J1a_2_-_APR8070.2.pdf

Ames Procedural Requirements

APR 8070.2

Effective Date: September 17, 2008 Expiration Date: September 17, 2013

COMPLIANCE IS MANDATORY

Responsible Office: D/Ames Office of the Chief Engineer, Ext. 4-6757, M/S 207-1

Directive Title: Class D Spacecraft Design and Environmental Test

Table of Contents

1.0 Introduction

1.1 Scope

1.2 Applicability

1.3 Precedence

1.4 Authority

1.5 Tailoring

1.6 Responsibilities

2.0 Documents

2.1 Applicable Documents

2.2 Reference Documents

3.0 Spacecraft Design

3.1 Design Margins and Fault Tolerance

3.1.1 Mechanical

3.1.2 Electrical

3.1.3 Thermal

3.1.4 Propulsion

3.1.5 Telemetry and Command

3.1.6 Software and Information System

3.1.7 Safety

3.2 Mechanical

3.2.1 Structural and Mechanical

3.2.2 Fasteners and Pins

3.2.3 Mechanisms

3.2.4 Materials

3.2.5 Fluids Systems

3.2.6 Mechanical Integration

3.3 Electrical

3.3.1 General

3.3.2 EEE Parts

3.3.3 Digital Design

3.3.4 Use of Plastic Encapsulated Microcircuits (PEMs)

3.3.5 Power Systems Design

3.3.6 Power Converters and Supply

3.3.7 Circuit Protection

3.3.8 Grounding, Bonding, and Isolation

3.3.9 Connectors

3.3.10 Radiation Tolerance

3.3.11 Pyrotechnic Functions

3.4 Thermal

3.4.1 General Design Approach

3.4.2 Use of Heaters

3.4.3 Thermal Environments

3.5 Propulsion

3.5.1 Design, Test, and Analysis Requirements

3.5.2 Sizing

3.5.3 Pressurized Components

3.5.4 Propulsion Safety

3.6 Telemetry and Command

3.6.1 Mission Critical Telemetry and Commands

3.6.2 Spacecraft State Management

3.7 Software and Information System

3.7.1 Software Requirements, Plans, and Definition

3.7.2 Flight Software Design

3.7.3 On-Board Data Management

3.7.4 Fault Protection

3.8 Ground Support Equipment

3.8.1 General

3.9 Parts, Materials, and Processes

3.9.1 General

3.9.2 Parts, Materials, and Processes List

3.9.3 Limited Shelf-Life Materials

3.9.4 Workmanship

3.9.5 Parts and Material Traceability

3.9.6 Outgassing

3.9.7 Mechanical Parts Selection

3.9.8 Finishes

3.9.9 Prohibited Materials

3.9.10 Dissimilar Materials

4.0 Integration and Environmental Test

4.1 Integration and Test Plan

4.1.1 System Level Functional and Performance Verification

4.1.2 General I&T Requirements

4.1.3 Qualification of Heritage Hardware

4.2 Environmental Verification

4.2.1 Level of Assembly Test Requirements

4.2.2 Test Sequence

4.2.3 Test Tolerances

4.2.4 Test Article Types

4.3 Structural and Mechanical

4.3.1 Test Factors and Durations

4.3.2 Random Vibration

4.3.3 Acoustic Vibration

4.3.4 Sine Vibration Test

4.3.5 Shock

4.3.6 Pressure Decay/Venting

4.4 Electrical

4.4.1 Electrical Functional and Performance Tests

4.4.2 Electromagnetic Interference and Compatibility (EMI/EMC)

4.5 Thermal

4.5.1 Thermal Vacuum/Cycling

4.6 Contamination Control

4.6.1 Applicability

4.6.2 Summary of Verification Process

4.6.3 Contamination Sensitivity

4.6.4 Contamination Allowance

4.6.5 Contamination Budget

4.6.6 Contamination Control Plan

4.6.7 Other Considerations

5.0 Verification

5.1 Requirements Applicability

5.2 Verification Method

6.0 Definitions and Acronyms

6.1 Definitions

6.2 Acronyms

1.0 Introduction

1.1 Scope

This document defines engineering design and environmental test requirements and guidelines for Class D Spacecraft. This document also incorporates sound engineering practices and lessons learned, to ensure uniformity and consistency of the design and interfaces.

This document is not intended to be all-inclusive. Project specific functional and performance requirements are above and beyond the requirements established in this document.

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1.2 Applicability

This procedure applies to Class D spacecraft projects managed in-house by Ames as well as spacecraft, or spacecraft components, procured by Ames. For spacecraft, and spacecraft components, procured from well-established aerospace contractors, the best practices established by those companies are likely acceptable. The acceptability of individual contractor best practices should be addressed during the tailoring process of the requirements of this document.

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1.3 Precedence

This document takes precedence over APR 8070.1, Engineering Design Requirements for Space Flight Systems, for Class D Spacecraft.

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1.4 Authority

NPD 1280.1, NASA Management Systems Policy

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1.5 Tailoring

Project should review the requirements of this document and develop a tailoring approach. The tailoring approach should define which requirements are applicable as stated, which aren't applicable, and those for which an alternate approach is being requested. It is envisioned that this is a one-time process that occurs very early in the project development life-cycle and requires approval by the Ames Chief Engineer's Office. Additional tailoring beyond the initial tailoring can occur if required by the project.

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1.6 Responsibilities

Project Manager - The project manager is responsible for understanding the requirements and properly reflecting them in resource and schedule decisions. He, or she, is also responsible for specifying and communicating these requirements to government and contractor project personnel.

Lead Systems Engineer - The Lead Systems Engineer (LSE) is responsible for developing the tailoring approach to the requirements of this document and is also the owner of these requirements for the project. The LSE is responsible for ensuring their implementation and verification. The LSE must accept and approve any requested changes or deviations. The LSE must document these decisions in waivers and report the waivers through the technical authority organization at the Center.

