Attachment J-LNEXT- LNERD Rev-.pdf

LNERD LNEXT-SYS-REQ-0005

Revision -ii

Landsat Next Environmental Requirements Document (LNERD)

Signature/Approval Page

Prepared by:

Electronic Signature in TDMS

12/15/2022 Daniel Helfrich Date Landsat Next Mission Mechanical Lead NASA/GSFC, Code 543

Approved by:

Mark Edison Date Landsat Next Deputy Observatory Manager

NASA/GSFC, Code 426

Wen-Ting Hsieh Date Landsat Next Payload Manager NASA/GSFC, Code 426

Samilet Lee Date Landsat Next Chief Safety and Mission Assurance Officer

NASA/GSFC, Code 383 iii

Evan Webb Date Landsat Next Systems Manager NASA/GSFC, Code 599

James Pontius Date Landsat Next Project Manager NASA/Goddard, Code 426 iv

CM Foreword This document is a Landsat Next Project Configuration Management (CM)-controlled document.

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

Questions or comments concerning this document should be addressed to:

NASA/Goddard Space Flight Center Landsat Next Project Office, Code 426 Attention: Configuration Management Office Greenbelt, Maryland 20771 v

Change History Log

Revision Effective Date Description of Changes

- 12/15/2022 LNEXT-CCR-0025 – Initial Release vi

List of TBDs/TBRs Hyperlink to TBx Location Summary Ind.

Name/Org.

Due Date

TBx-1

Section 1.2.4.3

The Launch Segment will provide the assets and services associated with the Launch Vehicle (LV) and the constellation-to-LV integration. This will include the LV; all Launch Vehicle-Ground Support Equipment (LV-GSE), property, and facilities to integrate the constellation to the LV and verify their integration; and prelaunch testing with ground-based functions. The launch vehicle and launch site are TBD. The three observatories comprising the Landsat Next constellation will be launched together on the same launch vehicle.

J. Satrom/

31 Jun

TBx-2

Section 2.5.2.5.1

LNERD-2929 Radiated Susceptibility testing shall be conducted over a frequency range of 14 kHz (TBR) to 40 GHz IAW GSFC-STD- 7000B and MIL-STD-461G RS103

R. Vernot/

31 Jan

TBx-3

Section 2.5.2.5.1

LNERD-2931 Observatories, subsystems and components that are powered on for launch shall exhibit no observable change of state while powered on in launch configuration and exposed to electric fields in accordance with RS103, at the test levels defined in Table 2.5.2.5.1-1. (TBR Pending LV and launch site selection) .

R. Vernot/

31 Jan

TBx-4

Section 2.5.2.5.1

LNERD-4213 Observatories, subsystems and components that are powered on for launch shall survive without damage or degradation following exposure to electric fields in accordance with RS103, at the test levels defined in Table 2.5.2.5.1-1. (TBR Pending LV and launch site selection)

R. Vernot/

31 Jan

TBx-5

Section 2.7.2

LNERD-1359 Flight and flight spare components shall be subjected to a minimum of eight (TBR) cold cycles prior to integration onto the observatory/spacecraft.

D. Patel/

31 Jan vii

Table of Contents SIGNATURE/APPROVAL PAGE ................................................................................................ II

CM FOREWORD ......................................................................................................................... IV

CHANGE HISTORY LOG ............................................................................................................ V

LIST OF TBDS/TBRS .................................................................................................................. VI

TABLE OF CONTENTS ............................................................................................................. VII

LIST OF FIGURES ...................................................................................................................... IX

LIST OF TABLES ........................................................................................................................ IX

