4.3Verification

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

Last Updated 29 Oct 2026

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[MPR 209] Verification of compliance with the requirements of this NASA Technical Standard shall consist of the following steps, as a minimum:

a. NASA approval of the contractor Materials and Processes Selection, Control, and Implementation Plan and other applicable materials data requirements documents, such as the Contamination Control Plan and NDE Plan.
b. Contractor M&P signature on engineering drawings to verify compliance with the requirements of this NASA Technical Standard or the Materials and Processes Selection, Control, and Implementation Plan.
c. NASA audits of contractor M&P activities relating to hardware design and manufacturing.
d. Establishment and operation of the M&P control panel in accordance with section 4.1.4 of this NASA Technical Standard.
e. NASA approval of MUAs.
f. NASA approval of MIULs. NOTE: Per section 4.1.6: The use of M&P that do not comply with the requirements of this NASA Technical Standard may be acceptable in the actual hardware applications. MUAs are required for all M&P that are technically acceptable but do not meet the requirements of this NASA Technical Standard, as implemented by the approved Materials and Processes Selection, Control, and Implementation Plan.

NASA-STD-6016B

Additional aspects of the verification process should be documented in the Materials and Processes Selection, Control, and Implementation Plan.

NOTE: Per section 4.1.6: The use of M&P that do not comply with the requirements of this NASA Technical Standard may be acceptable in the actual hardware applications. MUAs are required for all M&P that are technically acceptable but do not meet the requirements of this NASA Technical Standard, as implemented by the approved Materials and Processes Selection, Control, and Implementation Plan.

NASA-STD-6016B APPENDIX A STANDARD MATERIALS AND PROCESSES REQUIREMENTS FOR SPACECRAFT REQUIREMENTS COMPLIANCE MATRIX WITH REQUIREMENTS JUSTIFICATIONS A.1 PURPOSE

NASA-STD-6016A defines the minimum requirements for Materials and Processes (M&P) and provides a general control specification for incorporation in NASA program/project hardware procurements and technical programs. Incorporating a general control specification provides procedural requirements for programmatic and M&P oversight to ensure that the M&P minimum requirements are implemented, controlled, and verified. These M&P minimum requirements are derived mostly from a history of successful use by NASA and the aerospace industry of best practices and lessons learned; their imposition better ensures uniformity, interoperability, a high likelihood of reliable hardware, and ultimately, mission success.

The purpose of this appendix is to provide the justification for each of the mandatory minimum requirements that are imposed by NASA-STD-6016A. Requirement (“shall”) statements indicate contractually binding requirements that have to be implemented and their implementation verified. In some cases, the justification is obvious in the requirement statement itself. In many others, the justification results from prior failures or successes with a material or a material process, i.e., lessons learned, or based on a best practice. The section/paragraph numbers and titles are also presented for easy reference. NASA-STD-6016A has a significant amount of italicized text that provides additional requirements-related guidance and information in many of the sections.

(Note: Enter “Yes” to describe the requirement’s applicability to the program or project; or enter “No” if the intent is to tailor, and enter how tailoring is to be applied in the “Rationale” column.)

NASA-STD-6016B NOTE: Per section 4.1.6: The use of M&P that do not comply with the requirements of this NASA Technical Standard may be acceptable in the actual hardware applications. MUAs are required for all M&P that are technically acceptable but do not meet the requirements of this NASA Technical Standard, as implemented by the approved Materials and Processes Selection, Control, and Implementation Plan.

70 of 155 APPENDIX A

STANDARD MATERIALS AND PROCESSES REQUIREMENTS FOR SPACECRAFT

REQUIREMENTS COMPLIANCE MATRIX WITH REQUIREMENTS JUSTIFICATIONS

A.1 PURPOSE

NASA-STD-6016B

NOTE: Per section 4.1.6: The use of M&P that do not comply with the technical requirements of this NASA Technical Standard may be acceptable in the actual hardware applications. MUAs are required for all M&P that are technically acceptable but do not meet the technical requirements of this NASA Technical Standard, as implemented by the approved Materials and Processes Selection, Control, and Implementation Plan.

	Section Title	NASA-STD-6016B Requirement Statement	Justification
1.2a	Scope, Applicability	[MPR 1] Programs shall apply these controls to program/project hardware.	It is NASA policy to provide clear, concise, verifiable requirements for processes, products, and components; and it is NASA policy to track and implement requirements using requirements management and configuration management control processes.
1.2b	Scope, Applicability	[MPR 2] Programs shall be responsible for demonstrating compliance with these requirements.	It is NASA policy to implement and track requirements.
1.3a	Scope, Tailoring	[MPR 3] Tailoring of this NASA Technical Standard’s requirements in the Materials and Processes Selection, Control, and Implementation Plan for specific programs/projects shall be formally documented as part of program or project requirements and approved by the responsible program/project NASA M&P organization, the responsible project/program, and the responsible Technical Authority in accordance with NPR 7120.5, NASA Space Flight Program and Project Management Requirements, or NPR 7120.8, NASA Research and Technology Program and Project Management Requirements.	Requirements management processes require that M&P requirement modifications be tracked and approved by responsible and knowledgeable M&P personnel and program organizations.
2.1.1	Applicable Documents, General	[MPR 4] The latest issuances of cited documents shall apply unless specific versions are designated.	Current NASA Office of the Chief Engineer policy is to include this statement in all NASA standards.
2.1.2	Applicable Documents, General	[MPR 5] Non-use of specifically designated versions shall be approved by	Current NASA Office of the Chief Engineer policy is to include this

NASA-STD-6016B

84 of 155

NASA-STD-6016B

4.2.3.9 | Section Title: NASA-STD-6016B Requirement Statement=Glycols, NASA-STD-6016B Requirement Statement=[MPR 99] When solutions containing glycols (aliphatic dihydric alcohols) are used aboard spacecraft that have electrical or electronic circuits containing silver or silver-coated copper, a silver chelating agent, such as benzotriazole (BZT), shall, Justification=Required to prevent this unusual spontaneous ignition reaction.