SRB/ACMC - The ACMC, or its designate Standing Review Board (SRB), is responsible for reviewing the project plan, approach, and technical specification to ensure compliance to the requirements contained herein.

Ames Chief Engineer's Office - The Ames Chief Engineer's Office is responsible for reviewing and approving the proposed tailoring to the requirements of this document.

2.0 Documents

The following documents of the latest issue or revision shall be used unless a specific document version is given.

2.1 Applicable Documents

(document has no number) Aerospace Structural Metals Handbook

ANSI/AIAA S-080-1998 Space Systems - Metallic Pressure Vessels, Pressurized Structures, and Pressure Components

ANSI/AIAA S-081-2006 Space Systems - Composite Overwrapped Pressure Vessels

ANSI/ESD S20.20 Protection of Electrical and Electronic Parts, Assemblies and Equipment for the Development of an Electrostatic Discharge Control Program

NPR 7150.2 NASA Software Engineering Requirements

GSFC PPL-21 Preferred Parts List

J-STD-001CS Space Applications Electronic Hardware Addendum to Requirements for Soldered Electrical and Electronic Assemblies

IPC A-600 Acceptability of Printed Boards

IPC-A-610 Acceptability of Electronic Assemblies

IPC D-275 Design Standard for Rigid Printed Boards and Rigid Printed Board Assemblies

IPC-2221 Generic Standard on Printed Board Design

IPC-2222 Sectional Design Standard for Rigid Organic Printed Boards

IPC-2223 Sectional Design Standard for Flexible Printed Boards

IPC-6011 Generic Performance Specification for Printed Boards

IPC-6012 Qualification and Performance Specification for Rigid Printed Boards

IPC-6013 Qualification and Performance Specification for Flexible Printed Boards

IPC-6018 Microwave End Product Board Inspection and Test

MIL-HDBK-17 Plastics for Aerospace Vehicles

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

MIL-STD-462 Measurement of Electromagnetic Interference Characteristics

MMPDS-01 Metallic Materials Properties Development and Standardization

MSFC-STD-486 Standard Torque Limits for Threaded Fasteners

MSFC-STD-3029 Guidelines for Selection of Metallic Materials for Stress Corrosion Cracking Resistance

NASA-STD-4005 Low Earth Orbit Spacecraft Charging Design Standard

NASA-STD-5001 Structural Design and Test factors of Safety for Spaceflight Hardware

NASA-STD-5006 General Fusion Welding Requirements for Aerospace Materials Used in Flight hardware

NASA-STD-5017 Design and Development Requirements for Mechanisms

NASA-STD-8739.1 Workmanship Standard for Staking and Conformal Coating of Printed Wiring Boards and Electronic Assemblies

NASA-STD-8739.2 Surface Mount Technology

NASA-STD-8739.3 Soldered Electrical Connections

NASA-STD-8739.4 Crimping, Interconnecting Cables, Harnesses, and Wiring

NASA-STD-8739.5 Fiber Optic Terminations, Cable Assemblies, and Installation

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2.2 Reference Documents

500-PG-8700.2.2 GSFC Electronics Design and Development Guidelines

500-PG-8700.2.7 GSFC Design of Space Flight Field Programmable Gate Arrays

AFSPC MAN91-710 Range Safety User Requirements Manual Volume 3 - Launch Vehicles, Payloads, and Ground Support Systems Requirements

AIAA S-111-2005 Qualification & Quality Requirements for Space Solar Cells

AIAA S-112-2005 Qualification & Quality Requirements for Space Solar Panels

AIAA S-122 Electrical Power Systems for Unmanned Spacecraft

ASTM E595 Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment

DOD-A-83577 General Specification for Moving Mechanical Assemblies for Space Vehicles

DOD-E-83578 General Specification for Explosive Ordnances for Space Vehicles

DOD-W-83575 General Specification for Wiring Harness, Space Vehicle, Design and Testing

DOD-STD-1578 Nickel Cadmium Battery Usage Practices for Space Vehicles

GSFC-STD-7000 General Environment Verification Specification

JSC-07700, Vol XIV Space Shuttle System Payload Accommodations

MIL-A-8625 Anodic Coatings for Aluminum and Aluminum Alloys

MIL-C-5541 Chemical Conversion Coatings on Aluminum and Aluminum Alloys

MIL-F-7179 General Specification for Finishes & Coatings for Protection of Aerospace Weapons System, Structures, & Parts

MIL-M-3171 Processes for Pretreatment and Prevention of Corrosion on Magnesium Alloy

MIL-STD-463 Electromagnetic Interference and Electromagnetic Compatibility Technology Definitions and Systems of Units

MIL-STD-889 Dissimilar Metals

MIL-STD-975 NASA Standard Electrical, Electronic, and Electromechanical Parts List

MIL-HDBK-1568 Materials & Processes for Corrosion Prevention and Control in Aerospace Weapon Systems

MS33611 Tube Bend Radii

MSFC-HDBK-527 Materials Selection List for Space Hardware Systems

NASA-HDBK-4006 Low Earth Orbit Spacecraft Charging Design Handbook

NASA-STD-4003 Electrical Bonding for NASA Launch Vehicles, Spacecraft, Payloads and Flight Equipment

NASA-STD-5002 Load Analyses of Spacecraft and Payload

NASA-STD-6016 Standard Materials and Processes Requirements for Spacecraft

NASA-STD-7001 Payload Vibroacoustic Test Criteria

NASA-STD-7002A Payload Test Requirements

NASA-STD-7003 Pyroshock Test Criteria

QQ-N-290 Nickel Plating

QQ-C-320 Chromium Plating

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3.0 Spacecraft Design

Section 3 provides both requirements and design guidance. Requirements are provided in the form of shall statements and design guidance is in the form of should statements. All shall statements are to be verified while should statements do not require verification.

However, at each milestone review the project team should be able to explain how the design guidance was considered if asked by the review team.