1.0 GENERAL INFORMATION

1.1 Purpose

1.2 Mission Overview

1.2.1 Mission Statement

1.2.2 Mission Background

1.2.3 Mission Objectives

1.2.4 Mission Implementation

1.2.4.1 Space Segment

1.2.4.2 Ground Segment

1.2.4.3 Launch Segment

1.3 Applicability and Limitations

1.4 The Landsat Next Verification Approach

1.5 Other Assurance Requirements

1.6 Responsibility for Administration

1.7 Landsat Next Configuration Control and Distribution

1.8 Applicable and Reference Documents

1.8.1 Applicable Documents

1.8.2 Reference Documents

1.9 Criteria for Unsatisfactory Performance

1.10 Testing of Spare Hardware

1.10.1 "One-Shot" Items

1.11 Test Facilities, Calibration

1.12 Test Condition Tolerances

1.13 Test Measurement Considerations

2.0 VERIFICATION PROGRAM

2.1 System Performance Verification

2.2 General Verification Requirements

2.2.1 Test Sequence and Level of Assembly

2.2.2 Verification Program Tailoring

2.2.3 Qualification of Hardware by Similarity

2.2.4 Test Factors/Durations

2.2.5 Structural Analysis/Design Factors of Safety

2.3 Function and Performance Test Requirements

viii

2.3.1 Hardware Performance Testing

2.3.2 Long Duration and Failure Free System Level Test of Flight Software

2.4 Structural and Mechanical Verification Requirements

2.4.1 Structural Loads Qualification

2.4.1.1 Coupled Load Analysis

2.4.1.2 Modal Survey

2.4.1.3 Design Strength Qualification

2.4.1.4 Structural Reliability (Residual Strength Verification)

2.4.1.5 Acceptance Requirements

2.4.2 Vibroacoustic Qualification

2.4.2.1 Fatigue Life Considerations

2.4.2.2 Acoustic Testing

2.4.2.3 Observatory Random Vibration Tests

2.4.2.4 Subsystem/Instrument Vibroacoustic Tests

2.4.2.5 Component/Unit Vibroacoustic Tests

2.4.2.6 Acceptance Requirements

2.4.2.7 Retest of Reworked Hardware

2.4.3 Sinusoidal Sweep Vibration Qualification

2.4.3.1 Observatory Sine Sweep Vibration Tests

2.4.3.2 Subsystem and Component Sine Sweep Vibration Tests

2.4.3.3 Sine Vibration Acceptance Requirements

2.4.4 Mechanical Shock Qualification

2.4.4.1 Subsystem Mechanical Shock Tests

2.4.4.2 Observatory Mechanical Shock Tests

2.4.4.3 Acceptance Requirements

2.4.5 Mechanical Function Verification

2.4.5.1 Life Testing of Mechanisms

2.4.5.2 Demonstration of Mechanical Function

2.4.5.3 Torque/Force Margin

2.4.5.4 Acceptance Requirements

2.4.6 Pressure Profile Qualification

2.4.6.1 Pressure Profile Test Demonstration

2.4.6.2 Acceptance Requirements

2.4.7 Mass Properties Verification

2.4.7.1 Mass Properties Test Demonstration

2.4.7.2 Acceptance Requirements

2.5 Electromagnetic Compatibility (EMC) Requirements

2.5.1 General EMC Requirements

2.5.1.1 Requirements Applicability

2.5.1.2 Safety and Controls

2.5.1.3 Test Facility and Equipment Requirements

2.5.1.4 Class R Bonding Verification

2.5.1.5 Power Grounding Verification

2.5.2 Detailed EMC Requirements

2.5.2.1 Conducted Emissions, Time Domain, Transients

2.5.2.2 Conducted Emissions Requirements

ix

2.5.2.3 Conducted Susceptibility Requirements

2.5.2.4 Radiated Emissions Requirements

2.5.2.5 Radiated Susceptibility Requirements

2.6 Thermal Verification Testing

2.6.1 General Requirements

2.6.2 Unit Test

2.6.3 Subsystem/Instrument Test

2.6.4 System/Satellite Test

2.6.5 Flight Spares

2.6.6 Flight Unit Retest

2.7 Cryogenic Requirements

2.7.1 Cryogenic Systems Vacuum and Thermal Verification Requirements

2.7.2 Cryo-Thermal-Vacuum Qualification

2.7.2.1 Special Considerations

2.7.2.2 Level of Testing

2.7.2.3 Test Parameters

2.7.2.4 Test Setup

2.7.2.5 Demonstration

2.7.2.6 Special Tests

2.7.3 Leak Checking

2.7.4 Thermal Balance Qualification

APPENDIX A ABBREVIATIONS AND ACRONYMS

APPENDIX B DEFINITIONS

List of Figures Figure 1.2.4-1 Landsat Next Operations Concept Figure 2.4.2.5-1 Determination of Qualification and Acceptance Random Verification

Test Levels Figure 2.4.4.1.3-1 Shock Response Spectrum (SRS) for assessing Component Test

Requirements Figure 2.5.2.2.1-1 Conducted Emission Limits on Power Lines Figure 2.5.2.2.2-1 Common Mode Bulk Cable Emission Limits Figure 2.5.2.3.1-1 CS101 Power Line Conducted Susceptibility Voltage Limits Figure 2.5.2.3.1-2 CS101 Power Line Conducted Susceptibility Power Limits Figure 2.5.2.3.3-1 Conducted Susceptibility, Power Leads, 150 kHz to 50 MHz Voltage

Limit Figure 2.5.2.3.3-2 Conducted Susceptibility, Power Leads, 150 kHz to 50 MHz Current

Limits Figure 2.5.2.4.1-1 Radiated Emissions Electric Field Limit with Suggested Notches

List of Tables

Table 1.12-1 Test Condition Tolerances Table 2.2.2-1 Flight System Hardware Levels of Assembly Table 2.2.4-1 Test Factors/Durations x

Table 2.2.5-1 Flight Hardware Design/Analysis Factors of Safety Applied to Limit Loads

Table 2.4.1-1 Structural and Mechanical Verification Test Requirements Table 2.4.1-2 Minimum Probability-Level Requirements for Flight Limit (maximum expected) Level Table 2.4.2.6-1 Generalized Random Vibration Test Levels Components 22.7-kg (50-lb) or less Table 2.4.2.6-2 Component Minimum Workmanship Random Vibration Test Levels45.4-kg (100-lb) or less Table 2.4.5.3-1 Torque/Force Margins By Project Phase Table 2.5.1.1-1 EMC Requirements Applicability Table 2.5.2.5.1-1 Launch Site Survival Radiated Susceptibility Electric Field Exposure .82 Table 2.5.2.5.1-2 Operational Radiated Susceptibility Electric Field Environment Table 2.6.1-1 Environmental Verification Test Level Temperature Ranges Table 2.7.1-1 Cryogenic Vacuum, Thermal, Electrical and Mechanical Requirements Table 2.7.2.3-1 Margin as a function of Phase In Project Life Cycle

1.0 GENERAL INFORMATION

1.1 PURPOSE

This document provides requirements and guidelines for the environmental verification program for the Landsat Next (LNext) constellation, payload, observatories, instruments, spacecraft buses, subsystems and components and describes methods for implementing those requirements.

It contains a baseline for demonstrating by test or analysis the satisfactory performance of hardware in the expected mission environments, and that minimum workmanship standards have been met. It elaborates on those requirements, gives guideline test levels, provides guidance in the choice of test options, and describes acceptable test and analytical methods for implementing the requirements. For contractual requirements purposes, the instrument is treated as a subsystem, and the spacecraft bus is treated as a module. There will be environmental verification plans to address requirements specified within this document at the instrument and observatory levels. (The term payload as used in this document refers to the ready to launch collection of Landsat Next observatories, their separation adapters, and any other hardware that is together mounted atop or within the upper stage of the launch vehicle.)

