Volume 13 - Corrosion Part 15 pptx

36 Hydrogen embrittlement of a type 440A stainless steel valve seat from an orbiter solenoid latching valve.. Because these springs are made of high-strength steel alloys, stress corros

the spacecraft The pin failed by delayed fracture under the low installation preload used with a shear fastener (170 MPa,

or 25 ksi, tension) (Fig 33)

Fig 33 Hydrogen embrittlement failure of a 300M steel orbiter nose landing gear steering collar pin The pin

was heat treated to a 1895-MPa (275-ksi) strength level The part was plated with chromium and cadmium (a) Pin showing location of failure Actual size (b) Failure origin (arrow) 9× (c) Brittle intergranular fracture face characteristic of hydrogen embrittlement Parts did not receive a hydrogen embrittlement relief bake due to processing error 1380×

titanium-Metallurgical analyses revealed that cracks occurred in three locations at the radius of the head of the pin An embrittled microstructure (rock candy in appearance) was located in these areas with no evidence of corrosion A significant portion

of the fracture face was characteristic of hydrogen embrittlement, although final failure was ductile under more rapid fracture Investigation revealed this pin was part of a lot that had been reworked to correct a plating error The pin had both chromium plating and titanium-cadmium plating on the part Chromium was plated on the pin shank and head Cadmium-titanium was plated on the head radius and internal hole A review of the records at the vendor failed to disclose any evidence of a hydrogen embrittlement relief baking step

Corrective action consisted of a reinspection and rebaking of all reworked parts Future reworks required full manufacturing planning, not just a material review disposition

The second part that failed was the lower drag brace of the main landing gear It was also made of 300M steel at the same heat-treat level, had chromium-plated wear surfaces, and had cadmium-titanium plating on other surfaces for corrosion control The brace failed under a 2-h sustained load of 950 MPa (138 ksi), or 50% of its ultimate tensile strength, during static load qualifications testing Failure analysis again determined that crack propagation was by hydrogen The initiation sites were arc burns on the part caused by accidental contact with a hand-held electrode used to ensure more uniform plating (Fig 34) All parts that had been plated with a hand-held electrode had to be restripped and inspected For these reworked parts and future parts requiring a hand-held electrode, the electrode was adequately protected by wrapping with nylon or other approved organic web fabrics This incident was also an indication that baking in this heavy section was marginal for the removal of hydrogen A complete review of baking procedures, times, and temperatures was made to ensure that no deficiencies existed

Fig 34 Hydrogen embrittlement of an orbiter landing gear lower drag brace made of 300M steel Steel was

heat treated to 1895 MPa (275 ksi) The part was plated with titanium-cadmium Wear surfaces were chromium plated (a) Drag brace showing tensile failure that occurred at 50% of ultimate tensile strength during a qualification test (b) Close-up of failure (c) A section through the initiation site showing an arc burn A, area melted by arc burn; B, untempered martensite; C, overtempered martensite; D, tempered martensite of the base material 145× (d) Fracture face away from initiation area showing intergranular failure characteristic of hydrogen embrittlement 870×

The third part to fail was the trunnion pin of the nose landing gear The pin failed again under static load and was made from 300M steel heat treated to the same levels as the other parts This part was chromium plated on the shank area and had a titanium-cadmium plating applied to threads and to other surfaces requiring corrosion protection The thread plating ranged in thickness from 5 to 7.5 m (0.2 to 0.3 mil), while other corrosion protection plating was 12.5 to 17.5 m (0.5

to 0.7 mil) thick

Failure analysis disclosed the same grain-boundary fracture characteristic of hydrogen embrittlement again with no corrosion on fracture faces The part had a few local areas of untempered martensite from grinding burns (Fig 35), and some areas had chromium plating as thick as 0.3 mm (12 mils) The drawings called for a nominal thickness of 0.06 mm (2.5 mil) Some cadmium plating solution had entered the fracture surface and had both plated and deposited salts It was believed that overheating due to grinding resulted in untempered martensite, which cracked either before or upon immersion in the plating bath Residual hydrogen left in the part after plating migrated to the crack areas when the parts were under sustained load, resulting in slow crack growth leading to failure Fracture mechanics analyses were able to show that the initial flaw size would grow by subcritical crack growth to the final size (2.5 mm deep × 2.4 mm long, or 0.100 × 0.095 in.) that failed under load