3.1 Design Margins and Fault Tolerance

3.1.1 Mechanical

3.1.1.1 Mass

The Spacecraft minimum mass margin shall be as specified in Table 3.1.1.1-1.

Table 3.1.1.1-1 Mass Margin

SRR PDR CDR TRR

30% 20% 15% 5%

Margin = Allocation - Current Best Estimate (CBE)

% Margin = (Margin / Allocation) x 100

Note: Mass margins less than those indicated in Table 3.1.1.1-1 may be acceptable for systems using heritage hardware. However, the amount of reduced margin will depend on the maturity of the hardware in question and the risk of other components and subsystems impacting its physical and functional requirements.

3.1.1.2 Deployment Systems

Mission critical deployment and separation systems (e.g., solar arrays and other spacecraft appendages, spacecraft-to-launch vehicle separations, etc.) should demonstrate a functional force or torque margin of at least 50% for the entire range of motion or show a margin of 100% by analysis.

Note: The margin applies under worst-case conditions, including restart from any position within the range of motion including incipient latching events.

3.1.1.3 Actuator Design

Mission critical mechanisms and actuators (e.g., electromechanical motors and solenoids, phase-change and state-change actuators, and spring-energized devices, or mechanisms driven by these devices) should demonstrate at least 100% torque/force margin for the entire range of motion.

Note: The margin applies under worst-case conditions at the end-of-life, including restart from any position within the range of motion. Mechanisms and actuators should be kept to a minimum.

3.1.1.4 Stroke for Linear Actuators

Linear actuators implemented for mission critical mechanisms should demonstrate at least 10% stroke margin above the stroke requirements of the mechanism.

3.1.1.5 Mechanism Cycle Life

Mission critical mechanisms that function in a cyclic manner should demonstrate a minimum life capability according to Table 3.1.1.5-1 by test. The mission critical mechanism should operate within specified performance at the end of the life test, and the test unit should be disassembled and inspected for unacceptable wear or debris generation.

Table 3.1.1.5-1 Mechanisms cycle life design margins

Mission Critical Mechanism Element Type Minimum Cycle Life

Margin Requirement

Wet-lubricated low friction elements (e.g., rolling element bearings, involute gearing, etc.)

50%

Dry-lubricated elements, and wear-life limited elements (e.g., brush motors, slip rings, worm gearing, etc.)

100%

Wet-lubricated low friction elements, operating below the temperature rating of the lubricant

100%

Note: Vendor data may be used to demonstrate compliance with this requirement.

Note: A guideline for the life test is 50% of the cycle life test duration be conducted under nominal temperature conditions, 25% at the Protoflight cold temperature, and 25% at the Protoflight hot temperature.

Note: The margin applies under worst-case conditions at the end-of-life.

3.1.1.6 Enclosed Volume

A vented area of greater than 2 times the required area to accommodate the launch profile venting rate shall be provided for all enclosed volumes.

Note: The accommodation of a vented area eliminates the need for structural analysis due to pressure decay loads on structures like electronics box enclosures.

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3.1.2 Electrical

3.1.2.1 Power

The Spacecraft minimum power margin at the end of the mission shall be as specified in Table 3.1.2-1.

Table 3.1.2.1-1 Power Margin

SRR PDR CDR TRR

30% 20% 15% 10%

Margin = Allocation - Current Best Estimate (CBE)

% Margin = (Margin / Allocation) x 100

Note: Margins greater than those above may be indicated by project-specific circumstances. For example, highly complex mission and/or system designs, development of low TRL technology, uncertainty of heritage designs, tight performance margins, low budget reserves, and tight schedule margins might be reason to require higher than the above margins.

Note: Margins less than those above may be acceptable in certain cases. For example, re-use of a known system design with like payload in a mission application previously flown; other circumstances where the unknown factors are fewer and/or mature hardware is by project policy not to be changed; and where ample margins in other technical and programmatic resources may be reason for lesser power/energy margins than is indicated above.

3.1.2.2 Depth of Discharge

Depth Of Discharge (DOD) limits for cyclic operations - For cyclic operations that use battery energy at intervals, the maximum DOD, which depends on the number of battery cycles, shall be in accordance with Table 3.3.5.5.1-1, applicable to both NiH2 and Li-Ion battery types.

Table 3.3.5.5.1-1 DOD Limits for Cyclic Operations

Number of cycles Allowable DOD

< 100 < 70%

100 < # cycles < 5,000 < 60%

5,000 < # cycles < 30,0000 < 40%

> 30,0000 < 20%

Note: A battery cycle is defined as any time the battery is discharged to 50% or more of the allowable DOD and recharged to near full.

3.1.2.3 Power Distribution Circuit Margin

At PDR, there shall be a minimum margin of 30% on spare power switches and circuit count, including wiring, connector pins, and backplane insertion slots to accommodate late identified needs with minimum cost, schedule impact.

3.1.2.4 Pyrotechnic Systems

3.1.2.4.1 Pyrotechnic Circuit Margin

At PDR, there shall be 30% margin on the spare pyro firing circuits (i.e, wires and connector pins) to accommodate late identified needs with minimum cost, schedule impact.

3.1.2.4.2 Pyrotechnic Circuit Fault Protection

Pyrotechnic initiator electronics should be protected from inrush currents and overvoltage conditions.

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3.1.3 Thermal

3.1.3.1 Temperature

The design temperature range for spacecraft components shall be the maximum predicted flight temperature range for the specific mission extended by 10°C.

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3.1.4 Propulsion

3.1.4.1 Propellant Volume

The minimum propellant volume shall be based on a 3-sigma variation in spacecraft properties and include any propellant required for disposal at the end of the mission. The 3-sigma variation is due to the following:

Worst case spacecraft mass properties Low launch vehicle performance Low propulsion subsystem performance (thruster performance/alignment, propellant residuals) Flight dynamics errors and constraints Thruster failure (applies only to single-fault-tolerant systems)

3.1.4.2 Propellant Freezing

Hardware in contact with propellant or propellant vapor shall be thermally controlled to remain safely (>10°C) above propellant freezing temperature.

