TSIS2-SC-REQ-0005_Rev-B_Draft-1_Observatory Requirements Document_Redline (1).pdf

Effective Date: To Be Entered by CMO

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

TSIS2-SC-REQ-0005, Revision AB

Total and Spectral solar Irradiance Sensor (TSIS)-2, Code 424

TSIS-2 Observatory Requirements Document

DRAFT 1: January 23, 2020

DRAFT

TSIS-2 Observatory Requirements Document TSIS2-SC-REQ-0005 Revision AB

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

TSIS-2 Observatory Requirements Document Signature/Approval Page

Prepared by:

Sergey KrimchanskyMichael Rhee Date Deputy Project Manager, TSIS-2 Mission Systems Engineer NASA Goddard Space Flight Center

Reviewed by:

Susan Breon Date Project Manager, TSIS-2 NASA Goddard Space Flight Center

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

Preface

This document is under the TSIS-2 project configuration control. Once this document is approved, TSIS-2-approved changes are handled in accordance with Class I and Class II change control requirements as described in the TSIS-2 Configuration Management Procedures, and changes to this document shall be made by complete revision.

Any questions should be addressed to:

TSIS-2 Configuration Management Office

NASA/ GSFC

Code 424 Greenbelt, MD 20771

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

Change History Log

Revision Effective Date Description of Changes

(Reference the CCR & CCB/ERB Approve Date) Rev - 08/06/2019 TSIS2-CCR-0013, approved 8/6/2019 Rev A 11/26/2019 TSIS2-CCR-0028, approved 11/26/2019 Rev B Entered by CMO

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

Table of TBDs/TBRs/TBSs

Action Item No.

Location Summary

1 3.10.1 Availability of operating in Science Mode 2 4.4.11 Spacecraft battery charge 3 4.5.1.3 The Spacecraft shall be capable of propagating the on-board state for at least 1 week 4 4.5.1.4 The Spacecraft shall support science operations for up to 1 week from an on- board propagated ephemeris 5 4.5.2.1.1 As a primary response, the spacecraft shall off-point when the spacecraft velocity vector is within 15 degrees of the TIM boresight

6 4.5.2.1.2 the spacecraft shall send a safing command to TIM (to close its shutters) when the spacecraft velocity vector is within degrees of the TIM boresight.

7 4.5.2.1.3 As a secondary response, the spacecraft shall send a safing command to SIM (to close its shutters) when the spacecraft velocity vector is within 5 degrees of the SIM boresight.

8 4.5.2.2.3 The spacecraft Cruciform Calibration maneuver scans shall consist of 200 second dwells at each step in the maneuver.

9 4.5.2.2.4 The spacecraft Cruciform Calibration maneuver shall consist of four scans in orthogonal axes across the face of the sun with 200 second dwells at sun center prior to each scan

10 4.5.2.3.3 The spacecraft FOV map shall consist of 400 second dwells at each step

11 4.5.2.3.4 The spacecraft FOV map shall consist of 400 second dwells at sun center at the beginning and end of each solar orbit

12 4.6.2.11 The Spacecraft shall be capable of storing at least 2 weeks of Observatory housekeeping data at the maximum onboard telemetry rate

13 4.6.2.12 The Spacecraft shall be capable of storing at least 2 weeks of TSIS-2 Observatory data

14 6.3.8.1 The TSIS-2 Spacecraft shall be designed to meet all performance requirements after exposure to a maximum depressurization rate of 4.1 x 103 N/m2/s (0.6 psi/sec) during launch and ascent.

vi Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

Table of Contents

1 Introduction

1.1 Scope

1.2 Terminology

1.3 Mission Overview

Mission Elements 2 Applicable and Reference Documents

2.1 Applicable Documents

3 Mission System Requirements

3.1 Payload Accommodations

The spacecraft shall be integrated with the Government Furnished Equipment

(GFE) TSIS-2 payload

3.1.2 The TIM and SIM shall be mounted to the same mounting plate on the Spacecraft, mounting

3.2 On-Orbit Coordinate System and System of Units

A right hand, orthogonal, body-fixed on-orbit coordinate system shall be used for the spacecraft coordinate system

The Spacecraft and Observatory interface control documentation with the payload, launch vehicle (LV) and ground system shall be produced using the International System of Units (SI) as the primary system of units, unless if previously designed hardware is in English units

The Spacecraft software and firmware shall be produced using the International System of Units (SI) as the primary system of units, unless if previously designed software or firmware is in English units

3.3 Mission Lifetime

3.4 Ground Storage

The Observatory shall be capable of being placed in a ground storage state that requires minimal intervention by personnel for up to 30-day periods

The Observatory shall meet all mission requirements, including design lifetime at the completion of a Ground Storage Phase lasting up to 24 months

3.5 Launch Vehicle (LV)

The Spacecraft shall be designed for a 15 inch diameter interface to the Launch

Vehicle separation system, the LV will provide an adaptor if needed The Spacecraft and its Ground Support Equipment used at the launch site shall comply with the NASA-STD 8719.24 NASA ELV Payload Safety Requirements.