NASA-STD-6016B

		NASA-STD-6016B Requirement Statement	Justification	Section Title
		the responsible Technical Authority.	statement in all NASA standards.
2.4.2		[MPR 6] Conflict between this NASA Technical Standard and other requirements documents shall be resolved by the responsible Technical Authority.	Current NASA Office of the Chief Engineer policy is to include this statement in all NASA standards.	Applicable Documents, Order of Precedence
4a		[MPR 7] Materials shall meet the worst- case useful-life requirements for the particular application.	Common sense requirement that materials used in an application be assessed for all of the environmental extremes that the materials will encounter during the life cycle.	Requirements
4b		[MPR 8] M&P used in interfacing GSE, test equipment, hardware processing equipment, hardware packaging, and hardware shipment shall be controlled to prevent damage to or contamination of spaceflight hardware.	Ensures spaceflight hardware will not be damaged or contaminated by GSE, test equipment, or other M&P that comes in contact with the spaceflight hardware.	Requirements
4.1.1a		[MPR 9] Each organization that is responsible for the design and fabrication of spaceflight hardware shall provide a Materials and Processes Selection, Control, and Implementation Plan.	The Materials and Processes Selection, Control, and Implementation Plan represents a best practice for implementing, controlling, and verifying that the M&P requirements of this NASA Technical Standard are imposed and tracked through the program/project/element life cycle.	Materials and Processes Selection, Control, and Implementation Plan
4.1.1b		[MPR 10] The Materials and Processes Selection, Control, and Implementation Plan shall also describe the methods used to control compliance with these requirements by subcontractors and vendors.	The Materials and Processes Selection, Control, and Implementation Plan needs to provide to NASA insight into how the M&P requirements are imposed and controlled at the subcontractor and vendor level.	Materials and Processes Selection, Control, and Implementation Plan
4.1.1c		[MPR 11] Upon approval by the procuring activity, the Materials and Processes Selection, Control, and Implementation Plan shall become the M&P implementation document used for verification.	Since the Materials and Processes Selection, Control, and Implementation Plan contains the details of how a company implements and controls the requirements imposed by this NASA Technical Standard, it follows that the Plan is the document to be used to verify that the requirements have been met.	Materials and Processes Selection, Control, and Implementation Plan
4.1.1.1		[MPR 12] The Materials and Processes Selection, Control, and Implementation Plan shall identify the method of coordinating, approving, and tracking all engineering drawings, engineering orders, and other documents that establish or modify materials and/or processes usage.	A process needs to be in place for coordinating, approving, and tracking all M&P usage and M&P changes. The Materials and Processes Selection, Control, and Implementation Plan is the logical place for such a process to be identified.	Coordination, Approval, and Tracking
4.1.1.2		[MPR 13] The Materials and Processes Selection, Control, and Implementation Plan shall include a requirement that all hardware design drawings and revisions for spaceflight hardware that provides mission-critical functions contain an M&P approval block, or equivalent, to ensure that the design has been reviewed by the responsible M&P authority and complies with that document.	The M&P signature block on spaceflight hardware drawings ensures that M&P called out on drawings, and changes thereto, have been approved by M&P knowledgeable personnel.	Approval Signature
4.1.1.3		[MPR 14] The Materials and Processes Selection, Control, and Implementation Plan shall identify how the responsible M&P organization participates in manufacturing and inspection/verification	M&P expertise is required in the selection and processing of materials during manufacturing. How that expertise is applied and utilized needs to be described in the	Manufacturing Planning
		planning to ensure compliance with M&P requirements.	Materials and Processes Selection, Control, and Implementation Plan.
4.1.2a		[MPR 15] All M&P for spaceflight hardware that provides mission-critical functions shall be defined by standards and specifications selected from government, industry, and company specifications and standards.	Manufacturing spaceflight hardware to consensus, Government, or company M&P standards and specifications ensures reliable, repeatable, and trackable processes and spaceflight hardware. Manufacturing spaceflight hardware to standards and specifications has been a mandatory practice for a long time.	M&P Controls
4.1.2b		[MPR 16] All M&P for spaceflight hardware that provides mission-critical functions shall be identified directly on the appropriate engineering drawing.	Identifying all M&P directly on the drawing ensures that the M&P usage has been reviewed and approved. It also ensures that manufacturing follows the correct M&P processes, and that hardware built to print is repeatable.	M&P Controls
4.1.2c		[MPR 17] Company M&P specifications shall be identified by document number in the Materials and Processes Selection, Control, and Implementation Plan.	Ensures that NASA’s oversight responsibility has complete insight into all of the M&P that will be used to manufacture the spaceflight hardware.	M&P Controls
4.1.2d		[MPR 18] All M&P specifications used to produce spaceflight hardware that provides mission-critical functions shall be made available to the responsible NASA Program or Project Office and M&P organization.	NASA’s oversight role of approving and verifying M&P usage on mission-critical spaceflight hardware requires that all standards and specifications be available for review.	M&P Controls
4.1.2e		[MPR 19] Process specifications for spaceflight hardware that provides mission-critical functions shall define	Necessary to ensure a consistent and reliable product.	M&P Controls
		process steps at a level of detail that ensures a repeatable/controlled process that produces a consistent and reliable product.
4.1.2f		[MPR 20] Process qualification shall be conducted to demonstrate the repeatability of all processes for spaceflight hardware that provides mission-critical functions where the quality of the product cannot be directly verified by subsequent monitoring or measurement.	Necessary to ensure a consistent and reliable product.	M&P Controls
4.1.2.1		[MPR 21] A process shall be established for ensuring that updated, alternate, or new material or process standards or specifications from those identified in the Materials and Processes Selection, Control, and Implementation Plan meet or exceed the technical requirements identified in the original material or process standards or specifications.	The use of new material or material process standards or specifications in ongoing programs requires that the changes be reviewed and approved by the M&P and program authorities. A process for ensuring such a review and approval occurs needs to be identified in the Materials and Processes Selection, Control, and Implementation Plan. Also, engineering drawings do not call out company specifications by revision; so changes to such specifications can affect the repeatability of hardware built to print.	Standard and Specification Obsolescence
4.1.3		[MPR 22] A procedure shall be established to ensure that all vendor- designed, off-the-shelf, and vendor- furnished items are covered by the M&P requirements of this document, with the	COTS hardware can present or contain unknown hazards unless the M&P used to manufacture such hardware meets the requirements of this NASA Technical Standard, or	Commercial Off-The- Shelf (COTS) Hardware
		exception that M&P requirements for off- the-shelf hardware and other spaceflight hardware that do not provide mission- critical functions may be limited to those required to ensure safety of flight (for example, flammability, toxic offgassing) and vehicle compatibility (for example, vacuum outgassing).	the COTS M&P usage is reviewed and approved by the responsible NASA M&P and program authorities. A process needs to be established to ensure that COTS hardware is properly reviewed.
4.1.4a		[MPR 23] An M&P control panel shall be established by each contractor hardware provider (excluding subcontractors).	Each hardware provider needs to establish a material control panel to provide quick M&P expertise, review, and approval for hardware manufacturing problems and M&P changes. In addition, the panel provides the NASA M&P oversight responsibility with insight into company M&P implementation and control issues.	M&P Control Panel
4.1.4b		[MPR 24] The M&P control panel’s scope and membership shall be described in the Materials and Processes Selection, Control, and Implementation Plan.	The NASA M&P oversight responsibility requires understanding of who and how the contractor implements and controls requirements.	M&P Control Panel
4.1.5a		[MPR 25] M&P usage shall be documented in an electronic, searchable parts list or separate electronic searchable Materials Identification and Usage List (MIUL) with the following exceptions: (1) Electrical, electronic, and electromechanical (EEE) parts other than wire, cable, and exposed surfaces of connectors.	The MIUL is a best practice that has a long history of use in NASA that provides a quick, reliable, and easy review process for tracking and verifying M&P usage for both NASA and the contractors.	M&P Usage Documentation
		(2) Materials used in hermetically sealed electronic containers (maximum leak rate less than 1 x 10-4 cm3/sec).
4.1.5b		[MPR 26] The documentation of M&P usage shall cover the final design as delivered.	Common sense requirement that the end product reflects the actual M&P utilized in the as-built hardware (including M&P changes resulting from repairs to nonconforming hardware).	M&P Usage Documentation
4.1.5c		[MPR 27] The documentation approach shall be defined in the Materials and Processes Selection, Control, and Implementation Plan.	Ensures documentation of M&P usage and allows organizations to use their own documentation system.	M&P Usage Documentation
4.1.6a		[MPR 28] MUAs containing sufficient information to demonstrate that the application is acceptable shall be submitted by the responsible M&P organization for all M&P that are technically acceptable but do not meet the requirements of this NASA Technical Standard, as implemented by the approved Materials and Processes Selection, Control, and Implementation Plan.	The MUA process is a long used NASA M&P best practice that enables acceptance of materials and material processes that do not meet the requirements of this NASA Technical Standard but, through the MUA review process, are determined to be acceptable. The MUA process also provides the benefit of satisfactorily resolving design non-conformances at the engineering level as opposed to the program level.	Material Usage Agreements (MUAs)
4.1.6b		[MPR 29] MUAs shall not be used to change the M&P requirements for a nonconforming product.	The purpose of the MUA process is to justify the acceptance of specific material or material process design	Material Usage Agreements (MUAs)
			non-conformances, not the acceptance of a nonconforming product.
4.1.7a		[MPR 30] All parts or materials for spaceflight hardware that provides mission-critical functions shall be certified as to composition, properties, and requirements as identified by the procuring document.	Ensures that consistent and reliable materials are used in spaceflight hardware and further ensures that materials and their properties can be repeated from program to program and from company to company.	Materials Certification and Traceability
4.1.7b		[MPR 31] With the exception of off-the-shelf parts, parts and materials used in critical applications, such as life-limited materials and/or safety- and fracture-critical parts, shall be traceable through all processing steps defined in the engineering drawing to the end-item application.	M&P traceability to the end item of critical parts ensures that each processing step was accomplished and that a processing record exists which often aids in understanding a problem or failure.	Materials Certification and Traceability
4.1.7c		[MPR 32] Distributors or other processors shall not heat treat, hot work, or cold work metal stock unless they take the responsibilities as the producer of the metal and produce a new certification.	Required to control procurement of raw metallic materials. The processing changes the material properties and invalidates the original vendor’s certification.	Materials Certification and Traceability
4.1.8.1a		[MPR 33] A, B, or S-basis statistical values for mechanical properties of materials shall be utilized for the design and analysis of hardware for all applications where structural analysis is required.	To ensure structural integrity of spacecraft parts and structures, materials with statistically based minimum mechanical properties are required.	Requirements for Design Values
4.1.8.1b		[MPR 34] The sampling for other mechanical properties, such as dynamic properties like fatigue and fracture, and verification of design values shall be	Needed to ensure design values for other required mechanical properties are representative of the material, product form, and state	Requirements for Design Values
		representative of the material, product form, and state used in the design.	used in the design.
4.1.8.1c		[MPR 35] A, B, or S-basis statistical methods shall be defined by, and values for mechanical properties in their design environment taken from MMPDS, Metallic Materials Properties Development and Standardization, or SAE CMH-17, Composite Materials Handbook.	The handbooks imposed here have long been NASA’s and the aerospace industries’ most trusted and reliable sources of statistically derived mechanical design properties. MMPDS is the replacement for the old MIL- HDBK-5 and SAE CMH-17 is the replacement for the old MIL- HDBK-17. SAE CMH-17-5 is scheduled for publication in 2017 and MIL-HDBK-17-5 remains active until it is released.	Requirements for Design Values
4.1.8.1d		[MPR 36] For metallic materials, the alloy, heat treatment, product specification, product form, and thickness shall match the alloy, heat treatment, product specification, product form, and thickness in MMPDS.	Since the testing to develop design values is specific to alloy, heat treatment, specification and form, the statistical relationship is relevant only to that same combination. Thickness affects strength of metal because of metallurgical factors that influence strength, like heat transfer during heat treatment and variability of reduction during metal working. This requirement is stated in MMPDS but frequently ignored.	Requirements for Design Values
4.1.8.1e		[MPR 37] When statistical design values for new or existing structural materials are not available, they shall be determined by methods described in MMPDS or SAE	The statistical methods used in these handbooks are the NASA- and aerospace industry-accepted standard for determining	Requirements for Design Values
		CMH-17 and a report documenting the derivation of the new design values be made available to NASA for review.	mechanical design properties.
4.1.8.1f		[MPR 38] The implementation of an approach for generating statistical design values that deviates from the sampling methodology and statistical methods in MMPDS or SAE CMH-17 and the use of design values that deviate from those published in MMPDS or SAE CMH-17 shall be approved with an MUA.	Not all materials, product forms, and sizes are available in the cited handbooks. Development of design values by the analytical methods described in MMPDS or SAE CMH-17 is sometimes prohibitively expensive for the application. When the test approach deviates from the published values in MMPDS or the guidance in SAE CMH-17, the MUA review process by responsible and competent M&P engineering personnel assures that acceptable minimum design properties will be derived and verified.	Requirements for Design Values
4.1.8.1g		[MPR 39] All contractor-developed mechanical and physical property data shall be made available to the responsible NASA M&P organization.	NASA needs to have the capability to verify that contractor-developed mechanical and physical property data have been developed adequately and determine when an MUA is required per 4.1.8.1 g. NASA recognizes that such data may be proprietary and that the contractor can place limitations on how the data are made available.	Requirements for Design Values
4.1.8.2a		[MPR 40] Material “B” design values shall not be used except in redundant structure in which the failure of a component would result in a safe	“B” design values are designed for use in redundant structure. They have a less statistically significant minimum design value than “A”	Design Value Implementation in Design
		redistribution of applied loads to other load-carrying members.	design values and are, therefore, less reliable and not to be used in single load path critical structure.
4.1.8.2b		[MPR 41] Minimum property acceptance values in material specifications (specification minimums) not explicitly published as “S” design values in MMPDS shall require an approved MUA for use as design values.	Not all materials, product forms, and sizes are available in the cited handbooks. In some cases, specification minimums or “S” design values published in other sources are not properly derived in accordance with MMPDS. Therefore, the MUA review process by responsible and competent M&P engineering personnel assures that acceptable minimum design properties will be derived and verified.	Design Value Implementation in Design
4.1.8.3		[MPR 42] Structural fastener design values shall be defined by minimum load test requirements in the applicable part and/or procurement specification (government, aerospace industry consensus, company-specific, or custom specification).	MMPDS design values are not transferrable to fasteners, because raw material used to make fasteners is reprocessed using various metallurgical practices, such as hot heading, thread rolling, and heat treating. These processes change the strength of the metal from the original bar stock. NASA has chosen to adopt lot tested design strength defined in the specific fastener part and procurement specifications as the design value instead of the certifications of the original bar stock.	Structural Fastener Design Values
4.2.1		[MPR 43] Materials shall meet the requirements of NASA-STD-6001B,	Ensures that the NASA-required tests and evaluations of materials	Flammability, Offgassing, and
		Flammability, Offgassing, and Compatibility Requirements and Test Procedures, as described below.	for flammability, offgassing, and compatibility are imposed.	Compatibility Requirements
4.2.1.1		[MPR 44] Materials that are nonflammable or self-extinguishing in their use configuration as defined by NASA-STD-6001B, Test 1 or an appropriate configurational flammability test per NASA-STD-6001B, shall be used for flammability control with the following exceptions: a. Ceramics, metal oxides, and inorganic glasses are exempt. b. Materials that are designed to ignite and burn in their use application (for example, pyrotechnic materials) are exempted from flammability requirements, provided that systems/experiments using such materials are designed so that the burning materials cannot act as an ignition source for other hardware. c. Materials used in minor quantities (dimensions controlled so the potential fire propagation path is less than 15 linear cm (6 linear in) and the surface area is no more than 80 cm² (12 square in)) in crew environments and 30 linear cm (12 linear in) for external materials) with no propagation path or ignition source.	The lessons learned from the Apollo 1 fire regarding the use of flammable materials in an oxygen- rich environment makes it mandatory that special control and testing be exercised by M&P engineering to ensure that flammable materials are minimized and controlled in spacecraft environments.	Flammability Control
		d. Materials used in sealed containers are exempt because insufficient oxygen is available to maintain combustion. e. Materials within vented electronics packages with metallic cases and no forced air convection are exempt because of the self-extinguishing effect of expanding combustion gases in a constrained volume. f. Materials that have been shown by test to meet the requirements of NASA- STD-6001B may be used in an environment with an oxygen concentration lower than the test level without further testing (provided that the oxygen partial pressure is no greater than the partial pressure at the test level or general test data exist to demonstrate that the higher oxygen partial pressure is outweighed by the lower percentage concentration for the environments in question). g. Materials are acceptable when used on a metal substrate that provides an adequate heat sink. A heat sink is considered adequate in the use configuration by test or analysis. When testing is conducted, materials passing the flammability test on a metal substrate are acceptable on metal substrates of the same thickness or greater. Materials that
		are flammable but have a thickness less than 0.25 mm (0.010 in) and are attached to a metallic surface greater than 1.6 mm (0.062 in) thick are acceptable by analysis. h. Materials are unexposed, overcoated with a verified fire-blocking material, or no more than 1/4 inch thick and sandwiched between nonflammable materials with only the edges exposed. i. Flammability configuration analysis in accordance with JSC 29353 shows hardware configuration to be self-extinguishing.
4.2.1.2a		[MPR 45] All nonmetallic materials used in habitable flight compartments, with the exception of ceramics, metal oxides, inorganic glasses, and materials used in sealed container (maximum leak rate less than 1 x 10-4 cm3/sec), shall meet the offgassing requirements of Test 7 of NASA-STD-6001B.	All materials offgas small quantities of trace gas contaminants. Materials that release toxic gases must be controlled in closed spacecraft-habitable compartments. The required testing ensures that the materials used are safe from a toxicity standpoint.	Toxic Offgassing
4.2.1.2b		[MPR 46] Spacecraft maximum allowable concentration (SMAC) values shall be obtained from JSC 20584, Spacecraft Maximum Allowable Concentrations for Airborne Contaminants, or from MAPTIS for compounds for which no SMAC values are found in JSC 20584.	Industrial maximum concentrations of airborne contaminants are based on an 8-hour workday exposure and are not appropriate for the continuous exposure that applies in spacecraft environments. The SMAC values in JSC 20584 are set by the NASA/JSC Toxicology	Toxic Offgassing
			Group in cooperation with the National Research Council Committee on Toxicology. Supplementary values in MAPTIS are set by the NASA/JSC Toxicology Group only for evaluation of offgassing test data.
4.2.1.3a		[MPR 47] Materials exposed to hazardous fluids shall be evaluated or tested for compatibility.	Group in cooperation with the National Research Council Committee on Toxicology. Supplementary values in MAPTIS are set by the NASA/JSC Toxicology Group only for evaluation of offgassing test data.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3b		[MPR 48] Appropriate long-term tests to simulate the worst-case use environment that would enhance reactions or degradation shall be conducted for materials with long-term exposure to liquid fuels, oxidizers, and other hazardous liquids.	Since some incompatible material/fluid combinations only manifest themselves after long-term exposure, it is necessary to conduct long-term exposure testing.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3c		[MPR 49] Materials degradation in long- term tests shall be characterized by post- test analyses of the material and fluid to determine the extent of changes in chemical and physical characteristics, including mechanical properties.	To determine in long- or short-term tests if material degradation has occurred, post-test analyses must be performed to determine, e.g., if mechanical properties have changed or if unacceptable levels of the test material have been dissolved into the test fluid.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3d		[MPR 50] The effect of material condition (for example, parent versus weld metal or heat-affected zone) shall be addressed in the compatibility determination.	It has long been known that a material’s condition, such as a weld, its heat-affected zone, and the parent metal, can have significantly different outcomes in corrosion/compatibility testing	Fluid Compatibility (Fluids Other Than Oxygen)
			which needs to be addressed when determining compatibility and any necessary testing.
4.2.1.3e		[MPR 210] Materials used in nitrogen tetroxide systems shall be evaluated for flammability in nitrogen tetroxide using promoted combustion tests similar to NASA-STD-6001B, Test 17, for metallic materials used in oxygen systems and Test 1 for polymeric materials.	Nitrogen tetroxide is a powerful oxidizer, and flight hardware failures have been caused by ignition of metallic materials (notably titanium).	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3(1)		[MPR 211] When materials are determined to be flammable, a nitrogen tetroxide compatibility assessment shall be conducted using the methodology described for oxygen systems in NASA- STD-6001B and the system safety rationale of this assessment documented in an MUA.	The rigorous Oxygen Compatibility Assessment (OCA) methodology described in NASA-STD-6001 for evaluating potential materials ignition in oxygen systems is also the most effective methodology for evaluating ignition in nitrogen tetroxide systems.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3f		[MPR 212] Materials used in other oxidizer systems shall be evaluated for potential ignition.	The possibility of similar ignition issues occurring in systems containing oxidizers other than oxygen and nitrogen tetroxide needs to be addressed.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.4a		[MPR 51] Materials and components, and systems used in liquid oxygen (LOX) and gaseous oxygen (GOX) environments, compressed air systems, and pressurized systems containing enriched oxygen shall be evaluated for oxygen compatibility in accordance with NASA-STD-6001B.	Because of the highly reactive nature of oxygen with metals and nonmetals, it is necessary that all materials that are in contact with or will come in contact with oxygen must be evaluated for compatibility with oxygen. NASA-STD-6001B governs how such evaluation is to be accomplished.	Oxygen Compatibility
4.2.1.4b		[MPR 52] When materials in such	Materials are commonly found to	Oxygen Compatibility
		systems are determined to be flammable, an oxygen compatibility assessment shall be conducted as described in NASA- STD-6001B and the system safety rationale of this assessment documented in an MUA.	be flammable in standard flammability tests but can often be found not to be a flammability hazard in their use configuration. Such a determination is made by conducting an oxygen compatibility assessment and documenting the acceptance rationale in an MUA.
4.2.1.4.1		[MPR 53] GOX and enriched air system components that operate at pressures above 1.83 MPa (265 psia), with the exception of metallic components, such as hard lines (rigid metallic tubing), metallic flex hoses, metallic fluid fitting with all metal seals, and metallic pressure vessels (including composite overwrapped pressure vessels with metallic liners), shall undergo oxygen component acceptance testing at maximum design pressure for a minimum of ten cycles from ambient pressure to maximum design pressure within 100 milliseconds to ensure that all oxygen system spaceflight hardware is exposed to oxygen prior to launch.	Ensures that material surfaces that can react with oxygen at relatively high pressures have been thoroughly exposed and conditioned to the system oxygen pressure and will be nonreactive to oxygen during use.	Oxygen Component Acceptance Test
4.2.1.5a		[MPR 54] Electrical wire insulation materials shall be evaluated for flammability in accordance with NASA- STD-6001B, Test 4 or Test 1.	The lessons learned from the Apollo 1 fire and the Apollo 13 oxygen tank explosion make it imperative that the flammability of wire insulation be evaluated. NASA-STD-6001 specifies how to conduct such tests.	Electrical Wire Insulation Materials
4.2.1.5b		[MPR 55] Arc tracking shall be evaluated in accordance with NASA-STD-6001B, Test 18, for all wire insulation constructions except polytetrafluoroethylene (PTFE), PTFE/polyimide insulation conforming to SAE AS22759C, Wire, Electrical, Fluoropolymer-Insulated, Copper or Copper Alloy, and used in conditions bounded by the dry arc-tracking test specified by SAE AS22759C, ethylene tetrafluoroethylene (ETFE), and silicone- insulated wires (the resistance of these materials to arc tracking has already been established).	Since arc tracking is a problem that is due to a breakdown of a wire’s insulating material that results in uncontrolled arcing to other wires and components, it is necessary to test wire insulating materials for their susceptibility to arc tracking.	Electrical Wire Insulation Materials
4.2.2		[MPR 56] MSFC-STD-3029A, Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments, shall be used to select metallic materials to control stress corrosion cracking of metallic materials in sea and air environments, with the exception that an MUA is not required for the following conditions: a. Parts in electrical/electronic assemblies with maximum tensile stress less than 50 percent of the yield strength. b. Martensitic or precipitation-hardening (PH) stainless steels used in ball bearing or similar applications where the loading is compressive.	Stress corrosion cracking in seacoast and air environments is a potential problem for the high- strength alloys typically used in spacecraft structures. MSFC-STD- 3029 provides data on the stress- corrosion sensitivity of most structural alloys and contains requirements to eliminate the potential for stress-corrosion cracking when stress-corrosion- sensitive materials are used. The standard also contains test methods for evaluating stress-corrosion sensitivity of alloys without existing data and defines stress corrosion cracking threshold values for many of the alloys used in spacecraft structures.	Metals
		c. Carbon and low alloy high-strength steels with strength greater than 1240 MPa (180 ksi) used in ball bearings, springs, or similar applications where the loading is compressive and there is a history of satisfactory performance. d. Hardware provides no mission-critical functions
4.2.2.1 4.2.2.2.1a		[MPR 57] The 5000-series alloys containing more than 3 percent magnesium shall not be used in spaceflight hardware that provides mission-critical functions where the temperature exceeds 66 °C (150 °F). [MPR 58] When drilling, grinding, reaming, or machining is performed on high-strength steels that can form martensite and are used in spaceflight hardware that provides mission-critical functions, a validated process for machining shall be used.	Under certain conditions, an intergranular aluminum-magnesium grain boundary precipitate can occur rendering such alloys susceptible to exfoliation corrosion or stress corrosion. A well-known lesson learned is that some metal removal processes, particularly aggressive ones, cause surface overheating in high-strength steels that produces the undesirable crack-prone condition of untempered martensite.	Aluminum Drilling and Grinding of High-Strength Steel
4.2.2.2.1b		[MPR 59] The absence of machining damage for high-strength steels used in spaceflight hardware that provides mission-critical functions shall be verified by microexamination of production parts (such as Nital etch inspection) or by a microhardness and metallurgical examination of sample parts for either of the following situations:	Tests and examinations cited are best practices for identifying the presence of untempered martensite.	Drilling and Grinding of High Strength Steel
		(1) When the material has very low toughness (such as martensitic steel above 200 ksi) and the part has low stress margins or is fatigue driven. (2) When the surface condition of the part is critical to the design (such as its ability to withstand hertzian stresses or remain perfectly flat).
4.2.2.2.2a		[MPR 60] Unstabilized, austenitic steels shall not be processed or used above 371 °C (700 °F) in spaceflight hardware that provides mission-critical functions.	Eliminates the potential of a corrosion-susceptible condition due to grain boundary precipitation of chromium carbide.	Corrosion-Resistant Steel
4.2.2.2.2b		[MPR 61] Welded assemblies used in spaceflight hardware that provides mission-critical functions shall be solution heat-treated and quenched after welding except for the stabilized or low carbon grades, such as 321, 347, 316L, and 304L.	Eliminates the potential of a corrosion-susceptible condition of weld heat-affected zone grain boundary precipitation of chromium carbide.	Corrosion-Resistant Steel
4.2.2.2.2c		[MPR 62] Service-related corrosion issues are common for free-machining alloys, such as 303, and these alloys shall not be used in spaceflight hardware that provides mission-critical functions when they can be exposed to moisture other than transient condensation, or to nitrogen tetroxide	Eliminates the exposure of the more corrosion-susceptible stainless steel alloys to moisture or to nitrogen tetroxide, for which corrosion of these alloys has also been observed.	Corrosion-Resistant Steel
4.2.2.2.3		[MPR 63] Steels shall not be used in tension below their ductile-brittle transition temperature.	Prevents the use of alloys that are too impact- and crack-sensitive and prone to unexpected failure at low	Ductile-Brittle Transition Temperature
			temperature.
4.2.2.3.1a		[MPR 64] All cleaning fluids and other chemicals used during manufacturing and processing of titanium hardware for spaceflight hardware that provides mission-critical functions shall be verified to be compatible and not detrimental to performance before use.	An Apollo Program lesson learned occurred when a titanium pressure vessel failed due to stress corrosion in anhydrous methanol that was used as a substitute pressurant for Aerozine-50 in a ground test. Methanol, a commonly used solvent for cleaning metals, was considered benign to titanium.	Titanium Contamination
4.2.2.3.1b		[MPR 65] The surfaces of titanium and titanium alloy mill products used for spaceflight hardware that provides mission-critical functions shall be 100 percent machined, chemically milled, or pickled to a sufficient depth to remove all contaminated zones and layers formed while the material was at elevated temperature.	A lesson learned from a titanium pressure vessel that was severely cracked from an alpha case that formed during heat treatment.	Titanium Contamination
4.2.2.3.2		[MPR 66] All regions of titanium alloys for spaceflight hardware that provides mission-critical functions that are subject to fretting or wear shall be anodized per SAE AMS2488D, Anodic Treatment – Titanium and Titanium Alloys Solution, pH 13 or Higher, or hard-coated utilizing a wear-resistance material, such as tungsten carbide/cobalt thermal spray.	Since titanium is subject to wear and fretting which can lead to crack initiation, it is necessary to harden the titanium surfaces by anodizing per the cited specification or per a hard-coating process.	Titanium Wear
4.2.2.3.3a		[MPR 67] Titanium alloys shall not be used with LOX or GOX at any pressure or with air at oxygen partial pressures above 35 kPa (5 psia).	Titanium is highly flammable in oxygen-rich environments, and its use in such environments must be avoided.	Titanium Flammability
4.2.2.3.3b		[MPR 68] Titanium alloys shall not be	Because of titanium’s sensitivity to	Titanium Flammability
		machined inside spacecraft modules during ground processing or in flight, because machining operations can ignite titanium turnings and cause fire.	ignition in oxygen-containing environments, fine turnings and particles with large surface areas from machining and grinding processes can even ignite in air. Obviously, such fires in a spacecraft must be prevented.
4.2.2.4a		[MPR 69] Magnesium alloys shall not be used in primary structure or in other areas of spaceflight hardware that provides mission-critical functions that are subject to wear, abuse, foreign object damage, abrasion, erosion, or where fluid or moisture entrapment is possible.	Magnesium’s low strength and hardness renders it relatively sensitive to mechanical damage and abuse and is relatively corrosion sensitive. Therefore, its use in primary structure needs to be avoided.	Magnesium
4.2.2.4b		[MPR 70] Magnesium alloys shall not be machined inside spacecraft modules during ground processing or in flight, because machining operations can ignite magnesium turnings and cause fire.	Magnesium can be ignited in air at room temperature from machining operations. Obviously, such fires in a spacecraft must be prevented.	Magnesium
4.2.2.5a		[MPR 71] Alloys containing more than 4 percent beryllium by weight shall not be used for primary structural applications.	Beryllium alloys have an inherent low ductility and brittle nature, and their use in primary structure needs to be avoided.	Beryllium
4.2.2.5b		[MPR 72] Alloys containing more than 4 percent beryllium by weight shall not be used for any application within spacecraft crew compartments unless suitably protected to prevent erosion or formation of salts or oxides.	Beryllium containing alloys, even those containing as little as 4 percent beryllium, present a problem in that oxides and salts of beryllium are highly toxic and the use thereof requires that the metal be protected from abrasion or conditions that could form	Beryllium
4.2.2.5c		[MPR 73] Design of beryllium parts for spaceflight hardware that provides mission-critical functions shall address its low-impact resistance and notch sensitivity, particularly at low temperatures, and its directional material properties (anisotropy) and sensitivity to surface finish requirements.	beryllium particles, beryllium oxides, or salts. When beryllium is required to be used in mission-critical spaceflight hardware, it is important that the low ductility and brittle nature of beryllium be critically reviewed. When analyzing the stress state in beryllium, it is important to recognize the low Poisson’s Ratio of beryllium. This low Poisson’s Ratio can lead to very high stresses.	Beryllium
4.2.2.5d		[MPR 74] All beryllium parts used in spaceflight hardware that provides mission-critical functions shall be processed to ensure complete removal of the damaged layer (twins and microcracks) produced by surface-metal-working operations, such as machining and grinding.	When beryllium is required to be used in mission-critical hardware, the detrimental surface damage that occurs in beryllium from mechanical metal removal methods, such as machining and grinding, must be removed. The damage removal must be by chemical etch to remove the surface damage. The technique of removing successively shallower cuts will reduce the depth of the damage but will never completely eliminate the damage. The nature of the damage is a crystal twinning in the grains which pins dislocations and subsequently reduces material ductility.	Beryllium
4.2.2.5e		[MPR 75] Beryllium-containing alloys (including alloys containing less than 4	Due to the highly toxic nature of beryllium dust and oxides formed	Beryllium
		percent beryllium by weight) and oxides of beryllium shall not be machined inside spacecraft crew compartments at any stage of manufacturing, assembly, testing, modification, or operation.	during machining or grinding, such operations must be highly controlled and never allowed to occur in spacecraft crew compartments.
4.2.2.5f		[MPR 76] All beryllium parts used in spaceflight hardware that provides mission-critical functions shall be penetrant-inspected for crack-like flaws with a high-sensitivity fluorescent dye penetrant in accordance with section 4.2.5.	Ensures that the crack-sensitive beryllium parts are crack free.	Beryllium
4.2.2.6a		[MPR 77] Cadmium shall not be used in crew or vacuum environments.	Because of its toxic nature and relatively high vapor pressure, the use of cadmium in spacecraft crew compartments and in the space vacuum environment needs to be prevented.	Cadmium
4.2.2.6b		[MPR 78] Cadmium-plated tools and other hardware shall not be used in the manufacture or testing of spaceflight hardware that provides mission-critical functions.	Cadmium is relatively soft and could be readily transferred to spacecraft parts through contact between the parts and cadmium- plated tools. Therefore, the use of such tools must be prevented.	Cadmium
4.2.2.7		[MPR 79] Owing to zinc’s ability to grow whiskers, zinc plating other than black zinc-nickel plating shall not be used in spaceflight hardware that provides mission-critical functions.	The use of zinc-plated parts in spaceflight hardware needs to be prevented because zinc has a tendency to grow whiskers that cause electrical shorts and are also a source of contamination. Zinc also has a relatively high vapor pressure and, in the vacuum of space, it presents the risk of	Zinc
4.2.2.8a		[MPR 80] Equipment containing mercury shall not be used where the mercury could come in contact with spaceflight equipment during manufacturing, assembly, test, checkout, and flight.	volatilizing and redepositing on critical surfaces. Mercury in contact with some metals can result in severe embrittlement; therefore, such contact must be prevented.	Mercury
4.2.2.8b		[MPR 81] Spaceflight hardware (including fluorescent lamps) containing mercury shall have three levels of containment to prevent mercury leakage.	Because of the high toxicity of mercury vapor and liquid mercury’s potential for causing embrittlement of other metals, its safe containment in a spacecraft environment must be extraordinarily redundant.	Mercury
4.2.2.9		[MPR 82] For refractory alloys (alloys with a melting point above 2000 °C (3600 °F), plus osmium and iridium) used in mission-critical applications, tests shall be performed to characterize critical design properties for the intended application and the data documented in an MUA.	Refractory metals are used in high- temperature applications where the material properties at the intended use-temperature are lacking or limited. Testing must be performed to obtain the necessary design properties for the application. Documentation of the design properties is also required.	Refractory Metals
4.2.2.10a		[MPR 83] High-nickel content alloys are susceptible to sulfur embrittlement; therefore, for spaceflight hardware that provides mission-critical functions, any foreign material which could contain sulfur, such as oils, grease, and cutting lubricants, shall be removed by suitable means prior to heat treatment, welding, or high temperature service.	Since superalloys are susceptible to sulfur embrittlement from contact with sulfur containing materials at elevated temperatures, it is mandatory that such materials be removed from the surface of the superalloys before they are exposed to high temperatures.	Superalloys (Nickel- Based and Cobalt- Based)
4.2.2.10b		[MPR 84] The reduction to design	Since superalloys experience alloy	Superalloys (Nickel-
		properties of alloying element depletion at the surface in a high temperature, oxidizing environment shall be evaluated when a thin sheet of one of these alloys is used for spaceflight hardware that provides mission-critical functions, since a slight amount of depletion could involve a considerable proportion of the effective cross section of the material.	depletion at elevated temperatures, it is mandatory that the depth of such depletion on a superalloy’s design properties be assessed, particularly in the case of thin sheets.	Based and Cobalt- Based)
4.2.2.11a		[MPR 85] For spaceflight hardware that provides mission-critical functions, tin and tin plating shall be alloyed with at least 3 percent lead by weight or other proven alloying element(s) to prevent tin whisker growth.	Self-explanatory, in that 3 percent lead in tin and tin plating is necessary to prevent tin-whisker growth. Several other alloying elements in tin and tin plating have also been shown to prevent tin whiskers.	Tin
4.2.2.11b		[MPR 86] Tin and tin plating alloyed with less than 3 percent lead by weight and used in electrical/electronic applications shall comply with GEIA-STD-0005-1A, Performance Standard for Aerospace and High Performance Electronic Systems Containing Lead-Free Solder, and Control Level 2C requirements of GEIA-STD- 0005-2A, Standard for Mitigating the Effects of Tin Whiskers in Aerospace and High Performance Electronic Systems, with the following exceptions: (1) Solder alloy Sn96.3Ag3.7 (Sn96) used for high- temperature applications. (2) Solder alloy Au80Sn20 used as	Since the cited standards specifically address how to mitigate the tin whisker problems for tin and tin-plating containing less than 3 percent lead, it is necessary that they be used for applications that use tin and tin-plating with less than 3 percent lead.	Tin
		a die attach material or as a package sealing material. (3) Tin alloys containing less than 20 percent tin by weight.
4.2.2.11c		[MPR 87] When high-purity tin and tin plating are used for spaceflight hardware that provides mission-critical functions and will also be exposed to temperatures below 13 ˚C for periods longer than 6 months, the method for preventing tin pest formation shall be documented in the materials control plan.	Since tin pest is a problem that results from tin and tin plating being exposed to a low temperature for a long period of time (as can occur for spacecraft external materials), it is necessary to address how the tin pest problem will be controlled. Tin pest has been observed on ISS external hardware.	Tin
4.2.2.11d		[MPR 88] Tin plating shall not be used for contacts in electrical interconnects (connectors, sockets, switches, etc.) for spaceflight hardware that provides mission-critical functions.	Because of the many problems that have been encountered with tin plating, such as oxidation, wear, fretting, and whiskers, it is necessary to limit the use of tin plating in spaceflight hardware unless the potential problems are suitably mitigated and its use approved through an MUA review process.	Tin
4.2.3.1a		[MPR 89] Elastomeric materials used in spaceflight hardware that provides mission-critical functions shall be selected to operate within design parameters for the useful life of the hardware.	Since elastomers have age-life limitations, it is necessary to select elastomeric materials that can function and endure the space environment for the expected life of the hardware.	Elastomeric Materials
4.2.3.1b		[MPR 90] Elastomeric materials used in spaceflight hardware that provides mission-critical functions, other than	The age-life of elastomers begin when they are cured; therefore, because of age-life limitations, their	Elastomeric Materials
		those used in off-the-shelf parts, such as cable clamps, shall be cure-dated for tracking purposes.	cure dates need to be recorded so the ages of the elastomers can be tracked.
4.2.3.1c		[MPR 91] Room temperature vulcanizing (RTV) silicones that liberate acetic acid during cure shall not be used because they can cause corrosion.	Requirement justifies itself, in that the release of a corrosive acid is unacceptable.	Elastomeric Materials
4.2.3.1d		[MPR 92] When rubbers or elastomers are used at low temperatures in spaceflight hardware that provides mission-critical functions, the ability of these materials to maintain and provide required elastomeric properties shall be verified.	The lesson learned from the Challenger O-ring failure due to its inelastic behavior at a moderately cold temperature strongly emphasizes the need to verify the low temperature elastic properties of rubbers and elastomers. Fluorocarbon elastomers are particularly susceptible to embrittlement.	Elastomeric Materials
4.2.3.2		[MPR 93] Use of polyvinyl chloride on spaceflight hardware shall be limited to applications in pressurized areas where temperatures do not exceed 49 °C (120 °F).	Polyvinyl chloride (PVC) cannot be used in space vacuum because it contains large quantities of plasticizers that outgas in space vacuum, making the PVC brittle and acting as a source of contamination. PVC also offgasses excessively above 120 oF, so its use in hardware that experiences such temperatures must be restricted. PVC is flammable in spacecraft cabin environments, and its use is subject to the same controls as other flammable materials.	Polyvinyl Chloride
4.2.3.5		[MPR 94] All materials shall be selected to meet the useful life of the hardware	If a material cannot meet the useful life requirements of the hardware of	Limited-Life Items
		with no maintenance or be identified as limited-life items requiring maintainability.	which it is a part, then it must be identified as having limited-life and a maintenance plan identified to address its limited-life.
4.2.3.6a		[MPR 95] Nonmetallic materials that are exposed to space vacuum, with the exception of ceramics, metal oxides, inorganic glasses, and cetyl alcohol lubricants used on fasteners outside closed compartments, shall be tested using the technique of ASTM E595-15, Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment, with acceptance criteria as follows: (1) ≤0.1 percent collected volatile condensable materials (CVCM). (2) ≤1.0 percent total mass loss (TML) less water vapor recovery (WVR), except that a higher mass loss is permitted if this mass loss has no effect on the functionality of the material itself and no effect on the functionality of any materials, components, or systems that could be adversely affected by the subject mass loss.	Since nonmetallic materials outgas in the vacuum of space, it is necessary to quantify the nature and amount of outgassing that can occur. ASTM E595 is the consensus standard for conducting such tests and quantifying mass loss and CVCM. The traditional ≤1.0 percent total mass loss (TML) spacecraft requirement is relaxed in this standard. Vacuum outgassing causes contamination only from condensable materials and the TML is unimportant except in special cases where cryogenic surfaces are present and in cases where the mass loss causes unacceptable degradation of the material. The vast majority of materials with a TML ≤5 percent are acceptable for normal spacecraft applications, provided the ≤0.1 percent CVCM requirement is met.	Thermal Vacuum Stability
4.2.3.6b		[MPR 96] With the following exceptions, hardware items (components, assemblies, etc.) containing materials that fail the	Since vacuum baking can sufficiently outgas materials that fail the CVCM requirements	Thermal Vacuum Stability
		CVCM requirement and/or having unidentified materials shall be vacuum baked at the maximum tolerable temperature of the component, 10 ℃ above the maximum predicted operating temperature, or an alternate temperature selected by the program/project, to meet the program/project acceptance outgassing criteria: (1) Materials that are not near a critical surface and have a CVCM between 0.1 and 1.0 percent and an exposed surface area less than 13 cm² (2 in²) are exempt. (2) Materials with an exposed surface area less than 1.6 cm² (0.25 in²) are exempt. (3) Materials that are unexposed, overcoated, or encapsulated with approved materials are exempt. (4) Materials enclosed in a sealed container (maximum leak rate less than 1 x 10⁻⁴ cm³/sec) are exempt.	established by this NASA Technical Standard, it is necessary to vacuum bake hardware items that contain materials that do not meet the CVCM requirements.
4.2.3.7		[MPR 97] The critical properties of materials exposed in the spacecraft external environment shall meet	The spacecraft external environment includes ultraviolet radiation, atomic oxygen in low-	External Environment Survivability
		operational requirements for their intended life-cycle exposure.	Earth orbit, plasma, and high- energy radiation, all of which can cause materials degradation. For example, ultraviolet radiation causes darkening of many materials, leading to loss of thermal control, and atomic oxygen erodes most polymeric materials, causing physical degradation. Since repair capabilities are very limited or nonexistent, it is imperative that materials so exposed meet their operational life cycle.
4.2.3.8		[MPR 98] Materials that are non-nutrient to fungi, as identified in MIL-HDBK- 454B, General Guidelines for Electronic Equipment, Requirement 4, Fungus-Inert Materials, Table 4-I, Group I, shall be used in launch vehicles and pressurized flight compartments, except when one of the following criteria is met: a. Materials have been tested to demonstrate acceptability per MIL-STD- 810G, Department of Defense Test Method Standard for Environmental Engineering Considerations and Laboratory Tests, Method 508. b. Materials are used in crew areas where fungus would be visible and easily removed.	Materials that resist the growth of fungus are required in launch vehicles that are exposed to weather environments and in pressurized habitable compartments. The requirement statement identifies many exceptions where such controls are unnecessary.	Fungus Resistance
4.2.3.9	Section Title	Glycols	Required to prevent this unusual spontaneous ignition reaction.
		be added to the solution to prevent spontaneous ignition from the reaction of silver with the glycol.
4.2.3.10a		[MPR 100] The etching of PTFE, perfluoroalkoxy (PFA), and fluorinated ethylene propylene (FEP) shall meet the requirements of SAE AMS2491F, Surface Treatment of Polytetrafluoroethylene, Preparation for Bonding, when adhesion to the fluorocarbon surface is required, except that for insulated wire or cable a pull test on co-produced specimens may be performed in lieu of the tensile and shear strength tests in AMS2491 section 3.5.2.	It is NASA policy to use voluntary consensus standards where possible. The cited specification is the industry standard for preparing polytetrafluoroethylene for bonding and has been reviewed and approved for use by NASA.	Etching Fluorocarbons
4.2.3.10b		[MPR 101] Etched surfaces shall be processed within 24 hours or within 1 year if packaged per SAE AMS2491F.	Ensures that the etched surface remains active and ready for bonding.	Etching Fluorocarbons
4.2.4.1a		[MPR 102] Heat treatment of aluminum alloys used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS2772G, Heat Treatment of Aluminum Alloy Raw Materials; SAE AMS2770N, Heat Treatment of Wrought Aluminum Alloy Parts; or SAE AMS2771E, Heat Treatment of Aluminum Alloy Castings.	It is NASA policy to use voluntary consensus standards where possible. The cited standards are the widely accepted consensus standards for heat treating aluminum alloys, and they have been reviewed and approved for use by NASA.	Heat Treatment
4.2.4.1b		[MPR 103] Heat treatment of steel alloys used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS-H-6875B, Heat Treatment of Steel Raw Materials,	It is NASA policy to use voluntary consensus standards where possible. The cited standards are the widely accepted consensus standards for heat treating steel	Heat Treatment
		or SAE AMS2759E, Heat Treatment of Steel Parts, General Requirements.	alloys, and they have been reviewed and approved for use by NASA.
4.2.4.1c		[MPR 104] Heat treatment of titanium alloys used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS-H- 81200D, Heat Treatment of Titanium and Titanium Alloys, for raw stock and SAE AMS2801B, Heat Treatment of Titanium Alloy Parts, for parts requiring heat treatment during fabrication.	It is NASA policy to use voluntary consensus standards where possible. The cited standards are the widely accepted consensus standards for heat treating titanium alloys, and they have been reviewed and approved for use by NASA.	Heat Treatment
4.2.4.1d		[MPR 105] Heat treatment of nickel- and cobalt-based alloy parts used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS2774E, Heat Treatment, Wrought Nickel Alloy and Cobalt Alloy Parts, or SAE AMS2773E, Heat Treatment, Cast Nickel Alloy and Cobalt Alloy Parts.	It is NASA policy to use voluntary consensus standards where possible. The cited standards are the widely accepted consensus standards for heat treating nickel- and cobalt-based alloys and have been reviewed and approved for use by NASA.	Heat Treatment
4.2.4.1e		[MPR 106] For spaceflight hardware that provides mission-critical functions, process-control tensile-test coupons shall be taken from the production part (or from the same material lot, having the same thickness as and processed identically to the production part) to verify the adequacy of the heat treatment process for the following conditions: (1) Aluminum alloys are solution	Heat treat variations can compromise the strength and ductility of the alloys for the heat- treat conditions cited. In addition, the commonly used hardness test to verify the adequacy of the heat treat process is not a good predictor of strength and ductility.	Heat Treatment
		heat-treated. (2) High-strength steels (>200 ksi (1380 MPa) UTS), tool steels, and maraging steel alloys are heat-treated to high strength levels. (3) A286 or MP35N alloys (which have poor correlation between hardness and tensile strength) are heat treated. (4) Titanium alloys are annealed or solution heat treated and aged. (5) Nickel- and cobalt-based alloys are work strengthened before age hardening resulting in age- hardened tensile strengths greater than 1030 MPa (150 ksi) UTS. (6) Precipitation hardenable nickel- and cobalt-based alloys are solution heat treated.
4.2.4.1f		[MPR 107] When process-control tensile- test coupons are required, the requirement for the coupons shall be specified on the engineering drawing for the part.	Ensures compliance.	Heat Treatment
4.2.4.1g		[MPR 108] If no tensile values are available in MMPDS for a specific alloy, tensile-test acceptance values shall be specified on the engineering drawing for	Ensures that accept/reject criteria are established.	Heat Treatment
		the part.
4.2.4.1h		[MPR 109] Materials shall not be used in spaceflight hardware that provides mission-critical functions outside the limits of their procurement specification, heat treat specification, or MMPDS specification.	Ensures that the materials that are used have defined and dependable design properties.	Heat Treatment
4.2.4.2a		[MPR 110] Where forgings are used in mission-critical applications, first-article (preproduction) approval shall be obtained from the procuring authority.	Ensures that a NASA M&P technical expert(s) has reviewed and approved the first-article cut-up process.	Forging
4.2.4.2b		[MPR 111] First-article approval and the controls to be exercised in producing subsequent production forgings shall be in accordance with SAE AMS2375D, Control of Forgings Requiring First Article Approval.	It is NASA policy to use voluntary consensus standards where possible. The cited standard is the widely accepted consensus standard for control of the first-article forging cut-up and has been reviewed and approved for use by NASA.	Forging
4.2.4.2c		[MPR 112] After the forging technique, including degree of working, is established, the first production forging shall be sectioned to show the grain-flow patterns and to determine mechanical properties at control areas and the trim ring/protrusion specimens (prolongations).	Ensures that the forging process has broken up the cast structure and that expected microstructure and mechanical properties are produced by the forging process.	Forging
4.2.4.2d		[MPR 113] The mechanical properties of the trim ring/protrusion specimens (prolongations) for the first article shall be compared to the control coupons to show they are predictive of the properties	The mechanical properties of the trim ring/protrusion specimens (prolongations) for subsequent articles will be used to verify that the mechanical properties of	Forging
		in the body of the first article.	subsequent articles are acceptable. As such, they have to be representative of the properties in the body of each subsequent article. For the first article, they are compared with the mechanical properties of the control coupons from the body of the article to verify that they are truly representative.
4.2.4.2e		[MPR 114] Sectioning to show the grain- flow patterns and to determine mechanical properties at control areas shall be repeated after any substantive change in the forging technique, as determined by M&P analysis.	Since any changes in the forging process can significantly alter the grain flow, microstructure, and mechanical properties, it is necessary to conduct a first-article- like cut-up to ensure the resulting properties are acceptable.	Forging
4.2.4.2f		[MPR 115] These data and results of tests on the redesign shall be retained and made available for review by the procuring activity.	Since forging process changes can significantly alter the mechanical and metallurgical properties of the forging, it is necessary to conduct a first-article-like cut-up and have the results reviewed and approved by NASA M&P engineering.	Forging
4.2.4.2g		[MPR 116] Trim ring or protrusion specimens (prolongations) shall be obtained for each production forging used in safety-critical applications and tested for required minimum mechanical properties.	It is standard practice in the forging industry to obtain and test trim ring or protrusion specimens to ensure that the mechanical properties of each forging are achieved.	Forging
4.2.4.2h		[MPR 117] Surface and volumetric nondestructive inspection (NDI) shall be performed on all safety-critical forgings.	Obviously, the severe plastic deformation involved in a forging process can introduce both surface and internal discontinuities and defects.	Forging