3.1.4.3 Propellant Condensation

Hardware that will come in contact with propellant vapor shall be thermally controlled over the entire mission to remain >10°C above the temperature at which propellant condensation will occur.

3.1.4.4 Cryogenic Design Margin

The total energy load margin for passive coolers, mechanical coolers and stored cryogen systems designed to operate below -70°C shall be greater than 25%.

Note: The cryogenic margin is to be used in conjunction with the realistic (non-anomalous) stacked worst-case thermal analysis of the cryogenic system enclosure and adjacent spacecraft hardware.

Note: Small additional detector heat loads in the range of a few milli-watts to tens of milli-watts can have large adverse thermal impacts on some cryogenic systems. The total load is comprised of the active and parasitic heat loads.

3.1.4.5 Component Cycle Life

Propulsion components (e.g. chemical thrusters, catalyst beds, engine coatings, etc.)

that function in a cyclic manner shall demonstrate a life capability with greater than 50% margin beyond the worst-case planned mission usage.

Note: Based on the hardware heritage, prior mission use or qualification testing with a dedicated test unit.

3.1.4.6 Attitude Control System

Note: The effects of spacecraft-generated and external magnetic fields need to be considered in the design of the attitude control system including impacts on ACS energy management in counteracting magnetic torques.

3.1.4.6.1 Active Attitude Control System

An active Attitude Control System (ACS) shall have a stability margin of at least 6db for rigid body stability with 30 degrees phase margin and 12db of gain margin for flexible modes.

3.1.4.6.2 Passive Attitude Control System

A passive ACS such as permanent magnet, gravity gradient, and aero stabilized systems shall meet their stability requirements with a 30% margin relative to mission critical parameters, which may include:

Peak or rms microgravity levels Nutation angle limits Pointing accuracy angle

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3.1.5 Telemetry and Command

3.1.5.1 Telemetry and Command Hardware Data Channels

The telemetry and command hardware data channel margin shall be as specified in Table 3.1.5.1-1.

Table 3.1.5.1-1 Telemetry and Command Hardware Channel Margin

Resource SRR PDR CDR TRR

Telemetry and Command Hardware Channels

≥20% ≥15% ≥10% 0

RF Link 6 dB 6 dB 3 dB 3 dB

Margin = Allocation - Current Best Estimate (CBE)

% Margin = (Margin / Allocation) x 100

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3.1.6 Software and Information System

3.1.6.1 Use of Analysis in Lieu of Measurement

Before computer design/procurement, analysis shall be employed to establish margins for critical performance resource parameters such as CPU speed, control cycle rates, interrupt rates and durations, communications bandwidth, and size of RAM, PROM, and EEPROM. Analysis results are documented as the Current Best Estimate (CBE).

3.1.6.2 Information System

The information system (hardware and software) design margin shall be as shown below for margin at critical development milestones to accommodate post-launch fixes, new capabilities, and to maintain adequate in-flight operating margins:

At computer selection 75% Margin At PDR 60% Margin At CDR 50% Margin At launch 20% Margin

Note: The information system margin should consider constrained resources, e.g.

computing capacity, memory, throughput, bus bandwidth, etc. If reused code is used, then lower margins are acceptable. The acceptability of specific margins should be addressed in the tailoring process.

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3.1.7 Safety

3.1.7.1 Catastrophic Hazards

For failures that may lead to a catastrophic hazard, the system shall have three independent, verifiable inhibits (dual fault tolerant).

Note: For example, propellant leakage is a catastrophic hazard.

3.1.7.2 Critical Hazards

For failures that may lead to a critical hazard, the system shall have two independent, verifiable inhibits (single fault tolerant).

Note: For example, stored cryogen systems (and related GSE) are critical hazards for over-pressurization due to blockage or failure of a relief path.

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3.2 Mechanical

3.2.1 Structural and Mechanical

3.2.1.1 Loads

The spacecraft shall have positive margins for all ground, launch, and on-orbit loading conditions specified in Section 4.3.

3.2.1.2 Stiffness Requirements

3.2.1.2.1 Primary Structure

The first fundamental frequency of the primary structure shall be equal to, or greater than, the value specified by the launch vehicle provider. In the event that the spacecraft must be developed in the absence of an identified launch vehicle, then the first fundamental frequency shall be greater than 35 Hz when subjected to the environments specified in Section 4.3.

Note: The launch vehicle provider should be contacted early in the development process to determine the first fundamental frequency requirement. An assessment of the design against the requirement should be provided by PDR.

3.2.1.2.2 Secondary Structure and Components

The first fundamental resonant frequency of the secondary structure or a component weighing less than 23 kg shall be greater than 50 Hz when mounted to its interfacing structure.

3.2.1.3 Factors of Safety

Structural design factors of safety shall be applied in accordance with NASA-STD-5001.

A summary of structural factors of safety, as applied to limit loads, is provided in Table 3.2.1.3-1 for reference only. Refer to NASA-STD-5001 for all safety factors.

Table 3.2.1.3-1 Flight Hardware Design Factors of Safety Applied to Limit Loads

(See Notes 1 and 2)

Type Static Sine Vibe

Random/Acoustic Vibration

(Note 3)

Metallic Yield 1.25 1.25 1.6

Metallic Ultimate 1.4 1.4 1.8

Stability Ultimate 1.4 1.4 1.8

Composite Ultimate 1.5 1.5 1.9

Bonded Inserts/Joints Ultimate

1.5 1.5 1.9

Notes:

1. Factors of Safety for pressurized systems also need to consider appropriate range safety requirements.

2. Factors of Safety for glass & structural glass bonds are specified in NASA-STD- 5001.

3. Factors shown are to be applied to statistically derived peak responses based on RMS level.

3.2.1.3.1 Joint slip factor of safety

The minimum factor of safety (yield or ultimate basis) for all spacecraft structures for non-detrimental instances of bolt slippage, bolt preload exceedance, contact, or encroachment shall be 1.00. The minimum factor of safety for detrimental instances (e.g.

critical instrument alignment, etc.) shall be 1.25.

