Attachment B- PSPEC.pdf

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

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 13Pre-Launch Satellite Testing

3.2.14 13Spacecraft Telemetry During Launch and Orbit Raising

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: 1.0

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 17Power 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 19Battery Design and Construction Requirements

3.4.1.5.3 21Battery Launch Vehicle Payload Safety Requirements

3.4.2 21Command and Data Handling

3.4.2.1 21Command Verification

3.4.2.2 21Spacecraft Command Status Reporting

3.4.2.3 21Command Security

3.4.2.4 21Watchdog Timer

3.4.2.5 21Command Storage

3.4.2.6 22Command Order

3.4.2.7 22Packetized Telemetry

3.4.2.8 22Non-Packetized Telemetry

3.4.2.9 22Diagnostic Telemetry Mode

3.4.2.10 22Telemetry and Sensor Data Formatting

3.4.2.11 22Transfer Frame Routing

3.4.2.12 22Telemetry Monitor

3.4.2.13 23On-Board Computer (OBC)

3.4.2.14 23Time and Timing

3.4.3 23Communications

3.4.3.1 23Spacecraft Telemetry and Command Data Communications

3.4.3.1.1 23Orbit Raising, Tracking, Telemetry, and Commanding (ORTT&C) Communications

3.4.3.1.2 24Command and Data Acquisition Station (CDAS) Communications

3.4.3.2 24Spacecraft Raw Data Transmission

3.4.3.3 24Data Collection Platform Relay (DCPR) Communications

3.4.3.4 24Spacecraft Tracking

3.4.4 24Mechanical and Structural

3.4.4.1 24Structural Integrity

3.4.4.2 24Instrument Payload Mass

3.4.4.3 24Structural Analysis

Contents iii

3.4.4.4 25Design Limit Loads

3.4.4.5 25Structural Nonlinear Loads

3.4.4.6 25Structural Strength Design Factors

3.4.4.7 25Structural Deflections

3.4.4.8 25Material Properties

3.4.4.9 25Critical Member Design Values

3.4.4.10 25Redundant Member Design Values

3.4.4.11 25Selective Design Values

3.4.4.12 25Structural Reliability

3.4.4.13 26Mechanical Attachment

3.4.4.14 26Fluid Filter Performance Under Contamination

3.4.4.15 26Pressurized System

3.4.4.16 26Mechanisms

3.4.5 28Thermal

3.4.5.1 28Thermal Control

3.4.5.2 28Thermal Design Requirements

3.4.5.2.1 28Heaters

3.4.5.2.2 28Radiators

3.4.5.2.3 28Heat Pipes

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

3.4.6.1 28Orbit Determination

3.4.6.2 29Attitude Determination

3.4.6.3 29Attitude Control

3.4.6.4 30Spacecraft to Instrument Disturbances

3.4.6.5 31Instrument to Spacecraft Disturbances

3.4.6.6 31Instrument Interface Outages

3.4.6.7 31Momentum Management

3.4.6.8 32Spacecraft Contingency Modes

3.4.6.8.1 32Coarse Earth-Pointing Mode

3.4.6.8.2 33Sun-Pointing Safe-Hold Mode

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

3.4.6.10.1 34Rigid-Body Stability Margins

3.4.6.10.2 34Time Delay

3.4.6.10.3 34Flexible-Body Modes

3.4.6.10.4 35Reaction Wheel Momentum Margin

3.4.6.11 35Propulsion System

Contents iv

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 35Unintended Propellant Vapor Ignition

3.4.6.11.5 35Propellant Leakage Dual Fault Tolerant

3.4.6.11.6 35Thruster and Venting Impingement

3.4.7 36Flight Software

3.4.7.1 36Language and Methodology

3.4.7.2 36Flight Software Upload

3.4.7.3 36Flexibility and Ease of Software Modification

3.4.7.4 36Version Identifiers

3.4.7.5 36Flight Processor Resource Sizing

3.4.7.6 37Software Event Logging

3.4.7.7 37Soft Reset

3.4.7.8 37Memory Integrity

3.4.7.9 38Memory Dump

3.4.7.10 38Telemetry

3.4.7.11 38Stored Command Processing

3.4.7.11.1 38Command Execution

3.4.7.11.2 38Stored Command Buffer Management

3.4.7.12 38OBC Bootstrap Software

3.5 39Spacecraft Ground Support Equipment (GSE)