	Section Title	NASA-STD-6016B Requirement Statement	Justification and internal defects for which inspection is needed.
4.2.4.3a	Castings	[MPR 118] Castings used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS2175A (2010), Castings, Classification and Inspection of.	It is NASA policy to use voluntary consensus standards. SAE AMS2175 is the industry specification for classifying and inspecting castings and has been reviewed and approved for use by NASA.
4.2.4.3b	Castings	[MPR 119] Where castings are used in mission-critical applications, pre- production castings shall be subjected to first-article inspection to verify proper material flow, proper material integrity, minimum required mechanical properties, proper grain size, and macro/microstructure.	For mission-critical castings, it is considered mandatory to subject the first casting(s) to destructive testing to verify the mechanical properties and metallurgical characteristics of the casting process.
4.2.4.3c	Castings	[MPR 120] The mechanical properties in trim ring/protrusion of the first article shall be compared to the control coupons to show they are predictive of the properties in the body of the first article.	Trim rings/protrusions will be obtained for all production castings so the first article testing needs to verify that they are representative of the cast parts.
4.2.4.3d	Castings	[MPR 121] The same casting practice and heat-treating procedure shall be used for the production castings as for the approved first-article castings.	Ensures to a reasonable degree that the production castings will have the same mechanical properties and metallurgical characteristics as the first-article casting(s) that validated the casting process.
4.2.4.3e	Castings	[MPR 122] For Class 1 and Class 2 castings (classes as defined by SAE AMS2175A (2010)), mechanical property testing of integrally cast or excised tensile bars at critical locations shall be	Tensile bars cast at the same time as the production casting(s) ensure that the casting process is reproducing the same mechanical properties. Integrally casting tensile

NASA-STD-6016B

	NASA-STD-6016B Requirement Statement	Justification	Section Title
	conducted to ensure foundry control of cast lots.	bars with each production lot is a standard practice in the casting industry.
4.2.4.3f	[MPR 123] Periodic cut-ups or functional testing shall be conducted for Class 1 and Class 2 castings (classes as defined by SAE AMS2175A (2010)).	For the most critical castings, periodic cut-ups or functional testing ensures that the casting process is under control and is reproducible.	Castings
4.2.4.3g	[MPR 124] Surface and volumetric nondestructive inspection shall be performed on all safety-critical castings.	Castings frequently contain unacceptable defects that cannot be detected except by these NDE techniques.	Castings
4.2.4.4a	[MPR 125] Where formed panels are used in mission-critical applications, pre- production panels shall be subjected to first-article inspection to verify proper material integrity, minimum required mechanical properties, proper grain size, and macro/microstructure.	The process for forming panels needs to be verified as producing panels with proper material integrity, mechanical properties, grain size, and macro/microstructure.	Formed Panels
4.2.4.4b	[MPR 126] The mechanical properties of the first production article shall be compared to control coupons to show they are predictive of the properties in the body of the first article.	Control coupons will be used for verification of the properties of each production formed panel, so the first article testing needs to verify that they are representative of the formed panels.	Formed Panels
4.2.4.4c	[MPR 127] The same forming practice and heat-treating procedure shall be used for the production formed panels as for the approved first-article panels.	Ensures to a reasonable degree that the production formed panels will have the same mechanical properties and metallurgical characteristics as the first-article formed panel that validated the process.	Formed Panels
4.2.4.4d	[MPR 128] Sectioning to determine mechanical properties at control areas shall be repeated after any substantive change in the forming technique, as determined by M&P analysis.	Since any changes in the forming process can significantly alter the grain flow, microstructure, and mechanical properties of the panel, it is necessary to conduct a first- article-like cut-up to ensure the resulting properties are still acceptable.	Formed Panels
4.2.4.4e	[MPR 129] Surface and volumetric NDI shall be performed on all safety-critical formed panels.	Formed panels may contain unacceptable defects that cannot be detected except by these NDE techniques.	Formed Panels
4.2.4.5a	[MPR 130] Structural adhesive bonding shall meet MSFC-SPEC-445A, Adhesive Bonding, Process and Inspection, Requirements for.	MSFC-SPEC-445 is the NASA- wide, accepted specification for adhesive bonding and inspection.	Adhesive Bonding
4.2.4.5b	[MPR 131] Structural adhesive bonding processes shall be controlled to prevent contamination that would cause structural failure that could affect the safety of the mission, crew, or vehicle or affect mission success.	The strength of a structural adhesive bond can be significantly degraded by bond surface contamination, particularly from silicones. Therefore, it is imperative that bond surface cleanliness be maintained and verified.	Adhesive Bonding
4.2.4.5c	[MPR 132] Bonded primary structural joints shall demonstrate cohesive failure modes in shear.	In bonded joints, cohesive failure is the desired failure mode that best ensures a reliable and repeatable shear strength design value, whereas an adhesive failure is almost always the result of inadequate surface preparation and is indicative of an improper bonding process.	Adhesive Bonding
4.2.4.6a	[MPR 133] If alternative specifications to	The AWS D17 welding	Welding
	those cited in this section are utilized or developed, those specifications shall meet the requirements of NASA-STD-5006A, General Welding Requirements for Aerospace Materials.	requirements are commonly used by aerospace companies, but are far more detailed than is required by NASA and also have some deficiencies. If an aerospace company chooses to not follow the AWS D17 requirements, plus the supplemental requirements identified in this document, they may develop their own standard, provided that it complies with the less detailed but more stringent requirements in NASA-STD-5006.
4.2.4.6b	[MPR 134] Material Review Board disposition shall be required for weld repair/rework/processing activities that are not in accordance with the approved weld process specification (WPS).	Weld procedure variances that have not been qualified by prior test or experience need to be reviewed and approved by a Material Review Board and/or M&P knowledgeable personnel.	Welding
4.2.4.6c	[MPR 135] A weld development and certification plan shall be developed for large structural welded components such as crew modules and welded cryogenic tanks.	The welding of large structural welded components can have major problems if the weld development and certification plan isn’t established well before welding flight hardware. The plan is expected to include full scale pathfinder weldments, tooling development, and a design values program including sensitivity testing.	Welding
4.2.4.6.1	[MPR 136] The processing and quality assurance requirements for manual, automatic, and semiautomatic welding for	It is NASA policy to use voluntary consensus standards where possible. The cited specification is	Fusion Welding
	spaceflight applications that provide mission-critical functions shall meet the requirements of AWS D17.1/D17.1M (2010) AMD 1 (2012), Specification for Fusion Welding for Aerospace Applications, with the following modifications/additions:	the aerospace industry specification for fusion welding of aerospace structures and components and has been reviewed and approved for use by NASA with modifications/additions to address minor deficiencies in the baseline standard.
4.2.4.6.1a	[MPR 137] Mission-critical structural welds shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012), Class A requirements.	Structural welds that are safety- critical are required to meet the most stringent set of requirements, as is the case for the cited Class A requirements. Note that the weld classes in AWS D17.1 are based on nondestructive inspection requirements (with the most stringent inspection requirements for Class A welds), whereas traditional NASA weld classes were based on weld criticality (with Class A welds being the highest criticality level).	Fusion Welding
4.2.4.6.1b	[MPR 138] Other structural welds shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012), Class A or Class B requirements.	Structural welds that are not safety- critical are allowed to meet a less stringent set of requirements, as is the case for the cited Class B requirements. The more stringent Class A requirements may also be used.	Fusion Welding
4.2.4.6.1c	[MPR 139] Non-structural welds (including seal welds) shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012), Class C requirements.	Non-structural, non-safety-critical welds are allowed to meet the least stringent set of requirements, as is the case for the cited Class C	Fusion Welding

	Section Title	NASA-STD-6016B Requirement Statement	Justification requirements.
4.2.4.6.1d	Fusion Welding	[MPR 140] All Class A and Class B welds (including manual welds), as defined by AWS D17.1/D17.1M (2010) AMD 1 (2012), shall be qualified in accordance with AWS D17.1/D17.1M (2010) AMD 1 (2012).	Ensures that Class A and B manual weld processes are also qualified, not just the operators who perform the welds.
4.2.4.6.1e	Fusion Welding	[MPR 141] Titanium welds shall be light/dark straw or better (Ref. AWS D17.1/D17.1M (2010) AMD 1 (2012), Table 7.1).	Titanium welds with worse coloration than light/dark straw (as listed in AWS D17.1, Table 7.1) indicate improper inert gas control during welding and presents the possibility that weld contamination/oxidation occurred.
4.2.4.6.1f	Fusion Welding	[MPR 142] Titanium and its alloys shall be welded with alloy-matching or metallurgically compatible fillers or autogenously.	Prevents the creation of a completely unknown titanium alloy in the weldment by welding together, for example, an alpha alloy and an alpha-beta alloy or other incompatible combinations of titanium alloys.
4.2.4.6.1g	Fusion Welding	[MPR 143] Extra low interstitial (ELI) filler wires shall be used for titanium cryogenic applications and are preferred for general applications.	ELI weld wire produces a more ductile weld at cryogenic temperatures.
4.2.4.6.1h	Fusion Welding	[MPR 144] Commercially pure (CP) titanium filler shall not be used on Ti- 6Al-4V or other alloyed base material.	Welding Ti-6Al-4V with CP weld wire results in a significant alloy content and interstitial solubility difference between the parent metal and the weld metal, which can lead to interstitial migration and precipitation of embrittling species, such as titanium hydride, at the