Note: The required factor of safety is lower for joints that have room to slip in a benign way because the structure will not be strained, and the alignment isn't critical to performance. Verification: Analysis to demonstrate that friction joints will not slip.

3.2.1.3.2 Composite fittings factor of safety

An additional factor of safety of 1.15 shall be applied for composite joints, bonds, and fittings.

Note: This additional fitting factor of safety can be reduced to 1.00 if a comprehensive development test program has been performed for the specific joint configuration, tested under representative worst-case loading and environmental conditions.

3.2.1.3.3 Thermally-induced loading factor of safety

The spacecraft shall have positive margins of safety for yield and ultimate across the qualification/protoflight temperature range for all combined mechanical and thermal stresses.

3.2.1.4 Proof Testing Non-Metallic Structures

Primary and secondary structures fabricated from nonmetallic composites, beryllium, or containing bonded joints or bonded inserts shall be proof tested in accordance with

NASA-STD-5001.

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3.2.2 Fasteners and Pins

3.2.2.1 Torque Limits and Preload

All threaded fasteners shall be installed subject to torque limits and preload requirements of MSFC-STD-486.

3.2.2.2 Locking Requirements

All threaded fasteners for flight hardware shall utilize a locking method. The locking method should be in accordance with MSFC-STD-486.

3.2.2.3 Lubrication

All threaded fasteners shall be installed using appropriate lubrication in accordance with

MSFC-STD-486.

3.2.2.4 Fastener Materials

All threaded fasteners and mating surfaces should be carefully selected to prevent galling.

Series 300 CRES threaded fasteners shall not be used for structural applications.

3.2.2.5 Flat Head Screws

For sheet metal or similar applications, flat head screws should have 100° heads.

3.2.2.6 Use of Alignment Pins

Where pins are used for alignment of mating components, two sets of pins should be used to ensure accurate alignment. Where possible, double eccentric fasteners should be used in lieu of pins.

3.2.2.7 Use of Double Eccentric Fasteners

For transfer of shear loads between components, the use of double eccentric fasteners is preferred. Where shear pins are used to transfer loads, they should be made of hardened steel.

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3.2.3 Mechanisms

The design of mechanisms including deployment, sensor, pointing, drive, despin, separation mechanisms, and other moving mechanical assemblies shall be in accordance with NASA-STD-5017.

DOD-A-83577, General Specification for Moving Mechanical Assemblies for Space Vehicles, should be used as a guide.

3.2.3.1 Spring Energized Bolt Release for Pyrotechnic Separation Nuts

Where pyrotechnically-actuated separation nuts are utilized in the flight system, a spring should be incorporated in the design to ensure positive retraction of the bolt from the separation interface. The spring stroke should meet or exceed the bolt stroke necessary for complete withdrawal of the bolt from the separation interface.

3.2.3.2 Explosive Ordnances

Explosive ordnance should use DOD-E-83578 as a guide.

3.2.4 Materials

3.2.4.1 Environmental Effects on Material Selection

Thorough evaluation of the effect of the trajectory paths/orbits/planetary environment shall be addressed for the impact on materials selection and design as applicable.

Understanding the trajectory and orbital environmental effects (e.g., ESD, radiation, temperature, atomic oxygen, compatibility, debris, etc.) on the spacecraft will eliminate costly redesign and fixes, as well as minimize the on-orbit failures due to environmental interaction with spacecraft materials.

3.2.4.2 Metallic Materials

The composition and associated heat treatment of metallic components shall be in accordance with one of the following documents MSFC-STD-3029 (Table I), Metallic Materials Properties Development and Standardization, DOT/FAA/AR-MMPDS-01, or the Aerospace Structural Metals Handbook.

3.2.4.3 Composite Materials

The composition and lay-up of composite components shall be in accordance with either MIL-HDBK-17 or ANSI/AIAA S-081.

3.2.4.4 Galvanic Corrosion

To avoid the creation of galvanic corrosion couples, the use of dissimilar materials should be in accordance with MIL-STD-889.

3.2.4.5 Stress Corrosion Cracking

Materials susceptible to stress corrosion cracking (SCC), brittle fracture modes, liquid metal embrittlement, and hydrogen embrittlement should be avoided where possible and used in accordance with MIL-HDBK-1568.

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3.2.5 Fluids Systems

3.2.5.1 Materials Compatibility

A materials compatibility assessment shall be performed for all hardware containing fluids by PDR for a preliminary assessment and CDR for a final assessment.

Components containing aqueous solutions should consider the use of Teflon lining or stainless steel. Nickel-plated aluminum shall not be used in place of stainless steel.

3.2.5.2 Leak Before Burst (LBB) & Maximum Expected Pressure (MEP) Criteria

Fluid systems design shall ensure Leak Before Burst (LBB) and Maximum Expected Pressure (MEP) are in conformance with NASA-STD-5001.

3.2.6 Mechanical Integration

3.2.6.1 Component Precision Location

When precise location of a component is required, the design shall use a stable, positive location system (not relying on friction) as the primary means of attachment.

3.2.6.2 Sensor and Antenna Blockage

When a spacecraft is in its stowed (launch) configuration, it shall not obscure visibility of any attitude sensors required for acquisition nor block any antennas required for command and telemetry.

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3.3 Electrical

GSFC document, "Electronics Design and Development Guidelines", 500-PG-8700.2.2, should be used as a reference.

3.3.1 General

3.3.1.1 Burn-In

A minimum of 200 hours of operation shall be required for burn-in testing for vacuum sensitive components or subsystems, with at least 50 hours in a vacuum environment (10-4 Torr or less), and the last 100 hours shall be free of failures. The last 24 hours is intended to cover final functional testing.