3.5.3 The Spacecraft shall be designed to perform aliveness tests of the spacecraft and integrated payload while mated to the launch vehicle

3.5.4 The Observatory shall be designed to perform battery reconditioning while mated to the launch vehicle and encapsulated

3.6 Mission Phases

Launch Injection Orbit and Early Orbit Commissioning vii Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

Operational Decommissioning

3.7 Spacecraft Modes

The Spacecraft shall autonomously transition from Launch Mode to Safe Hold

Mode upon separation from the launch vehicle The On-Orbit Spacecraft shall not be capable of being commanded into Off Mode, Ground Mode, or Launch Mode (minimum 3 fault tolerant) The On-Orbit Spacecraft shall be commandable into Safe Hold Mode and Science

Mode The Spacecraft shall autonomously transition from Science Mode to the Safe Hold

Mode upon detection of faults that threaten the health and safety of the spacecraft and/or Payload

The Spacecraft shall transition out of the Safe Hold Mode only via ground command

3.8 Redundancy

The spacecraft shall provide available housekeeping telemetry for both primary and redundant sides at all times (if redundancy exists)

3.9 Autonomy

The Observatory shall be designed to operate safely for an indefinite period of time without ground intervention while in Safe Hold Mode

The Observatory shall be capable of autonomous Science Mode operations for a minimum of 1 week

The Observatory shall be capable of autonomously switching to the backup component of redundant systems while in Safe Hold Mode (if redundancy exists).

The Observatory shall not autonomously switch to a known failed redundant component

The Observatory shall monitor the telemetry of pre-defined components, instruments, and subsystems for limit violations and autonomously safe the spacecraft and instruments when necessary

The Observatory shall be capable of disabling autonomous functions via ground command

The Observatory shall report autonomous reconfigurations in housekeeping telemetry

The Observatory shall report detection of out of limit conditions for on-board monitored systems in Observatory housekeeping telemetry

The Spacecraft subsystems that perform self diagnostics shall report the results in Observatory housekeeping telemetry

The Spacecraft subsystems that support self diagnostics shall accept ground commands to run diagnostics and report the results in Observatory housekeeping telemetry

3.10 Availability

The Spacecraft shall be capable of operating in Science mode (pointing to the sun)

> 98% (TBR) of the time, after commissioning is complete, as measured on a monthly basis viii Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

The spacecraft shall have the ability to recover from Single Event Upsets (SEU).

3.11 Resource Margins

4 Spacecraft Subsystem Requirements

4.1 Structure

The Spacecraft shall have sufficient first fundamental mode separation from the

Payload The Spacecraft structure shall be sufficiently stiff to meet the axial and lateral frequency requirements as defined in the LV User’s Guide or LV ICD The Spacecraft in the stowed/launch configuration shall meet the dynamic envelope for the launch vehicle fairing as specified in the LV User’s Guide or LV

ICD

The Spacecraft shall accommodate a total payload mass per the IARD or SC to Payload ICD

The Spacecraft shall be designed to permit full range deployment of mechanisms after full integration during ground processing

The observatory mass shall not exceed 139 kg, to include the mass of the separation system between the Spacecraft and Launch Vehicle

4.2 Thermal

The spacecraft shall monitor and report in housekeeping telemetry the temperatures of subsystems on the Spacecraft The spacecraft shall maintain subsystems within their design survival temperature range when the Spacecraft is in Safe Hold Mode The Spacecraft shall maintain subsystems within their operational design temperature range when the Spacecraft is in Science Mode The Spacecraft shall maintain subsystems within their operational design temperature range when the Spacecraft is in Launch Mode The Spacecraft shall maintain payload temperatures as specified in Table 13

4.3 Flight Software

General Event Logging Initialization Stored Commands

4.4 Power

The Spacecraft shall provide a single point ground (SPG) as part of the power subsystem. The contractor may propose an alternate grounding scheme for an existing, proven design

All of the primary power returns shall be tied together at the SPG The Spacecraft shall provide for power generation, power control, energy storage, and power distribution The Spacecraft shall provide load shedding control and battery state of charge

(SOC) management The Spacecraft shall be power positive over each orbit while in any on-orbit mode including Safe Hold and Science The Spacecraft shall provide short circuit protection or current limitation for each ix Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

switchable power circuit The Spacecraft shall provide configurable protection against over voltage, under voltage and over current for the Spacecraft and Payload The Spacecraft shall be capable of being powered from Spacecraft Test

Equipment with or without flight or ground batteries installed The Spacecraft shall monitor and report in housekeeping telemetry, the switch position of each switchable power circuit at the Spacecraft distribution point. 1211 The Spacecraft shall monitor and report in housekeeping telemetry, current and voltage for each power circuit at the Spacecraft power distribution point The Spacecraft battery shall be sized such that the state of charge remains above

60% (TBR) during normal operations and above 40% (TBR) at all times for the Spacecraft design life

The Spacecraft power system shall be capable of electrically monitoring and reporting all battery cells and bypassing failed battery cell(s)

The spacecraft bus shall accommodate an orbit average power consumption per IARD or ICD

The spacecraft bus shall accommodate a peak power consumption per IARD or

ICD

Each instrument power feed shall be fused

4.5 Attitude and Orbit Determination and Control

Ephemeris Science Modes and Calibration Maneuvers Pointing

4.6 Command and Data Handling

General Telemetry Command Mass Storage

4.7 Radio Frequency Telecommunications System

The Spacecraft shall comply with National Telecommunications and Information

Administration (NTIA) Spectrum Standards per “Manual of Regulations and Procedures for Federal Radio Frequency Management (Redbook)”, Chapter 5 Spectrum Standards and Chapter 8 Procedures and Principles for the Assignment & Coordination of Frequencies

The Spacecraft shall comply with International Telecommunication Union (ITU) spectrum utilization and sharing requirements per “ITU-R Radio Regulations”, The Spacecraft shall comply with Recommendation SFCG 21-2R4 The Spacecraft shall be capable of transmitting a pseudo random noise pattern from all spacecraft-provided transmitters The Spacecraft RF active components shall operate at nominal transmission power in ambient conditions on the ground The Spacecraft shall use S-band command uplink The Spacecraft shall maintain power to the S-band receiver from launch countdown and through the life of the mission x Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