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	Justification	Section Title	NASA-STD-6016B Requirement Statement
	weld fusion line. An Apollo SIV-B stage was destroyed due to inadvertently welding a Ti-6Al-4V pressure vessel with CP weld wire.
4.2.4.6.1i	Titanium welding needs to be performed in an inert gas environment. The cited gases are not considered inert to titanium and can react with titanium alloys at weld temperatures.	Fusion Welding	[MPR 145] Nitrogen, hydrogen, carbon dioxide, and mixtures containing these gases shall not be used in welding titanium and its alloys.
4.2.4.6.1i(1)	Ensures that the water content in the inert gas is low enough that the potential reaction of water with titanium at the welding temperature is not a problem.	Fusion Welding	[MPR 146] The inert gas shall have a dew point of -60º C (-76º F) or lower.
4.2.4.6.1j	Welding of titanium alloys may produce residual stresses that are high enough that the weld needs to be stress relieved. Such stress relief treatments are done at high enough temperatures that precautions against oxidation of the titanium must be exercised.	Fusion Welding	[MPR 147] Welded alpha and alpha plus beta titanium alloys shall be stress relieved in a vacuum or inert gas environment (Ar or He), or stress relieved in air with verification of oxide removal per SAE AMS-H-81200D or SAE AMS 2801B, or certified in the as-welded condition.
4.2.4.6.1k	Titanium beta alloys are less prone to residual stress during welding, but stress relieving of beta welds still needs to be considered.	Fusion Welding	[MPR 148] Titanium beta alloys that are welded shall be evaluated on a case-by- case basis with respect to stress relief.
4.2.4.6.1l	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are industry standards for laser welding that have been reviewed and approved for use by NASA.	Fusion Welding	[MPR 149] Laser welding for spaceflight hardware that provides mission-critical functions shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012) or AWS C7.4/C7.4M (2008), Process Specification and Operator Qualification
			for Laser Beam Welding.
4.2.4.6.1m	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are industry standards for electron beam welding that have been reviewed and approved for use by NASA, whereas SAE AMS2681 is another common industry standard for electron beam welding but is not considered adequate.	Fusion Welding	[MPR 150] Electron beam welding for spaceflight hardware that provides mission-critical functions shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012) or SAE AMS2680C, Electron- Beam Welding for Fatigue Critical Applications.
4.2.4.6.1n	Although permitted by AWS D17.1/D17.1M (2010) AMD 1 (2012), these practices have been determined to be unacceptable for NASA Program hardware unless approved through an MUA review process.	Fusion Welding	[MPR 151] The following welding practices permitted by AWS D17.1/D17.1M (2010) AMD 1 (2012) shall not be used without an approved MUA to document the acceptance rationale: (1) Welding from both sides if full penetration of the first pass is not verified (either by inspection of the back side or by grinding prior to welding on the opposite side). (2) Partial weld penetration in structural welds. (3) Straightening operation after welding. (4) Lap welds in structural applications.
4.2.4.6.1o	NASA review of the cited AWS specification has determined that these additional requirements need to be included in the WPS to ensure the adequacy and completeness of the welding process.	Fusion Welding	[MPR 152] The Welding Procedure Specification (WPS) shall include the following content in addition to that required by AWS D17.1/D17.1M (2010) AMD 1 (2012): (1) Prequalified rework welds in accordance with AWS D17.1/D17.1M (2010) AMD 1 (2012). (2) Testing and documentation of allowable parameter variations for automatic and semi- automatic welds. (3) Manual welding parameters. (4) An associated Procedure Qualification Record (PQR) with tension testing and macro- examination results as part of weld qualification requirements.
4.2.4.6.2	It is NASA policy to use voluntary consensus standards where possible. The cited specification is an industry standard for resistance welding that has been reviewed and approved for use by NASA. Resistance welding is not used frequently in NASA programs and then primarily for spot welding of batteries.	Resistance Welding	[MPR 153] Resistance welding for spaceflight hardware that provides mission-critical functions, including resistance spot welding (RSW), shall meet the requirements of AWS D17.2/D17.2M (2013), Specification for Resistance Welding for Aerospace Applications.
4.2.4.6.3	It is NASA policy to use voluntary consensus standards where possible	Friction-Stir Welding of	[MPR 154] Friction-stir welding of aluminum alloys that are
	consensus standards where possible. The cited specification is an industry standard for friction-stir welding that has been reviewed and approved for use by NASA.	Aluminum Alloys	aluminum alloys for spaceflight hardware that provides mission-critical functions shall meet the requirements of AWS D17.3/D17.3M (2016), Specification for Friction Stir Welding of Aluminum Alloys for Aerospace Applications.
4.2.4.6.4	Although this standard is no longer active, it has been determined by NASA that it is the best standard existing for establishing and controlling the inertia welding process. Other active inertia friction welding specifications were reviewed and found to be inadequate for NASA’s use or were only recommended practices, not requirements.	Inertia Welding	[MPR 155] Inertia welding for spaceflight hardware that provides mission-critical functions shall meet the requirements of MIL-STD-1252 (1975), Inertia Friction Welding Process, Procedure and Performance Qualification.
4.2.4.6.4a	Generally, inertia welds are performed on pressure-containing tube and valve components that require inspection for surface and subsurface defects that could develop into leaks.	Inertia Welding	[MPR 156] Surface inspection (penetrant) and volumetric inspection (radiography) shall be performed.
4.2.4.6.4b	Generally, inertia welds are performed on pressure-containing tube and valve components that need to be proof tested.	Inertia Welding	[MPR 157] All welds shall be proof tested.
4.2.4.6.4c	Generally, inertia welds are performed on pressure-containing tube and valve components that need to be helium leak tested.	Inertia Welding	[MPR 158] Inertia welds used in fluid systems shall be helium leak tested.
4.2.4.7a	It is NASA policy to use voluntary consensus standards where	Brazing	[MPR 159] Brazing for spaceflight hardware that provides mission-critical
	possible. The cited specification is the industry specification for the design, manufacture, and examination of critical brazed components and has been reviewed and approved for use by NASA.		functions shall be conducted in accordance with AWS C3.3 (2008), Recommended Practices for Design, Manufacture, and Examination of Critical Brazed Components.
4.2.4.7b	It is NASA policy to use voluntary consensus standards where possible. The cited specification is an industry specification for brazing of aluminum that has been reviewed and approved for use by NASA.	Brazing	[MPR 160] Brazing of aluminum alloys for spaceflight hardware that provides mission-critical functions shall meet the requirements of AWS C3.7M/C3.7 (2011), Specification for Aluminum Brazing.
4.2.4.7c	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are the industry specifications for the types of brazing cited, and they have been reviewed and approved for use by NASA.	Brazing	[MPR 161] Torch, induction, and furnace brazing for spaceflight hardware that provides mission-critical functions shall meet the requirements of AWS C3.4M/C3.4 (2016), Specification for Torch Brazing; AWS C3.5M/C3.5 (2016), Specification for Induction Brazing; and AWS C3.6M/C3.6 (2016), Specification for Furnace Brazing, respectively.
4.2.4.7d	Need to ensure that high- temperature processing near existing braze joints does not degrade the joint by re-melting and reflowing the braze metal.	Brazing	[MPR 162] Subsequent fusion-welding operations in the vicinity of brazed joints or other operations involving high temperatures that might affect the brazed joint shall be prohibited for spaceflight hardware that provides mission-critical functions unless it can be demonstrated that the fixturing, processes, methods, and/or procedures employed will preclude
			degradation of the brazed joint.
4.2.4.7e	Since transferring all of the load in a structural braze joint by tension through the braze material is not considered as reliable as transferring the load by shear through the braze material, tension loaded braze material has to be avoided. This is accomplished by using lap joint designs versus butt joints, where the large brazed surface area in a lap joint results in a relatively low shear stress in the braze material.	Brazing	[MPR 163] Brazed joints used in spaceflight hardware that provides mission-critical functions shall be designed for shear loading and not be relied upon for strength in axial loading for structural parts.
4.2.4.7f	It is NASA policy to use voluntary consensus standards where possible. The cited standard is an industry standard for how to evaluate the strength of brazed joints and has been reviewed and approved for use by NASA.	Brazing	[MPR 164] The shear strength of brazed joints used in spaceflight hardware that provides mission-critical functions shall be evaluated in accordance with AWS C3.2M/C3.2 (2008), Standard Method for Evaluating the Strength of Brazed Joints.
4.2.4.7g	Destructive testing ensures the adequacy of the braze flow and that the braze alloy and its adhesion to the brazed surfaces are continuous and metallurgically acceptable.	Brazing	[MPR 165] For furnace brazing of complex configurations of spaceflight hardware that provides mission-critical functions, such as heat exchangers and cold plates, destructive testing shall be conducted on pre-production brazed joints to verify that the braze layer that extends beyond the fillet area is continuous and forms a uniform phase.
4.2.4.8	The low strength and poor creep resistance of solders means	Structural Soldering	[MPR 166] Soldering shall not be used for structural applications.
	soldering cannot be used in structural applications.
4.2.4.9a	Since the high surface temperature associated with the EDM and LM machining processes causes undesirable surface damage, it is necessary to control the processes to ensure the depth of damage is predictable and manageable.	Electrical Discharge Machining and Laser Machining	[MPR 167] Electrical-discharge machining (EDM) and laser machining (LM) processes for spaceflight hardware that provides mission-critical functions shall be controlled to limit the depth of the oxide layer, the recast layer, and the heat-affected zone.
4.2.4.9a(1)	The oxide layer that is produced by the EDM and LM machining processes is generally brittle and crack-prone and must be removed.	Electrical Discharge Machining and Laser Machining	[MPR 168] The oxide layer shall be removed from the surface.
4.2.4.9a(2)	The recast layer and the heat-affected zone produced by the EDM and LM machining processes are generally undesirable conditions, and need to be removed from wear surfaces and fracture-critical surfaces.	Electrical Discharge Machining and Laser Machining	[MPR 169] In addition, the recast layer and the heat-affected zone shall be removed from bearing, wear, fatigue or fracture-critical surfaces, and from crack- or notch-sensitive materials.
4.2.4.9b	Qualification of the EDM and LM machining processes ensures that the depth of the detrimental surface conditions is predictable and the amount of material that must be removed is known.	Electrical Discharge Machining and Laser Machining	[MPR 170] EDM/LM schedules for spaceflight hardware that provides mission-critical functions shall be qualified to determine the maximum thickness of the affected layers when the depth of the affected material must be known for removal or analysis.
4.2.4.10a	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are industry standards for electroplating nickel that have been reviewed and approved for use by	Nickel Plating	[MPR 171] Electrodeposited nickel plating for spaceflight hardware that provides mission-critical functions shall be applied according to the requirements of SAE AMS2403N, Plating, Nickel General Purpose; SAE AMS2423E,
	NASA.		Plating, Nickel Hard Deposit; or ASTM B689-97, Standard Specification for Electroplated Engineering Nickel Coatings.
4.2.4.10b	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are industry standards for electroless nickel plating that have been reviewed and approved for use by NASA.	Nickel Plating	[MPR 172] Electroless nickel plate for spaceflight hardware that provides mission-critical functions shall be applied per SAE AMS2404G, Plating, Electroless Nickel, or ASTM B733-15, Standard Specification for Autocatalytic (Electroless) Nickel-Phosphorus Coatings on Metal.
4.2.4.10c	Since there is a strong galvanic couple between nickel and aluminum, the interface between them must be protected in corrosive environments.	Nickel Plating	[MPR 173] The nickel-aluminum interface in nickel-plated aluminum used in spaceflight hardware that provides mission-critical functions shall be protected from exposure to corrosive environments.
4.2.4.11	Ensures that the NASA M&P technical authority reviews and approves this relatively new manufacturing technique that is not yet controlled by NASA or consensus standards or specifications.	Additive Manufacturing	[MPR 174] When structural hardware is manufactured by additive manufacturing techniques, a manufacturing and qualification plan shall be submitted to NASA and approved by the responsible NASA M&P and design organizations.
4.2.5.1a	Since there are many NDE techniques that are available to inspect spaceflight hardware and such inspections are performed throughout the life cycle of the hardware, a Plan is needed to	Nondestructive Evaluation (NDE) Plan	[MPR 175] The NDE Plan shall address the process for establishment, implementation, execution, and control of NDE through design, manufacturing, operations, and maintenance of spaceflight hardware.
	ensure that the application and sequence of the inspections are proper and efficient.
4.2.5.1b	Obviously, since NASA-STD-5009 sets NDE requirements for fracture- critical components, it has to be included in the NDE Plan for spaceflight hardware. Although MIL-HDBK-6870 states that it is for guidance only, it is to be treated as a requirements document for non-fracture-critical components in the NDE Plan.	Nondestructive Evaluation (NDE) Plan	[MPR 176] The Plan shall meet the intent of MIL-HDBK-6870B, Nondestructive Inspection Program Requirements for Aircraft and Missile Materials and Parts, and, when fracture control is applicable, the requirements of NASA-STD-5009, Nondestructive Evaluation Requirements for Fracture-Critical Metallic Components.
4.2.5.1c	NAS-410 has long been the aerospace industry and NASA standard for certifying and qualifying NDE personnel. Although ASNT-TC-1A is also widely used for certifying and qualifying nondestructive test personnel, it is a recommended practice only.	Nondestructive Evaluation (NDE) Plan	[MPR 177] Qualification and certification of personnel involved in nondestructive testing shall comply with NAS 410 (Revision 4), NAS Certification and Qualification of Nondestructive Test Personnel.
4.2.5.2a	Machining metal surfaces causes a disturbed, smeared thin layer of metal that masks cracks, and the disturbed layer must be removed by etching prior to penetrant inspection.	NDE Etching	[MPR 178] All machined or otherwise mechanically disturbed surfaces on metallic parts that are to be fluorescent dye-penetrant inspected shall be adequately etched to assure removal of smeared, masking material prior to penetrant application, with the following exceptions: (1) Previously etched parts do not need etching if the surface has
			not been smeared since the last etching. (2) When supporting rationale is provided, close tolerance parts may be machined near-final and etched and penetrant inspected before final machining in lieu of etching and penetrant inspecting after final machining.
4.2.5.2b	Since the disturbed metal layer has a certain depth, it is required that at least an amount of metal equal to the thickness of the disturbed layer be removed to have an effective penetrant inspection.	NDE Etching	[MPR 179] The etching procedure shall specify the minimum amount of material to be removed to ensure that smeared metal does not mask cracks.
4.2.5.2c	Since not being able to etch means that the penetrant inspection process would be ineffective, it is necessary to employ other NDE methods, such as eddy current or ultrasonics that can detect the critical flaw size.	NDE Etching	[MPR 180] If etching is not feasible, it shall be demonstrated that the required flaw size can be reliably detected.
4.2.6.1a	Many metal processes produce residual stresses that can add to the stresses from the normally applied service stresses resulting in a higher stress state that needs to be addressed and accounted for in the analyses. An important lesson was learned in	Residual Stresses	[MPR 181] Estimates of residual stresses in structural or stress-corrosion-sensitive hardware shall be considered in structural analyses and corrosion/stress-corrosion assessments.
	repair welding of 2195 Aluminum Lithium. The repair weld material was very low strength compared to the parent material (it was 4043 weld wire in 2195 parent material), so the effect of the residual stress was not apparent (i.e., there wasn’t much warpage). However, when wide specimens containing only a short repair-welded area were tensile tested, the residual stress added to the discontinuity shear stress at the weld/parent metal interface, causing early failure.
4.2.6.1b	Since straightening of a warped part requires plastic deformation and creates the potential for mechanical damage and high residual stresses, it is necessary for M&P engineering to review and approve the process.	Residual Stresses	[MPR 182] The straightening of warped parts in structural hardware shall require an approved MUA.
4.2.6.2a	Vents on vented sandwich constructions must be large enough to prevent the pressure differential between the sandwich interior and the external environment during launch becoming large enough to cause face-sheet separation.	Sandwich Assemblies	[MPR 183] For vented sandwich architectures, the differential pressure on ascent shall be adequately relieved to preclude core-to-face sheet bond line “peel” failure modes that could result in face-sheet separation.
4.2.6.2b	Non-vented sandwich structure will develop a pressure differential during ascent of one atmosphere (14.7 psia) between the sandwich interior and space vacuum. The structure must be verified to	Sandwich Assemblies	[MPR 184] Sandwich architectures that are not vented shall be capable of withstanding pressure buildup without violating strength and stability requirements.
	withstand this pressure differential without causing face-sheet separation.
4.2.6.2c	Since moisture can accumulate in vented core sandwich assemblies and can migrate from cell to cell presenting a corrosion problem, and since it can also cause structural failure in a launch environment where the exterior pressure decrease and the rapid temperature increase can flash the water to vapor causing over pressurization of the face sheet, it is necessary to prevent such water entrapment and migration in sandwich assemblies. An Apollo lesson learned occurred when a vented core honeycomb interstage tank structure failed due to moisture accumulation and flash evaporation during launch. Water intrusion can occur from condensation / exposure of sandwich constructions to humid air below the dew point, “cryo- pumping” where sandwich structures are exposed to large thermal / humidity gradients induced by cryogenic fluid storage, gravitational migration of rain water, forced penetration of wind- driven rain, etc.	Sandwich Assemblies	[MPR 185] Structural sandwich assemblies shall be designed to prevent the entrance and entrapment of water vapor or other contaminants into the core structure.
4.2.6.2d	The use of metallic or glass- reinforced cores to minimize	Sandwich Assemblies	[MPR 186] Structural honeycomb sandwich assemblies that will be
	intrusion of moisture intrusion into the sandwich core and can result in vapor build-up over-pressurizing the sandwich assembly when it is heated by the launch environment.		subjected to heating shall be tested for the expected environments to show that the construction can withstand them.
4.2.6.2e	To prevent moisture accumulation that can cause internal corrosion and structural failure, sandwich assemblies need to be protected from water intrusion prior to launch in addition to their being designed to prevent moisture intrusion.	Sandwich Assemblies	[MPR 187] Sandwich assemblies using perforated and moisture-absorbing cores shall be protected from water intrusion during assembly and prelaunch activities.
4.2.6.2f	It is NASA policy to use voluntary consensus standards where possible. The cited standard is an industry-accepted standard for sandwich construction and core materials that has been reviewed and approved for use by NASA.	Sandwich Assemblies	[MPR 188] Test methods for structural sandwich constructions and core materials shall meet the requirements of SAE AMS-STD-401 (1999), Sandwich Constructions and Core Materials: General Test Methods.
4.2.6.3	Ensures that corrosion of space spaceflight hardware will be addressed and eliminated to the extent possible.	Corrosion Prevention and Control	[MPR 189] All parts, assemblies, and equipment used in spaceflight hardware that provides mission-critical functions, including spares, shall be finished to provide protection from corrosion in accordance with the requirements of NASA-STD-6012, Corrosion Protection for Space Flight Hardware, with the following exceptions: a. SAE AMS2404G is permitted for electroless nickel plating as an alternative to ASTM B733-15.
			b. Titanium fasteners may be used in contact with graphite composites, provided that they are wet installed with sealant or primer materials.
4.2.6.3.1	Ensures that the normally present passive layer that provides corrosion resistance is restored.	Passivation	[MPR 190] Corrosion-resistant steels used in spaceflight hardware that provides mission-critical functions shall be passivated after machining.
4.2.6.3.2	Ensures that corrosive environments cannot intrude into sealed compartments.	Sealing	[MPR 191] Removable panels and access doors in exterior or interior corrosive environments shall be sealed either by mechanical seals or by separable, faying- surface sealing.
4.2.6.4a	Ensures that the potential for hydrogen embrittlement of metallic structures containing hydrogen will be evaluated.	Hydrogen Embrittlement	[MPR 192] When designing liquid or gaseous hydrogen systems, the degradation of metallic materials properties by hydrogen embrittlement shall be addressed in the Materials and Processes Selection, Control, and Implementation Plan.
4.2.6.4b	Ensures that the potential for hydrogen embrittlement of metallic structures containing hydrogen has been evaluated.	Hydrogen Embrittlement	[MPR 193] An MUA shall be written rationalizing the selection of metallic materials for liquid or gaseous hydrogen systems to preclude cracking and to ensure system reliability and safety.
4.2.6.4c	Ensures that the potential for hydrogen embrittlement is minimized by controlling the time of exposure to electrochemical processes and to acids and bases or that the embrittlement relief treatments are carried out soon after exposure to such processes and	Hydrogen Embrittlement	[MPR 194] Electrochemical processes or exposure to acids or bases during manufacturing or processing of spaceflight hardware that provides mission-critical functions shall be controlled to prevent hydrogen embrittlement, or embrittlement relief treatment shall be performed promptly
	fluids.		after processing.
4.2.6.4c(1)	Baking ensures relief of the potential for hydrogen embrittlement in steels after exposure to acids or plating processes. The baking process in SAE AMS2759/9D has a long history and is the accepted way to relieve the potential for hydrogen embrittlement in steels.	Hydrogen Embrittlement	[MPR 195] When acid cleaning baths or plating processes are used on steel parts for spaceflight hardware that provides mission-critical functions, the part shall be baked in accordance with SAE AMS2759/9D, Hydrogen Embrittlement Relief (Baking) of Steel Parts, to alleviate potential hydrogen embrittlement problems.
4.2.6.5.1	A lesson learned when indium solder in a Goddard satellite contacted gold wires and the resulting gold-indium intermetallic resulted in severe embrittlement and failure of the gold wires.	Gold-Indium Intermetallic Formation	[MPR 196] Gold shall not be used in contact with indium or an indium alloy (such as indium solder) for spaceflight hardware that provides mission-critical functions.
4.2.6.5.2a	Ensures that the dissolution of gold into tin-based solder will be minimized, thereby minimizing the formation of brittle gold-tin intermetallic compounds.	Gold-Tin Intermetallic Formation	[MPR 197] When used with tin-based solder joints in mission-critical hardware, gold shall be removed from at least 95 percent of the surface to be soldered of all component leads, component terminations, and solder terminals.
4.2.6.5.2b	Ensures that the propensity for fretting corrosion between gold- coated contacts and mating tin- coated contacts will be avoided.	Gold-Tin Intermetallic Formation	[MPR 198] Gold-coated contacts shall not be mated with tin or tin alloy-coated contacts for separable contact interfaces in mission-critical hardware.
4.2.6.5.3a	An aluminum-rich intermetallic phase, the AuAl2 intermetallic known as “Purple Plague,” is inherent (and not necessarily harmful) to gold-aluminum brazing; however, if excessive heat is applied, too much of the	Gold-Aluminum Intermetallic Formation	[MPR 199] Gold-aluminum brazing processes shall be controlled to minimize formation of the AuAl2 intermetallic known as “Purple Plague” and prevent formation of the Au5Al2 intermetallic known as “White Plague.”
	intermetallic can form at the braze/part interface, causing joint failure. A gold-rich phase, the Au5Al2 intermetallic known as “White Plague,” is always detrimental. Its formation is catalyzed by silicon, so care should be taken to keep the braze joint zone free of contamination
4.2.6.5.3b	intermetallic can form at the braze/part interface, causing joint failure. A gold-rich phase, the Au5Al2 intermetallic known as “White Plague,” is always detrimental. Its formation is catalyzed by silicon, so care should be taken to keep the braze joint zone free of contamination	Gold-Aluminum Intermetallic Formation	[MPR 200] Gold-aluminum bonding processes shall be controlled to prevent the formation of the AuAl2 intermetallic known as “Purple Plague” and the Au5Al2 intermetallic known as “White Plague.”
4.2.6.6.1	Imposing NASA-STD-5020 ensures that issues and problems associated with the use of liquid- locking compounds will be avoided.	Liquid Locking Compounds	[MPR 201] If a liquid-locking compound is used as a locking feature where rotational loosening or disengagement would result in a critical or catastrophic hazard, its use shall comply with the design and quality requirements and best
			practices in NASA-STD-5020, Requirements for Threaded Fastening Systems in Spaceflight Hardware, sections 5.5 and 7.6 and Appendix C.
4.2.6.6.2	Ensures that the friable black oxide that forms on silver exposed to atomic oxygen will not contaminate other components or mechanisms or contaminate astronaut suits and gloves during EVAs.	Silver-Plated Fasteners	[MPR 202] Silver-plated fasteners shall not be used in external applications where the silver plating is directly exposed to atomic oxygen for a period longer than 2 weeks.
4.2.6.7a	Spacecraft performance can be severely degraded by the presence of excessive contamination, leading to loss of mission or mission objectives. A Contamination Control Plan covering the entire hardware life cycle is necessary to ensure that a “build clean” approach is properly implemented. The cited ASTM standard is a voluntary consensus standard that has been reviewed and approved for use by NASA.	Contamination Control	[MPR 203] A Contamination Control Plan shall be generated in accordance with the guidelines of ASTM E1548 (2009), Standard Practice for Preparation of Aerospace Contamination Control Plans.
4.2.6.7b	The cited areas of material processing, hardware packaging, and FOD prevention are the most important types of contamination that need to be controlled and addressed in the Contamination Control Plan.	Contamination Control	[MPR 204] The Contamination Control Plan shall include controls on contamination-sensitive manufacturing processes, such as adhesive bonding, controls on packaging for shipment and storage, and a foreign object damage/debris (FOD) prevention program.
4.2.6.7c	Since damage to spaceflight hardware from FOD can be very	Contamination Control	[MPR 205] The FOD prevention program shall be established for all ground
	expensive and impact schedules, it is mandatory that a FOD prevention program be established that addresses all sources and means of preventing FOD damage throughout the hardware’s life cycle.		operations of mechanical and electrical systems of spaceflight hardware, including the design, development, manufacturing, assembly, repair, processing, testing, maintenance, operation, and checkout of the equipment to ensure the highest practical level of cleanliness.
4.2.6.7d	It is NASA policy to use voluntary consensus standards where possible. NAS 412 is the aerospace industry standard for FOD prevention and has been reviewed and approved for use by NASA.	Contamination Control	[MPR 206] The FOD prevention program shall conform to NAS 412 (Revision 1), Foreign Object Damage/Foreign Object Debris (FOD) Prevention.
4.2.6.7e	Ensures that the cleanliness level of each hardware item has been specified.	Contamination Control	[MPR 207] Cleanliness levels for assembly- and subassembly-level hardware shall be identified on the engineering drawings.
4.2.6.8	Obviously, spaceflight hardware needs to be protected during shipment and storage.	Packaging	[MPR 208] Packaging shall protect spaceflight hardware from corrosion and contamination during shipping and storage.
4.3	Since it is NASA policy to provide clear, concise, verifiable requirements, it is necessary for NASA M&P to have oversight and insight into how the contractors and their subs and vendors implement and control spaceflight hardware M&P. The cited steps of reviewing and approving the contractors’ M&P plans, requiring M&P drawing signature approval,	Verification	[MPR 209] Verification of compliance with the requirements of this NASA Technical Standard shall consist of the following steps as a minimum: a. NASA approval of the contractor Materials and Processes Selection, Control, and Implementation Plan and other applicable materials data requirements documents, such as the Contamination Control Plan and NDE
	conducting audits, membership on the materials control panel, and approval of MUAs and MIULs provide NASA M&P with the oversight and insight necessary to verify that the requirements of NASA-STD-6016 have been met.		Plan. b. Contractor M&P signature on engineering drawings to verify compliance with the requirements of this NASA Technical Standard or the Materials and Processes Selection, Control, and Implementation Plan. c. NASA audits of contractor M&P activities relating to hardware design and manufacturing. d. Establishment and operation of the M&P control panel in accordance with section 4.1.4 of this NASA Technical Standard. e. NASA approval of MUAs. f. NASA approval of MIULs.

NASA-STD-6016B APPENDIX B TYPICAL MUA FORM

MATERIAL USAGE AGREEMENT C: <empty> USAGE AGREEMENT NO.: <empty> REV: <empty> PAGE: <empty> OF: <empty>

PROJECT:: <empty> SYSTEM:: <empty> CATEGORY:: <empty> ORIGINATOR:: <empty> ORGANIZATION/CONTRACTOR:: <empty>

PART NUMBER(S):: <empty> USING ASSEMBLY(S):: <empty> ITEM DESCRIPTION:: <empty> ISSUE:: <empty>

MATERIAL DESIGNATION:: <empty> MANUFACTURER:: <empty> SPECIFICATION:: <empty> PROPOSED EFFECTIVITY:: <empty>

MATERIAL CODE:: <empty> LOCATION:: <empty> ENVIRONMENT:: <empty>

THICKNESS:: <empty> WEIGHT:: <empty> EXPOSED AREA: <empty> HABITABLE [ ] NONHABITABLE [ ] PRESSURE PSIA:: <empty> TEMP.F:: <empty> MEDIA:: <empty>

APPLICATION:: <empty>

RATIONALE: (use second page if required.): <empty>

MATERIAL USAGE AGREEMENT DISPOSITION CONTRACTOR TIER 1: <empty> CONTRACTOR PRIME: <empty> NASA PROJECT MGR.: <empty> NASA M & P: <empty> DATE: <empty> APPROVE: <empty> REJECT: <empty> DEFER: <empty> MEMO NO.:: <empty>

DATE: <empty> APPROVE: <empty> REJECT: <empty> DEFER: <empty> EFFECTIVITY:: <empty>

ORIGINATING CONTRACTOR: <empty>

NASA-STD-6016B APPENDIX C

RETIRED CATEGORY III MUA RATIONALE CODES

The Category III MUA rationale codes in the interim and baseline release of NASA-STD-6016 have been incorporated (with modifications) into the body of this NASA Technical Standard as exceptions to the requirements or have been eliminated. Category III MUAs are no longer required for codes that were incorporated as exceptions. Codes were eliminated when the rationales were no longer applicable or a full MUA was considered appropriate for the deviation. The retired codes are listed here for continuity with programs working to the earlier releases. As described in section 4.1.6.1.3, contractors who want to use Category III rationale codes may do so through their approved Materials and Processes Selection, Control, and Implementation Plan.

RETIRED FLAMMABILITY RATIONALE CODES

	RATIONALE
101	Approved Materials Usage Agreement (MUA) Category I.
102	Approved Materials Usage Agreement (MUA) Category II.
103	Materials passed requirements when tested in configuration.
104	Unexposed, overcoated, or sandwiched between nonflammable materials and no ignition source or propagation path.
105	Minor usage (less than 45 g (0.1 lb) mass and 13 cm2 (2 in2) surface area); no propagation path or ignition source.
107	Passes test No. 10 of NASA-STD-6001, Flammability Test for Materials in Vented Containers, by test or analysis.
108	Off-the-shelf equipment having material acceptable in configuration; no ignition source or propagation path.
109	Material not exposed; totally immersed in fluid; evaluated for fluid compatibility only.
110	Material is acceptable when used on a metal substrate that provides a good heat sink. Material considered noncombustible in this configuration by test or analysis.
111	Material is flammable but is sandwiched between nonflammable materials with edges only exposed and is more than 5 cm (2 in) from an ignition source or more than 30 cm (12 in) from other flammable materials.
112	Material is flammable but is unexposed or is overcoated with a nonflammable material.
113	Material is flammable but has a thickness less than 0.25 mm (0.010 in) and is sprayed or bonded to a metallic surface greater than 1.6 mm (0.062 in) thick.
114	Material is flammable but is used in “small amounts” and is more than 5 cm (2 in) from an ignition source or more than 30.5 cm (12 in) from other flammable materials. “Small amounts” for flammability may be quantified as follows: total weight less than 45 g (0.1 lb) and less than 13 cm2 (2.0 in2) surface area.

RETIRED TOXICITY (OFFGASSING) RATIONALE CODES

	RATIONALE
201	Approved Material Usage Agreement (MUA) Category I.
202	Meets toxicity requirements with performed cure.
203	T value for material/component in usage weight is <0.5 for manned flight compartment volume.
204	Materials usage in hermetically sealed container.

RETIRED FLUID SYSTEM COMPATIBILITY RATIONALE CODES

	RATIONALE
301	Approved Material Usage Agreement (MUA) Category I.
302	Passes requirements in configuration.
303	Material is B-rated in MAPTIS (batch/lot testing required) but batch/lot used in hardware passed test.
304	Approved Material Usage Agreement (MUA) Category II.

RETIRED THERMAL VACUUM STABILITY RATIONALE CODES

	RATIONALE
401	Approved Material Usage Agreement (MUA) Category I.
402	Approved Material Usage Agreement (MUA) Category II.
403	VCM between 0.1 and 1.0 percent; exposed area is less than 13 cm2 (2 in2) and not near a critical surface.
404	VCM >1.0 percent; exposed area is less than 1.6 cm2 (0.25 in2).
405	Unexposed, overcoated, or encapsulated with approved material.
406	Material is B-rated in MAPTIS (batch/lot testing required) but batch/lot used in hardware cured to meet requirements.
407	Meets thermal vacuum stability requirements in configuration.
408	Materials usage in hermetically sealed container.
409	Material has VCM >0.1 percent but is enclosed in a sealed container (maximum leak rate less than 1 x 10-4 cm3/sec).

RETIRED STRESS CORROSION CRACKING RATIONALE CODES

	RATIONALE
501	Approved Material Usage Agreement (MUA) Category I.
502	Approved Material Usage Agreement (MUA) Category II.
503	Maximum tensile stress <50 percent of yield strength for part on electrical/electronic assemblies.
504	Martensitic or PH stainless steels used in ball bearing, race, or similar applications where the primary loading is compressive.
505	Metal not listed in table 1 of MSFC-STD-3029 for stress corrosion cracking is not exposed to a corrosive environment after final assembly through end-item use.
506	Carbon and low alloy high strength steels greater than 1240 MPa (180 ksi) used in ball bearings, springs, or similar applications where primary loading is compressive, low tensile stresses, or history of satisfactory performance.