Note: Burn-in hours may include hours at the spacecraft level of integration.

Components containing plastic encapsulated microcircuits should be subjected to 500 hours of burn-in testing.

3.3.1.2 Printed Wiring Boards

All flight printed wiring/circuit boards (PWB/PCB) should be coupon tested prior to population with flight components.

3.3.1.3 Electro-Static Discharge (ESD) Control

The project shall document and implement an ESD Control Program that is in compliance ESD Association Standard for the Development of an Electrostatic Discharge Control Program, ANSI/ESD S20.20, for protection of electrical and electronic parts, components, and equipment (excluding electrically initiated explosive devices).

Note: The project shall ensure that all ground support personnel utilize proper equipment and procedures for protection of spacecrafts electronics from damage to electrostatic discharge (ESD).

3.3.1.4 RF Component Immunity to Multipaction

Active high-energy RF components, such as radars, shall be designed and tested for immunity to multipaction before and after environmental testing.

Note: Refer to NASA Public Lessons Learned Entry 0770 for guidance regarding testing and evaluation for the presence of multipaction.

3.3.1.5 Voltage/Temperature Margin Test

A Voltage/Temperature Margin Test (VTMT) should be performed on newly designed electronic parts, boards, or components, which includes combinations of worst-case environments and operational parameters prior to integration at the next level of assembly. The purpose of this test is to demonstrate that an adequate margin exists in the design to compensate for part tolerance, aging and radiation effects, as well as environment and input variation. During this test the parts, board, or component is subjected to variation of voltage at the high and low protoflight temperatures. Table 3.3.1.4-1 provides guidelines for selecting the test parameters. VTMT is not required if a Worse-Case Analysis (WCA) is performed. Existing designs must show evidence that a VTMT or a WCA is not required.

Table 3.3.1.4-1 VTMT Test Parameters

Assembly Type Voltage Variation Temperature Limits

Digital ±10% of worst-case supply voltage

Qual/Protoflight Temperatures

Analog ±10% of worst-case supply voltage

Qual/Protoflight Temperatures

Power Supply Spacecraft Power Bus Variation, -1 Volt @ min. & +1

Volt @ max.

Qual/Protoflight Temperatures

3.3.1.6 Polarity Checks of Critical Components

All hardware shall be verified by test or inspection of the proper polarity, orientation, and position of all components (sensors, switches, and mechanisms).

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3.3.2 EEE Parts

3.3.2.1 EEE Parts Control

An Electrical, Electronic and Electromechanical (EEE) parts control program shall be planned for and implemented for all flight missions for the purpose of parts selection, derating, screening, and overall qualification. The use of COTS parts will be evaluated on a project specific basis. The use of COTS parts for mission critical functions should be avoided.

Note: A preliminary plan to be developed by SRR and a final plan by CDR.

3.3.2.2 Parts Selection

EEE parts selection should consider, at a minimum, the following:

EEE parts susceptible to SEE caused by high-energy ionized particle radiation should be avoided or minimized

Microcircuits should be evaluated for Single Event Latchup (SEL) and Single Event Upset (SEU) susceptibility.

Power bipolar transistors should be evaluated for single event burnout.

Power Hexagonal Field Effect Transistors (HEXFET) should be evaluated for single event gate rupture. Circuit designs containing SEE sensitive electronic parts should be analyzed to minimize the effect of SEE and to assure compliance with component level requirements.

Parts susceptible to single event burnout and gate rupture shall be reported to the Parts Control Board members within 15 days of identification of the part susceptibility.

3.3.2.3 Derating

All EEE parts and components shall be derated in accordance with the guidelines of

GSFC PPL-21.

3.3.2.4 Storage of EEE Parts

EEE parts shall be stored at room temperature, clean conditions, and protected from ESD and high humidity (nitrogen purged).

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3.3.3 Digital Design

3.3.3.1 Synchronous Designs

Synchronous designs shall be used for digital logic to guarantee the sequence of logical decisions and the validity of data transfer. (Asynchronous design may be used if techniques are employed and demonstrated to provide guarantees for sequence verification and validation to the same confidence level as used for a synchronous design.) 3.3.3.2 ASIC/FPGA Synchronous Designs

The synchronous design of an Application Specific Integrated Circuits (ASIC) and Field Programmable Gate Arrays (FPGA) shall be verified, as a minimum by post-route timing analyses using a place and route tool and test vector simulation with timing checkers performed at the primitive level. Timing of boundary conditions (pin-outs) shall be constrained for place, route, and test vector simulation.

Note: GSFC document "Design of Space Flight Field Programmable Gate Arrays", 500- PG-8700.2.7, should be used as a reference.

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3.3.4 Use of Plastic Encapsulated Microcircuits (PEMs)

The use of PEMs is permitted provided each use is thoroughly evaluated for thermal, mechanical, and radiation implications of the specific application and found to meet mission requirements. PEMs should be selected for their functional advantage and availability, not for cost saving; the steps necessary to ensure reliability usually negate any initial apparent cost advantage. A PEM shall not be substituted for a form, fit and functional equivalent, high reliability, hermetic device.

Screening of PEMs is essential before PEMs are inserted into most flight hardware.

Burn-in at the part level is to be employed to addresses infant mortality. If burn-in at the part level is not practical, board level burn-in or board/box level Environmental Stress Screening may be substituted. The use of PEMs is time sensitive; a PEM should be no more than three years old from date of manufacture to date of application.

Boards containing PEMs should be cleaned and dried using solvents and baking methods that will not impact the reliability of parts or boards. The terminations of PEMs should be pretinned using tin-lead solder to reduce the risk of tin whisker growth or to remove gold plating. PEMs typically have pure tin plated terminations, which are a risk for tin whisker growth and subsequent system failure due to shorting or plasma arcs.

