraft Bus shall have a Reliability of at least 0.81 after 10 years of on-orbit operation, preceded by up to 5 years of ground storage and up to 5 years of on-orbit storage.

The DCS payload shall have a reliability of at least 0.9 after 10 years of on-orbit operation, preceded by up to 5 years of ground storage and up to 5 years of on-orbit storage. (CCR-X00148A)

3.2.3 Compatibility Requirements

The Spacecraft shall simultaneously and continuously operate Spacecraft subsystems, DCS, and instrument payloads in all their operational modes through end of life, including eclipse periods. (CCR-X00148A)

The Spacecraft and its units shall not generate electromagnetic interference that could adversely affect its own performance or the performance of instruments and DCS on the Spacecraft, launch vehicle, or launch site. (CCR-X00148A)

3.2.4 Launch Vehicle

The Satellite shall be compatible with a Falcon Heavy, Vulcan Centaur, and New Glenn launch vehicles.

The Spacecraft shall be compatible with vertical and horizontal processing and transportation at the launch-site after the Spacecraft is fueled for launch.

The Spacecraft shall be designed to launch any day of the year.

Page 12 of 66 Printed Tuesday, August 8, 2023

ID

SCFPS56

SCFPS57

SCFPS58

SCFPS59

SCFPS60

SCFPS61

SCFPS62

SCFPS63

SCFPS64

SCFPS65

SCFPS66

SCFPS67

SCFPS68

SCFPS69

SCFPS798

SCFPS799

SCFPS800

SCFPS70

SCFPS71

Object Number

3.2.4.0-4

3.2.5

3.2.5.0-1

3.2.5.0-2

3.2.5.0-3

3.2.5.0-4

3.2.5.0-5

3.2.5.0-6

3.2.5.0-7

3.2.5.0-8

3.2.5.0-9

3.2.6

3.2.6.0-1

3.2.6.0-2

3.2.6.0-3

3.2.6.0-4

3.2.6.0-5

3.2.7

3.2.7.0-1

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification

(F&PS)

The Spacecraft shall achieve geosynchronous orbit at the post-launch test and checkout station within 30 days of launch.

3.2.5 Fault Management

The Spacecraft Fault Management System (FMS) shall provide detection, isolation, and recovery from single credible faults to maintain the health and safety of the Satellite and continuity of operations.

No single ground command shall permanently preclude the Spacecraft from meeting the requirements of this specification.

The Spacecraft shall preclude a credible failure in any component from damaging the Spacecraft and permanently degrading the Satellite performance.

The Fault Management System shall continually monitor the performance of the subsystems and components and detect failures and anomalies.

The Spacecraft shall downlink real-time Fault Management System telemetry identifying failures and anomalies.

The Spacecraft shall prioritize the indications of Satellite faults in telemetry over the delivery of ancillary Fault Management System telemetry so that vital indicators are transmitted as they occur.

Each autonomous Fault Management System detection test and corrective action shall be individually enabled and disabled by ground commands without affecting the health and safety of the Satellite.

The Spacecraft shall survive the occurrence of any single credible failure in the Spacecraft without ground intervention for an indefinite period, except for ground commanding required for reaction wheel desaturation and station keeping.

During fault recovery, science data collection will not be required.

3.2.6 Autonomous Operations

The Spacecraft shall operate autonomously within specification without ground contact for a period of 7 days.

The Spacecraft shall maintain Spacecraft and instrument health and safety without ground support.

During Normal Operations, the Spacecraft shall perform autonomous onboard maneuver planning for station keeping and momentum unloading.

Upon ground command, the Spacecraft shall downlink maneuver plans.

The Spacecraft shall autonomously execute planned station keeping and momentum unloading maneuvers unless inhibited by ground command.

3.2.7 Yaw Flip

The Spacecraft shall perform a yaw flip upon ground commanding.

