4.2Detailed Requirements

4.2.1.1 Flammability Control

[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.
• 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 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

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.

Many situations arise where flammable materials are used in an acceptable manner without test, using mitigation practices and the MUA approval system. Guidelines for hardware flammability assessment and mitigation can be found in JSC 29353, Flammability Configuration Analysis for Spacecraft Applications.

• i. Flammability configuration analysis in accordance with JSC 29353 shows hardware configuration to be self-extinguishing.

Material flammability ratings and tests based on NASA-STD-6001B for many materials are found in the MAPTIS database.

4.2.1.2 Toxic Offgassing

• a. [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.
• b. [MPR 46] Spacecraft maximum allowable concentration (SMAC) values shall be obtained from JSC 20584 (2008), Spacecraft Maximum Allowable Concentrations for Airborne Contaminants, or from MAPTIS for compounds for which no SMAC values are found in JSC 20584.

4.2.1.3 Fluid Compatibility (Fluids Other Than Oxygen)

• a. [MPR 47] Materials exposed to hazardous fluids¹ shall be evaluated or tested for compatibility.

NASA-STD-6001B, Test 15, is a 48-hour screening test for short-term exposure to fuels and oxidizers.

NASA-STD-6016B

• b. [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.

The test program needs to demonstrate that materials will not degrade unacceptably during the actual time of hazardous fluid exposure, including ground processing prior to flight, flight exposure, and post-flight processing prior to decontamination. Use of elevated temperatures to accelerate the test and/or extrapolation of minor degradation from the test duration to the actual time of hazardous fluid exposure are acceptable approaches.

NASA-STD-6001B, Supplemental Test A.7, identifies changes resulting from incidental exposure (minor amounts, such as a splash) to fuels or oxidizers.

• c. [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.
• d. [MPR 50] The effect of material condition (for example, parent versus weld metal or heat-affected zone) shall be addressed in the compatibility determination.
• e. [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.
• (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.
• f. [MPR 212] Materials used in other oxidizer systems shall be evaluated for potential ignition.

4.2.1.4 Oxygen Compatibility

• a. [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.

Material flammability ratings and tests based on NASA-STD-6001B for many materials are found in the MAPTIS database.

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

b. [MPR 52] When materials in such 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.

As described in NASA-STD-6001B, material/component configurational testing may be required to support the compatibility assessment.

Compressed air systems and pressurized systems containing enriched oxygen are inherently less hazardous than systems containing pure oxygen; the hazard increases with oxygen concentration and pressure.

Guidelines on the design of safe oxygen systems are contained in ASTM MNL 36, Safe Use of Oxygen and Oxygen Systems: Handbook for Design, Operation, and Maintenance; ASTM G88, Standard Guide for Designing Systems for Oxygen Service; ASTM G63, Standard Guide for Evaluating Nonmetallic Materials for Oxygen Service; ASTM G94, Standard Guide for Evaluating Metals for Oxygen Service; and NASA/TM-2007-213740, Guide for Oxygen Compatibility Assessments on Oxygen Components and Systems.

4.2.1.4.1 Oxygen Component Acceptance Test

[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 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.

This test is an acceptance test, not a qualification test, and is required for all sets of spaceflight hardware. Retesting is required if post-test actions (such as rework, repair, or interfacing with hardwate having uncontrolled cleaniness) invalidate the acceptance test.

LOX system components should also be acceptance tested, but the system hardwate acceptance test is normally adequate for this purpose. The number of test cycles should be as defined in the system hardware acceptance test (typically well below the ten cycles required for gaseous systems).

4.2.1.5 Electrical Wire Insulation Materials

• a. [MPR 54] Electrical wire insulation materials shall be evaluated for flammability in accordance with NASA-STD-6001B, Test 4 or Test 1.

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

Test 4 should be used for insulation on electrical power wiring where the maximum current passing through the wiring can raise the temperature above 49 °C (120 °F). Either test can be used for signal/data wiring and other wiring where the current is too small to raise the temperature above this value.

b. [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).

4.2.2.1 Aluminum

Aluminum alloys used in structural applications should be resistant to general corrosion, pitting, intergranular corrosion, and stress corrosion cracking.

[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).

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

Grain boundary precipitation above 66 °C (150 °F) can create stress-corrosion sensitivity.

4.2.2.2 Steel

4.2.2.2.1 Drilling and Grinding of High-Strength Steel

When performing drilling, grinding, reaming, or machining of steels, low-stress machining techniques with coolant should be used. Uncontrolled high-stress machining used without coolant is detrimental to the steel microstructure. Untempered martensite is formed that leads to cracking. Overheating can soften the steel due to overtempering or decarburization. Low-stress machining practices, such as those in SAE AMS2453, Low Stress Grinding of Steel Parts Heat Treated to 180 ksi or Over, and Low Stress Grinding of Chrome Plating Applied to Steel Parts Heat Treated to 180 ksi or Over, should be used to prevent damage to the microstructure.