The Spacecraft shall be capable of transmission and reception of ground commands while in any powered mode

The Spacecraft shall be capable of S-band transmissions in any Spacecraft orientation such that the link can be closed to line of sight ground station, having an elevation angle above 5 degrees, with a 3dB margin

4.8 Contamination and Outgassing

The spacecraft bus materials shall have a total mass loss (TML) less than 1.00% and a collected volatile condensable mass (CVCM) less than 0.10%. Material vacuum outgassing will be determined in accordance with ASTM E-595

The spacecraft shall not use: pure tin, cadmium, or zinc plating The Observatory shall maintain a surface cleanliness Level of 400 A/2 throughout

I&T in compliance with the Product Cleanliness Levels and Contamination Control Program, IEST-STD-CC1246

The Observatory shall be integrated, tested and delivered to the Launch Site in compliance with the Product Cleanliness Levels and Contamination Control Program, IEST-STD-CC1246, Level 450 A/1.5

The Observatory shall meet a surface cleanliness Level of 500A at Beginning of Life in compliance with the Product Cleanliness Levels and Contamination Control Program, IEST-STD-CC1246

To the maximum extent practical the spacecraft bus shall vent outgassed species in the direction that is opposite the payload boresights

4.9 Radiation Space Environment

The Spacecraft shall be compatible with the mission radiation environment as shown in Appendix B The Spacecraft shall be analyzed and designed to address surface charging on the spacecraft for the given flight altitude (see NASA HDBK-4002A for guidelines).

5 Ground Support Equipment

5.1 Electrical Ground Support Equipment (EGSE)

Spacecraft contractor shall provide all EGSE required to support Spacecraft design, integration and test (including TVAC), and launch campaign

The EGSE, specifically the telemetry and command system shall interface with the TSIS-2 OASIS EGSE to allow for bent pipe instrument commanding in local or remote (via internet) mode

EGSE shall interface with the MOC for End to End (ETE) testing

5.2 Mechanical Ground Support Equipment (MGSE)

The Spacecraft MGSE shall support all environmental testing, including TVAC, vibration, and EMI with the Observatory in its corresponding mission configuration

The Spacecraft MGSE shall enable Spacecraft mechanism deployment in a 1-g negated environment

The spacecraft shall provide access to test connectors and alignment cubes after integration with the spacecraft

The Observatory Shipping and Storage Container shall be airtight (up to 10kPa under/over pressure relative to internal pressure) with a filtered pressure control xi Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

system, when the Spacecraft/Observatory is within the shipping container during transportation

The Observatory Shipping & Storage Container shall incorporate means of measuring and recording shock, temperature and humidity within the container during transport

5.3 Spacecraft Instrument Interface Simulator (SIIS)

6 Environmental TEST requirements

6.1 Electromagnetic Interference (EMI)/Electromagnetic Compatibility (EMC)

6.1.1 EMI/EMC General

6.1.2 Conducted Emissions (CE)

6.1.3 Radiated Susceptibility (RS)

6.1.4 Radiated Emissions (RE)

6.2 Magnetic Requirements

6.3 Mechanical Requirements

6.3.1 Factors of Safety for Design and Analysis

6.3.2 Quasi-Static Design Load Limit Factors

6.3.3 Random Vibration Environment

6.3.4 Sine Vibration Environment

6.3.5 External Shock Environment

6.3.6 Acoustics Environment

6.3.7 Mechanisms

6.3.8 Pressure

6.4 Thermal Requirements

6.4.1 Spacecraft

6.4.2 Thermal Cycles

6.4.3 Thermal Survival

6.4.4 Thermal Operational Verification

Appendix A. Abbreviation/Acronym List Appendix B. Mission Radiation Environment xii Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

List of Tables Table 1. Applicable Documents Table 2. Radiated Electromagnetic Field, Launch Configuration Table 3. Radiated Electromagnetic Field, Operational Configuration Table 4. Radiated Emissions Notches Table 5. Design/Analysis Factors of Safety Applied to Limit Loads2 Table 6. Quasi-Static Design Limit Loads Table 7. Sine Burst1 Test Levels Table 8. Random Vibration Environment for the TSIS-2 Spacecraft (All Axes) Table 9. Random Vibration Test Durations Table 10. Interface Sine Sweep Vibration Levels (All Axes) Table 11. Shock Environment at Spacecraft Interface to the LV (Q=10) Table 12. Acoustic Environment Table 13. Temperature Range Table B-1. Mission Radiation Environment Table B-2. SPENVIS Settings

List of Figures Figure 1. TSIS-2 Mission Elements Figure 2. Radiated Emissions Figure B-1. Total Dose vs. Shielding at 600 km Polar Orbit

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

1 INTRODUCTION

1.1 Scope

This specification establishes the top-level requirements for the TSIS-2 Observatory, which consists of a mission-unique spacecraft and the TSIS-2 payload. The TSIS-2 payload consists of two instruments, the Total Irradiance Monitor (TIM) and the Spectral Irradiance Monitor (SIM).

The TSIS-2 payload is Government Furnished Equipment (GFE) to the spacecraft developer.

Top-level functional, performance and interface requirements for the spacecraft are defined herein. The contractor is responsible for a hierarchy of lower-tier specifications that derive and allocate requirements to subsystems and components.

Instrument accommodation requirements contained in the Instrument Accommodations Requirements Documents (IARD) or Interface Control Documents (ICD) shall be considered for the spacecraft design.