RETIRED CORROSION RATIONALE CODES

	RATIONALE
601	Approved Material Usage Agreement (MUA) Category I.
602	Approved Material Usage Agreement (MUA) Category II.
603	Adequately finished for corrosion protection.
604	Acceptable in use environment.
606	Electrical grounding required, cladding plus conversion coating adequate.
607	Thermal conductance and electrical bonding requirements preclude painting. Conversion coating is adequate (for aluminum only).
608	Finished on a higher assembly.
609	Laminated shim - minimum exposure of corrosion resistant material.
610	Material does not meet the requirements of MSFC-SPEC-250, Class II, but is treated or coated in a manner which meets or exceeds the requirements of MSFC-SPEC-250. Actual surface treatment shall be listed.
611	Material does not meet the requirements of MSFC-SPEC-250, Class II, but is not exposed to a corrosive environment after final assembly through end-item use.
612	Welding of titanium alloy-to-alloy or commercially pure-to-alloy using commercially pure filler metal in mixed alloy welds where hydrogen embrittlement is not predicted in service.

RETIRED GENERAL CODES

	RATIONALE
702	Generic materials controlled by military or industry specification using MAPTIS averages for ratings or test results. Material codes for generic material shall be used.
703	Military specification or industry specification allowing several material options where all options have acceptable ratings.

End of chapter · NextAppendix D Recommended data requirements documents

4.3Verification

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

Last Updated 29 Oct 2026

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[MPR 209] Verification of compliance with the requirements of this NASA Technical Standard shall consist of the following steps, as a minimum:

a. NASA approval of the contractor Materials and Processes Selection, Control, and Implementation Plan and other applicable materials data requirements documents, such as the Contamination Control Plan and NDE Plan.
b. Contractor M&P signature on engineering drawings to verify compliance with the requirements of this NASA Technical Standard or the Materials and Processes Selection, Control, and Implementation Plan.
c. NASA audits of contractor M&P activities relating to hardware design and manufacturing.
d. Establishment and operation of the M&P control panel in accordance with section 4.1.4 of this NASA Technical Standard.
e. NASA approval of MUAs.
f. NASA approval of MIULs. NOTE: Per section 4.1.6: The use of M&P that do not comply with the requirements of this NASA Technical Standard may be acceptable in the actual hardware applications. MUAs are required for all M&P that are technically acceptable but do not meet the requirements of this NASA Technical Standard, as implemented by the approved Materials and Processes Selection, Control, and Implementation Plan.

NASA-STD-6016B

Additional aspects of the verification process should be documented in the Materials and Processes Selection, Control, and Implementation Plan.

NASA-STD-6016B APPENDIX A STANDARD MATERIALS AND PROCESSES REQUIREMENTS FOR SPACECRAFT REQUIREMENTS COMPLIANCE MATRIX WITH REQUIREMENTS JUSTIFICATIONS A.1 PURPOSE

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STANDARD MATERIALS AND PROCESSES REQUIREMENTS FOR SPACECRAFT

REQUIREMENTS COMPLIANCE MATRIX WITH REQUIREMENTS JUSTIFICATIONS

A.1 PURPOSE

NASA-STD-6016B

	Section Title	NASA-STD-6016B Requirement Statement	Justification
1.2a	Scope, Applicability	[MPR 1] Programs shall apply these controls to program/project hardware.	It is NASA policy to provide clear, concise, verifiable requirements for processes, products, and components; and it is NASA policy to track and implement requirements using requirements management and configuration management control processes.
1.2b	Scope, Applicability	[MPR 2] Programs shall be responsible for demonstrating compliance with these requirements.	It is NASA policy to implement and track requirements.
1.3a	Scope, Tailoring	[MPR 3] Tailoring of this NASA Technical Standard’s requirements in the Materials and Processes Selection, Control, and Implementation Plan for specific programs/projects shall be formally documented as part of program or project requirements and approved by the responsible program/project NASA M&P organization, the responsible project/program, and the responsible Technical Authority in accordance with NPR 7120.5, NASA Space Flight Program and Project Management Requirements, or NPR 7120.8, NASA Research and Technology Program and Project Management Requirements.	Requirements management processes require that M&P requirement modifications be tracked and approved by responsible and knowledgeable M&P personnel and program organizations.
2.1.1	Applicable Documents, General	[MPR 4] The latest issuances of cited documents shall apply unless specific versions are designated.	Current NASA Office of the Chief Engineer policy is to include this statement in all NASA standards.
2.1.2	Applicable Documents, General	[MPR 5] Non-use of specifically designated versions shall be approved by	Current NASA Office of the Chief Engineer policy is to include this