• a. [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.
• b. [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:
• (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).

4.2.2.2.2 Corrosion-Resistant Steel

• a. [MPR 60] Unstabilized, austenitic steels shall not be processed or used above 371 °C (700 °F) in spaceflight hardware that provides mission-critical functions.
• b. [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.
• c. [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.

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

4.2.2.2.3 Ductile-Brittle Transition Temperature

[MPR 63] Steels shall not be used in tension below their ductile-brittle transition temperature.

Alloy and carbon steels, as well as ferritic, martensitic, and duplex steels tend to become brittle as the temperature is reduced. Balls in ball bearings are generally acceptable because they are in compression.

4.2.2.3 Titanium

4.2.2.3.1 Titanium Contamination

• a. [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.

Hydrochloric acid, chlorinated solvents, chlorinated cutting fluids, fluorinated hydrocarbons, and anhydrous methyl alcohol can all produce stress corrosion cracking. Mercury, cadmium, silver, and gold have been shown to cause liquid-metal-induced embrittlement and/or solid-metal-induced embrittlement in titanium and its alloys. Liquid-metal-induced embrittlement of titanium alloys by cadmium can occur as low as 149 °C (300 °F), and solid-metal-induced embrittlement of titanium alloys by cadmium can occur as low as room temperature.

b. [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.

Contaminated zones and layers may be formed as a result of mill processing, heat treating, and elevated temperature-forming operations.

4.2.2.3.2 Titanium Wear

Titanium and its alloys exhibit very poor resistance to wear. Fretting that occurs at interfaces with titanium and its alloys has often contributed to crack initiation, especially fatigue initiation. The preferred policy is a design that avoids fretting and/or wear with titanium and its alloys. Bolted joints are not considered to fret.

[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.

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

4.2.2.3.3 Titanium Flammability

• a. [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).
• b. [MPR 68] Titanium alloys shall not be machined inside spacecraft modules during ground processing or in flight, because machining operations can ignite titanium turnings and cause fire.

4.2.2.4 Magnesium

• a. [MPR 69] Magnesium alloys shall not be used in primary structure or in other areas of spaceflight hardwate that provides mission-critical functions that are subject to wear, abuse, foreign object damage, abrasion, erosion, or where fluid or moisture entrapment is possible.
• b. [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.

4.2.2.5 Beryllium

Beryllium alloys are exceptionally lightweight but are not often used because of the extreme toxicity of beryllium salts and oxides, and because it is unusually susceptible to damage. Alloys containing up to 4 percent beryllium by weight are not an issue, provided they are not machined inside spacecraft crew compartments.

• a. [MPR 71] Alloys containing more than 4 percent beryllium by weight shall not be used for primary structural applications.
• b. [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.
• c. [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.
• d. [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.

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

Chemical/milling and etching are recognized successful processes for removal of the damaged layer.

• e. [MPR 75] Beryllium-containing alloys (including alloys containing less than 4 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.

Stripping, crimping, and cutting electrical wire are not considered to be machining operations.

• f. [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.

4.2.2.6 Cadmium

Cadmium is highly toxic, can sublime and cause outgassing contamination at elevated temperatures in vacuum, and can cause liquid metal embrittlement of titanium and some steel alloys. Cadmium plating can spontaneously form cadmium whiskers.

• a. [MPR 77] Cadmium shall not be used in crew or vacuum environments.
• b. [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.

4.2.2.7 Zinc

[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.

Metallic zinc is less volatile than cadmium but should not be used in vacuum environments where the temperature/pressure environment could cause contamination of optical surfaces or electrical devices.

4.2.2.8 Mercury

a. [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.

Mercury has the potential for causing liquid-metal embrittlement.

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

• b. [MPR 81] Spaceflight hardware (including fluorescent lamps) containing mercury shall have three levels of containment to prevent mercury leakage.

Nonflight lamps containing mercury, such as those used in hardware ground processing and fluorescent dye penetrant inspection of flight parts, are acceptable, provided that the bulbs are protected by a nonshatterable, leak-proof outer container.

4.2.2.9 Refractory Metals

[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.

Engineering data on refractory alloys are limited, especially under extreme environmental conditions of spacecraft.

4.2.2.10 Superalloys (Nickel-Based and Cobalt-Based)

In these requirements, a superalloy refers to a nickel- or cobalt- based alloy that retains all or most of its strength at usage temperatures approaching 538 °C (1000 °F) and higher. Examples include both wrought products such as Inconel 718 and castings such as Mar-M-247 and Inconel 625.

• a. [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.