This document, nicknamed the LNERD, is directly based on the General Environmental Verification Standard (GEVS, GSFC-STD-7000B). GEVS has been tailored to create the many Landsat Next Project specific environmental verification requirements within this document. For the purposes of environmental verification of Landsat Next Project flight hardware, GEVS is not a formally applicable document in its entirety, but instead, portions are referenced herein for numerous purposes including the clarification of testing rationale and for specific test guidance.

1.2 MISSION OVERVIEW

1.2.1 Mission Statement

Landsat Next, consistent with United States (U.S.) law and government policy, will continue the Landsat program's acquisition, archival, and distribution of multi-spectral imagery affording global, synoptic, and repetitive coverage of the Earth's land surfaces at a scale where natural and human-induced changes can be detected, differentiated, characterized, and monitored over time.

1.2.2 Mission Background

Following the successful launch of Landsat 8 (formally named Landsat Data Continuity Mission, LDCM) in February 2013, and during the development of Landsat 9, the United States Geological Survey (USGS) and National Aeronautics and Space Administration (NASA) recognized the need to assemble a team of experts from within both agencies to evaluate how to inform an acquisition strategy for the Landsat mission to follow Landsat 9. The NASA-USGS Joint Agency Sustainable Land Imaging (SLI) Architecture Study Team (AST) was formed in September 2018 and was tasked with investigating how to best satisfy the diverse set of user needs collected in the USGS “User Needs for the Sustainable Land Imaging Program – Release 2.0.” These investigations resulted in a set of recommendations to the headquarters of both agencies, delivered in December 2019. The highest-recommended “Roadmap 1” architecture described a small constellation of “superspectral” space-based sensors that would substantially improve the spectral, spatial, and temporal capabilities of previous Landsat missions, while continuing to satisfy the primary goal of ensuring a highly calibrated data set that maintains compatibility with the legacy data of the Earth’s land mass held in the National Satellite Land

Remote Sensing Data Archive (NSLRSDA) at USGS’s Earth Resources and Observation Science (EROS) Center. In April 2020 NASA/Goddard Space Flight Center (GSFC) received authorization to initiate the Landsat Next project consistent with the AST’s Roadmap 1 recommendation.

The goal of Landsat Next is to continue the acquisition, archival, and distribution of multi-spectral imagery affording global, synoptic, and repetitive coverage of the Earth's land surfaces at a scale where natural and human-induced changes can be detected, differentiated, characterized, and monitored over time. This goal is in keeping with the Landsat programmatic goals stated in both the Commercial Space Act of 1998 (Public Law 105-303) and the Land Remote Sensing Policy Act of 1992 (Public Law 102-555). This policy requires that the Landsat Program provide data into the future that is sufficiently consistent with previous Landsat data to allow the detection and quantitative characterization of changes in or on the land surface of the globe.

Landsat Next continues the long-running partnership of NASA and USGS, with NASA providing the space and launch segments and USGS providing the ground system, providing the longest continuous global record of the Earth’s surface. The Landsat series of satellites have continuously acquired multispectral images of the global land surface since the launch of the Earth Resources Technology Satellite (ERTS, later renamed Landsat 1) in 1972. The Landsat data archive constitutes the longest continuous moderate-resolution record of the global land surface as viewed from space.

1.2.3 Mission Objectives

Landsat Next has these major mission objectives:

• Collect and archive moderate resolution multispectral image data and thermal image data, affording seasonal coverage of the global landmass for a continuous period of not less than five (5) years.

• Ensure that Landsat Next data are sufficiently consistent with data from the earlier Landsat missions in terms of relative acquisition geometry, calibration, coverage characteristics, spectral characteristics, output product quality, and data availability to permit studies of land cover and land use change over multi-decadal periods.

• Ensure Landsat Next is responsive to critical emerging user needs and applications as characterized by periodic assessment, currently the User Needs for the Sustainable Land Imaging Program July 2018, Release 2.0, and identified by the operational requirements for collection, processing, archiving, and distribution of land surface data to the United States Government and other users.

• Distribute Landsat Next data products to the general public on a nondiscriminatory basis.

1.2.4 Mission Implementation

NASA and USGS each have specific responsibilities for Landsat Next and will deliver the major elements to the overall mission. NASA will provide the Space and Launch Segments of Landsat Next, and USGS will provide the Ground System and mission operations. NASA/GSFC will provide overall Landsat Next project management, mission system engineering, and mission assurance during development and will transition the mission to USGS following on-orbit commissioning.

The Landsat Next post-launch nominal mission operations concept is shown graphically in the Figure below.

Figure 1.2.4-1 Landsat Next Operations Concept

1.2.4.1 Space Segment

NASA/GSFC will provide the Space Segment via competitive procurements for the science instruments and the spacecraft bus. The Space Segment will consist of a constellation of three observatories flying in coordinated sun-synchronous orbits at 653km altitude, each with nominally identical spacecraft and instrument suites. The observatories will be equally spaced in the orbit, providing in aggregate a six-day ground repeat period at the equator.

The instrument suites will each provide 26 spectral bands with a maximum ground sampling distance of 10m-60m dependent on the spectral band, covering a swath on the ground of approximately 164km on a ground track defined by a world-wide reference system known as

WRS-3.

1.2.4.2 Ground Segment

The Landsat Ground Segment currently supports mission operations for Landsats 8 and 9.

Landsat Next takes advantage of developments on these missions and will upgrade and expand the current systems to accommodate Landsat Next. The Landsat Ground Segment consists of the Ground System (GS) and its external interfaces, including NASA institutional services. The GS includes the Mission Operations Center (MOC), the Ground Network (GN), and the Data Processing and Archive System (DPAS). External interfaces include NASA's Near Space Network (NSN) and NASA ACCESS Space relay (i.e. TDRSS) and government/commercial ground station, NASA/GSFC Conjunction Assessment and Risk Analysis (CARA) and its NASA/GSFC Flight Dynamics Facility, along with other external interfaces.