Fig 35 Hydrogen embrittlement of an orbiter nose landing gear trunnion pin Pin was made from 300M steel

heat treated to 1895 MPa (275 ksi) Wear surfaces were chromium plated, and nonwear surfaces were plated with titanium-cadmium (a) Failed trunnion pin showing fracture (arrow) Pin is loaded in shear and bending

× (b) Fracture surface and very thick chromium plating 85× (c) Fracture face showing intergranular failure propagated by hydrogen 875× (d) Localized grinding burn and untempered martensite (arrows) where cracking initiated 95×

In this case, special controls had to be placed on grinding, including lubricants, pressures, speeds, and feeds It is significant that all failures occurred during a 4- to 5-month period and that, since then (9 additional years), no subsequent failures have been noted

High-Pressure Hydrogen Valve Seat. Early shuttle orbiter flights required high-pressure hydrogen valves for the fuel cell system A solenoid valve, used successfully in the Apollo program for high-pressure helium in the reaction control system, was evaluated as a candidate valve for use with the high-pressure hydrogen Testing consisted of exposing the valve to a 16.5-MPa (2400-psi) pressure for 24 h, followed by 200 actuation cycles at 2.4 MPa (350 psi) After hydrogen testing, the valve was disassembled for metallurgical examination Hardness tests were performed and cross sections were made A valve seat made of type 440A stainless steel was found to be cracked (Fig 36)

Fig 36 Hydrogen embrittlement of a type 440A stainless steel valve seat from an orbiter solenoid latching

valve Seat is hardened to 52 HRC (a) Sectioned valve seat showing area of cracking (inside box) 9× (b) Cracking (arrows) appears to originate in hardness indentation 75× (c) Fracture surface of crack showing

intergranular nature of the failure Failure was caused by hydrogen in etching solution and residual stress at the hardness indentation, not by previous high-pressure hydrogen exposure 1440×

The crack originated from a hardness indentation, and concern was expressed regarding whether the hydrogen gaseous exposure or the metallographic examination caused the failure Using other available valve seats, unexposed to gaseous hydrogen, it could be demonstrated that the combination of the hardness indentation and the acid etchant (a mixture of HNO3, HCl, and water) was the cause of the failure Again, hydrogen was a culprit, but this time it was not from high-pressure hydrogen gas but from metallographic preparation procedures The valve was qualified for use

A Belleville spring is a convex-concave washer that stores energy when flattened It is widely used in aerospace in

bungee applications When a Belleville spring is compressed, very high tensile forces are put on its periphery Because these springs are made of high-strength steel alloys, stress corrosion and hydrogen embrittlement become real concerns Depending on the application, the springs are stacked in series, parallel, or series-parallel stackings (Fig 37)

Fig 37 Hydrogen embrittlement of alloy steel Belleville springs for the space shuttle orbiter program (a)

Illustration of spring design and stacking arrangements (b) Belleville spring that failed in service (c) Fracture face of a cadmium-plated Vascomax 300 maraging steel spring that failed from hydrogen embrittlement in saltwater immersion 1080× (d) Fracture face of cadmium-plated 6150 alloy steel Belleville spring that failed

by hydrogen embrittlement due to inadequate baking 1440×

The Belleville springs used on the space shuttle orbiter may be made from Vascomax 300 maraging steel or 6150 steel The springs have been plated with cadmium Cadmium plating is permitted because the springs are totally contained and will not be exposed to the space vacuum (see the discussion "Structural Joints and Fasteners" in this section) Testing has shown that cadmium-plated maraging steel springs will withstand 30 days of salt fog without failure, even with breaches

in the cadmium plating, but the springs fail in a 30-day saltwater exposure because of hydrogen embrittlement as a result

of cadmium cathodically protecting the steel

On several occasions, Belleville springs made of cadmium-plated 6150 steel have failed within minutes after loading In these cases, hydrogen embrittlement from the plating process is suspected On one occasion, 40 springs were replaced in a bungee When reloaded, 17 new springs failed within a short period of time Repeated baking at 190 °C (375 °F) for 23 h has not completely solved the problem The literature indicates that cadmium-plated springs should be baked at 260 °C (500 °F) for 1 h or at 230 °C (450 °F) for 4 h The current approach used on the orbiter has been to use vacuum plating, thus avoiding any exposure to hydrogen pickup during electroplating

Nickel-Tin-Plated Steel Parts. Three weeks before the first manned flight of the Apollo, a 4340 steel parachute fitting failed Metallurgical examination of the fracture face revealed a rock candy intergranular fracture typical of hydrogen embrittlement Investigation disclosed that the parachute system subcontractor specified a 3-h hydrogen embrittlement relief bake at 190 °C (375 °F) instead of the 23 h required by the Apollo contractor after application of the plated nickel-tin coating