1.2 Terminology

The term "TBD" (to be determined) means that the contractor should determine the missing requirement in coordination with the government. The term "TBS" (To Be Supplied) means that the government will clarify or supply the missing information during the course of the delivery order. The term "TBR" (To Be Resolved) means that the requirement may be reviewed for appropriateness by the contractor or the government and may be changed by the government during the course of the delivery order.

1.3 Mission Overview

See the TSIS-2 Observatory Statement of Work (TSIS2-SC-SOW-0002) for the Mission Overview.

Mission Elements

The architecture mission elements of the program, including Space Segment and Ground Segment, will be organized as shown in Figure 1.

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

Figure 1. TSIS-2 Mission Elements

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

2 APPLICABLE AND REFERENCE DOCUMENTS

2.1 Applicable Documents

The following documents are directly applicable, or contain policies or other directive matters that are binding upon the content of this document to the extent specified herein.

Table 1. Applicable Documents Document Number Document Title TSIS2-SC-TN-0003 TSIS-2 Spacecraft Instrument Accommodations Requirements

Document (TSIS-2 IARD) TSIS2-SC-REQ-0006 TSIS-2 Spacecraft Mission Assurance Requirements (MAR) TSIS2-SC-SOW-0002 TSIS-2 Spacecraft Statement of Work (SOW) TSIS2-SC-CDRL-0002 TSIS-2 Spacecraft Contract Data Requirements List

(CDRL)

Rocket-Lab-Payload-Users- Guide-6.42 https://www.rocketlabusa.com/assets/Uploads/Rocket-Lab- Payload-Users-Guide- 6.4.pdfhttps://www.rocketlabusa.com/assets/Uploads/Rock et-Lab-Payload-Users-Guide-6.2.pdf

Virgin-Orbit-Service- Guide_Sept2019 https://virginorbit.com/wp-content/uploads/2019/09/ServiceGuide_Sept2019.pdfhttps:/ /virginorbit.com/wp-content/uploads/2018/11/Virgin-Orbit- Service-Guide.pdf moog-espa-users-guide-datasheetSAIS_ESPA_Ring_Rev _0712 https://www.moog.com/content/dam/moog/literature/Space _Defense/spaceliterature/structures/moog-espa-users-guide-datasheet.pdfhttps://www.csaengineering.com/content/dam/ moog/literature/Space_Defense/spaceliterature/structures/S AIS_ESPA_Ring_Rev_0712.pdf

453-UG-002905 NEN User Guide 453-UG-002905 ASTM E595 - 07 Standard Test Method for Total Mass Loss and Collected Volatile

Condensable Materials from Outgassing in a Vacuum Environment

CCSDS 232.1-B-21 Recommendation for Space Data Systems Standards.

Communications Operations Procedure-1. Blue Book. Issue 21.

April 2019.

CCSDS 701.0-B-3 Advanced Orbiting Systems, Networks and Data Links:

Architectural Specification

IEST-STD-CC1246 Product Cleanliness Levels and Contamination Control Program

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

NPR-8715.6 NASA Procedural Requirements for Limiting Orbital Debris NTIA Spectrum Standards Manual of Regulations and Procedures for Federal Radio Frequency

Management (Redbook) SFCG 21-2R4 Efficient Spectrum Utilization for Space Research Service and Earth

Exploration-Satellite Service on Space-To-Earth Links GSFC-STD-1000 RevG GSFC Rules for the Design, Development, Verification, and Operation of Flight Systems NASA HDBK-4002A Mitigating In-Space Charging Effects - A Guideline

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

FIPS 140 Federal Information Processing Standards (FIPS) Publication 140 IS-GPS-200 GPS Directorate Systems Engineering and Integration Interface

Specification

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

3 MISSION SYSTEM REQUIREMENTS

3.1 Payload Accommodations

The spacecraft shall be integrated with the Government Furnished Equipment (GFE) TSIS-2 payload.

3.1.2 The TIM and SIM shall be mounted to the same mounting plate on the Spacecraft, mounting to opposite sides of the same plate is acceptable.

3.2 On-Orbit Coordinate System and System of Units

A right hand, orthogonal, body-fixed on-orbit coordinate system shall be used for the spacecraft coordinate system.

The Spacecraft and Observatory interface control documentation with the payload, launch vehicle (LV) and ground system shall be produced using the International System of Units (SI) as the primary system of units, unless if previously designed hardware is in English units.

The Spacecraft software and firmware shall be produced using the International System of Units (SI) as the primary system of units, unless if previously designed software or firmware is in English units.

3.3 Mission Lifetime

The spacecraft bus shall be designed to support the science mission for a minimum of 3 years after commissioning.

3.4 Ground Storage

The Observatory shall be capable of being placed in a ground storage state that requires minimal intervention by personnel for up to 30-day periods.

The Observatory shall meet all mission requirements, including design lifetime at the completion of a Ground Storage Phase lasting up to 24 months.

3.5 Launch Vehicle (LV)

The Spacecraft shall be designed for a 15 inch diameter interface to the Launch Vehicle separation system, the LV will provide an adaptor if needed.

The Spacecraft and its Ground Support Equipment used at the launch site shall comply with the NASA-STD 8719.24 NASA ELV Payload Safety Requirements.

3.5.3 The Spacecraft shall be designed to perform aliveness tests of the spacecraft and integrated payload while mated to the launch vehicle.

3.5.4 The Observatory shall be designed to perform battery reconditioning while mated to the launch vehicle and encapsulated.

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

3.6 Mission Phases

Launch

3.6.1.1 The Observatory shall be capable of recycling from a launch abort and be launch-ready within 24 hours of abort.

3.6.1.2 The Observatory and Ground Support Equipment (GSE) shall be designed to interface through launch vehicle harnessing to perform testing and monitoring of the spacecraft and integrated payload (instruments) while mated to the launch vehicle.