Page 13 of 66 Printed Tuesday, August 8, 2023

ID

SCFPS72

SCFPS73

SCFPS74

SCFPS75

SCFPS76

SCFPS77

SCFPS78

SCFPS79

SCFPS80

SCFPS81

SCFPS82

SCFPS83

SCFPS84

SCFPS826

SCFPS85

SCFPS88

SCFPS89

SCFPS90

Object Number

3.2.7.0-2

3.2.7.0-3

3.2.8

3.2.8.0-1

3.2.9

3.2.9.0-1

3.2.10

3.2.10.0-1

3.2.10.0-2

3.2.11

3.2.11.0-1

3.2.12

3.2.12.0-1

3.2.12.0-2

3.2.13

3.2.14

3.2.14.0-1

3.2.14.0-2

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification

(F&PS)

The Spacecraft shall meet all performance requirements without the need of yaw flip maneuvers. If performed, yaw flip maneuvers are conducted in service to the payload instruments.

Spacecraft yaw flip maneuvers shall be initiated by single command sequence from the ground and performed autonomously by the Spacecraft.

3.2.8 Orbital Debris

The Spacecraft shall limit the generation of orbital debris in compliance with NASA- STD-8719.14B [22] and NPR 8715.6B [23].

3.2.9 Natural and Induced Environment

The Spacecraft shall meet all requirements during and after exposure to both natural and induced environments in orbit as defined in the GIRD [1] and the space radiation environment defined in 418-XO-RPT-0042 [16].

3.2.10 Contamination Requirements

The Spacecraft design, including instrument layout, integration, test, ground handling, storage, and transportation shall comply with the Spacecraft and instrument contamination requirements as specified in GIRD [1] and Unique Instrument Interface Requirements Documents (UIIDs) [2][3][4][5][6].

Satellite degradation of hardware due to contamination on ground and during all mission phases shall not prevent the Satellite hardware from meeting mission requirements during the 15 years on-orbit lifetime.

3.2.11 Measurement Units

The International System of Units (SI) shall be used in accordance with IEEE/ASTM SI 10-2016 [24].

3.2.12 Spacecraft Access

The Spacecraft design shall accommodate the installation and removal of any instrument and its associated electronics without requiring the removal of other instruments’ hardware, during all pre-launch phases. (CCR-X00148A)

The Spacecraft shall leave adequate clearance between electronic units and surrounding structures to provide access to mounting hardware, access to connectors including a view during mating, and space for interfacing harness service loops. (CCR- X00148A)

3.2.13 Reserved (CCR-X00148A)

3.2.14 Spacecraft Command & Telemetry During Launch and

Orbit Raising (CCR-X00148A)

The Spacecraft shall provide telemetry during the execution of all-mission critical events.

The Spacecraft shall provide telemetry during launch, and orbit raising.

Page 14 of 66 Printed Tuesday, August 8, 2023

ID

SCFPS91

SCFPS92

SCFPS93

SCFPS94

SCFPS95

SCFPS96

SCFPS97

Object Number

3.2.14.0-3

3.2.14.0-4

3.2.14.0-5

3.2.15

3.2.15.0-1

3.2.16

3.2.16.0-1

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification

(F&PS)

When the Spacecraft is in stowed (launch) configuration, it shall not obscure visibility of any attitude sensor required for acquisition.

The Spacecraft shall support a command and telemetry link while in any Spacecraft orientation in stowed and deployed configurations.

The Spacecraft shall accept valid commands after completion of the flight software start up process.

Note: This includes ground test, pre-launch, launch vehicle separation, orbit raising, and on-orbit operations.

3.2.15 Data Latency

The Spacecraft shall transmit instrument data within one second of receipt.

3.2.16 Spacecraft Resource Margins

The Spacecraft resource margins at key Spacecraft reviews shall be in accordance with Table SCFPS97.

Note: Per Definitions, Spacecraft is the GeoXO Satellite without the Instrument Payload.

Note: The instrument payload resource margins are controlled by the government as described in the PRAD [9].