NASA-STD-6016B

84 of 155

NASA-STD-6016B

		NASA-STD-6016B Requirement Statement	Justification	Section Title
		the responsible Technical Authority.	statement in all NASA standards.
2.4.2		[MPR 6] Conflict between this NASA Technical Standard and other requirements documents shall be resolved by the responsible Technical Authority.	Current NASA Office of the Chief Engineer policy is to include this statement in all NASA standards.	Applicable Documents, Order of Precedence
4a		[MPR 7] Materials shall meet the worst- case useful-life requirements for the particular application.	Common sense requirement that materials used in an application be assessed for all of the environmental extremes that the materials will encounter during the life cycle.	Requirements
4b		[MPR 8] M&P used in interfacing GSE, test equipment, hardware processing equipment, hardware packaging, and hardware shipment shall be controlled to prevent damage to or contamination of spaceflight hardware.	Ensures spaceflight hardware will not be damaged or contaminated by GSE, test equipment, or other M&P that comes in contact with the spaceflight hardware.	Requirements
4.1.1a		[MPR 9] Each organization that is responsible for the design and fabrication of spaceflight hardware shall provide a Materials and Processes Selection, Control, and Implementation Plan.	The Materials and Processes Selection, Control, and Implementation Plan represents a best practice for implementing, controlling, and verifying that the M&P requirements of this NASA Technical Standard are imposed and tracked through the program/project/element life cycle.	Materials and Processes Selection, Control, and Implementation Plan
4.1.1b		[MPR 10] The Materials and Processes Selection, Control, and Implementation Plan shall also describe the methods used to control compliance with these requirements by subcontractors and vendors.	The Materials and Processes Selection, Control, and Implementation Plan needs to provide to NASA insight into how the M&P requirements are imposed and controlled at the subcontractor and vendor level.	Materials and Processes Selection, Control, and Implementation Plan
4.1.1c		[MPR 11] Upon approval by the procuring activity, the Materials and Processes Selection, Control, and Implementation Plan shall become the M&P implementation document used for verification.	Since the Materials and Processes Selection, Control, and Implementation Plan contains the details of how a company implements and controls the requirements imposed by this NASA Technical Standard, it follows that the Plan is the document to be used to verify that the requirements have been met.	Materials and Processes Selection, Control, and Implementation Plan
4.1.1.1		[MPR 12] The Materials and Processes Selection, Control, and Implementation Plan shall identify the method of coordinating, approving, and tracking all engineering drawings, engineering orders, and other documents that establish or modify materials and/or processes usage.	A process needs to be in place for coordinating, approving, and tracking all M&P usage and M&P changes. The Materials and Processes Selection, Control, and Implementation Plan is the logical place for such a process to be identified.	Coordination, Approval, and Tracking
4.1.1.2		[MPR 13] The Materials and Processes Selection, Control, and Implementation Plan shall include a requirement that all hardware design drawings and revisions for spaceflight hardware that provides mission-critical functions contain an M&P approval block, or equivalent, to ensure that the design has been reviewed by the responsible M&P authority and complies with that document.	The M&P signature block on spaceflight hardware drawings ensures that M&P called out on drawings, and changes thereto, have been approved by M&P knowledgeable personnel.	Approval Signature
4.1.1.3		[MPR 14] The Materials and Processes Selection, Control, and Implementation Plan shall identify how the responsible M&P organization participates in manufacturing and inspection/verification	M&P expertise is required in the selection and processing of materials during manufacturing. How that expertise is applied and utilized needs to be described in the	Manufacturing Planning
		planning to ensure compliance with M&P requirements.	Materials and Processes Selection, Control, and Implementation Plan.
4.1.2a		[MPR 15] All M&P for spaceflight hardware that provides mission-critical functions shall be defined by standards and specifications selected from government, industry, and company specifications and standards.	Manufacturing spaceflight hardware to consensus, Government, or company M&P standards and specifications ensures reliable, repeatable, and trackable processes and spaceflight hardware. Manufacturing spaceflight hardware to standards and specifications has been a mandatory practice for a long time.	M&P Controls
4.1.2b		[MPR 16] All M&P for spaceflight hardware that provides mission-critical functions shall be identified directly on the appropriate engineering drawing.	Identifying all M&P directly on the drawing ensures that the M&P usage has been reviewed and approved. It also ensures that manufacturing follows the correct M&P processes, and that hardware built to print is repeatable.	M&P Controls
4.1.2c		[MPR 17] Company M&P specifications shall be identified by document number in the Materials and Processes Selection, Control, and Implementation Plan.	Ensures that NASA’s oversight responsibility has complete insight into all of the M&P that will be used to manufacture the spaceflight hardware.	M&P Controls
4.1.2d		[MPR 18] All M&P specifications used to produce spaceflight hardware that provides mission-critical functions shall be made available to the responsible NASA Program or Project Office and M&P organization.	NASA’s oversight role of approving and verifying M&P usage on mission-critical spaceflight hardware requires that all standards and specifications be available for review.	M&P Controls
4.1.2e		[MPR 19] Process specifications for spaceflight hardware that provides mission-critical functions shall define	Necessary to ensure a consistent and reliable product.	M&P Controls
		process steps at a level of detail that ensures a repeatable/controlled process that produces a consistent and reliable product.
4.1.2f		[MPR 20] Process qualification shall be conducted to demonstrate the repeatability of all processes for spaceflight hardware that provides mission-critical functions where the quality of the product cannot be directly verified by subsequent monitoring or measurement.	Necessary to ensure a consistent and reliable product.	M&P Controls
4.1.2.1		[MPR 21] A process shall be established for ensuring that updated, alternate, or new material or process standards or specifications from those identified in the Materials and Processes Selection, Control, and Implementation Plan meet or exceed the technical requirements identified in the original material or process standards or specifications.	The use of new material or material process standards or specifications in ongoing programs requires that the changes be reviewed and approved by the M&P and program authorities. A process for ensuring such a review and approval occurs needs to be identified in the Materials and Processes Selection, Control, and Implementation Plan. Also, engineering drawings do not call out company specifications by revision; so changes to such specifications can affect the repeatability of hardware built to print.	Standard and Specification Obsolescence
4.1.3		[MPR 22] A procedure shall be established to ensure that all vendor- designed, off-the-shelf, and vendor- furnished items are covered by the M&P requirements of this document, with the	COTS hardware can present or contain unknown hazards unless the M&P used to manufacture such hardware meets the requirements of this NASA Technical Standard, or	Commercial Off-The- Shelf (COTS) Hardware
		exception that M&P requirements for off- the-shelf hardware and other spaceflight hardware that do not provide mission- critical functions may be limited to those required to ensure safety of flight (for example, flammability, toxic offgassing) and vehicle compatibility (for example, vacuum outgassing).	the COTS M&P usage is reviewed and approved by the responsible NASA M&P and program authorities. A process needs to be established to ensure that COTS hardware is properly reviewed.
4.1.4a		[MPR 23] An M&P control panel shall be established by each contractor hardware provider (excluding subcontractors).	Each hardware provider needs to establish a material control panel to provide quick M&P expertise, review, and approval for hardware manufacturing problems and M&P changes. In addition, the panel provides the NASA M&P oversight responsibility with insight into company M&P implementation and control issues.	M&P Control Panel
4.1.4b		[MPR 24] The M&P control panel’s scope and membership shall be described in the Materials and Processes Selection, Control, and Implementation Plan.	The NASA M&P oversight responsibility requires understanding of who and how the contractor implements and controls requirements.	M&P Control Panel
4.1.5a		[MPR 25] M&P usage shall be documented in an electronic, searchable parts list or separate electronic searchable Materials Identification and Usage List (MIUL) with the following exceptions: (1) Electrical, electronic, and electromechanical (EEE) parts other than wire, cable, and exposed surfaces of connectors.	The MIUL is a best practice that has a long history of use in NASA that provides a quick, reliable, and easy review process for tracking and verifying M&P usage for both NASA and the contractors.	M&P Usage Documentation
		(2) Materials used in hermetically sealed electronic containers (maximum leak rate less than 1 x 10-4 cm3/sec).
4.1.5b		[MPR 26] The documentation of M&P usage shall cover the final design as delivered.	Common sense requirement that the end product reflects the actual M&P utilized in the as-built hardware (including M&P changes resulting from repairs to nonconforming hardware).	M&P Usage Documentation
4.1.5c		[MPR 27] The documentation approach shall be defined in the Materials and Processes Selection, Control, and Implementation Plan.	Ensures documentation of M&P usage and allows organizations to use their own documentation system.	M&P Usage Documentation
4.1.6a		[MPR 28] MUAs containing sufficient information to demonstrate that the application is acceptable shall be submitted by the responsible M&P organization for all M&P that are technically acceptable but do not meet the requirements of this NASA Technical Standard, as implemented by the approved Materials and Processes Selection, Control, and Implementation Plan.	The MUA process is a long used NASA M&P best practice that enables acceptance of materials and material processes that do not meet the requirements of this NASA Technical Standard but, through the MUA review process, are determined to be acceptable. The MUA process also provides the benefit of satisfactorily resolving design non-conformances at the engineering level as opposed to the program level.	Material Usage Agreements (MUAs)
4.1.6b		[MPR 29] MUAs shall not be used to change the M&P requirements for a nonconforming product.	The purpose of the MUA process is to justify the acceptance of specific material or material process design	Material Usage Agreements (MUAs)
			non-conformances, not the acceptance of a nonconforming product.
4.1.7a		[MPR 30] All parts or materials for spaceflight hardware that provides mission-critical functions shall be certified as to composition, properties, and requirements as identified by the procuring document.	Ensures that consistent and reliable materials are used in spaceflight hardware and further ensures that materials and their properties can be repeated from program to program and from company to company.	Materials Certification and Traceability
4.1.7b		[MPR 31] With the exception of off-the-shelf parts, parts and materials used in critical applications, such as life-limited materials and/or safety- and fracture-critical parts, shall be traceable through all processing steps defined in the engineering drawing to the end-item application.	M&P traceability to the end item of critical parts ensures that each processing step was accomplished and that a processing record exists which often aids in understanding a problem or failure.	Materials Certification and Traceability
4.1.7c		[MPR 32] Distributors or other processors shall not heat treat, hot work, or cold work metal stock unless they take the responsibilities as the producer of the metal and produce a new certification.	Required to control procurement of raw metallic materials. The processing changes the material properties and invalidates the original vendor’s certification.	Materials Certification and Traceability
4.1.8.1a		[MPR 33] A, B, or S-basis statistical values for mechanical properties of materials shall be utilized for the design and analysis of hardware for all applications where structural analysis is required.	To ensure structural integrity of spacecraft parts and structures, materials with statistically based minimum mechanical properties are required.	Requirements for Design Values
4.1.8.1b		[MPR 34] The sampling for other mechanical properties, such as dynamic properties like fatigue and fracture, and verification of design values shall be	Needed to ensure design values for other required mechanical properties are representative of the material, product form, and state	Requirements for Design Values
		representative of the material, product form, and state used in the design.	used in the design.
4.1.8.1c		[MPR 35] A, B, or S-basis statistical methods shall be defined by, and values for mechanical properties in their design environment taken from MMPDS, Metallic Materials Properties Development and Standardization, or SAE CMH-17, Composite Materials Handbook.	The handbooks imposed here have long been NASA’s and the aerospace industries’ most trusted and reliable sources of statistically derived mechanical design properties. MMPDS is the replacement for the old MIL- HDBK-5 and SAE CMH-17 is the replacement for the old MIL- HDBK-17. SAE CMH-17-5 is scheduled for publication in 2017 and MIL-HDBK-17-5 remains active until it is released.	Requirements for Design Values
4.1.8.1d		[MPR 36] For metallic materials, the alloy, heat treatment, product specification, product form, and thickness shall match the alloy, heat treatment, product specification, product form, and thickness in MMPDS.	Since the testing to develop design values is specific to alloy, heat treatment, specification and form, the statistical relationship is relevant only to that same combination. Thickness affects strength of metal because of metallurgical factors that influence strength, like heat transfer during heat treatment and variability of reduction during metal working. This requirement is stated in MMPDS but frequently ignored.	Requirements for Design Values
4.1.8.1e		[MPR 37] When statistical design values for new or existing structural materials are not available, they shall be determined by methods described in MMPDS or SAE	The statistical methods used in these handbooks are the NASA- and aerospace industry-accepted standard for determining	Requirements for Design Values
		CMH-17 and a report documenting the derivation of the new design values be made available to NASA for review.	mechanical design properties.
4.1.8.1f		[MPR 38] The implementation of an approach for generating statistical design values that deviates from the sampling methodology and statistical methods in MMPDS or SAE CMH-17 and the use of design values that deviate from those published in MMPDS or SAE CMH-17 shall be approved with an MUA.	Not all materials, product forms, and sizes are available in the cited handbooks. Development of design values by the analytical methods described in MMPDS or SAE CMH-17 is sometimes prohibitively expensive for the application. When the test approach deviates from the published values in MMPDS or the guidance in SAE CMH-17, the MUA review process by responsible and competent M&P engineering personnel assures that acceptable minimum design properties will be derived and verified.	Requirements for Design Values
4.1.8.1g		[MPR 39] All contractor-developed mechanical and physical property data shall be made available to the responsible NASA M&P organization.	NASA needs to have the capability to verify that contractor-developed mechanical and physical property data have been developed adequately and determine when an MUA is required per 4.1.8.1 g. NASA recognizes that such data may be proprietary and that the contractor can place limitations on how the data are made available.	Requirements for Design Values
4.1.8.2a		[MPR 40] Material “B” design values shall not be used except in redundant structure in which the failure of a component would result in a safe	“B” design values are designed for use in redundant structure. They have a less statistically significant minimum design value than “A”	Design Value Implementation in Design
		redistribution of applied loads to other load-carrying members.	design values and are, therefore, less reliable and not to be used in single load path critical structure.
4.1.8.2b		[MPR 41] Minimum property acceptance values in material specifications (specification minimums) not explicitly published as “S” design values in MMPDS shall require an approved MUA for use as design values.	Not all materials, product forms, and sizes are available in the cited handbooks. In some cases, specification minimums or “S” design values published in other sources are not properly derived in accordance with MMPDS. Therefore, the MUA review process by responsible and competent M&P engineering personnel assures that acceptable minimum design properties will be derived and verified.	Design Value Implementation in Design
4.1.8.3		[MPR 42] Structural fastener design values shall be defined by minimum load test requirements in the applicable part and/or procurement specification (government, aerospace industry consensus, company-specific, or custom specification).	MMPDS design values are not transferrable to fasteners, because raw material used to make fasteners is reprocessed using various metallurgical practices, such as hot heading, thread rolling, and heat treating. These processes change the strength of the metal from the original bar stock. NASA has chosen to adopt lot tested design strength defined in the specific fastener part and procurement specifications as the design value instead of the certifications of the original bar stock.	Structural Fastener Design Values
4.2.1		[MPR 43] Materials shall meet the requirements of NASA-STD-6001B,	Ensures that the NASA-required tests and evaluations of materials	Flammability, Offgassing, and
		Flammability, Offgassing, and Compatibility Requirements and Test Procedures, as described below.	for flammability, offgassing, and compatibility are imposed.	Compatibility Requirements
4.2.1.1		[MPR 44] Materials that are nonflammable or self-extinguishing in their use configuration as defined by NASA-STD-6001B, Test 1 or an appropriate configurational flammability test per NASA-STD-6001B, shall be used for flammability control with the following exceptions: a. Ceramics, metal oxides, and inorganic glasses are exempt. b. Materials that are designed to ignite and burn in their use application (for example, pyrotechnic materials) are exempted from flammability requirements, provided that systems/experiments using such materials are designed so that the burning materials cannot act as an ignition source for other hardware. c. Materials used in minor quantities (dimensions controlled so the potential fire propagation path is less than 15 linear cm (6 linear in) and the surface area is no more than 80 cm² (12 square in)) in crew environments and 30 linear cm (12 linear in) for external materials) with no propagation path or ignition source.	The lessons learned from the Apollo 1 fire regarding the use of flammable materials in an oxygen- rich environment makes it mandatory that special control and testing be exercised by M&P engineering to ensure that flammable materials are minimized and controlled in spacecraft environments.	Flammability Control
		d. Materials used in sealed containers are exempt because insufficient oxygen is available to maintain combustion. e. Materials within vented electronics packages with metallic cases and no forced air convection are exempt because of the self-extinguishing effect of expanding combustion gases in a constrained volume. f. Materials that have been shown by test to meet the requirements of NASA- STD-6001B may be used in an environment with an oxygen concentration lower than the test level without further testing (provided that the oxygen partial pressure is no greater than the partial pressure at the test level or general test data exist to demonstrate that the higher oxygen partial pressure is outweighed by the lower percentage concentration for the environments in question). g. Materials are acceptable when used on a metal substrate that provides an adequate heat sink. A heat sink is considered adequate in the use configuration by test or analysis. When testing is conducted, materials passing the flammability test on a metal substrate are acceptable on metal substrates of the same thickness or greater. Materials that
		are flammable but have a thickness less than 0.25 mm (0.010 in) and are attached to a metallic surface greater than 1.6 mm (0.062 in) thick are acceptable by analysis. h. Materials are unexposed, overcoated with a verified fire-blocking material, or no more than 1/4 inch thick and sandwiched between nonflammable materials with only the edges exposed. i. Flammability configuration analysis in accordance with JSC 29353 shows hardware configuration to be self-extinguishing.
4.2.1.2a		[MPR 45] All nonmetallic materials used in habitable flight compartments, with the exception of ceramics, metal oxides, inorganic glasses, and materials used in sealed container (maximum leak rate less than 1 x 10-4 cm3/sec), shall meet the offgassing requirements of Test 7 of NASA-STD-6001B.	All materials offgas small quantities of trace gas contaminants. Materials that release toxic gases must be controlled in closed spacecraft-habitable compartments. The required testing ensures that the materials used are safe from a toxicity standpoint.	Toxic Offgassing
4.2.1.2b		[MPR 46] Spacecraft maximum allowable concentration (SMAC) values shall be obtained from JSC 20584, Spacecraft Maximum Allowable Concentrations for Airborne Contaminants, or from MAPTIS for compounds for which no SMAC values are found in JSC 20584.	Industrial maximum concentrations of airborne contaminants are based on an 8-hour workday exposure and are not appropriate for the continuous exposure that applies in spacecraft environments. The SMAC values in JSC 20584 are set by the NASA/JSC Toxicology	Toxic Offgassing
			Group in cooperation with the National Research Council Committee on Toxicology. Supplementary values in MAPTIS are set by the NASA/JSC Toxicology Group only for evaluation of offgassing test data.
4.2.1.3a		[MPR 47] Materials exposed to hazardous fluids shall be evaluated or tested for compatibility.	Group in cooperation with the National Research Council Committee on Toxicology. Supplementary values in MAPTIS are set by the NASA/JSC Toxicology Group only for evaluation of offgassing test data.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3b		[MPR 48] Appropriate long-term tests to simulate the worst-case use environment that would enhance reactions or degradation shall be conducted for materials with long-term exposure to liquid fuels, oxidizers, and other hazardous liquids.	Since some incompatible material/fluid combinations only manifest themselves after long-term exposure, it is necessary to conduct long-term exposure testing.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3c		[MPR 49] Materials degradation in long- term tests shall be characterized by post- test analyses of the material and fluid to determine the extent of changes in chemical and physical characteristics, including mechanical properties.	To determine in long- or short-term tests if material degradation has occurred, post-test analyses must be performed to determine, e.g., if mechanical properties have changed or if unacceptable levels of the test material have been dissolved into the test fluid.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3d		[MPR 50] The effect of material condition (for example, parent versus weld metal or heat-affected zone) shall be addressed in the compatibility determination.	It has long been known that a material’s condition, such as a weld, its heat-affected zone, and the parent metal, can have significantly different outcomes in corrosion/compatibility testing	Fluid Compatibility (Fluids Other Than Oxygen)
			which needs to be addressed when determining compatibility and any necessary testing.
4.2.1.3e		[MPR 210] Materials used in nitrogen tetroxide systems shall be evaluated for flammability in nitrogen tetroxide using promoted combustion tests similar to NASA-STD-6001B, Test 17, for metallic materials used in oxygen systems and Test 1 for polymeric materials.	Nitrogen tetroxide is a powerful oxidizer, and flight hardware failures have been caused by ignition of metallic materials (notably titanium).	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3(1)		[MPR 211] When materials are determined to be flammable, a nitrogen tetroxide compatibility assessment shall be conducted using the methodology described for oxygen systems in NASA- STD-6001B and the system safety rationale of this assessment documented in an MUA.	The rigorous Oxygen Compatibility Assessment (OCA) methodology described in NASA-STD-6001 for evaluating potential materials ignition in oxygen systems is also the most effective methodology for evaluating ignition in nitrogen tetroxide systems.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.3f		[MPR 212] Materials used in other oxidizer systems shall be evaluated for potential ignition.	The possibility of similar ignition issues occurring in systems containing oxidizers other than oxygen and nitrogen tetroxide needs to be addressed.	Fluid Compatibility (Fluids Other Than Oxygen)
4.2.1.4a		[MPR 51] Materials and components, and systems used in liquid oxygen (LOX) and gaseous oxygen (GOX) environments, compressed air systems, and pressurized systems containing enriched oxygen shall be evaluated for oxygen compatibility in accordance with NASA-STD-6001B.	Because of the highly reactive nature of oxygen with metals and nonmetals, it is necessary that all materials that are in contact with or will come in contact with oxygen must be evaluated for compatibility with oxygen. NASA-STD-6001B governs how such evaluation is to be accomplished.	Oxygen Compatibility
4.2.1.4b		[MPR 52] When materials in such	Materials are commonly found to	Oxygen Compatibility
		systems are determined to be flammable, an oxygen compatibility assessment shall be conducted as described in NASA- STD-6001B and the system safety rationale of this assessment documented in an MUA.	be flammable in standard flammability tests but can often be found not to be a flammability hazard in their use configuration. Such a determination is made by conducting an oxygen compatibility assessment and documenting the acceptance rationale in an MUA.
4.2.1.4.1		[MPR 53] GOX and enriched air system components that operate at pressures above 1.83 MPa (265 psia), with the exception of metallic components, such as hard lines (rigid metallic tubing), metallic flex hoses, metallic fluid fitting with all metal seals, and metallic pressure vessels (including composite overwrapped pressure vessels with metallic liners), shall undergo oxygen component acceptance testing at maximum design pressure for a minimum of ten cycles from ambient pressure to maximum design pressure within 100 milliseconds to ensure that all oxygen system spaceflight hardware is exposed to oxygen prior to launch.	Ensures that material surfaces that can react with oxygen at relatively high pressures have been thoroughly exposed and conditioned to the system oxygen pressure and will be nonreactive to oxygen during use.	Oxygen Component Acceptance Test
4.2.1.5a		[MPR 54] Electrical wire insulation materials shall be evaluated for flammability in accordance with NASA- STD-6001B, Test 4 or Test 1.	The lessons learned from the Apollo 1 fire and the Apollo 13 oxygen tank explosion make it imperative that the flammability of wire insulation be evaluated. NASA-STD-6001 specifies how to conduct such tests.	Electrical Wire Insulation Materials
4.2.1.5b		[MPR 55] Arc tracking shall be evaluated in accordance with NASA-STD-6001B, Test 18, for all wire insulation constructions except polytetrafluoroethylene (PTFE), PTFE/polyimide insulation conforming to SAE AS22759C, Wire, Electrical, Fluoropolymer-Insulated, Copper or Copper Alloy, and used in conditions bounded by the dry arc-tracking test specified by SAE AS22759C, ethylene tetrafluoroethylene (ETFE), and silicone- insulated wires (the resistance of these materials to arc tracking has already been established).	Since arc tracking is a problem that is due to a breakdown of a wire’s insulating material that results in uncontrolled arcing to other wires and components, it is necessary to test wire insulating materials for their susceptibility to arc tracking.	Electrical Wire Insulation Materials
4.2.2		[MPR 56] MSFC-STD-3029A, Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments, shall be used to select metallic materials to control stress corrosion cracking of metallic materials in sea and air environments, with the exception that an MUA is not required for the following conditions: a. Parts in electrical/electronic assemblies with maximum tensile stress less than 50 percent of the yield strength. b. Martensitic or precipitation-hardening (PH) stainless steels used in ball bearing or similar applications where the loading is compressive.	Stress corrosion cracking in seacoast and air environments is a potential problem for the high- strength alloys typically used in spacecraft structures. MSFC-STD- 3029 provides data on the stress- corrosion sensitivity of most structural alloys and contains requirements to eliminate the potential for stress-corrosion cracking when stress-corrosion- sensitive materials are used. The standard also contains test methods for evaluating stress-corrosion sensitivity of alloys without existing data and defines stress corrosion cracking threshold values for many of the alloys used in spacecraft structures.	Metals
		c. Carbon and low alloy high-strength steels with strength greater than 1240 MPa (180 ksi) used in ball bearings, springs, or similar applications where the loading is compressive and there is a history of satisfactory performance. d. Hardware provides no mission-critical functions
4.2.2.1 4.2.2.2.1a		[MPR 57] The 5000-series alloys containing more than 3 percent magnesium shall not be used in spaceflight hardware that provides mission-critical functions where the temperature exceeds 66 °C (150 °F). [MPR 58] When drilling, grinding, reaming, or machining is performed on high-strength steels that can form martensite and are used in spaceflight hardware that provides mission-critical functions, a validated process for machining shall be used.	Under certain conditions, an intergranular aluminum-magnesium grain boundary precipitate can occur rendering such alloys susceptible to exfoliation corrosion or stress corrosion. A well-known lesson learned is that some metal removal processes, particularly aggressive ones, cause surface overheating in high-strength steels that produces the undesirable crack-prone condition of untempered martensite.	Aluminum Drilling and Grinding of High-Strength Steel
4.2.2.2.1b		[MPR 59] The absence of machining damage for high-strength steels used in spaceflight hardware that provides mission-critical functions shall be verified by microexamination of production parts (such as Nital etch inspection) or by a microhardness and metallurgical examination of sample parts for either of the following situations:	Tests and examinations cited are best practices for identifying the presence of untempered martensite.	Drilling and Grinding of High Strength Steel
		(1) When the material has very low toughness (such as martensitic steel above 200 ksi) and the part has low stress margins or is fatigue driven. (2) When the surface condition of the part is critical to the design (such as its ability to withstand hertzian stresses or remain perfectly flat).
4.2.2.2.2a		[MPR 60] Unstabilized, austenitic steels shall not be processed or used above 371 °C (700 °F) in spaceflight hardware that provides mission-critical functions.	Eliminates the potential of a corrosion-susceptible condition due to grain boundary precipitation of chromium carbide.	Corrosion-Resistant Steel
4.2.2.2.2b		[MPR 61] Welded assemblies used in spaceflight hardware that provides mission-critical functions shall be solution heat-treated and quenched after welding except for the stabilized or low carbon grades, such as 321, 347, 316L, and 304L.	Eliminates the potential of a corrosion-susceptible condition of weld heat-affected zone grain boundary precipitation of chromium carbide.	Corrosion-Resistant Steel
4.2.2.2.2c		[MPR 62] Service-related corrosion issues are common for free-machining alloys, such as 303, and these alloys shall not be used in spaceflight hardware that provides mission-critical functions when they can be exposed to moisture other than transient condensation, or to nitrogen tetroxide	Eliminates the exposure of the more corrosion-susceptible stainless steel alloys to moisture or to nitrogen tetroxide, for which corrosion of these alloys has also been observed.	Corrosion-Resistant Steel
4.2.2.2.3		[MPR 63] Steels shall not be used in tension below their ductile-brittle transition temperature.	Prevents the use of alloys that are too impact- and crack-sensitive and prone to unexpected failure at low	Ductile-Brittle Transition Temperature
			temperature.
4.2.2.3.1a		[MPR 64] All cleaning fluids and other chemicals used during manufacturing and processing of titanium hardware for spaceflight hardware that provides mission-critical functions shall be verified to be compatible and not detrimental to performance before use.	An Apollo Program lesson learned occurred when a titanium pressure vessel failed due to stress corrosion in anhydrous methanol that was used as a substitute pressurant for Aerozine-50 in a ground test. Methanol, a commonly used solvent for cleaning metals, was considered benign to titanium.	Titanium Contamination
4.2.2.3.1b		[MPR 65] The surfaces of titanium and titanium alloy mill products used for spaceflight hardware that provides mission-critical functions shall be 100 percent machined, chemically milled, or pickled to a sufficient depth to remove all contaminated zones and layers formed while the material was at elevated temperature.	A lesson learned from a titanium pressure vessel that was severely cracked from an alpha case that formed during heat treatment.	Titanium Contamination
4.2.2.3.2		[MPR 66] All regions of titanium alloys for spaceflight hardware that provides mission-critical functions that are subject to fretting or wear shall be anodized per SAE AMS2488D, Anodic Treatment – Titanium and Titanium Alloys Solution, pH 13 or Higher, or hard-coated utilizing a wear-resistance material, such as tungsten carbide/cobalt thermal spray.	Since titanium is subject to wear and fretting which can lead to crack initiation, it is necessary to harden the titanium surfaces by anodizing per the cited specification or per a hard-coating process.	Titanium Wear
4.2.2.3.3a		[MPR 67] Titanium alloys shall not be used with LOX or GOX at any pressure or with air at oxygen partial pressures above 35 kPa (5 psia).	Titanium is highly flammable in oxygen-rich environments, and its use in such environments must be avoided.	Titanium Flammability
4.2.2.3.3b		[MPR 68] Titanium alloys shall not be	Because of titanium’s sensitivity to	Titanium Flammability
		machined inside spacecraft modules during ground processing or in flight, because machining operations can ignite titanium turnings and cause fire.	ignition in oxygen-containing environments, fine turnings and particles with large surface areas from machining and grinding processes can even ignite in air. Obviously, such fires in a spacecraft must be prevented.
4.2.2.4a		[MPR 69] Magnesium alloys shall not be used in primary structure or in other areas of spaceflight hardware that provides mission-critical functions that are subject to wear, abuse, foreign object damage, abrasion, erosion, or where fluid or moisture entrapment is possible.	Magnesium’s low strength and hardness renders it relatively sensitive to mechanical damage and abuse and is relatively corrosion sensitive. Therefore, its use in primary structure needs to be avoided.	Magnesium
4.2.2.4b		[MPR 70] Magnesium alloys shall not be machined inside spacecraft modules during ground processing or in flight, because machining operations can ignite magnesium turnings and cause fire.	Magnesium can be ignited in air at room temperature from machining operations. Obviously, such fires in a spacecraft must be prevented.	Magnesium
4.2.2.5a		[MPR 71] Alloys containing more than 4 percent beryllium by weight shall not be used for primary structural applications.	Beryllium alloys have an inherent low ductility and brittle nature, and their use in primary structure needs to be avoided.	Beryllium
4.2.2.5b		[MPR 72] Alloys containing more than 4 percent beryllium by weight shall not be used for any application within spacecraft crew compartments unless suitably protected to prevent erosion or formation of salts or oxides.	Beryllium containing alloys, even those containing as little as 4 percent beryllium, present a problem in that oxides and salts of beryllium are highly toxic and the use thereof requires that the metal be protected from abrasion or conditions that could form	Beryllium
4.2.2.5c		[MPR 73] Design of beryllium parts for spaceflight hardware that provides mission-critical functions shall address its low-impact resistance and notch sensitivity, particularly at low temperatures, and its directional material properties (anisotropy) and sensitivity to surface finish requirements.	beryllium particles, beryllium oxides, or salts. When beryllium is required to be used in mission-critical spaceflight hardware, it is important that the low ductility and brittle nature of beryllium be critically reviewed. When analyzing the stress state in beryllium, it is important to recognize the low Poisson’s Ratio of beryllium. This low Poisson’s Ratio can lead to very high stresses.	Beryllium
4.2.2.5d		[MPR 74] All beryllium parts used in spaceflight hardware that provides mission-critical functions shall be processed to ensure complete removal of the damaged layer (twins and microcracks) produced by surface-metal-working operations, such as machining and grinding.	When beryllium is required to be used in mission-critical hardware, the detrimental surface damage that occurs in beryllium from mechanical metal removal methods, such as machining and grinding, must be removed. The damage removal must be by chemical etch to remove the surface damage. The technique of removing successively shallower cuts will reduce the depth of the damage but will never completely eliminate the damage. The nature of the damage is a crystal twinning in the grains which pins dislocations and subsequently reduces material ductility.	Beryllium
4.2.2.5e		[MPR 75] Beryllium-containing alloys (including alloys containing less than 4	Due to the highly toxic nature of beryllium dust and oxides formed	Beryllium
		percent beryllium by weight) and oxides of beryllium shall not be machined inside spacecraft crew compartments at any stage of manufacturing, assembly, testing, modification, or operation.	during machining or grinding, such operations must be highly controlled and never allowed to occur in spacecraft crew compartments.
4.2.2.5f		[MPR 76] All beryllium parts used in spaceflight hardware that provides mission-critical functions shall be penetrant-inspected for crack-like flaws with a high-sensitivity fluorescent dye penetrant in accordance with section 4.2.5.	Ensures that the crack-sensitive beryllium parts are crack free.	Beryllium
4.2.2.6a		[MPR 77] Cadmium shall not be used in crew or vacuum environments.	Because of its toxic nature and relatively high vapor pressure, the use of cadmium in spacecraft crew compartments and in the space vacuum environment needs to be prevented.	Cadmium
4.2.2.6b		[MPR 78] Cadmium-plated tools and other hardware shall not be used in the manufacture or testing of spaceflight hardware that provides mission-critical functions.	Cadmium is relatively soft and could be readily transferred to spacecraft parts through contact between the parts and cadmium- plated tools. Therefore, the use of such tools must be prevented.	Cadmium
4.2.2.7		[MPR 79] Owing to zinc’s ability to grow whiskers, zinc plating other than black zinc-nickel plating shall not be used in spaceflight hardware that provides mission-critical functions.	The use of zinc-plated parts in spaceflight hardware needs to be prevented because zinc has a tendency to grow whiskers that cause electrical shorts and are also a source of contamination. Zinc also has a relatively high vapor pressure and, in the vacuum of space, it presents the risk of	Zinc
4.2.2.8a		[MPR 80] Equipment containing mercury shall not be used where the mercury could come in contact with spaceflight equipment during manufacturing, assembly, test, checkout, and flight.	volatilizing and redepositing on critical surfaces. Mercury in contact with some metals can result in severe embrittlement; therefore, such contact must be prevented.	Mercury
4.2.2.8b		[MPR 81] Spaceflight hardware (including fluorescent lamps) containing mercury shall have three levels of containment to prevent mercury leakage.	Because of the high toxicity of mercury vapor and liquid mercury’s potential for causing embrittlement of other metals, its safe containment in a spacecraft environment must be extraordinarily redundant.	Mercury
4.2.2.9		[MPR 82] For refractory alloys (alloys with a melting point above 2000 °C (3600 °F), plus osmium and iridium) used in mission-critical applications, tests shall be performed to characterize critical design properties for the intended application and the data documented in an MUA.	Refractory metals are used in high- temperature applications where the material properties at the intended use-temperature are lacking or limited. Testing must be performed to obtain the necessary design properties for the application. Documentation of the design properties is also required.	Refractory Metals
4.2.2.10a		[MPR 83] High-nickel content alloys are susceptible to sulfur embrittlement; therefore, for spaceflight hardware that provides mission-critical functions, any foreign material which could contain sulfur, such as oils, grease, and cutting lubricants, shall be removed by suitable means prior to heat treatment, welding, or high temperature service.	Since superalloys are susceptible to sulfur embrittlement from contact with sulfur containing materials at elevated temperatures, it is mandatory that such materials be removed from the surface of the superalloys before they are exposed to high temperatures.	Superalloys (Nickel- Based and Cobalt- Based)
4.2.2.10b		[MPR 84] The reduction to design	Since superalloys experience alloy	Superalloys (Nickel-
		properties of alloying element depletion at the surface in a high temperature, oxidizing environment shall be evaluated when a thin sheet of one of these alloys is used for spaceflight hardware that provides mission-critical functions, since a slight amount of depletion could involve a considerable proportion of the effective cross section of the material.	depletion at elevated temperatures, it is mandatory that the depth of such depletion on a superalloy’s design properties be assessed, particularly in the case of thin sheets.	Based and Cobalt- Based)
4.2.2.11a		[MPR 85] For spaceflight hardware that provides mission-critical functions, tin and tin plating shall be alloyed with at least 3 percent lead by weight or other proven alloying element(s) to prevent tin whisker growth.	Self-explanatory, in that 3 percent lead in tin and tin plating is necessary to prevent tin-whisker growth. Several other alloying elements in tin and tin plating have also been shown to prevent tin whiskers.	Tin
4.2.2.11b		[MPR 86] Tin and tin plating alloyed with less than 3 percent lead by weight and used in electrical/electronic applications shall comply with GEIA-STD-0005-1A, Performance Standard for Aerospace and High Performance Electronic Systems Containing Lead-Free Solder, and Control Level 2C requirements of GEIA-STD- 0005-2A, Standard for Mitigating the Effects of Tin Whiskers in Aerospace and High Performance Electronic Systems, with the following exceptions: (1) Solder alloy Sn96.3Ag3.7 (Sn96) used for high- temperature applications. (2) Solder alloy Au80Sn20 used as	Since the cited standards specifically address how to mitigate the tin whisker problems for tin and tin-plating containing less than 3 percent lead, it is necessary that they be used for applications that use tin and tin-plating with less than 3 percent lead.	Tin
		a die attach material or as a package sealing material. (3) Tin alloys containing less than 20 percent tin by weight.
4.2.2.11c		[MPR 87] When high-purity tin and tin plating are used for spaceflight hardware that provides mission-critical functions and will also be exposed to temperatures below 13 ˚C for periods longer than 6 months, the method for preventing tin pest formation shall be documented in the materials control plan.	Since tin pest is a problem that results from tin and tin plating being exposed to a low temperature for a long period of time (as can occur for spacecraft external materials), it is necessary to address how the tin pest problem will be controlled. Tin pest has been observed on ISS external hardware.	Tin
4.2.2.11d		[MPR 88] Tin plating shall not be used for contacts in electrical interconnects (connectors, sockets, switches, etc.) for spaceflight hardware that provides mission-critical functions.	Because of the many problems that have been encountered with tin plating, such as oxidation, wear, fretting, and whiskers, it is necessary to limit the use of tin plating in spaceflight hardware unless the potential problems are suitably mitigated and its use approved through an MUA review process.	Tin
4.2.3.1a		[MPR 89] Elastomeric materials used in spaceflight hardware that provides mission-critical functions shall be selected to operate within design parameters for the useful life of the hardware.	Since elastomers have age-life limitations, it is necessary to select elastomeric materials that can function and endure the space environment for the expected life of the hardware.	Elastomeric Materials
4.2.3.1b		[MPR 90] Elastomeric materials used in spaceflight hardware that provides mission-critical functions, other than	The age-life of elastomers begin when they are cured; therefore, because of age-life limitations, their	Elastomeric Materials
		those used in off-the-shelf parts, such as cable clamps, shall be cure-dated for tracking purposes.	cure dates need to be recorded so the ages of the elastomers can be tracked.
4.2.3.1c		[MPR 91] Room temperature vulcanizing (RTV) silicones that liberate acetic acid during cure shall not be used because they can cause corrosion.	Requirement justifies itself, in that the release of a corrosive acid is unacceptable.	Elastomeric Materials
4.2.3.1d		[MPR 92] When rubbers or elastomers are used at low temperatures in spaceflight hardware that provides mission-critical functions, the ability of these materials to maintain and provide required elastomeric properties shall be verified.	The lesson learned from the Challenger O-ring failure due to its inelastic behavior at a moderately cold temperature strongly emphasizes the need to verify the low temperature elastic properties of rubbers and elastomers. Fluorocarbon elastomers are particularly susceptible to embrittlement.	Elastomeric Materials
4.2.3.2		[MPR 93] Use of polyvinyl chloride on spaceflight hardware shall be limited to applications in pressurized areas where temperatures do not exceed 49 °C (120 °F).	Polyvinyl chloride (PVC) cannot be used in space vacuum because it contains large quantities of plasticizers that outgas in space vacuum, making the PVC brittle and acting as a source of contamination. PVC also offgasses excessively above 120 oF, so its use in hardware that experiences such temperatures must be restricted. PVC is flammable in spacecraft cabin environments, and its use is subject to the same controls as other flammable materials.	Polyvinyl Chloride
4.2.3.5		[MPR 94] All materials shall be selected to meet the useful life of the hardware	If a material cannot meet the useful life requirements of the hardware of	Limited-Life Items
		with no maintenance or be identified as limited-life items requiring maintainability.	which it is a part, then it must be identified as having limited-life and a maintenance plan identified to address its limited-life.
4.2.3.6a		[MPR 95] Nonmetallic materials that are exposed to space vacuum, with the exception of ceramics, metal oxides, inorganic glasses, and cetyl alcohol lubricants used on fasteners outside closed compartments, shall be tested using the technique of ASTM E595-15, Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment, with acceptance criteria as follows: (1) ≤0.1 percent collected volatile condensable materials (CVCM). (2) ≤1.0 percent total mass loss (TML) less water vapor recovery (WVR), except that a higher mass loss is permitted if this mass loss has no effect on the functionality of the material itself and no effect on the functionality of any materials, components, or systems that could be adversely affected by the subject mass loss.	Since nonmetallic materials outgas in the vacuum of space, it is necessary to quantify the nature and amount of outgassing that can occur. ASTM E595 is the consensus standard for conducting such tests and quantifying mass loss and CVCM. The traditional ≤1.0 percent total mass loss (TML) spacecraft requirement is relaxed in this standard. Vacuum outgassing causes contamination only from condensable materials and the TML is unimportant except in special cases where cryogenic surfaces are present and in cases where the mass loss causes unacceptable degradation of the material. The vast majority of materials with a TML ≤5 percent are acceptable for normal spacecraft applications, provided the ≤0.1 percent CVCM requirement is met.	Thermal Vacuum Stability
4.2.3.6b		[MPR 96] With the following exceptions, hardware items (components, assemblies, etc.) containing materials that fail the	Since vacuum baking can sufficiently outgas materials that fail the CVCM requirements	Thermal Vacuum Stability
		CVCM requirement and/or having unidentified materials shall be vacuum baked at the maximum tolerable temperature of the component, 10 ℃ above the maximum predicted operating temperature, or an alternate temperature selected by the program/project, to meet the program/project acceptance outgassing criteria: (1) Materials that are not near a critical surface and have a CVCM between 0.1 and 1.0 percent and an exposed surface area less than 13 cm² (2 in²) are exempt. (2) Materials with an exposed surface area less than 1.6 cm² (0.25 in²) are exempt. (3) Materials that are unexposed, overcoated, or encapsulated with approved materials are exempt. (4) Materials enclosed in a sealed container (maximum leak rate less than 1 x 10⁻⁴ cm³/sec) are exempt.	established by this NASA Technical Standard, it is necessary to vacuum bake hardware items that contain materials that do not meet the CVCM requirements.
4.2.3.7		[MPR 97] The critical properties of materials exposed in the spacecraft external environment shall meet	The spacecraft external environment includes ultraviolet radiation, atomic oxygen in low-	External Environment Survivability
		operational requirements for their intended life-cycle exposure.	Earth orbit, plasma, and high- energy radiation, all of which can cause materials degradation. For example, ultraviolet radiation causes darkening of many materials, leading to loss of thermal control, and atomic oxygen erodes most polymeric materials, causing physical degradation. Since repair capabilities are very limited or nonexistent, it is imperative that materials so exposed meet their operational life cycle.
4.2.3.8		[MPR 98] Materials that are non-nutrient to fungi, as identified in MIL-HDBK- 454B, General Guidelines for Electronic Equipment, Requirement 4, Fungus-Inert Materials, Table 4-I, Group I, shall be used in launch vehicles and pressurized flight compartments, except when one of the following criteria is met: a. Materials have been tested to demonstrate acceptability per MIL-STD- 810G, Department of Defense Test Method Standard for Environmental Engineering Considerations and Laboratory Tests, Method 508. b. Materials are used in crew areas where fungus would be visible and easily removed.	Materials that resist the growth of fungus are required in launch vehicles that are exposed to weather environments and in pressurized habitable compartments. The requirement statement identifies many exceptions where such controls are unnecessary.	Fungus Resistance
4.2.3.9	Section Title	Glycols	Required to prevent this unusual spontaneous ignition reaction.
		be added to the solution to prevent spontaneous ignition from the reaction of silver with the glycol.
4.2.3.10a		[MPR 100] The etching of PTFE, perfluoroalkoxy (PFA), and fluorinated ethylene propylene (FEP) shall meet the requirements of SAE AMS2491F, Surface Treatment of Polytetrafluoroethylene, Preparation for Bonding, when adhesion to the fluorocarbon surface is required, except that for insulated wire or cable a pull test on co-produced specimens may be performed in lieu of the tensile and shear strength tests in AMS2491 section 3.5.2.	It is NASA policy to use voluntary consensus standards where possible. The cited specification is the industry standard for preparing polytetrafluoroethylene for bonding and has been reviewed and approved for use by NASA.	Etching Fluorocarbons
4.2.3.10b		[MPR 101] Etched surfaces shall be processed within 24 hours or within 1 year if packaged per SAE AMS2491F.	Ensures that the etched surface remains active and ready for bonding.	Etching Fluorocarbons
4.2.4.1a		[MPR 102] Heat treatment of aluminum alloys used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS2772G, Heat Treatment of Aluminum Alloy Raw Materials; SAE AMS2770N, Heat Treatment of Wrought Aluminum Alloy Parts; or SAE AMS2771E, Heat Treatment of Aluminum Alloy Castings.	It is NASA policy to use voluntary consensus standards where possible. The cited standards are the widely accepted consensus standards for heat treating aluminum alloys, and they have been reviewed and approved for use by NASA.	Heat Treatment
4.2.4.1b		[MPR 103] Heat treatment of steel alloys used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS-H-6875B, Heat Treatment of Steel Raw Materials,	It is NASA policy to use voluntary consensus standards where possible. The cited standards are the widely accepted consensus standards for heat treating steel	Heat Treatment
		or SAE AMS2759E, Heat Treatment of Steel Parts, General Requirements.	alloys, and they have been reviewed and approved for use by NASA.
4.2.4.1c		[MPR 104] Heat treatment of titanium alloys used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS-H- 81200D, Heat Treatment of Titanium and Titanium Alloys, for raw stock and SAE AMS2801B, Heat Treatment of Titanium Alloy Parts, for parts requiring heat treatment during fabrication.	It is NASA policy to use voluntary consensus standards where possible. The cited standards are the widely accepted consensus standards for heat treating titanium alloys, and they have been reviewed and approved for use by NASA.	Heat Treatment
4.2.4.1d		[MPR 105] Heat treatment of nickel- and cobalt-based alloy parts used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS2774E, Heat Treatment, Wrought Nickel Alloy and Cobalt Alloy Parts, or SAE AMS2773E, Heat Treatment, Cast Nickel Alloy and Cobalt Alloy Parts.	It is NASA policy to use voluntary consensus standards where possible. The cited standards are the widely accepted consensus standards for heat treating nickel- and cobalt-based alloys and have been reviewed and approved for use by NASA.	Heat Treatment
4.2.4.1e		[MPR 106] For spaceflight hardware that provides mission-critical functions, process-control tensile-test coupons shall be taken from the production part (or from the same material lot, having the same thickness as and processed identically to the production part) to verify the adequacy of the heat treatment process for the following conditions: (1) Aluminum alloys are solution	Heat treat variations can compromise the strength and ductility of the alloys for the heat- treat conditions cited. In addition, the commonly used hardness test to verify the adequacy of the heat treat process is not a good predictor of strength and ductility.	Heat Treatment
		heat-treated. (2) High-strength steels (>200 ksi (1380 MPa) UTS), tool steels, and maraging steel alloys are heat-treated to high strength levels. (3) A286 or MP35N alloys (which have poor correlation between hardness and tensile strength) are heat treated. (4) Titanium alloys are annealed or solution heat treated and aged. (5) Nickel- and cobalt-based alloys are work strengthened before age hardening resulting in age- hardened tensile strengths greater than 1030 MPa (150 ksi) UTS. (6) Precipitation hardenable nickel- and cobalt-based alloys are solution heat treated.
4.2.4.1f		[MPR 107] When process-control tensile- test coupons are required, the requirement for the coupons shall be specified on the engineering drawing for the part.	Ensures compliance.	Heat Treatment
4.2.4.1g		[MPR 108] If no tensile values are available in MMPDS for a specific alloy, tensile-test acceptance values shall be specified on the engineering drawing for	Ensures that accept/reject criteria are established.	Heat Treatment
		the part.
4.2.4.1h		[MPR 109] Materials shall not be used in spaceflight hardware that provides mission-critical functions outside the limits of their procurement specification, heat treat specification, or MMPDS specification.	Ensures that the materials that are used have defined and dependable design properties.	Heat Treatment
4.2.4.2a		[MPR 110] Where forgings are used in mission-critical applications, first-article (preproduction) approval shall be obtained from the procuring authority.	Ensures that a NASA M&P technical expert(s) has reviewed and approved the first-article cut-up process.	Forging
4.2.4.2b		[MPR 111] First-article approval and the controls to be exercised in producing subsequent production forgings shall be in accordance with SAE AMS2375D, Control of Forgings Requiring First Article Approval.	It is NASA policy to use voluntary consensus standards where possible. The cited standard is the widely accepted consensus standard for control of the first-article forging cut-up and has been reviewed and approved for use by NASA.	Forging
4.2.4.2c		[MPR 112] After the forging technique, including degree of working, is established, the first production forging shall be sectioned to show the grain-flow patterns and to determine mechanical properties at control areas and the trim ring/protrusion specimens (prolongations).	Ensures that the forging process has broken up the cast structure and that expected microstructure and mechanical properties are produced by the forging process.	Forging
4.2.4.2d		[MPR 113] The mechanical properties of the trim ring/protrusion specimens (prolongations) for the first article shall be compared to the control coupons to show they are predictive of the properties	The mechanical properties of the trim ring/protrusion specimens (prolongations) for subsequent articles will be used to verify that the mechanical properties of	Forging
		in the body of the first article.	subsequent articles are acceptable. As such, they have to be representative of the properties in the body of each subsequent article. For the first article, they are compared with the mechanical properties of the control coupons from the body of the article to verify that they are truly representative.
4.2.4.2e		[MPR 114] Sectioning to show the grain- flow patterns and to determine mechanical properties at control areas shall be repeated after any substantive change in the forging technique, as determined by M&P analysis.	Since any changes in the forging process can significantly alter the grain flow, microstructure, and mechanical properties, it is necessary to conduct a first-article- like cut-up to ensure the resulting properties are acceptable.	Forging
4.2.4.2f		[MPR 115] These data and results of tests on the redesign shall be retained and made available for review by the procuring activity.	Since forging process changes can significantly alter the mechanical and metallurgical properties of the forging, it is necessary to conduct a first-article-like cut-up and have the results reviewed and approved by NASA M&P engineering.	Forging
4.2.4.2g		[MPR 116] Trim ring or protrusion specimens (prolongations) shall be obtained for each production forging used in safety-critical applications and tested for required minimum mechanical properties.	It is standard practice in the forging industry to obtain and test trim ring or protrusion specimens to ensure that the mechanical properties of each forging are achieved.	Forging
4.2.4.2h		[MPR 117] Surface and volumetric nondestructive inspection (NDI) shall be performed on all safety-critical forgings.	Obviously, the severe plastic deformation involved in a forging process can introduce both surface and internal discontinuities and defects.	Forging