Injection Orbit and Early Orbit

3.6.2.1 The Observatory shall be designed for an altitude of 600 +/-15 km.

3.6.2.2 The Observatory shall be designed for a circular, sun-synchronous, 98 degree inclination orbit with a 16:00 ± :15 LTAN.

3.6.2.3 The Observatory shall autonomously execute all actions necessary to enter a power positive safe state after separation from the launch vehicle.

3.6.2.4 The Observatory shall be capable of autonomously activating its telemetry transmitter immediately or some predetermined delay after Launch Vehicle fairing separation.

Commissioning

3.6.3.1 The SC bus shall be capable of completing its commissioning within 15 days of launch.

3.6.3.2 The Observatory shall support payload commission for 45 days following SC bus commissioning.

Operational

The Observatory shall be commanded by the MOC.

Decommissioning

The Observatory shall be decommissioned in accordance with NASA Procedural Requirements for Limiting Orbital Debris, NPR 8715.6 and NASA Process for Limiting Orbital Debris, NASA-STD 8719.14A.

3.7 Spacecraft Modes

This document uses the following definition of Spacecraft modes to maintain consistency between the mission system architecture and other external interfaces with the Spacecraft.

Though the Spacecraft contractor is not required to use the same definitions; the contractor shall provide a mapping of the contractor’s Spacecraft modes to these for requirement verification.

OFF Mode Spacecraft is off.

Ground Mode provides a spacecraft that allows integration and testing (I&T) on the ground to verify and troubleshoot all modes of the observatory in orbit.

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

Launch Mode provides a launch-ready observatory integrated with the launch vehicle, and includes launch, ascent, and vehicle separation. The instruments’ vacuum doors are closed through launch and early spacecraft commissioning to allow the observatory to outgas; once opened on-orbit the vacuum doors remain open for the remainder of the mission.

Safe Hold Mode provides a maximum power positive attitude, and an observatory configuration that maintains the Spacecraft in a power positive, thermally safe, and instrument safe configuration requiring no ground contact for an indefinite period of time. The instruments’ doors shutters shall be closed.

Science Mode provides fine solar pointing for TSIS-2 data acquisition or special pointing for TSIS-2 calibrations. During the time the instruments are powered on and collecting science data, the instruments’ doors will remain open.shutters are cycled open/closed by the instrument electronics.

The Spacecraft shall autonomously transition from Launch Mode to Safe Hold Mode upon separation from the launch vehicle.

The On-Orbit Spacecraft shall not be capable of being commanded into Off Mode, Ground Mode, or Launch Mode (minimum 3 fault tolerant).

The On-Orbit Spacecraft shall be commandable into Safe Hold Mode and Science Mode.

The Spacecraft shall autonomously transition from Science Mode to the Safe Hold Mode upon detection of faults that threaten the health and safety of the spacecraft and/or Payload.

The Spacecraft shall transition out of the Safe Hold Mode only via ground command.

3.8 Redundancy

The spacecraft shall provide available housekeeping telemetry for both primary and redundant sides at all times (if redundancy exists).

3.9 Autonomy

The Observatory shall be designed to operate safely for an indefinite period of time without ground intervention while in Safe Hold Mode.

The Observatory shall be capable of autonomous Science Mode operations for a minimum of 1 week.

The Observatory shall be capable of autonomously switching to the backup component of redundant systems while in Safe Hold Mode (if redundancy exists).

The Observatory shall not autonomously switch to a known failed redundant component.

The Observatory shall monitor the telemetry of pre-defined components, instruments, and subsystems for limit violations and autonomously safe the spacecraft and instruments when necessary.

The Observatory shall be capable of disabling autonomous functions via ground command.

The Observatory shall report autonomous reconfigurations in housekeeping telemetry.

Check the TSIS-2 TDMS Server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

The Observatory shall report detection of out of limit conditions for on-board monitored systems in Observatory housekeeping telemetry.

The Spacecraft subsystems that perform self diagnostics shall report the results in Observatory housekeeping telemetry.

The Spacecraft subsystems that support self diagnostics shall accept ground commands to run diagnostics and report the results in Observatory housekeeping telemetry.

3.10 Availability

The Spacecraft shall be capable of operating in Science mode (pointing to the sun) > 98% (TBR) of the time, after commissioning is complete, as measured on a monthly basis.

The spacecraft shall have the ability to recover from Single Event Upsets (SEU).

3.11 Resource Margins

Resource margins shall be maintained in accordance with GSFC-STD-1000 Rev G Gold Rules

1.06 and 3.07. The contractor may propose to meet only the final margin requirement for an existing, proven design.

Check the TSIS-2 Technical Data Management System server at https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use.

4 SPACECRAFT SUBSYSTEM REQUIREMENTS

4.1 Structure

The Spacecraft shall have sufficient first fundamental mode separation from the Payload.

The Spacecraft structure shall be sufficiently stiff to meet the axial and lateral frequency requirements as defined in the LV User’s Guide or LV ICD.

The Spacecraft in the stowed/launch configuration shall meet the dynamic envelope for the launch vehicle fairing as specified in the LV User’s Guide or LV ICD.

The Spacecraft shall accommodate a total payload mass per the IARD or SC to Payload

ICD.

The Spacecraft shall be designed to permit full range deployment of mechanisms after full integration during ground processing.

The observatory mass shall not exceed 139 kg, to include the mass of the separation system between the Spacecraft and Launch Vehicle.