The MOC provides the primary means to control and monitor the Landsat Next constellation.

The Landsat Next Flight Operations Team (FOT) at the MOC performs mission planning and scheduling, command and control, health and status monitoring, orbit and attitude maintenance, performance analysis, onboard memory management, and flight and ground software maintenance. The FOT utilizes MOC functionality to detect, investigate and resolve spacecraft anomalies and monitor the instrument image collections from the onboard constellation and generate special image collections. The MOC ingests, processes and archives data via the GN.

The GN includes geographically dispersed ground station resources for mission execution, and includes both the Landsat Ground Network (LGN) and a wideband or cloud-based data routing capability to transfer both mission data to DPAS, and TT&C data to the MOC. The LGN consists of US Government-owned, international, and commercial ground stations, and provides communication capability for each observatory of the constellation for commanding and housekeeping data via S-Band two-way links. The LGN will also receive mission data from each observatory via high rate Ka-band downlinks.

The DPAS ingests, processes, and archives LNext mission data from the GN. The DPAS also provides a long-term archive capability for raw data and allows the user community to query, download, and directly interact in the cloud with Landsat Next science products, via a public-facing web portal for receiving data products. The DPAS is located at USGS Earth Resources Observation and Science (EROS) near Sioux Falls, South Dakota.

The USGS will lead overall Landsat Next GS development. The USGS will also lead integration of the GS and ensure timely completion of ground readiness testing in preparation for NASA-led mission readiness activities. The MOC will perform planning, scheduling, and observatory operations activities during Landsat Next mission readiness testing.

1.2.4.3 Launch Segment

The Launch Segment will provide the assets and services associated with the Launch Vehicle (LV) and the constellation-to-LV integration. This will include the LV; all Launch Vehicle- Ground Support Equipment (LV-GSE), property, and facilities to integrate the constellation to the LV and verify their integration; and prelaunch testing with ground-based functions. The launch vehicle and launch site are TBD. The three observatories comprising the Landsat Next constellation will be launched together on the same launch vehicle.

Section 1.2 was generated using Landsat Next Common Boilerplate (LNEXT-MGMT-DESC- 0005) Rev B

1.3 APPLICABILITY AND LIMITATIONS

These requirements apply to LNext hardware and associated software that is to be launched on an LNext Project designated launch vehicle.

These requirements apply to all space flight hardware, including Ground Support interface hardware that is developed as part of the LNext Project.

These requirements are written in accordance with the current GSFC practice of using a single protoflight observatory to qualify for space flight (see definition of hardware, 1.8). The protoflight verification program, therefore, is given as the nominal test program. Unless granted relief by the Landsat Next Project, the second and third observatories will also be tested as protoflight hardware.

1.4 THE LANDSAT NEXT VERIFICATION APPROACH

The entire LNext Constellation, Observatories, instruments, subsystems, and components must be verified under conditions that simulate the launch operations, flight operations and flight environment as realistically as possible.

Since testing at the component (or unit) level, or lower level of assembly for large components, often becomes a primary part of the verification program, all components should be operating and monitored during all environmental tests if practicable.

Environmental verification of hardware is only a portion of the total assurance effort for LNext that establishes confidence that the constellation will function correctly and fly a successful mission. The environmental test program provides confidence that the design will perform when subjected to environments more severe than expected during the mission, and provides environmental stress screening to uncover workmanship defects.

The total verification process also includes the development of models representing the hardware, tests to verify the adequacy of the models, analyses, alignments, calibrations, functional/performance tests to verify proper operation, and finally end-to-end tests and simulations to show that the total system will perform as specified.

The level, procedure, and decision criteria for performing any additional tests will be included in the LNext system verification plan.

1.5 OTHER ASSURANCE REQUIREMENTS

In addition to the verification program, the assurance effort includes parts and materials selection and control, reliability assessment, quality assurance, software assurance, design reviews, and system safety per the LNext Instrument Mission Assurance Requirements (IMAR) and Spacecraft Mission Assurance Requirements (SMAR).

1.6 RESPONSIBILITY FOR ADMINISTRATION

The responsibility and authority for decisions in waving the requirements rest with the GSFC LNext Project Manager.

The requirements thus derived and deviations from the requirements of this document are subject to review and approval by the GSFC Landsat Next Project.

1.7 LANDSAT NEXT CONFIGURATION CONTROL AND DISTRIBUTION

This document is controlled and maintained by the GSFC Landsat Next Project.

1.8 APPLICABLE AND REFERENCE DOCUMENTS

LNERD-25 The Contractor shall use the version of all Applicable Documents in effect at the time of procurement throughout development, except as identified in specific requirement(s).

1.8.1 Applicable Documents

Document Number Title Revision Applicable Section(s) GSFC-STD-1000 Rules for the Design, Development, Verification, and Operation of Flight Systems

G GR 2.01

NASA-STD-5001 Structural Design and Test Factors of Safety for Space Flight Hardware

B w/CHANGE

Section 4.2.4

NASA-STD-5009 Nondestructive Evaluation Requirements for Fracture Critical Metallic Components

B ALL

NASA-STD-5019 Fracture Control Requirements for Spaceflight Hardware

A w/CHANGE

FCR 6 through FCR 17

NASA-STD-7000 General Environmental Verification Standard

(GEVS)

B See requirements

NASA-STD-7001 Payload Vibroacoustic Test Criteria B Section 4.3.4

NASA-STD-

8719.24 (Annex)

Annex to NASA Expendable Launch Vehicle Payload Safety Requirements: Requirements Table

A w/Change

Chapters 11 and 12

MIL-STD-461 Requirements for the Control of Electromagnetic Interference Characteristics of Subsystems and Equipment

G Sections 5.4 thru 5.9, 5.12, 5.18, & 5.21

MSFC-STD-1249 Standard NDE Guidelines and Requirements for Fracture Control Programs

In this document, citations are assumed to be the latest version unless otherwise noted.