Some of the precipitation-hardening superalloys are susceptible to alloying element depletion at the surface in a high temperature, oxidizing environment. Nickel-based superalloys are susceptible to grain boundary cracking phenomena at elevated temperatures, such as Stress Assisted Grain Boundary Oxidation Embrittlement (SAGBOE), and to stress rupture during heat treatment.

• b. [MPR 84] The reduction to design 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.

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

4.2.2.11 Tin

High-purity tin and tin plating are susceptible to two degradation phenomena: tin whisker growth and tin pest (sometimes known as tin plague). The following requirements for tin whisker mitigation may not be adequate to prevent tin pest from occurring:

• a. [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.
• b. [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 a die attach material or as a package sealing material.
• (3) Tin alloys containing less than 20 percent tin by weight.
• c. [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.

At this time, insufficient data exist to define hard requirements for controlling tin pest. Alloying with more than 5 percent lead or smaller quantities of antimony or bismuth significantly slows tin pest formation but may not eliminate it completely. Alloying recommendations based on weight percentage are:

• (1) Not less than 5 percent lead.
• (2) Not less than 0.3 percent bismuth.
• (3) Not less than 0.5 percent antimony.
• (4) Not less than 3.5 percent silver.
• d. [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.

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

Tin plating has been demonstrated to have oxidation, wear, cold-weld, fretting, and tin whisker issues that reduce performance and reliability. Tin plating may be accepted through the MUA process, provided appropriate controls (contact lubricant, stabilant, etc.) are used to mitigate the performance and/or reliability risk.

4.2.3.1 Elastomeric Materials

• a. [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.
• b. [MPR 90] Elastomeric materials used in spaceflight hardware that provides mission-critical functions, other than those used in off-the-shelf parts, such as cable clamps, shall be cure-dated for tracking purposes.
• c. [MPR 91] Room temperature vulcanizing (RTV) silicones that liberate acetic acid during cure shall not be used because they can cause corrosion.
• d. [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.

Elastomers do not seal effectively below their glass-transition temperature, and the ability of elastomers to seal decreases significantly as the glass-transition temperature is approached. The high glass-transition temperature of Viton® rubbers was a major contributor to the loss of the Space Shuttle Challenger, and other elastomers may struggle to retain sealing performance at spacecraft temperatures.

4.2.3.2 Polyvinyl Chloride

[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 offgasses unacceptably at temperatures above 49 °C (120 °F). It is also normally flammable in pressurized areas of spacecraft; so, when it is used, its flammability has to be controlled in accordance with section 4.2.1.1.

Polyvinyl chloride is not vacuum-compatible.

Rigid polyvinyl chloride tubing is not acceptable for pressurized gas applications, because it can have a brittle failure when used for pressurized gas transport.

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

4.2.3.3 Composite Materials

Requirements and guidance on materials design values for polymer matrix, ceramic matrix, and metal matrix composites are in section 4.1.8. Requirements on sandwich assemblies are in section 4.2.6.2.

4.2.3.4 Lubricants

NASA-TM-86556, Lubrication Handbook for the Space Industry, Part A: Solid Lubricants, Part B: Liquid Lubricants, provides guidance on the evaluation and selection of lubricants for spaceflight systems and components. Guidelines on additional lubricants not listed in NASA-TM-86556 are contained in NASA/CR-2005-213424, Lubrication for Space Applications. Lubricants containing chloro-fluoro components may react with newly exposed rubbing surfaces of aluminum, magnesium or titanium alloys, especially at elevated temperatures.

4.2.3.5 Limited-Life Items

[MPR 94] All materials shall be selected to meet the useful life of the hardware with no maintenance or be identified as limited-life items requiring maintainability.

4.2.3.6 Thermal Vacuum Stability

• a. [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.

Many materials contain absorbed water, but the loss of absorbed water does not normally affect functionality; so the WVR (a measure of the total water vapor lost in the ASTM E595-15 test) is subtracted from the TML.

More stringent requirements may be needed for materials that are line of sight to contamination-sensitive surfaces on spacecraft or attached vehicles. Contamination-sensitive surfaces include windows, lenses, star trackers, solar arrays, radiators, and other surfaces with highly controlled

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

optical properties. The approach taken depends on the specific program needs but may include lowering the CVCM requirement to ≤0.01 percent CVCM, use of optical surfaces in testing to characterize the effects of deposition, or measurement of outgassing deposition rates as functions of source and target temperature (the standard test method is ASTM E1559, Standard Test Method for Contamination Outgassing Characteristics of Spacecraft Materials) and calculating deposition on critical surfaces using an integrated vehicle model. Spacecraft with cryogenic optics may be sensitive to water vapor deposition; in such cases, the WVR would not be subtracted from the TML.