4.2 Thermal

The spacecraft shall monitor and report in housekeeping telemetry the temperatures of subsystems on the Spacecraft.

The spacecraft shall maintain subsystems within their design survival temperature range when the Spacecraft is in Safe Hold Mode.

The Spacecraft shall maintain subsystems within their operational design temperature range when the Spacecraft is in Science Mode.

The Spacecraft shall maintain subsystems within their operational design temperature range when the Spacecraft is in Launch Mode.

The Spacecraft shall maintain payload temperatures as specified in Table 13.

4.3 Flight Software

General

On-board bootstrap shall reside in non-volatile memory that is not reprogrammable on-orbit.

The Spacecraft flight software contained in reprogrammable memory shall be reprogrammable on orbit.

The Spacecraft flight software shall verify the validity of all memory areas.

Spacecraft shall detect and correct single bit errors in on-board processor memory.

Spacecraft shall detect and report multiple bit errors in on-board processor memory.

The Spacecraft flight software shall accept code uploads without affecting loaded stored commands.

4.3.1.6.1 Spacecraft flight software code uploads shall be verified on-board prior to its use.

to verify that this is the correct version prior to use.

The Spacecraft shall store the software version identifier within the software embedded onboard.

The spacecraft shall report the software version identifier in telemetry.

The Spacecraft flight software shall provide the capability to load any location of on-board programmable memory by referencing its physical memory address.

Spacecraft flight software shall be capable of dumping any location in program memory.

The Spacecraft flight software shall be capable of dumping the entire memory of on-board processors.

The Spacecraft flight software memory dump capability shall not interfere with ongoing spacecraft operations.

Event Logging

The Spacecraft flight software shall time tag events logged in telemetry with an accuracy of 100 milliseconds or less to the spacecraft clock.

Spacecraft flight software shall report all event messages in the Spacecraft housekeeping telemetry.

Spacecraft reported events shall include information on the source processor, flight software task or function, severity level, message identifier and informational string that identifies the specific event.

The Spacecraft shall preserve contents of the event log after rebooting or power cycling a processor.

Initialization

The Spacecraft shall initialize flight software and begin operations without the need of a ground based command.

The Spacecraft flight software shall execute a restart warm boot of a processor’s software from Read Only Memory in response to a ground command. Note: A warm boot is defined as an executable code refresh from non-volatile memory and estart of the code. Data is preserved wherever possible (i.e., re-initialization of memory is not performed).

The Spacecraft shall execute a restart cold boot of a processor’s software from Read Only Memory following a power cycle or hardware reset.

The Spacecraft flight software shall default to a known telemetry configuration following a restart.

Stored Commands

The Spacecraft flight software shall provide a capability to store and execute absolute-time command sequences.

The Spacecraft flight software absolute-time stored command sequences shall be capable of invoking other absolute or relative-time stored command sequences.

to verify that this is the correct version prior to use.

Spacecraft flight software shall provide a capability to store and execute relative-time command sequences.

The Spacecraft flight software shall execute real-time commands before executing stored commands.

The Spacecraft flight software shall provide the capability to store multiple time-tagged command sequences.

The Spacecraft flight software stored command sequences shall be modifiable by ground command.

The Spacecraft flight software command sequences shall be enabled, disabled or canceled by ground command.

The Spacecraft flight software command sequences shall be verified onboard prior to use.

The Spacecraft flight software shall be capable of placing the contents of the stored command buffer(s) into the housekeeping telemetry by a single command.

The Spacecraft flight software shall uniquely identify stored command sequences.

The Spacecraft flight software shall use the unique command sequence identifier to report the status of each actively executing command sequence in telemetry.

The Spacecraft flight software time resolution for stored commands shall be 1 second or less.

The Spacecraft flight software shall be capable of executing commands within 1 second of the execution time tag.

4.4 Power

The Spacecraft shall provide a single point ground (SPG) as part of the power subsystem.

The contractor may propose an alternate grounding scheme for an existing, proven design.

All of the primary power returns shall be tied together at the SPG.

The Spacecraft shall provide for power generation, power control, energy storage, and power distribution.

The Spacecraft shall provide load shedding control and battery state of charge (SOC) management.

The Spacecraft shall be power positive over each orbit while in any on-orbit mode including Safe Hold and Science.

The Spacecraft shall provide short circuit protection or current limitation for each switchable power circuit.

The Spacecraft shall provide configurable protection against over voltage, under voltage and over current for the Spacecraft and Payload.

to verify that this is the correct version prior to use.

The Spacecraft shall be capable of being powered from Spacecraft Test Equipment with or without flight or ground batteries installed.

The Spacecraft shall monitor and report in housekeeping telemetry, the switch position of each switchable power circuit at the Spacecraft distribution point.

The Spacecraft shall monitor and report in housekeeping telemetry, current and voltage for each power circuit at the Spacecraft power distribution point.

The Spacecraft battery shall be sized such that the state of charge remains above 60% (TBR) during normal operations and above 40% (TBR) at all times for the Spacecraft design life.

The Spacecraft power system shall be capable of electrically monitoring and reporting all battery cells and bypassing failed battery cell(s).

The spacecraft bus shall accommodate an orbit average power consumption per IARD or

ICD.

The spacecraft bus shall accommodate a peak power consumption per IARD or ICD.

Each instrument power feed shall be switchable with overcurrent protectionhall be fused.

4.5 Attitude and Orbit Determination and Control

Ephemeris

The Spacecraft shall provide ephemeris data to the ground segment.

The Spacecraft shall be capable of receiving a ground based ephemeris update. This requirement is optional for on-board Global Positioning System (GPS)-based architecture.Requirement deleted.