This document table was generated from the Landsat Next Referenced Documents List Draft Rev -

1.8.2 Reference Documents

Document Number Title Revision GPR 8070.4 Administration and Application of Goddard Open-Learning

Design (GOLD) Rules for the Design, Development, Verification and Operation of Flight Systems

D

GPR 8730.1 Meterology: Control of Measuring and Test Equipment N NASA-HDBK-7004 Force Limited Vibration Testing D NASA-HDBK-7005 Dynamic Environmental Criteria A NASA-STD-5002 Load Analyses of Spacecraft and Payloads A NASA-STD 7002 Payload Test Requirements B NASA-STD-7003 Pyroshock Test Criteria A ANSI/NCSL z540.1 Requirements for the Calibration of Measuring and Test

Equipment

TBD

ANSI/NCSL z540.3 Requirements for the Calibration of Measuring and Test Equipment

TBD

ISO/IEC 17025 General requirements for the competence of testing and calibration laboratories

Third Edition

MIL-HDBK-728 Nondestructive Testing. 1 Notice 1 - Validation

MIL-HDBK-6870 Inspection Program Requirements, Non-Destructive Testing for Aircraft and Missile Materials and Parts.

C

NAS-410 Certification and Qualification of Nondestructive Test Personnel

In this document, citations are assumed to be the latest version unless otherwise noted.

This document table was generated from the Landsat Next Referenced Documents List Draft Rev -

1.9 CRITERIA FOR UNSATISFACTORY PERFORMANCE

Deterioration or any change in performance of any test item that does or could in any manner prevent the item from meeting its functional, operational, or design requirements throughout its mission is reason to consider the test item as having failed. Further elaboration of project requirements regarding failed testing may be found in the LNext IMAR or SMAR, as appropriate.

If during a test sequence a test item is operated in excess of design life and wears out or becomes unsuitable for further testing from causes other than deficiencies, a spare may be substituted.

LNERD-3492 If the substitution of a test item that has exceeded its design life affects the significance of test results, the test during which the item was replaced and any previously completed tests that are affected shall be repeated to the extent necessary to demonstrate satisfactory performance.

1.10 TESTING OF SPARE HARDWARE

A supply of selected spares is often maintained in case of the failure of flight hardware. Sparing requirements are documented in the Instrument and Spacecraft Statements of Work.

LNERD-3496 Spares shall undergo a verification program equal to that required for follow-on hardware.

LNERD-261 The extent and type of testing shall be determined as part of the flight hardware test program.

LNERD-262 If a flight element is replaced for reasons of failure and is then repaired and re-designated as a spare, appropriate retesting shall be conducted as agreed to by the Landsat Next Project.

LNERD-263 When the need for a spare arises, immediate analysis and review of the failed hardware shall be made. If failure occurs in a hardware item of which there are others of identical design, the fault may be generic and may affect all hardware of that design.

1.10.1 "One-Shot" Items Some items may be degraded or expended during the integration and test period and replaced by spares.

LNERD-3497 The spare that is used to replace an expended “One-Shot” item shall have met the required quality control standards and passed all required flight qualification tests for items. Such as pyrotechnic gas generators and release devices, low-shock energy release mechanisms, and elements that absorb impact energy by plastic yielding.

LNERD-3498 When the replacement of a 'One-Shot' item entails procedures that could jeopardize mission success, the replacement procedure shall be successfully demonstrated with the hardware in the same configuration that it will be in when final replacement is to be accomplished.

1.11 TEST FACILITIES, CALIBRATION

The facilities and fixtures used in conducting tests must be capable of producing and maintaining the test conditions prescribed with the test specimen installed and operating or not operating, as required.

LNERD-3499 Facility performance shall be verified prior to the test.

LNERD-3500 All measurements that require a specified level of accuracy shall be taken using equipment that has been properly calibrated, with documentation available, using one of the standards: ANSI/NCSL z540.1, ANSI/NCSL z540.3, or ISO 17025, IAW GPR 8730.1.

1.12 TEST CONDITION TOLERANCES

LNERD-270 The values and measurement uncertainties in Table 1.12-1 shall be used unless a rationale is provided for using other test condition tolerances, subject to the approval of the Landsat Next Project Office.

Table 1.12-1 Test Condition Tolerances

Acoustics Overall Level: ≤ 1 dB Tolerance: +3/-0 dB l/3 Octave Band Tolerance: Frequency (Hz) Tolerance (dB) f ≤ 40

40 < F < 3150 f ≥ 3150

+3, -6 ±3

+3, -6 Duration: +10%, -0% Antenna Pattern Determination ± 2 dB Electromagnetic Compatibility Voltage Magnitude: ± 5% of the peak value Current Magnitude: ± 5% of the peak value RF Amplitudes: ± 2 dB Frequency: ± 2 % Distance: ± 5% of specified distance or

± 5 cm, whichever is greater Humidity ± 5% RH Loads Steady-State (Acceleration): ± 5% Sine Burst Amplitude: ± 5% Static: ± 5% Magnetic Properties Mapping Distance Measurement: ± 1 cm Displacement of assembly center of gravity (cg) from rotation axis:

± 5 cm

Vertical displacement of single probe centerline from cg of assembly:

± 5 cm

Mapping turntable angular displacement: ± 3 degrees Magnetic Field Strength: ± 1 nT Repeatability of magnetic measurements (short term): ± 5% or ± 2 nT, whichever is greater Demagnetizing and Magnetizing Field Level: ±5% of nominal Mass Properties Weight: ± 0.2% Center of Gravity: ± 0.15 cm (± 0.06 in.)