Table SCPFS97 Spacecraft Technical Resource Margins

Spacecraft Resource Margins

Resource SDR PDR CDR PSR

Spacecraft Mass (dry)

20% 15% 10% 0

Satellite Mass (wet)1

<= Launch Vehicle

Capacity

<= Launch Vehicle

Capacity

<= Launch Vehicle

Capacity

<= Launch Vehicle

Capacity

Spacecraft Power (wrt EOL

Capacity) 20% 15% 10% 5%2

Propellant 3σ 3σ 3σ 3σ

Telemetry and Command hardware channels

20% 15% 10% 5%

Notes:

1. Wet mass applies to the Satellite and Spacecraft will assume SCFPS mass

Page 16 of 66 Printed Tuesday, August 8, 2023

ID

SCFPS107

SCFPS109

SCFPS110

SCFPS111

SCFPS112

SCFPS801

SCFPS113

SCFPS114

SCFPS115

SCFPS116

Object Number

3.3.1.0-2

3.4

3.4.1

3.4.1.1

3.4.1.1.0-1

3.4.1.1.0-2

3.4.1.2

3.4.1.2.0-1

3.4.1.3

3.4.1.3.0-1

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification

(F&PS)

The Spacecraft shall meet all instrument resource allocations defined in the UIIDs [2] [3][4][5][6].

Note: PRAD [9] takes precedence over UIIDs [2][3][4][5][6] for Spacecraft resources.

3.4 Spacecraft Requirements

3.4.1 Electrical Power

3.4.1.1 Instrument Payload with Power Margin

The East/West Spacecraft shall be designed for the instrument power allocations given in the following PRAD [9] tables.

a) Table PRAD24, GeoI Instrument Average (300 sec avg Operational) Power Allocation

b) Table PRAD27, GeoI Instrument Maximum (20 ms avg Operational) Power Allocation

c) Table PRAD30, GeoI Instrument Average (72 min avg Survival) Power Allocation

d) Table PRAD33, GeoI Instrument Maximum (20 ms avg Survival) Power Allocation Table

Note: PRAD [9] takes precedence over UIIDs [2][3][4][5][6] for Spacecraft resources.

The Central Spacecraft shall be designed for the instrument power allocations given in the following PRAD [9] tables.

a) Table PRAD37, Instrument Average (300 sec avg) Operational Power Allocation

b) Table PRAD40, GeoS Instrument Maximum (20 ms avg) Operational Power Allocation

c) Table PRAD43, GeoS Instrument Average (72 min avg) Survival Power Allocation

d) Table PRAD46, GeoS Instrument Maximum (20 ms avg) Survival Power Allocation

Note: PRAD [9] takes precedence over UIIDs [2][3][4][5][6] for Spacecraft resources.

3.4.1.2 Energy

The Spacecraft solar array and batteries shall provide energy to perform all operating modes from LOR until the end of life.

3.4.1.3 Power Conditioning and Distribution

The gain and phase margins for Spacecraft power regulator and power supplies shall be in accordance with Table SCFPS116 at the indicated project milestone.

Note: These margins can be adjusted based on flight heritage hardware, if approved by the government.

Page 17 of 66 Printed Tuesday, August 8, 2023

ID

SCFPS116

SCFPS117

SCFPS118

SCFPS119

SCFPS120

SCFPS121

SCFPS122

SCFPS123

SCFPS124

SCFPS125

SCFPS126

SCFPS129

SCFPS133

Object Number

3.4.1.3.0-1

3.4.1.3.0-2

3.4.1.4

3.4.1.4.0-1

3.4.1.4.0-2

3.4.1.5

3.4.1.5.0-1

3.4.1.5.1

3.4.1.5.1.0-1

3.4.1.5.1.0-2

3.4.1.5.1.0-3

3.4.1.5.1.0-4

3.4.1.5.1.0-5

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification

(F&PS)

Table SCPS116: Power Regulator and Power Supply Stability Margins

Milestone Gain Margin (dB) Phase Margin (deg)

PDR 12 45

CDR 9 35

PSR 6 30

The Spacecraft shall connect and disconnect all non-essential loads by ground command.