Alternatively PEMs may be available with gold plated terminations, which are at risk for failure due to gold embrittlement. After installation and cleaning, the application of conformal coating to the devices is recommended to minimize re-absorption of moisture and to further reduce the risk of tin whisker growth.

Note: It is important to bake out Plastic Encapsulated Microcircuits prior to storage and prior to use in order to drive out absorbed moisture from the plastic molding material.

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3.3.5 Power Systems Design

3.3.5.1 Wiring Design

The electrical wiring harnesses between components should be in accordance with DOD-W-83575 and MIL-STD-975.

Note: Commercial wiring should be acceptable as long as assembled harnesses are baked out and hipot tested for space environment.

3.3.5.2 Power Quality

Power quality should be per AIAA S-122 as a guide.

3.3.5.3 Solar Array

Solar arrays shall be designed to meet the system average and peak power load required during daylight and eclipse operations. Design considerations should include orbit altitude, eclipse duration, design lifetime, the type of solar cells and power output, the beginning-of-life (BOL) power production capability, and end-of-life (EOL) power production capability.

Solar arrays should be designed using AIAA S-111 and AIAA S-112 as guides.

3.3.5.4 Batteries

Batteries shall be designed to meet the system power requirements during both nominal and contingency operations. Design considerations should include the required voltage, current loading, duty cycles, activation time, storage time, mission length, primary or secondary power storage, orbital parameters, power use profile, temperature and radiation environment, and battery charge/discharge cycle limits.

For systems using solar arrays, battery capacity requirements should account for periods from launch and array deployment up to nominal operational conditions, including appropriate margins for ground and/or flight anomalies.

Note: Contact the launch provider regarding design and/or certification requirements for use of Lithium-Ion batteries. Nickel-Cadmium Battery usage should be in accordance with DOD-STD-1578.

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3.3.6 Power Converters and Supply

3.3.6.1 Power Converter Synchronization

Subsystem DC-DC power converters shall be synchronized via an externally supplied sync frequency (preferred) or operated in a free-running mode subject to the following conditions:

If operating in a free-running mode, operating frequencies are spread.

Testing and/or analysis of operating frequencies/harmonics show that no interference will be generated.

3.3.6.2 Power-on-Reset

The electronics design shall include the capability to cycle the power on and off either through on-board hardware or software, or a commanded link.

Note: The design should consider the effect of power transients on the electronics. A time-delay between power resets should be considered to minimize the effect on electronics.

3.3.6.3 Power Supply Transient Analysis

A power supply transient analysis shall be performed to identify all transient conditions presented to the spacecraft power bus, including turn-on/off transients, in-rush current, component state changes, etc. The supplier shall maintain the transient analysis on file and available for review by the buyer upon request.

3.3.6.4 Power System Charging

The electrical power system charging design shall be in accordance with NASA-STD- 4005, Low-Earth Orbit Spacecraft Charging Design Standard, when voltages beyond ±55 volts are present.

Note: NASA-HDBK-4006 should be used as design guidance.

Note: Surge protection and power clamping should be considered in the design for the protection of power sources.

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3.3.7 Circuit Protection

Circuit protection shall be provided for the spacecraft system, subsystems, and components to adequately protect the hardware. The use of fuses in mission-critical applications should be minimized.

Note: There is no recovery from an in-flight blown fuse, thus their use in mission-critical applications must be carefully considered. The design should consider the use of, or lack of use of, circuit protection for mission critical hardware when tripped circuit breakers would result in the loss of space vehicle control.

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3.3.8 Grounding, Bonding, and Isolation

3.3.8.1 Local Single Point Ground

Each subsystem ground tree shall have a local single point DC ground to spacecraft chassis via the shortest practical wire length.

Grounding should be in accordance with AIAA S-122.

3.3.8.2 Primary Circuit Return Path

Wires should be used for the primary circuit return path. Structure or shields should not be used.

3.3.8.3 Bonding

All components shall be bonded to the structure of the spacecraft with a resistance across the interface of less than 2.5 milliohms.

Note: NASA-STD-4003, Electrical Bonding for NASA Launch Vehicles, Spacecraft, Payloads and Flight Equipment, should be used as a guide.

3.3.8.4 Electrical Isolation

All component interface circuits shall be electrically isolated to not less than 1 mega-ohm DC between power lead and chassis ground and between power lead and signal ground.

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3.3.9 Connectors

3.3.9.1 Inadvertent Mating

Keyed or unique connectors should be used to prevent inadvertent connector mating.

Note: If keyed connectors are not used then the design shall preclude inadvertent mating of the wrong connectors.

3.3.9.2 Mating of Powered Connection

Provisions shall be provided to prevent mating or de-mating of connectors at signal or power interfaces unless power has been turned off (dead facing).

3.3.9.3 Blind Mate Connection

Provisions shall be provided to allow verification of proper connection (e.g. alignment and depth of engagement) of blind mate connectors. However, the use of blind mate connectors should be avoided due to difficulties in verifying proper connection.

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3.3.10 Radiation Tolerance

Note: Verification by analysis is acceptable for the radiation tolerance requirements of Section 3.3.10.

3.3.10.1 Single Event Effect (SEE)

The spacecraft electronic devices shall be chosen such that the subsystem operates within performance specification during and after exposure to the high- energy radiation environments defined for the mission with an RDF of 1. The high-energy radiation environments are unique for each mission and need to be determined using NASA qualified models.

The Radiation Design Factor (RDF) is defined as:

RDF =

Radiation-resisting capability of a part or component in a given application

Radiation environment present at the location of the part or component

The component, system, and spacecraft level requirements regarding performance with respect to SEE during operation are as follows:

a. Temporary loss of function or loss of data shall be permitted provided that the loss does not compromise subsystem/system health, full performance can be recovered rapidly, and there is no time in the mission that the loss is mission critical.

b. Normal operation and function shall be restored via internal correction methods without external intervention in the event of an SEU.

c. Fault traceability shall be provided in the telemetry stream for anomalies involving SEE's.