The Spacecraft shall be capable of propagating the on-board state for at least 1 week (TBR).Requirement deleted. This requirement is optional for on-board GPS-based architecture.

The Spacecraft shall support science operations for up to 1 week (TBR) from an on-board propagated ephemeris.Requirement deleted. This requirement is optional for on-board GPS-based architecture.

The Spacecraft shall be capable of receiving a ground based Spacecraft reference time and time correction coefficients.

If applicable, the Spacecraft ephemeris data shall include GPS information. GPS implementation shall be compliant with the IS-GPS-200 Specification.

Science Modes and Calibration Maneuvers

Ram Avoidance

The TIM and SIM instruments are sensitive to degradation due to exposure to atomic oxygen when their boresights are pointed near the spacecraft velocity vector.

4.5.2.1.1 As a primary response, the spacecraft shall off-point when the spacecraft velocity vector is within 15 (TBR) degrees of the TIM boresight.

to verify that this is the correct version prior to use.

4.5.2.1.2 As a secondary response, the spacecraft shall send a safing command to TIM (to close its shutters) when the spacecraft velocity vector is within 15 (TBR) degrees of the TIM boresight.

4.5.2.1.3 As a secondary response, the spacecraft shall send a safing command to SIM (to close its shutters) when the spacecraft velocity vector is within 5 (TBR) degrees of the SIM boresight.

Cruciform Calibration

The SIM boresight alignment needs to be verified on-orbit every 6 months to account for any changes.

4.5.2.2.1 The spacecraft shall be capable of performing a Cruciform Calibration maneuver to assess SIM alignment.

4.5.2.2.2 The spacecraft Cruciform Calibration maneuver scans shall consist of steps at 0.1 degree intervals from 0.4 to 1.5 degrees of sun center.

4.5.2.2.3 The spacecraft Cruciform Calibration maneuver scans shall consist of 200 second (TBR) dwells at each step in the maneuver.

4.5.2.2.4 The spacecraft Cruciform Calibration maneuver shall consist of four scans in orthogonal axes across the face of the sun with 200 second (TBR) dwells at sun center prior to each scan (e.g., sun center, scan from +0.4 to +1.5 deg in X, sun center, scan from -0.4 to -1.5 deg in X, sun center, scan from +0.4 to +1.5 deg in Y, sun center, scan from -0.4 to -1.5 deg in Y, sun center). Note: XY nomenclature may change in final ICD.

Field of View Map

The TIM instrument needs to map its field of view every 6 months to account for any degradation of the cavities. Each FOV map will take multiple orbits to complete. Observations at sun center are required for each orbit to ensure that TIM is still able to meet its nominal TSI measurement requirement as well as to provide context for the off-pointing data.

4.5.2.3.1 The spacecraft shall be capable of performing a FOV map to assess potential TIM degradation.

4.5.2.3.2 The spacecraft FOV map shall consist of 25 equally spaced points ranging from -10 to +10 arcminutes (TBC) off sun center in orthogonal axes.

4.5.2.3.3 The spacecraft FOV map shall consist of 400 second (TBR) dwells at each step.

4.5.2.3.4 The spacecraft FOV map shall consist of 400 second (TBR) dwells at sun center at the beginning and end of each solar orbit (i.e., FOV map points in a given orbit are enveloped by observations at sun center).

Pointing

The spacecraft bus shall be 3-axis stabilized.

to verify that this is the correct version prior to use.

Precision

4.5.3.1.1 The spacecraft shall have the capability to point the instrument boresights within 6 arcmin (3) of sun center for nominal science observations.

4.5.3.1.2 The spacecraft shall have the capability to point the instrument boresights within 6 arcmin (3) of the commanded position during Cruciform Calibration and Field of View maneuvers.

4.5.3.1.3 The spacecraft shall have the capability to fine tune solar offsets based on measured solar offsets derived from on-orbit FSS data. This allows the spacecraft to null out any shifts in the boresights relative to the spacecraft master attitude reference.

4.5.3.1.4 The spacecraft solar offset tuning shall have a resolution of 1 arcmin (3) or better.

Short-Term Stability (Jitter)

While sun pointing for nominal science observations, the spacecraft shall maintain an average drift motion of less than 2 arcmin (3) over 1 second intervals during any single orbit.

Co-Alignment

The TIM and SIM shall be co-aligned to within 2 arcmin (3)

4.6 Command and Data Handling

General

The spacecraft shall be capable of supporting simultaneous uplink and downlink during ground contacts.

The spacecraft shall provide RS-422 command and telemetry services to the TSIS-2 payload.

The Spacecraft shall report in spacecraft housekeeping telemetry the health and safety of the spacecraft and spacecraft-monitored Payload telemetry.

The Spacecraft onboard processor(s) shall be capable of being reset via a discrete command.

The Spacecraft shall packetize spacecraft data into unique Consultative Committee for Space Data Systems (CCSDS) packets.

All Payload data shall be timestamped with an accuracy of 100 ms.

Telemetry

The Observatory shall process, monitor, and report spacecraft and payload housekeeping telemetry.

The Spacecraft shall be capable of continuously transmitting real-time housekeeping data to the ground when in contact with a ground station.

The Spacecraft shall record all housekeeping telemetry for downlink.

The Spacecraft shall time-tag housekeeping telemetry to the spacecraft time reference to verify that this is the correct version prior to use.

with an accuracy of 100 ms.

The Spacecraft clock shall be accurate to within 125 ms of UTC.

The Spacecraft shall report command acceptance or rejection in housekeeping telemetry upon receipt.