3.5.1 39GSE General Requirements

3.5.2 39Electrical Ground Support Equipment (EGSE)

3.5.2.1 39EGSE Command and Data Handling

3.5.2.2 39EGSE Command Interface to Mission Management Center

3.5.2.3 39EGSE Telemetry Interface to Mission Management Center

3.5.2.3.1 39EGSE Telemetry to Instrument EGSE

3.5.2.3.2 40EGSE Failure

3.6 40Flight Software Development Environment

3.7 40Spacecraft Hardware Simulator (SHS)

3.7.1 40SHS General Requirements

3.7.2 41Unique SHS Instantiations

3.7.3 41SHS Control

3.7.4 42SHS Status

3.7.5 42SHS Data Logging

3.7.6 42SHS Modeling Requirements

3.7.7 43SHS Performance

Contents v

3.7.8 43SHS Interface

3.8 43Satellite Software Simulator (S3)

3.8.1 43S3 General Requirements

3.8.2 44S3 Control

3.8.3 45S3 Status

3.8.4 45S3 Data Logging

3.8.5 45S3 Modeling Requirements

3.8.6 46S3 Performance

3.8.7 46S3 Interface

3.9 46Spacecraft to Instrument Interface Simulator

4 47Design Verification Requirements

4.1 47Electrical Functional and Performance Verification Requirements

4.1.1 47Electrical Interface Tests

4.1.2 47Comprehensive Performance Tests

4.1.3 47Limited Performance Tests

4.1.4 48Performance Operating Time and Trouble-Free Performance

4.2 48Structural and Mechanical Verification Requirements

4.2.1 48Mechanical 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)

4.4.4 55Heater 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)

Contents vi

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

4.4.8 57Unit/Component Level Thermal-Vacuum Testing

4.4.9 57Satellite System Level Thermal-Vacuum Testing

4.4.10 58Satellite System Level Thermal-Balance (TB)

4.4.10.1 58Thermal Control Objectives

4.4.10.2 58Thermal 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 60Battery Life Test

4.5.1.1 60Mission Life Charge/Discharge Cycles

4.5.1.2 61Battery Accelerated Geostationary-Earth-Orbit (GEO) Life Testing

4.6 61Solar Array Verification Requirements

4.6.1 61Solar Array Component Qualification

4.6.2 61Solar Array Panel Qualification

4.6.3 62Flight Solar Array Panel Testing

4.6.4 63Solar Array Spacecraft Level Tests

4.7 63Test Condition Tolerances

5 65Appendix A: Battery System Definitions

6 67Appendix 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 Platform Relay (DCPR)

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

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

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and estimated attitude.

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.

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 occurrence. 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.

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 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

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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 - Current Best 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 critical events include orbit raising, station keeping and appendage deployment.

Mission Allowable Temperature: Mission Allowable Temperature (MAT) is defined as the specific temperature range where a unit can operate and meet performance requirements over the mission lifetime. The allowable operational and non-operational ranges are based on design limits of the unit and should be as robust as possible. MAT ranges may be adjusted based upon flight predictions as thermal modelling matures.

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 Temperatures: Non-operational Temperatures (NOT) are the established range of temperatures that components are permitted to experience while dormant, not operating and not powered. NOT temperatures represent the permissible range while the hardware is off. During flight, survival heaters maintain hardware at or above the cold NOT limit and passive design maintains hardware below the upper NOT limit. It is also known as non-operating MAT.

Normal Operation: Operational states of the Spacecraft that exist or occur by design, according to the expectation of the mission designers and planners.

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

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

Random Errors: Unrepeatable statistical fluctuations in the measured data due to the

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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 DCPR auxiliary communication.

Spacecraft Bus: Satellite without instruments and DCPR 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.

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

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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.

Examples are electronics unit and 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.