Moments of Inertia: ± 1.5% Mechanical Shock Response Spectrum: Frequency (Hz) Tolerance (dB) Simulated Fn ≤ 3 kHz

Fn ≥ 3kHz ± 6 +9/-6

Shaker Fn ≤ 3 kHz ± 3 Overall Spectrum > 50% of SRS magnitude above nominal test level

Time History: ± 10% Pressure Greater than 1.3 X 104 Pa

(Greater than 100 mm Hg):

± 5%

1.3 X l04 to 1.3 X l02 Pa

(100 mm Hg to 1 mm Hg):

± 10%

1.3 X l02 to 1.3 X 101 Pa

(1 mm Hg to 1 micron):

± 25%

Less than 1.3 X 101 Pa (less than 1 micron):

± 80%

1.13 TEST MEASUREMENT CONSIDERATIONS

LNERD-348 From the moment a test article is excited or illuminated by an environmental source until the article is returned to an ambient condition, measurements shall be collected over time at sufficient temporal resolution to capture any relevant frequency- or time-dependent effects and to eliminate, in the cases of an anomaly or test failure, the possibility that the incident was caused by overtest due to an error in process or in the test equipment itself.

2.0 VERIFICATION PROGRAM

2.1 SYSTEM PERFORMANCE VERIFICATION

This section applies to the entire Landsat Next Constellation, each Observatory including its instrument/subsystems, spacecraft bus and components. The provisions apply to all flight hardware, and associated software.

The SOW provides the guidelines for an environmental verification program and the CDRL describes the many associated deliverables

2.2 GENERAL VERIFICATION REQUIREMENTS

Sections 2.3 through 2.7 cover the specific requirements for implementing the Landsat Next environmental verification program for verifying observatories, subsystems, and components as follows:

2.3 Function & Performance Test Requirements

2.4 Structural and Mechanical

2.5 EMC

2.6 Thermal

2.7 Cryogenics

For the purposes of this document, a spacecraft is considered an observatory, an instrument is considered to be a subsystem, and a spacecraft bus is considered to be a module when determining the environmental verification requirements. In general, the environmental tests are not performed at the spacecraft bus level, but they can be done where beneficial for risk reduction or other purposes. Functional and performance tests are what will generally be done at the spacecraft bus level prior to LandIS integration.

The basic provisions are written assuming protoflight hardware. They are, in general, also applicable to prototype hardware. Acceptance requirements are also given for the flight acceptance of previously qualified hardware. This applies to follow-on hardware (multiple copies of the same item) developed for the program, or hardware (from another program) qualified by similarity.

2.2.1 Test Sequence and Level of Assembly

The verification activities herein are grouped by discipline; they are not all in a recommended sequence of performance. The recommended order of testing is electrical, mechanical and thermal at the end of the testing sequence. No specific environmental test sequence is required, but the test program will be arranged in a way to best disclose problems and failures associated with the characteristics of the hardware and the mission objectives.

LNERD-492 In cases where the magnetic properties of the hardware need to be controlled, the dc magnetics testing shall be performed after vibration testing to provide an opportunity to correct for any magnetization of the flight hardware caused by fields associated with the vibration test equipment.

Table 2.2.2-1 provides a hierarchy of levels of assembly for the flight hardware, with examples.

These level designators are based on those used in the Space Systems Engineering Database developed by The Aerospace Corporation for the Air Force, and agreed to by NASA Headquarters, GSFC, and JPL. The LNERD environmental test requirements generally start at the “unit” level and end at the “system segment” level. However, screening and life-tests often occur at lower levels, and overall system verification continues beyond the “system segment” level.

2.2.2 Verification Program Tailoring

Table 2.2.2-1 provides a hierarchy of levels of assembly for the flight hardware, with examples.

The LNERD environmental test requirements generally start at the “unit” level and end at the “system segment” level. However, screening and life-tests often occur at lower levels, and overall system verification continues beyond the “system segment” level.

This document assumes that the observatory is of modular design and can be tested at the unit/component, subsystem/instrument, and observatory system levels of assembly.

The Contractor will develop a verification program as stated in the SOW that satisfies the intent of the required verification program while taking into consideration the specific characteristics of the mission and the hardware. For example:

An observatory subsystem, or instrument, may be a functional subdivision of the spacecraft, but it may be distributed throughout the spacecraft rather than being a physical entity.

In this case, the environmental tests, and associated functional tests, will be performed as a minimum at the component, instrument and observatory levels. Performance tests and calibrations may still be performed on the functional subsystem or instrument.

The physical size of the system may necessitate testing at other levels of assembly. Facility limitations may not allow certain environmental tests to be performed at the system level. In this case, testing should be performed at the highest practicable level. Also, for very large systems or subsystems/instruments, tests at additional levels of assembly may be added in order to adequately verify the hardware design, workmanship and/or performance.

In some cases, because of the hardware configuration it may be reasonable to test more than one component at a time. The components may be stacked in their flight configuration and may therefore be tested as a "section." Part of the decision process must consider the physical size and mass of the hardware. The test configuration must allow for adequate dynamic or thermal stress inputs to the hardware to uncover design errors and workmanship flaws.

Some test requirements stated as subsystem/instrument requirements may be satisfied at a higher level of assembly if approved by the GSFC project. For example, externally induced mechanical shock test requirements may be satisfied at the system level by firing the environment-producing pyro. A simulation of this environment is difficult, especially for large subsystems or instruments.