Note: This does not preclude disconnecting power to non-essential loads in response to the Fault Management System’s monitoring system. (CCR-X00148A)

3.4.1.4 Solar Array

At Pre-Ship Review, the solar array shall meet all operational power requirements with a minimum of 5% power margin between the solar array available power and the solar array required power at end-of-life design conditions which assume worst-case season and one solar array circuit failure.

Note (1): A solar array string is assumed to be a fixed number of solar cells in series. A solar array circuit is assumed to be a fixed number of solar array strings. Circuits are electrically connected in parallel throughout the solar array.

Note (2): Exceptions to this requirement are 1) during eclipse portions of the orbit and

2) electric propulsion firing at operational power. Other short-term margin relief for selected peak power conditions can be allowed if approved by the government.

The stowed outboard solar array panel(s) shall supply power to the Spacecraft when exposed to sun illumination.

3.4.1.5 Lithium-Ion Batteries

The Spacecraft shall utilize lithium-ion batteries for the battery system.

3.4.1.5.1 Battery Functional and Performance Requirements

The measured 20 °C beginning-of-life (BOL) battery actual capacity (Cba) shall be greater than or equal to the battery rated capacity (Cbr). (CCR-X00200)

The GeoXO mission battery cell operating temperatures shall be within their mission allowable temperatures. (CCR-X00200)

The battery cells shall meet all performance requirements after exposure to survival temperatures which are 10 °C above their hot mission allowable temperature and 10 °C below their cold mission allowable temperature. (CCR-X00200)

The 20 °C beginning-of-life (BOL) battery system actual capacity (Cbsysa) shall be ≥ 120% of the battery system nameplate capacity (Cbsysn). (CCR-X00200)

The difference between the maximum and minimum cell capacity within the battery system shall not exceed 3%.

Page 18 of 66 Printed Tuesday, August 8, 2023

ID

SCFPS138

SCFPS139

SCFPS140

SCFPS141

SCFPS142

SCFPS143

SCFPS144

SCFPS145

SCFPS147

SCFPS148

SCFPS150

SCFPS151

SCFPS157

SCFPS159

Object Number

3.4.1.5.1.0-6

3.4.1.5.1.0-7

3.4.1.5.1.0-8

3.4.1.5.1.0-9

3.4.1.5.1.0-10

3.4.1.5.1.0-11

3.4.1.5.2

3.4.1.5.2.0-1

3.4.1.5.2.0-2

3.4.1.5.2.0-3

3.4.1.5.2.0-4

3.4.1.5.2.0-5

3.4.1.5.2.0-6

3.4.1.5.2.0-7

418-XO-PSPEC-0051, RM Version, Geostationary eXtended Observations (GeoXO) Functional and Performance Specification

(F&PS)

The battery system shall be sized to operate the Satellite on-orbit through all eclipses up to 72 minutes of total duration, with the mission depth-of-discharge (DODMIS) of any battery not to exceed 65% of the battery nameplate capacity (Cbn).

The battery system shall be sized to operate the Satellite through launch and transfer orbit such that battery DOD at beginning-of-life (DODBOL) of any battery does not exceed 70% of the battery actual capacity (Cba).

The battery system shall be sized to operate the Satellite through all on-orbit operations including outgas, station-keeping, storage, contingency, and end-of-life boost modes such that battery mission depth-of-discharge (DODMIS) of any battery does not exceed 70% of the battery nameplate capacity (Cbn).

The battery system shall be sized to meet all performance requirements after sustaining any single cell failure with the depth-of-discharge (DOD) of any battery not to exceed 75% of the battery nameplate capacity, Cbn.

The Spacecraft shall recharge the batteries to full state-of-charge at least once during each 24-hour period without over stressing the batteries, including eclipse season.

The battery system telemetry shall include, as a minimum, the following status:

a) Each cell voltage

b) Cell temperatures

c) Battery voltage

d) Battery temperature

e) Battery charge/discharge current

f) Cell balancing status

g) Bypass switch status

(CCR-X00200)

3.4.1.5.2 Battery Design and Construction Requirements

The battery shall require no external adjustments after start of Acceptance or Qualification testing.

The battery system shall include a power disconnect relay which allows emergency…