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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. SS to Ground Located Command, Control, and Communications (SS-C3S)

Interface Requirements Document (IRD), 418-XO-IRD-0043

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

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

10. 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

11. 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

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

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

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

15. 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

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

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

18. AIAA S-112A-2013, Qualification and Quality Requirements for Electrical

Components on Space Solar Panels

19. Federal Information Processing Standards Publication 140-3, Security

Requirements for Cryptographic Modules, March 22, 2019.

20. GSFC-STD-7000B, General Environmental Verification Standard (GEVS) For GSFC

Flight Programs and Projects, April 28, 2021

21. Standard for the Design and Fabrication of Ground Support Equipment, NASA-

STD-5005D, Change 1, 14 June 2013

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

23. NPR 8715.6B, NASA Procedural Requirements for Limiting Orbital Debris and

Evaluating the Meteoroid and Orbital Debris Environments, February 16, 2017, NASA-STD-8719.14B, Process for Limiting Orbital Debris

Page 9 of 68 Printed Tuesday, February 22, 2022

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)

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

25. NIST Special Publication 800-53, Rev. 5, Security and Privacy Controls for

Information Systems and Organizations, September 2020

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 68 Printed Tuesday, February 22, 2022

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 Platform Relay (DCPR)

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 68 Printed Tuesday, February 22, 2022

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 DCPR 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.

3.2.3 Compatibility Requirements

The Spacecraft shall simultaneously and continuously operate Spacecraft subsystems, DCPR, and instrument payloads in all their operational modes through end of life, including eclipse periods.

The Spacecraft and its units shall not generate electromagnetic interference that could adversely affect its own performance or the performance of instruments and DCPR on the Spacecraft, launch vehicle, or launch site.

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 68 Printed Tuesday, February 22, 2022

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 68 Printed Tuesday, February 22, 2022

ID

SCFPS72

SCFPS73

SCFPS74

SCFPS75

SCFPS76

SCFPS77

SCFPS78

SCFPS79

SCFPS80

SCFPS81

SCFPS82

SCFPS83

SCFPS84

SCFPS85

SCFPS86

SCFPS87

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

3.2.13

3.2.13.0-1

3.2.13.0-2

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 insturments' hardware, during all pre-launch phases.

3.2.13 Pre-Launch Satellite Testing

Spacecraft and instrument aliveness tests shall be performed while on the launch pad to demonstrate Satellite health and safety.

The Spacecraft shall perform end-to-end Satellite Command and Data Handling (C&DH) and communication testing while at the launch site facility.

3.2.14 Spacecraft Telemetry During Launch and Orbit Raising

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 68 Printed Tuesday, February 22, 2022

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 MEV for

Page 15 of 68 Printed Tuesday, February 22, 2022

ID

SCFPS97

SCFPS98

SCFPS99

SCFPS100

SCFPS101

SCFPS102

SCFPS103

SCFPS104

SCFPS105

Object Number

3.2.16.0-1

3.2.17

3.2.17.0-1

3.2.18

3.2.18.0-1

3.2.19

3.2.19.0-1

3.3

3.3.1

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

(F&PS)

Payload as found in the Project PRAD [9].

2. At launch a 5% predicted power margin for mission critical orbit raising and safing operation modes as well as to accommodate in-flight operational uncertainties

3. The 3-sigma variations are due to all the following:

a) Worst-cast Satellite mass properties

b) 3-sigma low launch vehicle performance

c) 3-sigma low propulsion subsystem performance (thruster performance/alignment, propellant residuals)

d) 3-sigma flight dynamics errors and constraints

3.2.17 Electrostatic Arc-Discharge Susceptibility

The Spacecraft shall withstand both a radiated and direct arc as shown in Table SCFPS99 without sustaining permanent damage, verified by test on the first Protoflight Spacecraft with follow-on Spacecraft qualified by similarity. Note: Test verification at the unit/component level of assembly is acceptable.