• b. [MPR 96] With the following exceptions, hardware items (components, assemblies, etc.) containing materials that fail the CVCM requirement and/or having unidentified materials shall be vacuum baked at the maximum tolerable temperature of the component, 10 °C 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 CVM 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. Determination of acceptable molecular outgassing limits, selection of the bakeout method and temperature, and determination of the specific bakeout completion criteria are the responsibility of the program/project. ASTM E2900, Standard Practice for Spacecraft Hardware Thermal Vacuum Bakeout, is recommended as a guide for performing thermal vacuum bakeout.

A vacuum-bake temperature of 125 °C (257 °F) (the ASTM E595-15 screening temperature) may damage some spaceflight hardware. Temperature requirements for hardware thermal vacuum bakeout should be adjusted to prevent damage to components. Because bakeout time and efficiency are dependent on temperature, the chosen temperature should be the highest possible without damage to hardware.

A hardware functionality bench test should be performed to re-verify performance after baking.

4.2.3.7 External Environment Survivability

[MPR 97] The critical properties of materials exposed in the spacecraft external environment shall meet operational requirements for their intended life-cycle exposure.

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

Applicable space environments include atomic oxygen, solar ultraviolet (UV) radiation, ionizing radiation, plasma, vacuum, thermal cycling, and contamination. Applicable planetary environments, such as dust and planetary atmospheres, may also apply. Meteoroids and orbital debris should also be considered in the analysis of long-term degradation.

4.2.3.8 Fungus Resistance

[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.
• c. Materials are used inside sealed containers (maximum leak rate less than 1 x 10-4 cm3/sec) with internal container humidity less than 60 percent relative humidity (RH) at ambient conditions.
• d. Materials are used inside electrical boxes where the temperature is always greater than or equal to the ambient cabin temperature.
• e. Materials have edge exposure only.
• f. Materials are normally stowed with no risk of condensation in stowage locations.
• g. Materials are used on noncritical, off-the-shelf electrical/electronic hardware that is stowed and/or used in crew areas.
• h. Materials are fluorocarbon polymers (including ETFE) or silicones.
• i. Materials are used in crew clothing items.

When fungus-nutrient materials have to be used (and none of the above exception criteria are met) and an MUA is submitted for approval, supporting rationale should include how materials are treated to prevent fungus growth, using a fungus treatment that does not adversely affect unit performance or service life, does not constitute a health hazard to higher order life, and is not leached out by the use environment. 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

4.2.3.9 Glycols

[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 be added to the solution to prevent spontaneous ignition from the reaction of silver with the glycol.

This reaction is known to occur for ethylene and propylene glycol. Solutions containing other glycols may be exempted if testing is conducted to demonstrate that the spontaneous ignition reaction does not occur.

4.2.3.10 Etching Fluorocarbons

• a. [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.

Adhesion to the fluorocarbon surface is required for electrical wire or cable insulated or coated with fluorocarbons intended for potting if mechanical bond strength and environmental sealing are a design requirement.

b. [MPR 101] Etched surfaces shall be processed within 24 hours or within 1 year if packaged per SAE AMS2491F.

The open end of the wire should not be exposed to the etchant.

4.2.4.1 Heat Treatment

• a. [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.
• b. [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, or SAE AMS2759E, Heat Treatment of Steel Parts, General Requirements.
• c. [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

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

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.

• d. [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.

Care should be taken in design of heat treatment fixturing and in heat treatment procedures for nickel- and cobalt-based alloy parts to avoid the introduction of excessive tensile stress. Inconel 718 parts have failed through SAGBOE and stress rupture from stresses induced by as little as an abrupt change in part thickness combined with introduction of a “cold” part into a hot furnace.

These AMS heat treatment specifications require hardness and/or conductivity measurements to verify the adequacy of the heat treatment process. In many cases, hardness tests are inadequate for spacecraft hardware; and testing of process-control tensile-test coupons is required.

• e. [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-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.

Note: Representative tensile test coupons are preferred over hardness and conductivity measurements for aging of aluminum alloys.

• f. [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.

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

• g. [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 the part.

h. [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.

Procurement specifications, heat treat specifications, and MMPDS all limit product size of metal stock to ensure full consolidation of cast structure, product uniformity, and consistent mechanical properties.

4.2.4.2 Forging

Because mechanical properties are optimum in the direction of material flow during forging, forging techniques should be used that, whenever possible, produce an internal grain-flow pattern such that the direction of flow is essentially parallel to the principal stresses. If forging techniques do not allow for an internal grain-flow pattern such that the direction of the flow is parallel to the principal stresses, parts should be designed such that the weakest grain flow direction is not in line with the principal stresses. The forging pattern should be essentially free from re-entrant and sharply folded flow lines.