	Section Title	NASA-STD-6016B Requirement Statement	Justification and internal defects for which inspection is needed.
4.2.4.3a	Castings	[MPR 118] Castings used in spaceflight hardware that provides mission-critical functions shall meet the requirements of SAE AMS2175A (2010), Castings, Classification and Inspection of.	It is NASA policy to use voluntary consensus standards. SAE AMS2175 is the industry specification for classifying and inspecting castings and has been reviewed and approved for use by NASA.
4.2.4.3b	Castings	[MPR 119] Where castings are used in mission-critical applications, pre- production castings shall be subjected to first-article inspection to verify proper material flow, proper material integrity, minimum required mechanical properties, proper grain size, and macro/microstructure.	For mission-critical castings, it is considered mandatory to subject the first casting(s) to destructive testing to verify the mechanical properties and metallurgical characteristics of the casting process.
4.2.4.3c	Castings	[MPR 120] The mechanical properties in trim ring/protrusion of the first article shall be compared to the control coupons to show they are predictive of the properties in the body of the first article.	Trim rings/protrusions will be obtained for all production castings so the first article testing needs to verify that they are representative of the cast parts.
4.2.4.3d	Castings	[MPR 121] The same casting practice and heat-treating procedure shall be used for the production castings as for the approved first-article castings.	Ensures to a reasonable degree that the production castings will have the same mechanical properties and metallurgical characteristics as the first-article casting(s) that validated the casting process.
4.2.4.3e	Castings	[MPR 122] For Class 1 and Class 2 castings (classes as defined by SAE AMS2175A (2010)), mechanical property testing of integrally cast or excised tensile bars at critical locations shall be	Tensile bars cast at the same time as the production casting(s) ensure that the casting process is reproducing the same mechanical properties. Integrally casting tensile

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	NASA-STD-6016B Requirement Statement	Justification	Section Title
	conducted to ensure foundry control of cast lots.	bars with each production lot is a standard practice in the casting industry.
4.2.4.3f	[MPR 123] Periodic cut-ups or functional testing shall be conducted for Class 1 and Class 2 castings (classes as defined by SAE AMS2175A (2010)).	For the most critical castings, periodic cut-ups or functional testing ensures that the casting process is under control and is reproducible.	Castings
4.2.4.3g	[MPR 124] Surface and volumetric nondestructive inspection shall be performed on all safety-critical castings.	Castings frequently contain unacceptable defects that cannot be detected except by these NDE techniques.	Castings
4.2.4.4a	[MPR 125] Where formed panels are used in mission-critical applications, pre- production panels shall be subjected to first-article inspection to verify proper material integrity, minimum required mechanical properties, proper grain size, and macro/microstructure.	The process for forming panels needs to be verified as producing panels with proper material integrity, mechanical properties, grain size, and macro/microstructure.	Formed Panels
4.2.4.4b	[MPR 126] The mechanical properties of the first production article shall be compared to control coupons to show they are predictive of the properties in the body of the first article.	Control coupons will be used for verification of the properties of each production formed panel, so the first article testing needs to verify that they are representative of the formed panels.	Formed Panels
4.2.4.4c	[MPR 127] The same forming practice and heat-treating procedure shall be used for the production formed panels as for the approved first-article panels.	Ensures to a reasonable degree that the production formed panels will have the same mechanical properties and metallurgical characteristics as the first-article formed panel that validated the process.	Formed Panels
4.2.4.4d	[MPR 128] Sectioning to determine mechanical properties at control areas shall be repeated after any substantive change in the forming technique, as determined by M&P analysis.	Since any changes in the forming process can significantly alter the grain flow, microstructure, and mechanical properties of the panel, it is necessary to conduct a first- article-like cut-up to ensure the resulting properties are still acceptable.	Formed Panels
4.2.4.4e	[MPR 129] Surface and volumetric NDI shall be performed on all safety-critical formed panels.	Formed panels may contain unacceptable defects that cannot be detected except by these NDE techniques.	Formed Panels
4.2.4.5a	[MPR 130] Structural adhesive bonding shall meet MSFC-SPEC-445A, Adhesive Bonding, Process and Inspection, Requirements for.	MSFC-SPEC-445 is the NASA- wide, accepted specification for adhesive bonding and inspection.	Adhesive Bonding
4.2.4.5b	[MPR 131] Structural adhesive bonding processes shall be controlled to prevent contamination that would cause structural failure that could affect the safety of the mission, crew, or vehicle or affect mission success.	The strength of a structural adhesive bond can be significantly degraded by bond surface contamination, particularly from silicones. Therefore, it is imperative that bond surface cleanliness be maintained and verified.	Adhesive Bonding
4.2.4.5c	[MPR 132] Bonded primary structural joints shall demonstrate cohesive failure modes in shear.	In bonded joints, cohesive failure is the desired failure mode that best ensures a reliable and repeatable shear strength design value, whereas an adhesive failure is almost always the result of inadequate surface preparation and is indicative of an improper bonding process.	Adhesive Bonding
4.2.4.6a	[MPR 133] If alternative specifications to	The AWS D17 welding	Welding
	those cited in this section are utilized or developed, those specifications shall meet the requirements of NASA-STD-5006A, General Welding Requirements for Aerospace Materials.	requirements are commonly used by aerospace companies, but are far more detailed than is required by NASA and also have some deficiencies. If an aerospace company chooses to not follow the AWS D17 requirements, plus the supplemental requirements identified in this document, they may develop their own standard, provided that it complies with the less detailed but more stringent requirements in NASA-STD-5006.
4.2.4.6b	[MPR 134] Material Review Board disposition shall be required for weld repair/rework/processing activities that are not in accordance with the approved weld process specification (WPS).	Weld procedure variances that have not been qualified by prior test or experience need to be reviewed and approved by a Material Review Board and/or M&P knowledgeable personnel.	Welding
4.2.4.6c	[MPR 135] A weld development and certification plan shall be developed for large structural welded components such as crew modules and welded cryogenic tanks.	The welding of large structural welded components can have major problems if the weld development and certification plan isn’t established well before welding flight hardware. The plan is expected to include full scale pathfinder weldments, tooling development, and a design values program including sensitivity testing.	Welding
4.2.4.6.1	[MPR 136] The processing and quality assurance requirements for manual, automatic, and semiautomatic welding for	It is NASA policy to use voluntary consensus standards where possible. The cited specification is	Fusion Welding
	spaceflight applications that provide mission-critical functions shall meet the requirements of AWS D17.1/D17.1M (2010) AMD 1 (2012), Specification for Fusion Welding for Aerospace Applications, with the following modifications/additions:	the aerospace industry specification for fusion welding of aerospace structures and components and has been reviewed and approved for use by NASA with modifications/additions to address minor deficiencies in the baseline standard.
4.2.4.6.1a	[MPR 137] Mission-critical structural welds shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012), Class A requirements.	Structural welds that are safety- critical are required to meet the most stringent set of requirements, as is the case for the cited Class A requirements. Note that the weld classes in AWS D17.1 are based on nondestructive inspection requirements (with the most stringent inspection requirements for Class A welds), whereas traditional NASA weld classes were based on weld criticality (with Class A welds being the highest criticality level).	Fusion Welding
4.2.4.6.1b	[MPR 138] Other structural welds shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012), Class A or Class B requirements.	Structural welds that are not safety- critical are allowed to meet a less stringent set of requirements, as is the case for the cited Class B requirements. The more stringent Class A requirements may also be used.	Fusion Welding
4.2.4.6.1c	[MPR 139] Non-structural welds (including seal welds) shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012), Class C requirements.	Non-structural, non-safety-critical welds are allowed to meet the least stringent set of requirements, as is the case for the cited Class C	Fusion Welding

	Section Title	NASA-STD-6016B Requirement Statement	Justification requirements.
4.2.4.6.1d	Fusion Welding	[MPR 140] All Class A and Class B welds (including manual welds), as defined by AWS D17.1/D17.1M (2010) AMD 1 (2012), shall be qualified in accordance with AWS D17.1/D17.1M (2010) AMD 1 (2012).	Ensures that Class A and B manual weld processes are also qualified, not just the operators who perform the welds.
4.2.4.6.1e	Fusion Welding	[MPR 141] Titanium welds shall be light/dark straw or better (Ref. AWS D17.1/D17.1M (2010) AMD 1 (2012), Table 7.1).	Titanium welds with worse coloration than light/dark straw (as listed in AWS D17.1, Table 7.1) indicate improper inert gas control during welding and presents the possibility that weld contamination/oxidation occurred.
4.2.4.6.1f	Fusion Welding	[MPR 142] Titanium and its alloys shall be welded with alloy-matching or metallurgically compatible fillers or autogenously.	Prevents the creation of a completely unknown titanium alloy in the weldment by welding together, for example, an alpha alloy and an alpha-beta alloy or other incompatible combinations of titanium alloys.
4.2.4.6.1g	Fusion Welding	[MPR 143] Extra low interstitial (ELI) filler wires shall be used for titanium cryogenic applications and are preferred for general applications.	ELI weld wire produces a more ductile weld at cryogenic temperatures.
4.2.4.6.1h	Fusion Welding	[MPR 144] Commercially pure (CP) titanium filler shall not be used on Ti- 6Al-4V or other alloyed base material.	Welding Ti-6Al-4V with CP weld wire results in a significant alloy content and interstitial solubility difference between the parent metal and the weld metal, which can lead to interstitial migration and precipitation of embrittling species, such as titanium hydride, at the