The nickel-tin coating had originally been developed to replace cadmium plating on fasteners because cadmium plating sublimes in the vacuum of space The total coating is 5 to 10 m (0.2 to 0.4 mil) thick and is excellent for close-tolerance threads Investigation of the records at the plating shop, however, revealed that the plater performed no hydrogen relief bakeout, because the military specification for tin plating at that time did not require it The plater ignored the drawing callouts

Over 1000 different spacecraft part designs were analyzed to determine which parts needed to be inspected and/or replaced Extensive testing was performed to find the threshold stress levels of parts that had not been baked Efforts were concentrated on evaluating the highest strength, most highly loaded threaded parts first, because these have the least tolerance for hydrogen and the highest probability of failure One such configuration is shown in Fig 38 This part, made

of low-alloy steel heat treated to 1380 MPa (200 ksi) tensile strength and nickel-tin plated, failed within 7 h at a stress of

69 MPa (100 ksi) Through inspections of critical parts, torque level verification, and associated testing, the safety of flight parts was ensured No failure had been found on any flight safety critical parts

Fig 38 Hydrogen embrittlement of a low-alloy steel Apollo test part plated with nickel-tin and tested under

sustained load Nickel-tin plating is 5 to 10 m (0.2 to 0.4 mil) thick No hydrogen embrittlement relief baking was used in the test part The part was tested at 50% of its ultimate strength (1380 MPa, or 200 ksi) and failed

in less than 7 h, beginning at external threads

One may question why there were no previous part failures by hydrogen embrittlement given that the same plating procedures had been used for 3 to 4 years A plausible explanation is that hydrogen will find its way out of steel over a period of time by diffusion through plated coatings The nickel-tin coating was far more permeable to hydrogen than cadmium platings Therefore, parts that did not encounter sustained tensile loads shortly after plating were eventually relieved of hydrogen As it happened, a single supplier had plated nearly 9000 spacecraft parts by the time this problem surfaced Many of these parts were already installed into the Apollo vehicles under production Fortunately, the problem could be resolved without vehicle disassembly or scrapping of parts

Oxygen Ignition

Oxygen is widely used in spacecraft operations as either liquid oxygen or gaseous oxygen Liquid oxygen will react under certain impact conditions with nearly any metal Metals such as titanium and magnesium are relatively easy to ignite, while aluminum and stainless steels require considerably more energy to ignite and are used in LOX tubing, valves, and pressure vessel designs Inconel alloy 718 is one of the most resistant metals to LOX ignition and is widely used in orbiter LOX or GOX applications In gaseous oxygen, both mechanical impact and pneumatic impact can cause ignition (see the discussion "Main Propulsion System" in this section) Organic materials also readily react with oxygen under impact conditions When ignition occurs, the part is often so badly melted that no direct identification of the cause can be determined, and the causes of failure can only be inferred from detailed analyses of design and operating or test parameters

Extravehicular Mobility Unit. A fire destroyed an extravehicular test unit and space suit at the NASA Johnson Spacecraft Center in Houston Although the location of the ignition was pinpointed, the cause of the fire could never be positively identified Particle impact, as well as design and manufacturing defects, could not be ruled out Aluminum and 300-series stainless steel were used in the design The aluminum was severely burned

The oxygen flow control valve controls the flow of hot gaseous oxygen at 280 °C (540 °F) and at 31 MPa (4500 psi)

to the external tank of the space shuttle system Two explosions occurred during testing of this valve In both cases, the valve was a victim of the explosion, not the cause The first, in January 1977, took place 7 min into a flow test of the valve at a test facility A facility check valve, which was acting as a shutoff valve against 31-MPa (4500-psi) oxygen, failed at ambient temperatures The valve was made of type 316 stainless steel with a Stellite ball on the valve stem The ensuing ignition damaged the stainless steel flow control valve (Fig 39) The cause was suspected to be contamination in the system

Fig 39 Oxygen ignition of a type 316 stainless steel check valve that occurred during testing a shuttle orbiter

LOX flow control valve (a) The check valve (right) ignited during test at 31 MPa (4500 psi) oxygen pressure Also shown is the LOX flow control valve (left) (b) Close-up of failed check valve