• a. [MPR 110] Where forgings are used in mission-critical applications, first-article (preproduction) approval shall be obtained from the procuring authority.

First article requirements apply to all forgings, including hand forging and spin forming. The first article requirements may be waived for hand forging by review of the forging plan through the MUA process.

• b. [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.

Mechanical properties testing requirements may include fracture, durability, or damage tolerance testing.

• c. [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).
• d. [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 in the body of the first article.

NASA-STD-6016B

• e. [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.

The information gained from this effort is utilized to redesign the forging as necessary.

• f. [MPR 115] These data and results of tests on the redesign shall be retained and made available for review by the procuring activity.
• g. [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.
• h. [MPR 117] Surface and volumetric nondestructive inspection (NDI) shall be performed on all safety-critical forgings.

4.2.4.3 Castings

• a. [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.
• b. [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.
• c. [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.

Mechanical properties testing requirements may include fracture, durability, or damage tolerance testing.

• d. [MPR 121] The same casting practice and heat-treating procedure shall be used for the production castings as for the approved first-article castings.

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

• e. [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 conducted to ensure foundry control of cast lots.
• f. [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)).
• g. [MPR 124] Surface and volumetric nondestructive inspection shall be performed on all safety-critical castings.

4.2.4.4 Formed Panels

Barrel and gore panels with complex geometries and integral stiffeners are often formed by processes such as roll forming, brake forming, peen forming, stretch forming, and explosive forming and then heat treated. The mechanical and thermal forming processes can result in strength and toughness variations across the panel.

• a. [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.
• b. [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.

Mechanical properties testing requirements may include fracture, durability, or damage tolerance testing.

• c. [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.
• d. [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.
• e. [MPR 129] Surface and volumetric NDI shall be performed on all safety-critical formed panels.

Given the complexities of some formed panels, volumetric NDI may be performed as a raw stock inspection.

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

4.2.4.5 Adhesive Bonding

• a. [MPR 130] Structural adhesive bonding shall meet MSFC-SPEC-445A, Adhesive Bonding, Process and Inspection, Requirements for.

The Operator Certification Plan and the Adhesive Control Plan may be described in the implementing process specification in place of submittal to NASA for approval.

Retesting of adhesives used for production parts is not required if they are within shelf life.

• b. [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 sensitivity of structural adhesive bonds to contamination is of particular concern. Bond sensitivity studies should be conducted to verify that the required adhesive properties are maintained after exposure to potential contaminant species and concentrations, and in-process cleanliness inspections should be conducted as part of the bonding process. Silicone contamination is a particular concern because it severely degrades adhesive performance.

• c. [MPR 132] Bonded primary structural joints shall demonstrate cohesive failure modes in shear.

4.2.4.6 Welding

The design selection of parent materials and weld methods should be based on consideration of the weldments, including adjacent heat-affected zones, as they affect operational capability of the parts concerned. Peaking and mismatch limits should be considered in the mechanical properties.

Welding procedures should be selected to provide the required weld quality, minimum weld energy input, and protection of the heated metal from contamination.

The suitability of the equipment, processes, welding supplies, and supplementary treatments selected should be demonstrated through qualification testing of welded specimens representing the materials and joint configuration of production parts.

Working with the responsible Technical Authority and NASA M&P organization, Programs/Projects will assess if it is more beneficial to directly use welding specifications cited in this section or develop alternatives. The requirements in this section provide compliance with NASA-STD-5006A, General Welding Requirements for Aerospace Materials. If the requirements in this section are met, no further application of NASA-STD-5006A is needed.

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

• a. [MPR 133] If alternative specifications to 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.

Note: NASA-STD-5006A defines weld classes in terms of weld criticality (with Class A representing safety-critical structural welds) and AWS D17.1/D17.1M AMD 1 defines weld classes in terms of inspection requirements (with Class A representing the most stringent nondestructive inspection requirements).

• b. [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).

Examples include:

• Wrong filler metal used.
• • More than two attempts performed at the same location on heat-sensitive materials.
• • More than five attempts performed at the same location on materials that are not heat sensitive.
• Repair weld required after the weldment has been post-weld heat treated.
• Repair weld required after final machining has been completed.
• Repair extends outside the original weld zone.
• Weldment has been direct aged.
• • Repairs following proof or leak test.
• c. [MPR 135] A weld development and certification plan shall be developed for large structural welded components such as crew modules and welded cryogenic tanks.

This plan should include full scale pathfinder weldments, tooling development, and a design values program including sensitivity testing.

4.2.4.6.1 Fusion Welding

[MPR 136] The processing and quality assurance requirements for manual, automatic, and semiautomatic welding for 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:

Note: AWS D17.1/D17.1M (2010) AMD 1 (2012) defines weld classes in terms of inspection requirements (with Class A representing the most stringent nondestructive inspection requirements).