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	Justification	Section Title	NASA-STD-6016B Requirement Statement
	weld fusion line. An Apollo SIV-B stage was destroyed due to inadvertently welding a Ti-6Al-4V pressure vessel with CP weld wire.
4.2.4.6.1i	Titanium welding needs to be performed in an inert gas environment. The cited gases are not considered inert to titanium and can react with titanium alloys at weld temperatures.	Fusion Welding	[MPR 145] Nitrogen, hydrogen, carbon dioxide, and mixtures containing these gases shall not be used in welding titanium and its alloys.
4.2.4.6.1i(1)	Ensures that the water content in the inert gas is low enough that the potential reaction of water with titanium at the welding temperature is not a problem.	Fusion Welding	[MPR 146] The inert gas shall have a dew point of -60º C (-76º F) or lower.
4.2.4.6.1j	Welding of titanium alloys may produce residual stresses that are high enough that the weld needs to be stress relieved. Such stress relief treatments are done at high enough temperatures that precautions against oxidation of the titanium must be exercised.	Fusion Welding	[MPR 147] Welded alpha and alpha plus beta titanium alloys shall be stress relieved in a vacuum or inert gas environment (Ar or He), or stress relieved in air with verification of oxide removal per SAE AMS-H-81200D or SAE AMS 2801B, or certified in the as-welded condition.
4.2.4.6.1k	Titanium beta alloys are less prone to residual stress during welding, but stress relieving of beta welds still needs to be considered.	Fusion Welding	[MPR 148] Titanium beta alloys that are welded shall be evaluated on a case-by- case basis with respect to stress relief.
4.2.4.6.1l	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are industry standards for laser welding that have been reviewed and approved for use by NASA.	Fusion Welding	[MPR 149] Laser welding for spaceflight hardware that provides mission-critical functions shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012) or AWS C7.4/C7.4M (2008), Process Specification and Operator Qualification
			for Laser Beam Welding.
4.2.4.6.1m	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are industry standards for electron beam welding that have been reviewed and approved for use by NASA, whereas SAE AMS2681 is another common industry standard for electron beam welding but is not considered adequate.	Fusion Welding	[MPR 150] Electron beam welding for spaceflight hardware that provides mission-critical functions shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012) or SAE AMS2680C, Electron- Beam Welding for Fatigue Critical Applications.
4.2.4.6.1n	Although permitted by AWS D17.1/D17.1M (2010) AMD 1 (2012), these practices have been determined to be unacceptable for NASA Program hardware unless approved through an MUA review process.	Fusion Welding	[MPR 151] The following welding practices permitted by AWS D17.1/D17.1M (2010) AMD 1 (2012) shall not be used without an approved MUA to document the acceptance rationale: (1) Welding from both sides if full penetration of the first pass is not verified (either by inspection of the back side or by grinding prior to welding on the opposite side). (2) Partial weld penetration in structural welds. (3) Straightening operation after welding. (4) Lap welds in structural applications.
4.2.4.6.1o	NASA review of the cited AWS specification has determined that these additional requirements need to be included in the WPS to ensure the adequacy and completeness of the welding process.	Fusion Welding	[MPR 152] The Welding Procedure Specification (WPS) shall include the following content in addition to that required by AWS D17.1/D17.1M (2010) AMD 1 (2012): (1) Prequalified rework welds in accordance with AWS D17.1/D17.1M (2010) AMD 1 (2012). (2) Testing and documentation of allowable parameter variations for automatic and semi- automatic welds. (3) Manual welding parameters. (4) An associated Procedure Qualification Record (PQR) with tension testing and macro- examination results as part of weld qualification requirements.
4.2.4.6.2	It is NASA policy to use voluntary consensus standards where possible. The cited specification is an industry standard for resistance welding that has been reviewed and approved for use by NASA. Resistance welding is not used frequently in NASA programs and then primarily for spot welding of batteries.	Resistance Welding	[MPR 153] Resistance welding for spaceflight hardware that provides mission-critical functions, including resistance spot welding (RSW), shall meet the requirements of AWS D17.2/D17.2M (2013), Specification for Resistance Welding for Aerospace Applications.
4.2.4.6.3	It is NASA policy to use voluntary consensus standards where possible	Friction-Stir Welding of	[MPR 154] Friction-stir welding of aluminum alloys that are
	consensus standards where possible. The cited specification is an industry standard for friction-stir welding that has been reviewed and approved for use by NASA.	Aluminum Alloys	aluminum alloys for spaceflight hardware that provides mission-critical functions shall meet the requirements of AWS D17.3/D17.3M (2016), Specification for Friction Stir Welding of Aluminum Alloys for Aerospace Applications.
4.2.4.6.4	Although this standard is no longer active, it has been determined by NASA that it is the best standard existing for establishing and controlling the inertia welding process. Other active inertia friction welding specifications were reviewed and found to be inadequate for NASA’s use or were only recommended practices, not requirements.	Inertia Welding	[MPR 155] Inertia welding for spaceflight hardware that provides mission-critical functions shall meet the requirements of MIL-STD-1252 (1975), Inertia Friction Welding Process, Procedure and Performance Qualification.
4.2.4.6.4a	Generally, inertia welds are performed on pressure-containing tube and valve components that require inspection for surface and subsurface defects that could develop into leaks.	Inertia Welding	[MPR 156] Surface inspection (penetrant) and volumetric inspection (radiography) shall be performed.
4.2.4.6.4b	Generally, inertia welds are performed on pressure-containing tube and valve components that need to be proof tested.	Inertia Welding	[MPR 157] All welds shall be proof tested.
4.2.4.6.4c	Generally, inertia welds are performed on pressure-containing tube and valve components that need to be helium leak tested.	Inertia Welding	[MPR 158] Inertia welds used in fluid systems shall be helium leak tested.
4.2.4.7a	It is NASA policy to use voluntary consensus standards where	Brazing	[MPR 159] Brazing for spaceflight hardware that provides mission-critical
	possible. The cited specification is the industry specification for the design, manufacture, and examination of critical brazed components and has been reviewed and approved for use by NASA.		functions shall be conducted in accordance with AWS C3.3 (2008), Recommended Practices for Design, Manufacture, and Examination of Critical Brazed Components.
4.2.4.7b	It is NASA policy to use voluntary consensus standards where possible. The cited specification is an industry specification for brazing of aluminum that has been reviewed and approved for use by NASA.	Brazing	[MPR 160] Brazing of aluminum alloys for spaceflight hardware that provides mission-critical functions shall meet the requirements of AWS C3.7M/C3.7 (2011), Specification for Aluminum Brazing.
4.2.4.7c	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are the industry specifications for the types of brazing cited, and they have been reviewed and approved for use by NASA.	Brazing	[MPR 161] Torch, induction, and furnace brazing for spaceflight hardware that provides mission-critical functions shall meet the requirements of AWS C3.4M/C3.4 (2016), Specification for Torch Brazing; AWS C3.5M/C3.5 (2016), Specification for Induction Brazing; and AWS C3.6M/C3.6 (2016), Specification for Furnace Brazing, respectively.
4.2.4.7d	Need to ensure that high- temperature processing near existing braze joints does not degrade the joint by re-melting and reflowing the braze metal.	Brazing	[MPR 162] Subsequent fusion-welding operations in the vicinity of brazed joints or other operations involving high temperatures that might affect the brazed joint shall be prohibited for spaceflight hardware that provides mission-critical functions unless it can be demonstrated that the fixturing, processes, methods, and/or procedures employed will preclude
			degradation of the brazed joint.
4.2.4.7e	Since transferring all of the load in a structural braze joint by tension through the braze material is not considered as reliable as transferring the load by shear through the braze material, tension loaded braze material has to be avoided. This is accomplished by using lap joint designs versus butt joints, where the large brazed surface area in a lap joint results in a relatively low shear stress in the braze material.	Brazing	[MPR 163] Brazed joints used in spaceflight hardware that provides mission-critical functions shall be designed for shear loading and not be relied upon for strength in axial loading for structural parts.
4.2.4.7f	It is NASA policy to use voluntary consensus standards where possible. The cited standard is an industry standard for how to evaluate the strength of brazed joints and has been reviewed and approved for use by NASA.	Brazing	[MPR 164] The shear strength of brazed joints used in spaceflight hardware that provides mission-critical functions shall be evaluated in accordance with AWS C3.2M/C3.2 (2008), Standard Method for Evaluating the Strength of Brazed Joints.
4.2.4.7g	Destructive testing ensures the adequacy of the braze flow and that the braze alloy and its adhesion to the brazed surfaces are continuous and metallurgically acceptable.	Brazing	[MPR 165] For furnace brazing of complex configurations of spaceflight hardware that provides mission-critical functions, such as heat exchangers and cold plates, destructive testing shall be conducted on pre-production brazed joints to verify that the braze layer that extends beyond the fillet area is continuous and forms a uniform phase.
4.2.4.8	The low strength and poor creep resistance of solders means	Structural Soldering	[MPR 166] Soldering shall not be used for structural applications.
	soldering cannot be used in structural applications.
4.2.4.9a	Since the high surface temperature associated with the EDM and LM machining processes causes undesirable surface damage, it is necessary to control the processes to ensure the depth of damage is predictable and manageable.	Electrical Discharge Machining and Laser Machining	[MPR 167] Electrical-discharge machining (EDM) and laser machining (LM) processes for spaceflight hardware that provides mission-critical functions shall be controlled to limit the depth of the oxide layer, the recast layer, and the heat-affected zone.
4.2.4.9a(1)	The oxide layer that is produced by the EDM and LM machining processes is generally brittle and crack-prone and must be removed.	Electrical Discharge Machining and Laser Machining	[MPR 168] The oxide layer shall be removed from the surface.
4.2.4.9a(2)	The recast layer and the heat-affected zone produced by the EDM and LM machining processes are generally undesirable conditions, and need to be removed from wear surfaces and fracture-critical surfaces.	Electrical Discharge Machining and Laser Machining	[MPR 169] In addition, the recast layer and the heat-affected zone shall be removed from bearing, wear, fatigue or fracture-critical surfaces, and from crack- or notch-sensitive materials.
4.2.4.9b	Qualification of the EDM and LM machining processes ensures that the depth of the detrimental surface conditions is predictable and the amount of material that must be removed is known.	Electrical Discharge Machining and Laser Machining	[MPR 170] EDM/LM schedules for spaceflight hardware that provides mission-critical functions shall be qualified to determine the maximum thickness of the affected layers when the depth of the affected material must be known for removal or analysis.
4.2.4.10a	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are industry standards for electroplating nickel that have been reviewed and approved for use by	Nickel Plating	[MPR 171] Electrodeposited nickel plating for spaceflight hardware that provides mission-critical functions shall be applied according to the requirements of SAE AMS2403N, Plating, Nickel General Purpose; SAE AMS2423E,
	NASA.		Plating, Nickel Hard Deposit; or ASTM B689-97, Standard Specification for Electroplated Engineering Nickel Coatings.
4.2.4.10b	It is NASA policy to use voluntary consensus standards where possible. The cited specifications are industry standards for electroless nickel plating that have been reviewed and approved for use by NASA.	Nickel Plating	[MPR 172] Electroless nickel plate for spaceflight hardware that provides mission-critical functions shall be applied per SAE AMS2404G, Plating, Electroless Nickel, or ASTM B733-15, Standard Specification for Autocatalytic (Electroless) Nickel-Phosphorus Coatings on Metal.
4.2.4.10c	Since there is a strong galvanic couple between nickel and aluminum, the interface between them must be protected in corrosive environments.	Nickel Plating	[MPR 173] The nickel-aluminum interface in nickel-plated aluminum used in spaceflight hardware that provides mission-critical functions shall be protected from exposure to corrosive environments.
4.2.4.11	Ensures that the NASA M&P technical authority reviews and approves this relatively new manufacturing technique that is not yet controlled by NASA or consensus standards or specifications.	Additive Manufacturing	[MPR 174] When structural hardware is manufactured by additive manufacturing techniques, a manufacturing and qualification plan shall be submitted to NASA and approved by the responsible NASA M&P and design organizations.
4.2.5.1a	Since there are many NDE techniques that are available to inspect spaceflight hardware and such inspections are performed throughout the life cycle of the hardware, a Plan is needed to	Nondestructive Evaluation (NDE) Plan	[MPR 175] The NDE Plan shall address the process for establishment, implementation, execution, and control of NDE through design, manufacturing, operations, and maintenance of spaceflight hardware.
	ensure that the application and sequence of the inspections are proper and efficient.
4.2.5.1b	Obviously, since NASA-STD-5009 sets NDE requirements for fracture- critical components, it has to be included in the NDE Plan for spaceflight hardware. Although MIL-HDBK-6870 states that it is for guidance only, it is to be treated as a requirements document for non-fracture-critical components in the NDE Plan.	Nondestructive Evaluation (NDE) Plan	[MPR 176] The Plan shall meet the intent of MIL-HDBK-6870B, Nondestructive Inspection Program Requirements for Aircraft and Missile Materials and Parts, and, when fracture control is applicable, the requirements of NASA-STD-5009, Nondestructive Evaluation Requirements for Fracture-Critical Metallic Components.
4.2.5.1c	NAS-410 has long been the aerospace industry and NASA standard for certifying and qualifying NDE personnel. Although ASNT-TC-1A is also widely used for certifying and qualifying nondestructive test personnel, it is a recommended practice only.	Nondestructive Evaluation (NDE) Plan	[MPR 177] Qualification and certification of personnel involved in nondestructive testing shall comply with NAS 410 (Revision 4), NAS Certification and Qualification of Nondestructive Test Personnel.
4.2.5.2a	Machining metal surfaces causes a disturbed, smeared thin layer of metal that masks cracks, and the disturbed layer must be removed by etching prior to penetrant inspection.	NDE Etching	[MPR 178] All machined or otherwise mechanically disturbed surfaces on metallic parts that are to be fluorescent dye-penetrant inspected shall be adequately etched to assure removal of smeared, masking material prior to penetrant application, with the following exceptions: (1) Previously etched parts do not need etching if the surface has
			not been smeared since the last etching. (2) When supporting rationale is provided, close tolerance parts may be machined near-final and etched and penetrant inspected before final machining in lieu of etching and penetrant inspecting after final machining.
4.2.5.2b	Since the disturbed metal layer has a certain depth, it is required that at least an amount of metal equal to the thickness of the disturbed layer be removed to have an effective penetrant inspection.	NDE Etching	[MPR 179] The etching procedure shall specify the minimum amount of material to be removed to ensure that smeared metal does not mask cracks.
4.2.5.2c	Since not being able to etch means that the penetrant inspection process would be ineffective, it is necessary to employ other NDE methods, such as eddy current or ultrasonics that can detect the critical flaw size.	NDE Etching	[MPR 180] If etching is not feasible, it shall be demonstrated that the required flaw size can be reliably detected.
4.2.6.1a	Many metal processes produce residual stresses that can add to the stresses from the normally applied service stresses resulting in a higher stress state that needs to be addressed and accounted for in the analyses. An important lesson was learned in	Residual Stresses	[MPR 181] Estimates of residual stresses in structural or stress-corrosion-sensitive hardware shall be considered in structural analyses and corrosion/stress-corrosion assessments.
	repair welding of 2195 Aluminum Lithium. The repair weld material was very low strength compared to the parent material (it was 4043 weld wire in 2195 parent material), so the effect of the residual stress was not apparent (i.e., there wasn’t much warpage). However, when wide specimens containing only a short repair-welded area were tensile tested, the residual stress added to the discontinuity shear stress at the weld/parent metal interface, causing early failure.
4.2.6.1b	Since straightening of a warped part requires plastic deformation and creates the potential for mechanical damage and high residual stresses, it is necessary for M&P engineering to review and approve the process.	Residual Stresses	[MPR 182] The straightening of warped parts in structural hardware shall require an approved MUA.
4.2.6.2a	Vents on vented sandwich constructions must be large enough to prevent the pressure differential between the sandwich interior and the external environment during launch becoming large enough to cause face-sheet separation.	Sandwich Assemblies	[MPR 183] For vented sandwich architectures, the differential pressure on ascent shall be adequately relieved to preclude core-to-face sheet bond line “peel” failure modes that could result in face-sheet separation.
4.2.6.2b	Non-vented sandwich structure will develop a pressure differential during ascent of one atmosphere (14.7 psia) between the sandwich interior and space vacuum. The structure must be verified to	Sandwich Assemblies	[MPR 184] Sandwich architectures that are not vented shall be capable of withstanding pressure buildup without violating strength and stability requirements.
	withstand this pressure differential without causing face-sheet separation.
4.2.6.2c	Since moisture can accumulate in vented core sandwich assemblies and can migrate from cell to cell presenting a corrosion problem, and since it can also cause structural failure in a launch environment where the exterior pressure decrease and the rapid temperature increase can flash the water to vapor causing over pressurization of the face sheet, it is necessary to prevent such water entrapment and migration in sandwich assemblies. An Apollo lesson learned occurred when a vented core honeycomb interstage tank structure failed due to moisture accumulation and flash evaporation during launch. Water intrusion can occur from condensation / exposure of sandwich constructions to humid air below the dew point, “cryo- pumping” where sandwich structures are exposed to large thermal / humidity gradients induced by cryogenic fluid storage, gravitational migration of rain water, forced penetration of wind- driven rain, etc.	Sandwich Assemblies	[MPR 185] Structural sandwich assemblies shall be designed to prevent the entrance and entrapment of water vapor or other contaminants into the core structure.
4.2.6.2d	The use of metallic or glass- reinforced cores to minimize	Sandwich Assemblies	[MPR 186] Structural honeycomb sandwich assemblies that will be
	intrusion of moisture intrusion into the sandwich core and can result in vapor build-up over-pressurizing the sandwich assembly when it is heated by the launch environment.		subjected to heating shall be tested for the expected environments to show that the construction can withstand them.
4.2.6.2e	To prevent moisture accumulation that can cause internal corrosion and structural failure, sandwich assemblies need to be protected from water intrusion prior to launch in addition to their being designed to prevent moisture intrusion.	Sandwich Assemblies	[MPR 187] Sandwich assemblies using perforated and moisture-absorbing cores shall be protected from water intrusion during assembly and prelaunch activities.
4.2.6.2f	It is NASA policy to use voluntary consensus standards where possible. The cited standard is an industry-accepted standard for sandwich construction and core materials that has been reviewed and approved for use by NASA.	Sandwich Assemblies	[MPR 188] Test methods for structural sandwich constructions and core materials shall meet the requirements of SAE AMS-STD-401 (1999), Sandwich Constructions and Core Materials: General Test Methods.
4.2.6.3	Ensures that corrosion of space spaceflight hardware will be addressed and eliminated to the extent possible.	Corrosion Prevention and Control	[MPR 189] All parts, assemblies, and equipment used in spaceflight hardware that provides mission-critical functions, including spares, shall be finished to provide protection from corrosion in accordance with the requirements of NASA-STD-6012, Corrosion Protection for Space Flight Hardware, with the following exceptions: a. SAE AMS2404G is permitted for electroless nickel plating as an alternative to ASTM B733-15.
			b. Titanium fasteners may be used in contact with graphite composites, provided that they are wet installed with sealant or primer materials.
4.2.6.3.1	Ensures that the normally present passive layer that provides corrosion resistance is restored.	Passivation	[MPR 190] Corrosion-resistant steels used in spaceflight hardware that provides mission-critical functions shall be passivated after machining.
4.2.6.3.2	Ensures that corrosive environments cannot intrude into sealed compartments.	Sealing	[MPR 191] Removable panels and access doors in exterior or interior corrosive environments shall be sealed either by mechanical seals or by separable, faying- surface sealing.
4.2.6.4a	Ensures that the potential for hydrogen embrittlement of metallic structures containing hydrogen will be evaluated.	Hydrogen Embrittlement	[MPR 192] When designing liquid or gaseous hydrogen systems, the degradation of metallic materials properties by hydrogen embrittlement shall be addressed in the Materials and Processes Selection, Control, and Implementation Plan.
4.2.6.4b	Ensures that the potential for hydrogen embrittlement of metallic structures containing hydrogen has been evaluated.	Hydrogen Embrittlement	[MPR 193] An MUA shall be written rationalizing the selection of metallic materials for liquid or gaseous hydrogen systems to preclude cracking and to ensure system reliability and safety.
4.2.6.4c	Ensures that the potential for hydrogen embrittlement is minimized by controlling the time of exposure to electrochemical processes and to acids and bases or that the embrittlement relief treatments are carried out soon after exposure to such processes and	Hydrogen Embrittlement	[MPR 194] Electrochemical processes or exposure to acids or bases during manufacturing or processing of spaceflight hardware that provides mission-critical functions shall be controlled to prevent hydrogen embrittlement, or embrittlement relief treatment shall be performed promptly
	fluids.		after processing.
4.2.6.4c(1)	Baking ensures relief of the potential for hydrogen embrittlement in steels after exposure to acids or plating processes. The baking process in SAE AMS2759/9D has a long history and is the accepted way to relieve the potential for hydrogen embrittlement in steels.	Hydrogen Embrittlement	[MPR 195] When acid cleaning baths or plating processes are used on steel parts for spaceflight hardware that provides mission-critical functions, the part shall be baked in accordance with SAE AMS2759/9D, Hydrogen Embrittlement Relief (Baking) of Steel Parts, to alleviate potential hydrogen embrittlement problems.
4.2.6.5.1	A lesson learned when indium solder in a Goddard satellite contacted gold wires and the resulting gold-indium intermetallic resulted in severe embrittlement and failure of the gold wires.	Gold-Indium Intermetallic Formation	[MPR 196] Gold shall not be used in contact with indium or an indium alloy (such as indium solder) for spaceflight hardware that provides mission-critical functions.
4.2.6.5.2a	Ensures that the dissolution of gold into tin-based solder will be minimized, thereby minimizing the formation of brittle gold-tin intermetallic compounds.	Gold-Tin Intermetallic Formation	[MPR 197] When used with tin-based solder joints in mission-critical hardware, gold shall be removed from at least 95 percent of the surface to be soldered of all component leads, component terminations, and solder terminals.
4.2.6.5.2b	Ensures that the propensity for fretting corrosion between gold- coated contacts and mating tin- coated contacts will be avoided.	Gold-Tin Intermetallic Formation	[MPR 198] Gold-coated contacts shall not be mated with tin or tin alloy-coated contacts for separable contact interfaces in mission-critical hardware.
4.2.6.5.3a	An aluminum-rich intermetallic phase, the AuAl2 intermetallic known as “Purple Plague,” is inherent (and not necessarily harmful) to gold-aluminum brazing; however, if excessive heat is applied, too much of the	Gold-Aluminum Intermetallic Formation	[MPR 199] Gold-aluminum brazing processes shall be controlled to minimize formation of the AuAl2 intermetallic known as “Purple Plague” and prevent formation of the Au5Al2 intermetallic known as “White Plague.”
	intermetallic can form at the braze/part interface, causing joint failure. A gold-rich phase, the Au5Al2 intermetallic known as “White Plague,” is always detrimental. Its formation is catalyzed by silicon, so care should be taken to keep the braze joint zone free of contamination
4.2.6.5.3b	intermetallic can form at the braze/part interface, causing joint failure. A gold-rich phase, the Au5Al2 intermetallic known as “White Plague,” is always detrimental. Its formation is catalyzed by silicon, so care should be taken to keep the braze joint zone free of contamination	Gold-Aluminum Intermetallic Formation	[MPR 200] Gold-aluminum bonding processes shall be controlled to prevent the formation of the AuAl2 intermetallic known as “Purple Plague” and the Au5Al2 intermetallic known as “White Plague.”
4.2.6.6.1	Imposing NASA-STD-5020 ensures that issues and problems associated with the use of liquid- locking compounds will be avoided.	Liquid Locking Compounds	[MPR 201] If a liquid-locking compound is used as a locking feature where rotational loosening or disengagement would result in a critical or catastrophic hazard, its use shall comply with the design and quality requirements and best
			practices in NASA-STD-5020, Requirements for Threaded Fastening Systems in Spaceflight Hardware, sections 5.5 and 7.6 and Appendix C.
4.2.6.6.2	Ensures that the friable black oxide that forms on silver exposed to atomic oxygen will not contaminate other components or mechanisms or contaminate astronaut suits and gloves during EVAs.	Silver-Plated Fasteners	[MPR 202] Silver-plated fasteners shall not be used in external applications where the silver plating is directly exposed to atomic oxygen for a period longer than 2 weeks.
4.2.6.7a	Spacecraft performance can be severely degraded by the presence of excessive contamination, leading to loss of mission or mission objectives. A Contamination Control Plan covering the entire hardware life cycle is necessary to ensure that a “build clean” approach is properly implemented. The cited ASTM standard is a voluntary consensus standard that has been reviewed and approved for use by NASA.	Contamination Control	[MPR 203] A Contamination Control Plan shall be generated in accordance with the guidelines of ASTM E1548 (2009), Standard Practice for Preparation of Aerospace Contamination Control Plans.
4.2.6.7b	The cited areas of material processing, hardware packaging, and FOD prevention are the most important types of contamination that need to be controlled and addressed in the Contamination Control Plan.	Contamination Control	[MPR 204] The Contamination Control Plan shall include controls on contamination-sensitive manufacturing processes, such as adhesive bonding, controls on packaging for shipment and storage, and a foreign object damage/debris (FOD) prevention program.
4.2.6.7c	Since damage to spaceflight hardware from FOD can be very	Contamination Control	[MPR 205] The FOD prevention program shall be established for all ground
	expensive and impact schedules, it is mandatory that a FOD prevention program be established that addresses all sources and means of preventing FOD damage throughout the hardware’s life cycle.		operations of mechanical and electrical systems of spaceflight hardware, including the design, development, manufacturing, assembly, repair, processing, testing, maintenance, operation, and checkout of the equipment to ensure the highest practical level of cleanliness.
4.2.6.7d	It is NASA policy to use voluntary consensus standards where possible. NAS 412 is the aerospace industry standard for FOD prevention and has been reviewed and approved for use by NASA.	Contamination Control	[MPR 206] The FOD prevention program shall conform to NAS 412 (Revision 1), Foreign Object Damage/Foreign Object Debris (FOD) Prevention.
4.2.6.7e	Ensures that the cleanliness level of each hardware item has been specified.	Contamination Control	[MPR 207] Cleanliness levels for assembly- and subassembly-level hardware shall be identified on the engineering drawings.
4.2.6.8	Obviously, spaceflight hardware needs to be protected during shipment and storage.	Packaging	[MPR 208] Packaging shall protect spaceflight hardware from corrosion and contamination during shipping and storage.
4.3	Since it is NASA policy to provide clear, concise, verifiable requirements, it is necessary for NASA M&P to have oversight and insight into how the contractors and their subs and vendors implement and control spaceflight hardware M&P. The cited steps of reviewing and approving the contractors’ M&P plans, requiring M&P drawing signature approval,	Verification	[MPR 209] Verification of compliance with the requirements of this NASA Technical Standard shall consist of the following steps as a minimum: a. NASA approval of the contractor Materials and Processes Selection, Control, and Implementation Plan and other applicable materials data requirements documents, such as the Contamination Control Plan and NDE
	conducting audits, membership on the materials control panel, and approval of MUAs and MIULs provide NASA M&P with the oversight and insight necessary to verify that the requirements of NASA-STD-6016 have been met.		Plan. b. Contractor M&P signature on engineering drawings to verify compliance with the requirements of this NASA Technical Standard or the Materials and Processes Selection, Control, and Implementation Plan. c. NASA audits of contractor M&P activities relating to hardware design and manufacturing. d. Establishment and operation of the M&P control panel in accordance with section 4.1.4 of this NASA Technical Standard. e. NASA approval of MUAs. f. NASA approval of MIULs.

NASA-STD-6016B APPENDIX B TYPICAL MUA FORM

MATERIAL USAGE AGREEMENT C: <empty> USAGE AGREEMENT NO.: <empty> REV: <empty> PAGE: <empty> OF: <empty>

PROJECT:: <empty> SYSTEM:: <empty> CATEGORY:: <empty> ORIGINATOR:: <empty> ORGANIZATION/CONTRACTOR:: <empty>

PART NUMBER(S):: <empty> USING ASSEMBLY(S):: <empty> ITEM DESCRIPTION:: <empty> ISSUE:: <empty>

MATERIAL DESIGNATION:: <empty> MANUFACTURER:: <empty> SPECIFICATION:: <empty> PROPOSED EFFECTIVITY:: <empty>

MATERIAL CODE:: <empty> LOCATION:: <empty> ENVIRONMENT:: <empty>

THICKNESS:: <empty> WEIGHT:: <empty> EXPOSED AREA: <empty> HABITABLE [ ] NONHABITABLE [ ] PRESSURE PSIA:: <empty> TEMP.F:: <empty> MEDIA:: <empty>

APPLICATION:: <empty>

RATIONALE: (use second page if required.): <empty>

DATE: <empty> APPROVE: <empty> REJECT: <empty> DEFER: <empty> EFFECTIVITY:: <empty>

ORIGINATING CONTRACTOR: <empty>

NASA-STD-6016B APPENDIX C

RETIRED CATEGORY III MUA RATIONALE CODES

RETIRED FLAMMABILITY RATIONALE CODES

	RATIONALE
101	Approved Materials Usage Agreement (MUA) Category I.
102	Approved Materials Usage Agreement (MUA) Category II.
103	Materials passed requirements when tested in configuration.
104	Unexposed, overcoated, or sandwiched between nonflammable materials and no ignition source or propagation path.
105	Minor usage (less than 45 g (0.1 lb) mass and 13 cm2 (2 in2) surface area); no propagation path or ignition source.
107	Passes test No. 10 of NASA-STD-6001, Flammability Test for Materials in Vented Containers, by test or analysis.
108	Off-the-shelf equipment having material acceptable in configuration; no ignition source or propagation path.
109	Material not exposed; totally immersed in fluid; evaluated for fluid compatibility only.
110	Material is acceptable when used on a metal substrate that provides a good heat sink. Material considered noncombustible in this configuration by test or analysis.
111	Material is flammable but is sandwiched between nonflammable materials with edges only exposed and is more than 5 cm (2 in) from an ignition source or more than 30 cm (12 in) from other flammable materials.
112	Material is flammable but is unexposed or is overcoated with a nonflammable material.
113	Material is flammable but has a thickness less than 0.25 mm (0.010 in) and is sprayed or bonded to a metallic surface greater than 1.6 mm (0.062 in) thick.
114	Material is flammable but is used in “small amounts” and is more than 5 cm (2 in) from an ignition source or more than 30.5 cm (12 in) from other flammable materials. “Small amounts” for flammability may be quantified as follows: total weight less than 45 g (0.1 lb) and less than 13 cm2 (2.0 in2) surface area.

RETIRED TOXICITY (OFFGASSING) RATIONALE CODES

	RATIONALE
201	Approved Material Usage Agreement (MUA) Category I.
202	Meets toxicity requirements with performed cure.
203	T value for material/component in usage weight is <0.5 for manned flight compartment volume.
204	Materials usage in hermetically sealed container.

RETIRED FLUID SYSTEM COMPATIBILITY RATIONALE CODES

	RATIONALE
301	Approved Material Usage Agreement (MUA) Category I.
302	Passes requirements in configuration.
303	Material is B-rated in MAPTIS (batch/lot testing required) but batch/lot used in hardware passed test.
304	Approved Material Usage Agreement (MUA) Category II.

RETIRED THERMAL VACUUM STABILITY RATIONALE CODES

	RATIONALE
401	Approved Material Usage Agreement (MUA) Category I.
402	Approved Material Usage Agreement (MUA) Category II.
403	VCM between 0.1 and 1.0 percent; exposed area is less than 13 cm2 (2 in2) and not near a critical surface.
404	VCM >1.0 percent; exposed area is less than 1.6 cm2 (0.25 in2).
405	Unexposed, overcoated, or encapsulated with approved material.
406	Material is B-rated in MAPTIS (batch/lot testing required) but batch/lot used in hardware cured to meet requirements.
407	Meets thermal vacuum stability requirements in configuration.
408	Materials usage in hermetically sealed container.
409	Material has VCM >0.1 percent but is enclosed in a sealed container (maximum leak rate less than 1 x 10-4 cm3/sec).

RETIRED STRESS CORROSION CRACKING RATIONALE CODES

	RATIONALE
501	Approved Material Usage Agreement (MUA) Category I.
502	Approved Material Usage Agreement (MUA) Category II.
503	Maximum tensile stress <50 percent of yield strength for part on electrical/electronic assemblies.
504	Martensitic or PH stainless steels used in ball bearing, race, or similar applications where the primary loading is compressive.
505	Metal not listed in table 1 of MSFC-STD-3029 for stress corrosion cracking is not exposed to a corrosive environment after final assembly through end-item use.
506	Carbon and low alloy high strength steels greater than 1240 MPa (180 ksi) used in ball bearings, springs, or similar applications where primary loading is compressive, low tensile stresses, or history of satisfactory performance.

RETIRED CORROSION RATIONALE CODES

	RATIONALE
601	Approved Material Usage Agreement (MUA) Category I.
602	Approved Material Usage Agreement (MUA) Category II.
603	Adequately finished for corrosion protection.
604	Acceptable in use environment.
606	Electrical grounding required, cladding plus conversion coating adequate.
607	Thermal conductance and electrical bonding requirements preclude painting. Conversion coating is adequate (for aluminum only).
608	Finished on a higher assembly.
609	Laminated shim - minimum exposure of corrosion resistant material.
610	Material does not meet the requirements of MSFC-SPEC-250, Class II, but is treated or coated in a manner which meets or exceeds the requirements of MSFC-SPEC-250. Actual surface treatment shall be listed.
611	Material does not meet the requirements of MSFC-SPEC-250, Class II, but is not exposed to a corrosive environment after final assembly through end-item use.
612	Welding of titanium alloy-to-alloy or commercially pure-to-alloy using commercially pure filler metal in mixed alloy welds where hydrogen embrittlement is not predicted in service.

RETIRED GENERAL CODES

	RATIONALE
702	Generic materials controlled by military or industry specification using MAPTIS averages for ratings or test results. Material codes for generic material shall be used.
703	Military specification or industry specification allowing several material options where all options have acceptable ratings.

End of chapter · NextAppendix D Recommended data requirements documents