Aspects of the design and/or mission may negate certain test conditions to be imposed. For example, if the on-orbit temperature variations are small, less than 5°C, then consideration should be given to waiving the thermal-vacuum cycling at the system, or instrument, level of assembly in favor of increasing the hot and cold dwell times.

The same process must be applied when developing the test plan for an instrument. While testing is required at the instrument component and all-up instrument levels of assembly, additional test levels may be called for because of hardware complexity or physical size.

Table 2.2.2-1 Flight System Hardware Levels of Assembly

2.2.3 Qualification of Hardware by Similarity

There are cases in which hardware qualified for one flight program is to be built and used on another program. Hardware that has been previously qualified may be considered qualified for use on a new program by showing that the hardware is sufficiently similar to the original hardware and that the previous qualification program has adequately enveloped the new mission environments. The details for performing this comparison should be defined by the Contractor with approval from the LNext Project. As a minimum the following areas should be reviewed and documented:

LNERD-507 Design and test requirements of original hardware that was previously qualified shall be shown to envelope the design and test requirements of subsequent hardware that is to be qualified by similarity to include a detailed review of the test configuration and of all waivers and deviations that may have occurred during testing of the original hardware.

LNERD-508 Manufacturing information of hardware that is to be qualified by similarity shall be reviewed to determine if changes have been made that would invalidate the previous hardware qualification, covering parts, materials, packaging techniques as well as changes to the assembly process or procedures.

LNERD-509 Test experience with the previous flight build shall be reviewed to verify that no modifications were made to the hardware during testing to successfully complete the test program, or else to verify all changes to the heritage hardware as a result of testing have been implemented on the current flight hardware.

If the LNext Project review shows that the hardware is of sufficiently similar design as the first build and that the previous test requirements envelope any new environmental requirements, then the hardware can be, with LNext approval, treated as qualified and need only to be subjected to acceptance level test requirements. The review of the hardware for similarity must be documented and included as part of the Contractor verification package.

2.2.4 Test Factors/Durations

Test factors for prototype, protoflight, and acceptance are given in Table 2.2.4-1.

Table 2.2.4-1 Test Factors/Durations

Test Prototype

Qualification

Protoflight

Qualification

Acceptance

Structural Loads1 Level

Duration Centrifuge/Static Load4 Sine Burst

1.25 x Limit Load 1 minute

5 cycles @ full level per axis

1.25 x Limit Load 30 seconds

5 cycles @ full level per axis

1.0 x Limit Load 30 seconds

5 cycles @ full level per axis

Acoustics Level2 Duration

Limit Level + 3dB 2 minutes

Limit Level + 3dB 1 minute

Limit Level 1 minute

Random Vibration Level2 Duration

Limit Level + 3dB 2 minutes/axis

Limit Level + 3dB 1 minute/axis

Limit Level 1 minute/axis

Sine Vibration3 Level Sweep Rate

1.25 x Limit Level 2 oct/min

1.25 x Limit Level 4 oct/min

Limit Level 4 oct/min

Test Prototype

Qualification

Protoflight

Qualification

Acceptance

Mechanical Shock Actual Device Simulated

2 actuations

1.4 x Limit Level

2 x Each Axis

2 actuations

1.4 x Limit Level

1 x Each Axis

1 actuations Limit Level

1 x Each Axis Thermal-Vacuum Max./min. predict.

± 10°C Max./min. predict.

± 10°C Max./min. predict.

± 5°C EMC & Magnetics As Specified for

Mission Same Same

1 - If qualified by analysis only, positive margins are to be shown for factors of safety of 2.0 on yield and 2.6 on ultimate as directed in subsection 2.4.1.1.1. Beryllium and composite materials may not be qualified by analysis alone as covered in section 2.4.1.3.4. Test levels for weldments, beryllium, bonded and composite structure, including metal matrix, are to be 1.25 x Limit Level for acceptance testing as covered in section 2.4.1.5.

2 - As a minimum, the test level should be equal to or greater than the workmanship level.

3 - The sweep direction must be evaluated and chosen to minimize the risk of damage to the hardware. If a sine sweep is used to satisfy the loads or other requirements, rather than to simulate an oscillatory mission environment, a faster sweep rate may be considered, e.g., 6-8 oct/min to reduce the potential for over stress.

4 - Shorter durations may be used in static testing if necessary to protect the hardware from damage due to facility limitations. If a shorter duration is used then the dwell time at load must be sufficient to demonstrate that the target loading condition has been achieved within the specified tolerances, all test measurements have been recorded, and the structure is stable under the applied loading condition.

2.2.5 Structural Analysis/Design Factors of Safety

LNERD-514 Structural and mechanical verification testing shall be supported by structural analysis to provide confidence that the hardware will not experience failure or detrimental permanent deformation under test or launch conditions.

LNERD-3714 The factors of safety that shall be applied to limit loads in order to calculate structural margins are shown in Table 2.2.5-1.

These factors of safety have been selected to be consistent with the test factors shown in Table 2.2.4-1. The yield factor of safety ensures that a prototype or protoflight test can be conducted with low risk of the hardware experiencing detrimental yielding. The ultimate factor of safety provides adequate separation between yield and ultimate failure modes and ensures that the hardware will not experience an ultimate failure under expected loading conditions.

LNERD-515 In the case of thermally induced loads or stresses, the factors of safety shown in Table 2.2.5-1 for the static loading condition shall be used for calculating strength margins.

LNERD-3715 If the absolute value of the temperature differential between the stressed and un-stressed condition for the hardware using flight acceptance temperatures is equal to or greater than 20° C, then the static factors of safety from Table 2.2.5-1 shall be applied to the loads/stresses induced by acceptance temperatures.