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

• a. [MPR 137] Mission-critical structural welds shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012), Class A requirements.
• b. [MPR 138] Other structural welds shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012), Class A or Class B requirements.
• c. [MPR 139] Non-critical welds (including seal welds) shall comply with AWS D17.1/D17.1M (2010) AMD 1 (2012), Class C requirements.
• d. [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).
• AWS G2.4/G2.4M, Guide for the Fusion Welding of Titanium and Titanium Alloys, should be used for guidance on titanium welding.
• e. [MPR 141] Titanium welds shall be light/dark straw or better (Ref. AWS D17.1/D17.1M (2010) AMD 1 (2012), Table 7.1).
• f. [MPR 142] Titanium and its alloys shall be welded with alloy-matching or metallurgically compatible fillers or autogenously.

Ti-6Al-4V filler metal is considered to be metallurgically compatible with Ti-3Al-2.5V.

• g. [MPR 143] Extra low interstitial (ELI) filler wires shall be used for titanium cryogenic applications and are preferred for general applications.
• h. [MPR 144] Commercially pure (CP) titanium filler shall not be used on Ti-6Al-4V or other alloyed base material.

Hydride formation can occur in the weld, which can produce a brittle, catastrophic failure.

• i. [MPR 145] Nitrogen, hydrogen, carbon dioxide, and mixtures containing these gases shall not be used in welding titanium and its alloys.

Care needs to be exercised to ensure complete inert gas (argon or helium) coverage during welding.

• (1) [MPR 146] The inert gas shall have a dew point of -60° C (-76° F) or lower.
• j. [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

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

oxide removal per SAE AMS-H-81200D or SAE AMS2801B, or certified in the as-welded condition.

• k. [MPR 148] Titanium beta alloys that are welded shall be evaluated on a case-by-case basis with respect to stress relief.

Note: Stress relief of weldments does not require tensile test coupons.

1. [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.

• m. [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.

Compliance with SAE AMS2680C is preferred but compliance with AWS D17.1/D17.1M (2010) AMD 1 (2012) is acceptable.

• n. [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.
• o. [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.

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

• (4) An associated Procedure Qualification Record (PQR) with tension testing and macro-examination results as part of weld qualification requirements.

Qualification strength requirements should be included in weld qualification test acceptance. Qualification strength values are not the same as materials design values.

4.2.4.6.2 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 Friction-Stir Welding of Aluminum Alloys

[MPR 154] Friction-stir welding of 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 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.

• a. [MPR 156] Surface inspection (penetrant) and volumetric inspection (radiography) shall be performed.
• b. [MPR 157] All welds shall be proof tested.
• c. [MPR 158] Inertia welds used in fluid systems shall be helium leak tested.

Acceptable helium leak rates should be defined by the system.

4.2.4.7 Brazing

• a. [MPR 159] Brazing for spaceflight hardware that provides mission-critical functions shall be conducted in accordance with AWS C3.3 (2008), Recommended Practices for Design, Manufacture, and Examination of Critical Brazed Components.

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

• b. [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.
• c. [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.
• d. [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.
• e. [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.
• f. [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.
• g. [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 Structural Soldering

[MPR 166] Soldering shall not be used for structural applications.

4.2.4.9 Electrical Discharge Machining and Laser Machining

• a. [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.
• (1) [MPR 168] The oxide layer shall be removed from the surface.
• (2) [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.

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

The recast layer and heat-affected zone may be left on a part if an engineering evaluation shows that they are not of consequence to the required performance of the part.

• b. [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.

4.2.4.10 Nickel Plating

• a. [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, Plating, Nickel Hard Deposit; or ASTM B689-97, Standard Specification for Electroplated Engineering Nickel Coatings.
• b. [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.
• c. [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.

Nickel and aluminum form a strong galvanic cell at the nickel-aluminum interface, and exposure of the aluminum alloy to a corrosive environment can produce rapid disbonding of the nickel plate.

4.2.4.11 Additive Manufacturing

Guidelines documents and standards for additive manufacturing are in development at this time. The requirements of this NASA Technical Standard on M&P controls, materials design values, metallic and nonmetallic materials, and nondestructive inspection apply to hardware manufactured by additive techniques, just as they do for traditional manufacturing techniques. For nonstructural, nonmetallic 3-D printed hardware, controlled and verified processes are essential; but other M&P aspects like flammability, toxic offgassing, and vacuum outgassing also apply, just as for any other nonmetallic material.

[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.

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

Key aspects of producing structural metallic hardware by additive manufacturing techniques, such as direct metal laser sintering (DMLS) and selective laser melting (SLM), include proper development of structural design values and controlled processes, although other requirements, such as stress-corrosion resistance and corrosion control, also apply. Verification of appropriate process control should include first article inspection to verify proper material properties and macro/microstructure and mechanical property testing of integrally manufactured specimens from each hardware unit.