Table SCFPS99 ESD Characteristics Table

ESD Characteristics Table

Item Description Characteristics

1 Discharge Voltage 10 kv

2 Discharge Energy 3 millijoules, maximum

3 Peak Current 1 amp

4 Time Constant 600 nsec

5 Repetition Rate 1 sec

6 Quantity of Discharges per Surface 30

7 Distance of Radiated Discharge from Instrument Surface

30 cm

3.2.18 External Surface-to-Surface Direct Discharge

The Spacecraft shall not be impaired by differential charging between its external surfaces.

3.2.19 Deep Dielectric Charging

The Spacecraft shall withstand all direct discharges caused by deep dielectric charging (Internal Electrostatic Discharge, IESD).

3.3 Interfaces

3.3.1 Instrument Interfaces

Page 16 of 68 Printed Tuesday, February 22, 2022

ID

SCFPS106

SCFPS107

SCFPS108

SCFPS109

SCFPS110

SCFPS111

SCFPS112

SCFPS801

SCFPS113

SCFPS114

Object Number

3.3.1.0-1

3.3.1.0-2

3.3.1.0-3

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

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

(F&PS)

The Spacecraft shall meet all instrument interface requirements defined in the GIRD [1], and UIIDs [2][3][4][5][6].

Note: PRAD [9] takes precedence over GIRD and UIIDs for Spacecraft interface requirements.

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.

The Spacecraft shall provide cross-strapping for all interface signals between the Spacecraft and the GFE instruments.

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.

Page 17 of 68 Printed Tuesday, February 22, 2022

ID

SCFPS115

SCFPS116

SCFPS117

SCFPS118

SCFPS119

SCFPS120

SCFPS121

SCFPS122

SCFPS123

SCFPS124

SCFPS125

SCFPS126

Object Number

3.4.1.3

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

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

(F&PS)

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.

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.

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 or equal to the battery rated capacity (Cbr) when charged at rate Cbn/10 to maximum end-of-charge voltages of 4.1 Vdc per cell and discharged at Cbn/2until the first battery cell end-of-discharge voltage reaches 3.0 Vdc.

The Battery Operating Temperature at the cell level shall be 20°C ± 10°C.

The battery cells shall meet all performance requirements after exposure to the survival temperatures of 40°C for the hot case and 0°C for the cold case.

Page 18 of 68 Printed Tuesday, February 22, 2022

ID

SCFPS127

SCFPS128

SCFPS129

SCFPS130

SCFPS131

SCFPS132

SCFPS133

SCFPS134

SCFPS135

SCFPS136

SCFPS137

SCFPS138

Object Number

3.4.1.5.1.0-4

3.4.1.5.1.0-5

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

3.4.1.5.1.0-12

3.4.1.5.1.0-13

3.4.1.5.1.0-14

3.4.1.5.1.0-15

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

(F&PS)

The measured 20°C beginning-of-life (BOL) battery cell actual capacity (Ca) shall be 120% of the battery cell nameplate capacity (Cn) when charged at a rate Cn/10 to a maximum end-of-charge voltage of 4.1 Vdc and discharged at a rate Cn/2 until the end-of-discharge voltage reaches 3.0 Vdc.

The measured 20 °C beginning-of-life (BOL) battery capacity (Cba) shall be 120% of the battery nameplate capacity (Cbn) when charged at a rate Cbn/10 to maximum end-of-charge voltages of 4.1 Vdc per cell and discharged at a rate Cbn/2 until the first battery cell end-of-discharge voltage reaches 3.0 Vdc.

The 20 °C beginning-of-life (BOL) battery system actual capacity (Cbsysa) shall be 120% of the battery system nameplate capacity (Cbsysn) when charged at a rate Cbsysn/10 to maximum end-of-charge voltages of 4.1 Vdc per cell and discharged at a rate Cbsysn /2 until the first battery cell end-of-discharge voltage reaches 3.0 Vdc.

Maximum intercell temperature gradients of the battery shall not exceed 3˚ C.

Ground storage temperature of the cell/battery shall be 0°± 5˚C.

Transportation temperature of the cell/battery shall be 0°± 5˚C. Note: This does not apply to transportation of the battery once it is integrated into the Spacecraft.

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

The…