3.3.10.2 Total Ionizing Dose

3.3.10.2.1 General Shielding

The spacecraft's electronic components shall be chosen such that the component operates within performance specification during and after the TID exposure at an RDF of 2 times the TID level present at the location of the device. The TID is unique for each mission and needs to be determined using NASA qualified models.

3.3.10.2.2 Spot Shielding

Where spot shielding is to be applied, an RDF of 3 shall be required.

Note: The greater RDF with use of spot shielding is to account for uncertainties in part capabilities, space environment, and transport modeling. General shielding should be considered in the evaluation of shielding needs for an RDF of 3. Spot shielding alone does not need to account for an RDF of 3.

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3.3.11 Pyrotechnic Functions

3.3.11.1 Simultaneous Arming of Multiple Functions

The design of pyrotechnic firing circuits shall avoid simultaneous arming of multiple functions.

Note: It is important to avoid spurious unplanned pyrotechnic events caused by planned firings or other transient effects.

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3.4 Thermal

3.4.1 General Design Approach

3.4.1.1 Design Tailored to Specific Application

Thermal control design should be tailored to the specific applications of the mission, with consideration for both equipment reliability and temperature/ performance interactions.

Nominal and worst-case temperature ranges for mission critical hardware shall be determined and the impact to the thermal design from credible failure modes assessed.

3.4.1.2 Passive vs. Active Thermal Control Systems

When possible, passive thermal control systems should be utilized. Active thermal control systems shall be used only when necessary.

3.4.1.3 Motor and Actuator Self-Heating

Motor or actuator operating flight acceptance, protoflight, and qualification temperature ranges shall be defined by the interface or environmental temperature extremes, prior to energizing a motor or actuator. The maximum allowable temperature limit of the motor or actuator due to self-heating shall be defined separately.

Note: This requirement is to ensure consistency between the definition of operating and non-operating Allowable Flight Temperature limits, while accommodating the characteristic of motor or actuator self-heating during ground testing.

3.4.1.4 Thermal Coatings Properties

All thermal analysis shall employ thermal coatings properties validated to be accurate for materials and mission flight parameters over the lifecycle of the mission.

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3.4.2 Use of Heaters

The heater design should consider the survival and operational temperature ranges of the hardware.

3.4.2.1 Power Density Limit for Film Heaters

Kapton film heaters, when bonded to a metallic or composite substrate over 100% of the heater's active area, should be limited to a maximum power density of 3.0 watts per square centimeter

Kapton film heaters should be limited to a maximum power density of 0.5 watts per square centimeter when bonded over voids or other discontinuities of the heat-dissipating substrate.

3.4.2.2 Maximum Duty Cycle of Actively Controlled Heaters

The maximum duty cycle should be 80% or less under worst-case cold environmental conditions when actively controlled heaters are implemented in a flight system thermal control design.

Note: The purpose of this requirement is to ensure that positive heater control authority exists during the design phase of the flight system thermal control. Heater duty cycling can fall outside of the prescribed design range during the mission, but should generally meet the above maximum duty cycling criteria during ground testing to validate the thermal control design.

Note: This principle applies to systems above -70 degrees C. Active controllers include mechanical thermostats and PID or pulse wave modulated controllers. To conserve power consumption throughout the mission due to typically large spacecraft bus voltage ranges, the maximum 80% duty cycle applies to the condition of a minimum nominal voltage, and not a minimum failed voltage, e.g. failed battery cell or solar cell string.

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3.4.3 Thermal Environments

3.4.3.1 Thermal Radiation

3.4.3.1.1 Direct Solar Radiation

The direct solar radiation values to be used for the design of earth orbiting spacecraft shall be as defined in Table 3.4.3.1.1-1.

Table 3.4.3.1.1-1 Solar Radiation

Case Direct Solar Flux

(W/m2)

Cold 1322

Nominal 1367

Hot 1414

3.4.3.1.2 Earth Albedo

The reflected solar radiation from earth, albedo, to be used for the design of earth orbiting spacecraft shall be as defined in Table 3.4.3.1.2-1.

Table 3.4.3.1.2-1 Earth Albedo

Latitude Range (deg)

Annual Mean Albedo

Annual Range Albedo

90 - 80 0.67 0.44 - 0.75

80 - 70 0.57 0.49 - 0.83

70 - 60 0.46 0.39 - 0.78

60 - 50 0.41 0.37 - 0.56

50 - 40 0.36 0.32 - 0.46

40 - 30 0.31 0.26 - 0.37

30 - 20 0.26 0.25 - 0.30

20 - 10 0.24 0.20 - 0.27

10 - 0 0.25 0.24 - 0.26

0 - -10 0.23 0.21 - 0.25

-10 - -20 0.23 0.21 - 0.24

-20 - -30 0.24 0.23 - 0.25

-30 - -40 0.29 0.27 - 0.30

-40 - -50 0.35 0.33 - 0.39

-50 - -60 0.42 0.41 - 0.47

-60 - -70 0.51 0.46 - 0.77

-70 - -80 0.64 0.61 - 0.88

-80 - -90 0.70 0.40 - 0.80 global annual average

0.30 0.29 - 0.31

3.4.3.1.3 Earth Infrared Radiation

The earth infrared radiation value (global annual average) to be used for the design of earth orbiting spacecraft shall be 234 ±7 W/m2.

3.4.3.2 Deep Space Temperature

The deep space radiative sink temperature used for the design shall be 2.7°K.

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3.5 Propulsion

3.5.1 Design and Analysis Requirements

3.5.1.1 Fuses for Propulsion System

Flight fuses for wetted propulsion system components shall be selected such that overheating of propellant will not occur at the maximum current limit rating of the flight fuse.

3.5.1.2 Plume Impingement Analysis

Thruster or external venting plume impingement shall be analyzed and demonstrated to meet mission requirements.

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3.5.2 Sizing

3.5.2.1 Propellant Tanks

Propellant tank volume…