4.2.5.1 Nondestructive Evaluation (NDE) Plan

• a. [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.
• b. [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.

In case of conflict between the requirements of the two standards, the requirements of NASA-STD-5009 are applicable. It is expected that fracture-critical and safety-critical parts have surface and volumetric inspections unless there is rationale that it is not necessary. The need for internal (volumetric) inspection depends on the application and on materials characteristics such as thickness, product form, and other factors. Internal inspection requirements and methods should be determined early in the design process so that proper flaw screening is accomplished.

• c. [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.

4.2.5.2 NDE Etching

• a. [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.

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

• (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.
• b. [MPR 179] The etching procedure shall specify the minimum amount of material to be removed to ensure that smeared metal does not mask cracks.

Minimum material removal requirements for fracture-critical hardware are identified in NASA-STD-5009.

• c. [MPR 180] If etching is not feasible, it shall be demonstrated that the required flaw size can be reliably detected.

4.2.6.1 Residual Stresses

Residual tensile stresses are induced into manufactured parts as a result of forming, machining, heat treating, welding, special metal-removal processes, or the straightening of warped parts. Residual stresses may be harmful in structural applications when the part is subjected to fatigue loading, operation stresses, or corrosive environments. Special consideration should be given to the possibility of high residual stresses that might develop during off-nominal processing, such as repeated weld repairs in the same location. Residual stresses should be controlled or minimized during the fabrication sequence by special treatments, such as annealing and stress relieving. Bending and forming of tubing (for example, stainless steel and titanium tubing) is typically not stress relieved.

• a. [MPR 181] Estimates of residual stresses in structural or stress-corrosion-sensitive hardware shall be considered in structural analyses and corrosion/stress-corrosion assessments.

Residual stresses should be quantified by an appropriate technique (such as x-ray diffraction).

• b. [MPR 182] The straightening of warped parts in structural hardware shall require an approved MUA.

4.2.6.2 Sandwich Assemblies

Structural sandwich assemblies using honeycomb or open-cell core constructions which are exposed to ascent-aerodynamic heating and/or vacuum exposure may be either vented to relieve internal pressure, or non-vented.

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

• a. [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.

Associated material and process controls should demonstrate that production of full-scale flight articles produces venting performance that is both repeatable and reliable.

• b. [MPR 184] Sandwich architectures that are not vented shall be capable of withstanding pressure buildup without violating strength and stability requirements.
• c. [MPR 185] Structural sandwich assemblies shall be designed to prevent the entrance and entrapment of water vapor or other contaminants into the core structure.
• d. [MPR 186] Structural honeycomb sandwich assemblies that will be subjected to heating shall be tested for the expected environments to show that the construction can withstand them.
• e. [MPR 187] Sandwich assemblies using perforated and moisture-absorbing cores shall be protected from water intrusion during assembly and prelaunch activities.
• f. [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.

Because the ability of NDE to detect core-to-face sheet bond line defects is limited, the potential presence of defects makes it necessary to address both the static strength and cyclic loading capability under combined loading environments. Such environments can include pressure, temperature, and externally applied loads. Elevated temperatures from panel heating can soften adhesives and cause failure.

Acceptance / proof testing of each flight article should address the critical combined loading environment, including rapid decompression.

4.2.6.3 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.

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

• 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 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 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.

4.2.6.4 Hydrogen Embrittlement

Hydrogen embrittlement through hydrogen-metals interactions can be classified into three broad categories: hydrogen environmental embrittlement (HEE), internal hydrogen embrittlement (IHE), and hydrogen reaction embrittlement (HRE). In general, hydrogen environmental embrittlement represents the condition where the materials are exposed to a high-pressure gaseous hydrogen environment. The hydrogen for internal hydrogen embrittlement is usually not from a high-pressure gaseous system, but from an electrochemical process such as electroplating, corrosion, cathodic charging, and even from thermal charging. The hydrogen for internal hydrogen embrittlement can also come from moisture and enter the metals during welding, casting, and solidification processes from the foundry.

The hydrogen environmental embrittlement and internal hydrogen embrittlement effects are similar in many instances and they both require an external applied stress in order for the hydrogen embrittlement effects to occur. In contrast, hydrogen reaction embrittlement is usually irreversible hydrogen damage caused by a chemical reaction with hydrogen and such damage can occur without an external applied stress. This form of hydrogen damage can occur in materials such as titanium, zirconium, and some types of iron or steel-based alloys.