Flight acceptance temperatures are defined as the maximum predicted flight temperature plus acceptance margin which is typically 5° C for thermal vacuum or 20° C for thermal cycling at ambient pressure as defined in Table 2.2.4-1. Survival temperatures are treated as flight acceptance temperatures for assessing strength margins as they do not include qualification margin. If the absolute value of the temperature differential between the stressed and un-stressed condition using flight acceptance temperatures is < 20°C, then the static factors of safety from Table 2.2.5-1 can be applied to the loads/stresses induced by qualification temperatures for hardware which does not contain any heat generating components. Both hot and cold conditions should be evaluated and the factors of safety should be applied to the appropriate test temperature (acceptance or qualification). Thermally induced loads should be combined with mechanical loading due to launch loads, gravity, or external loads due to enforced deflection if these mechanical loads occur at the same time.

Table 2.2.5-1 Flight Hardware Design/Analysis Factors of Safety Applied to Limit Loads

Type Static Sine Random/Acoustic

Metallic Yield 1.25 1.25 1.6 Metallic Ultimate 1.4 1.4 1.8 Stability Ultimate 1.4 1.4 1.8 Beryllium Yield 1.4 1.4 1.8

Beryllium Ultimate 1.6 1.6 2.0 Composite Ultimate 1.5 1.5 1.9

Bonded Inserts/Joints Ultimate 1.5 1.5 1.9

LNERD-605 Factors of safety for pressurized systems shall be compliant with NASA-

STD-8719.24.

LNERD-606 Factors of safety for glass and structural glass bonds shall be as specified inNASA-STD-5001B, Section 4.2.4.

LNERD-608 The random/acoustic safety factors shown in Table 2.2.5-1 shall be applied to statistically derived peak response based on RMS level and calculated as a 3-sigma value as a minimum.

Factors shown in Table 2.2.5-1 assume that qualification/protoflight testing is performed at acceptance level plus 3dB. If difference between acceptance and qualification levels is less than 3dB, then these factors may be applied to qualification level minus 3dB instead of analyzing to acceptance level.

2.3 FUNCTION AND PERFORMANCE TEST REQUIREMENTS

The SOW defines functional, performance and confidence tests that are meant to verify and validate the observatory's operation before, during, and after environmental testing. These tests along with all other calibrations, measurements/ demonstrations, alignments (and alignment verifications), end-to-end tests, simulations, etc., that are part of the overall verification program are to be described in the System Performance Verification Plan. The following requirements address the minimum times for performance tests, and for failure-free testing/simulations that are needed for the Landsat Next observatories.

2.3.1 Hardware Performance Testing

LNERD-623 One-thousand (1000) hours of operating/power-on time shall be accumulated on all flight electronic hardware, including any spares, prior to launch, in accordance with GOLD Rule GR-2.01.

NOTE: For electronics consisting of a prime and redundant sides, i.e. A and B sides, 1000 hours should be accumulated on each side.

LNERD-624 At the conclusion of the performance verification program, the observatory shall have demonstrated failure-free performance testing for at least the last 350 hours of operation via any combination of CPT, LPT and Mission Simulation testing.

NOTE: Hardware changes during or after the verification program invalidate previous demonstration.

LNERD-2109 At the conclusion of the performance verification program, all flight electronic hardware shall have demonstrated failure-free performance testing for at least the last 350 hours of operation.

The demonstration hours may be accumulated at the subsystem level or system level of assembly. Hardware changes during or after the verification program invalidate previous demonstration.

LNERD-2110 At the conclusion of the performance verification program, flight electronic hardware shall have demonstrated failure-free performance testing for at least 200 hours of operation during the thermal-vacuum test exposure, with 100 hours of the trouble-free operation being logged at the hot-dwell temperatures and 100 hours being logged at the cold-dwell temperature.

NOTE: Hardware changes during or after the verification program invalidate previous demonstration.

The general intent of the above requirements is to accumulate 1000 hours of operating time on all flight hardware, and to demonstrate trouble-free performance at high-, low-, and nominal temperature. However, it is understood that under certain conditions this goal may not be met, requiring a waiver. For example hardware change-out, rework or repair just prior to launch may not provide sufficient time to fulfill these requirements. Also, the retest requirements following component failure during system level thermal vacuum, or other tests, must be evaluated on a case-by-case basis taking into account the criticality of the hardware element and the risk impact on achieving mission goals, as technical justification for any waiver approval.

LNERD-2111 The LandIS instrument shall accumulate at least 300 hrs of operation per side, prior to shipment for integration to the observatory.

LNERD-2112 The LandIS instrument shall accumulate at least 50 hours of failure-free operation, prior to shipment to the observatory for integration and test.

NOTE: Any failure or hardware change invalidates prior failure-free time.

LNERD-2113 The LandIS instrument shall accumulate at least 100 hrs of operation in vacuum per side or any redundant hardware, prior to shipment for integration to the observatory.

2.3.2 Long Duration and Failure Free System Level Test of Flight Software As required by the GOLD Rules, ground test of the fully integrated FSW system includes the demonstration of error-free operations-like scenarios over a continuous 72 hour period.

NOTE: The minimum duration uninterrupted FSW system-level test (on the highest fidelity FSW testbed) is 72 hours.

2.4 STRUCTURAL AND MECHANICAL VERIFICATION REQUIREMENTS

A series of tests and analyses should be conducted to demonstrate that the flight hardware is qualified for the expected mission environments and that the design of the hardware complies with the specified verification requirements such as factors of safety, interface compatibility, structural reliability, workmanship, and associated elements of system safety.

LNERD-632 Table 2.4.1-1 specifies the structural and mechanical verification activities that shall be completed.

Note: When the tests and analyses are planned, consideration must be given to the expected environments of…