The Spacecraft shall be capable of playing back stored housekeeping data while in any powered mode.

The Spacecraft shall be capable of changing the housekeeping telemetry data acquisition rates and have a high rate of at least once per second.

The Spacecraft shall collect housekeeping every 30 seconds.

Housekeeping telemetry data shall include flags for any attempted interference or unauthorized commanding to the Spacecraft.

The Spacecraft shall be capable of storing at least 2 weeks (TBR) of Observatory housekeeping data at the maximum onboard telemetry rate.

The Spacecraft shall be capable of storing at least 2 weeks (TBR) of TSIS-2 Observatory data (science plus housekeeping).

The spacecraft telemetry stream shall employ Grade 2 CCSDS service or better as defined by the CCSDS 701.0-B-3 specification at the downlink frame level.

The downlink frame protocol shall be CCSDS compliant.

The downlink frame protocol shall be compatible with NEN per User Guide 453-UG- 002905.

Command

The Spacecraft bus shall execute real time commands within 1 second of receipt of the command.

The Spacecraft shall support Communication Operations Procedure-1 protocol as defined in CCSDS 232.1-B-1.

The Spacecraft shall have command authentication capability without encryption..

Unencrypted command authentication is acceptable.

The Spacecraft shall reject invalid commands.

The Spacecraft shall be capable of receiving FSW table loads across multiple contacts.

The Spacecraft shall be capable of receiving flight software loads across multiple contacts.

The Spacecraft shall use unique commands to change state and condition, if applicable.

The commanding protocol shall be CCSDS compliant.

Mass Storage

General to verify that this is the correct version prior to use.

4.6.4.1.1 The mass storage overwrite mode shall be configurable to either halt recording or overwrite the oldest data when full.

4.6.4.1.2 The spacecraft mass storage device shall be capable of mapping around corrupt memory sectors.

4.6.4.1.3 The Spacecraft mass storage shall be power cycled independently from any other Spacecraft subsystem or instrument.

4.6.4.1.4 The Spacecraft shall retain all stored data when transitioning into and out of any operational mode, unless power is cycled off.

4.6.4.1.5 The Spacecraft shall have the capability to enable/disable sections of mass storage by ground command.

Recording and Playback

4.6.4.2.1 The Spacecraft shall concurrently record real-time Science and Housekeeping Data and playback stored Science and Housekeeping Data.

4.6.4.2.2 The Spacecraft shall be capable of mass storage playback by physical block address location.

4.6.4.2.3 The Spacecraft shall be capable of repeat playbacks of stored data (for previous 2 week period).

Mass Memory Diagnostics

4.6.4.3.1 The Spacecraft shall have autonomous internal fault detection and correction capabilities for single bit errors and fault detection for multiple bit errors in the mass storage.

4.6.4.3.2 The Spacecraft shall indicate in housekeeping telemetry that stored data has been corrected.

4.6.4.3.3 The Spacecraft shall indicate in housekeeping telemetry that non-correctable bit errors have been detected.

4.6.4.3.4 The Spacecraft mass storage shall be capable of playing back stored science data or test data to the downlink output interface while simultaneously recording science or test data.

4.6.4.3.5 The Spacecraft shall backfill mass storage partitions with a 0xAA byte pattern, prior to writing new data to that partition.

4.7 Radio Frequency Telecommunications System

The Spacecraft shall comply with National Telecommunications and Information Administration (NTIA) Spectrum Standards per “Manual of Regulations and Procedures for Federal Radio Frequency Management (Redbook)”, Chapter 5 Spectrum Standards and Chapter 8 Procedures and Principles for the Assignment & Coordination of to verify that this is the correct version prior to use.

Frequencies.

The Spacecraft shall comply with International Telecommunication Union (ITU) spectrum utilization and sharing requirements per “ITU-R Radio Regulations”, 2016.

The Spacecraft shall comply with Recommendation SFCG 21-2R4.

The Spacecraft shall be capable of transmitting a pseudo random noise pattern from all spacecraft-provided transmitters.

The Spacecraft RF active components shall operate at nominal transmission power in ambient conditions on the ground.

The Spacecraft shall use S-band command uplink.

The Spacecraft shall maintain power to the S-band receiver from launch countdown and through the life of the mission.

The Spacecraft shall be capable of transmission and reception of ground commands while in any powered mode.

The Spacecraft shall be capable of S-band transmissions in any Spacecraft orientation such that the link can be closed to line of sight ground station, having an elevation angle above 5 degrees, with a 3dB margin.

4.8 Contamination and Outgassing

The spacecraft bus materials shall have a total mass loss (TML) less than 1.00% and a collected volatile condensable mass (CVCM) less than 0.10%. Material vacuum outgassing will be determined in accordance with ASTM E-595.

The spacecraft shall not use: pure tin, cadmium, or zinc plating.

The Observatory shall maintain a surface cleanliness Level of 400 A/2 throughout I&T in compliance with the Product Cleanliness Levels and Contamination Control Program, IEST-STD-CC1246.

The Observatory shall be integrated, tested and delivered to the Launch Site in compliance with the Product Cleanliness Levels and Contamination Control Program, IEST-STD-CC1246, Level 450 A/1.5.

The Observatory shall meet a surface cleanliness Level of 500A at Beginning of Life in compliance with the Product Cleanliness Levels and Contamination Control Program, IEST-STD-CC1246.

To the maximum extent practical the spacecraft bus shall vent outgassed species in the direction that is opposite the payload boresights.

4.9 Radiation Space Environment

The Spacecraft shall be compatible with the mission radiation environment as shown in Appendix B.

The Spacecraft shall be analyzed and designed to address surface charging on the…