Overall, hydrogen embrittlement of materials is not very well understood and only limited materials property data have been generated. NASA/TM-2016-218602, Hydrogen Embrittlement, provides a detailed description of hydrogen embrittlement mechanisms, a comprehensive summary of existing hydrogen embrittlement data, and guidance on selection of materials and control of processes to prevent hydrogen embrittlement. It is a companion document to ANSI/AIAA G-095, Guide to Safety of Hydrogen and Hydrogen Systems. a. [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.

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

b. [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.

Test data may have to be generated in a simulated environment to support the rationale.

• c. [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 after processing.
• (1) [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 Intermetallic Compounds

When joining two or more metals together (soldering, wire bonding, etc.), intermetallic compounds can form that are problematic to the metal joint, leading to a weakened mechanical or electrical connection. Problems can arise from embrittlement, changes in density, or conductivity of the intermetallic compound. Gold-indium and gold-tin are two of the more common systems that form intermetallic compounds that will degrade a metal joint.

4.2.6.5.1 Gold-Indium Intermetallic Formation

Indium reacts with gold to form a succession of gold-indium intermetallic compounds. The gold-indium intermetallic compounds occupy a volume approximately four times the original volume of the consumed gold and are brittle. The original mechanical and thermal properties of gold are degraded by this intermetallic reaction possibly resulting in unreliable electrical interconnections. The gold-indium intermetallic formation progresses significantly even at room temperature and accelerates at elevated temperatures.

[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.

4.2.6.5.2 Gold-Tin Intermetallic Formation

The combination of gold and tin is subject to two distinct degradation mechanisms: gold embrittlement of tin-based solder joints and fretting corrosion of gold-coated contacts when mated with tin or tin alloy-coated contacts.

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

When tin-based solder is used to join gold-plated surfaces, the gold coatings will dissolve into the final solder joint. If the concentration of gold in the tin-based solder joint exceeds approximately 3 percent by weight, then a brittle gold-tin intermetallic compound may form within the joint that can impact the long-term reliability of the solder joint.

• a. [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.

A double tinning process or dynamic solder wave may be used for gold removal prior to mounting the component on the assembly.

The combination of one gold-coated surface and one tin-coated or tin-alloy-coated surface for separable contact interfaces (e.g., electrical connector contact pairs) is highly prone to fretting corrosion, which can build up an electrically insulating tin-oxide film at the contact interface, resulting in increased contact resistance.

• b. [MPR 198] Gold-coated contacts shall not be mated with tin or tin alloy-coated contacts for separable contact interfaces in mission-critical hardware.

4.2.6.5.3 Gold-Aluminum Intermetallic Formation

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 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.

• a. [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.”

Purple Plague is a significant problem in microelectronic applications, because as purple plague forms, it reduces in volume. This creates cavities in the metal surrounding the purple plague, which increases electrical resistance and structurally weakens the wire bonding. White plague is worse, because it has low electrical conductivity and, if enough of it forms, the resulting electrical resistance can cause a total failure of the component.

• b. [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.” 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

4.2.6.6 Fastener Installation

4.2.6.6.1 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 Silver-Plated Fasteners

Silver reacts rapidly with atomic oxygen to generate a loose, friable, black oxide that can cause contamination and affect the operation of mechanisms.

[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.

4.2.6.7 Contamination Control

Contamination, if not adequately anticipated and controlled, can result in loss of spacecraft, performance degradation, mission degradation, and/or loss or injury of flight crew. Examples of contamination sources and mechanisms include:

• • Particulate and molecular contamination from both ground processing and on-orbit migration may degrade the performance of optical devices, thermal control surfaces, and solar arrays.
• • Particulate contamination accumulated during ground processing or generated during operations may interfere with mechanisms, bearings, and seals; may plug or restrict fluid orifices or filters; and may pose a hazard to the crew.
• Particulate and nonvolatile residues pose an ignition hazard within systems containing oxidizers, such as oxygen or nitrogen tetroxide. Cleaning agents intended to remove contaminants from these systems may leave reactive residues if not adequately controlled and removed.
• • Microbial contamination within life support systems and ultrapure water systems may degrade these systems.
• • Control of terrestrial microbial contamination and extraterrestrial material may be required in accordance with NASA Procedural Requirements (NPR) 8020.12 on Planetary Protection Provisions for Robotic Extraterrestrial Missions.
• a. [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.

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

• b. [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.
• c. [MPR 205] The FOD prevention program shall be established for all ground 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.
• d. [MPR 206] The FOD prevention program shall conform to NAS 412 (Revision 1), Foreign Object Damage/Foreign Object Debris (FOD) Prevention.
• e. [MPR 207] Cleanliness levels for assembly- and subassembly-level hardware shall be identified on the engineering drawings.

4.2.6.8 Packaging

[MPR 208] Packaging shall protect spaceflight hardware from corrosion and contamination during shipping and storage.