Concern by NASA for the safety of the orbiter LOX flow control valves resulted in the decision to make these valves from Inconel alloy 718; this decision was based on tests conducted at NASA White Sands using high-velocity particle

impacts During the acceptance testing of a valve for the Atlantis spacecraft in June 1984, an ignition occurred in an

adjacent 300-series stainless steel fitting The ignition melted the stainless steel, aluminum base plate, and part of the Inconel alloy 718 valve (Fig 40) This failure was attributed to the ignition of a loose silicone rubber O-ring seal after approximately 600 s of flow at 195 °C (380 °F) and 27.6 MPa (4000 psi)

Fig 40 Damage to an Inconel alloy 718 shuttle LOX flow control valve and aluminum test fixture due to

ignition of a silicone O-ring seal in an austenitic stainless steel test fitting (a) Test article and fixture after ignition (b) Side view of valve showing localized melting Arrow indicates area where solenoid screws on (c) Bottom view of valve

High-Temperature Gaseous Reactions

High-temperature gaseous reactions occur during mill processing, heat treating, and surface hardening of metals During heat treating, detrimental reactions with metals take the form of carburizing or decarburizing in steel, intergranular oxidation in nickel-base superalloys, and formation of an case on titanium alloys Surface attack on aluminum alloys is cosmetic in nature and not particularly detrimental to its properties To prevent these reactions, furnaces with inert, controlled, or vacuum atmospheres can be used, the part can be coated or protected, or the detrimental surfaces can be machined off, grit blasted, pickled, and so on

During surface hardening, carburizing or nitriding atmospheres are used to achieve the desired surface hardnesses These are normally well controlled by specifications and quality control sampling to prevent detrimental surfaces from being accepted

High-temperature detrimental gas reactions become a major concern when they are unanticipated There may be insufficient allowance on raw material to remove unacceptable layers Reactions with hardware have occurred where parts, in final dimensions, have little or no allowance for property losses or embrittled surfaces The examples presented below describe such typical problems

Launch Escape Tower Tubular Members. The function of the launch escape tower in the Apollo program was to pull the command module free of the Saturn V launch system in the event of an abort The launch escape tower, a titanium tubular truss structure about 3 m (10 ft) high, was attached to a solid rocket motor case on the upper end and the command module on the lower end through the tower leg bolts The titanium tubing specified was Ti-6Al-4V with an 89-

mm (3.5-in.) outside diameter and a 3.2-mm (0.125-in.) wall The titanium was produced by a hot extrusion process in which the billet is coated with a glass layer, which not only lubricates the extrusion but also protects it from oxidation

Excessive surface roughness (85 to 190 RHR) and pitting on the inside of a lot of tubing prompted a destructive microsectioning to determine the cause of the problem (Fig 41) The inside surface was found to contain a brittle case and localized cracks indicative of a high-temperature reaction with oxygen (>705 °C, or 1300 °F) Because there was no practical way to rework the inside of the tubing at that time, the lot was scrapped Future lots provided for adequate material removal on the tube inside diameter

Fig 41 Oxygen embrittlement of an extruded Ti-6Al-4V launch escape tower tube for the Apollo spacecraft (a)

Cross section of tube inside diameter showing pitting and case (arrow) 30× (b) Cracking of the case (arrow) 260× The glassy coating used to protect the part during processing was not continuous, resulting in high-temperature oxidation in air

Reaction Control System Vernier Engine Chambers. The reaction control system on the space shuttle orbiter provides the rocket propulsion to change the attitude of the orbiter with regard to the sun or earth The RCS vernier engine chamber, made of niobium alloy C 103, must function to 1315 °C (2400 °F) The chambers are protected with an R512A silicide coating Localized failure of the coating was observed in a vernier RCS engine that had undergone an extensive number of firing cycles In one case, failure occurred at a slight (75 m, or 3 mil) mismatch between two machining cuts This resulted in an offset in the coating, accelerating localized failure To provide the greatest coating cyclic life capability, action was taken to ensure blending of all machine cuts, to use a dual coating thickness, and to ensure a minimum total coating thickness of 0.1 mm (4 mils)

The silicide coating used on the RCS engine chambers can also fail when exposed to a cyclic, low-temperature (650 to

815 °C, or 1200 to 1500 °F) oxidizing environment for extended periods The low-temperature failure is caused by the thermal expansion mismatch between the silicide coating and the niobium alloy C 103 substrate Cracks in the brittle coating fill with oxides, eventually causing spalling of the coating Once the coating spalls, oxygen can reach the niobium alloy substrate Oxidation of the niobium alloy at these temperatures is relatively slow, but as the substrate oxidizes, the adjacent coating is undermined This results in more spalling, which enlarges the coating failure site (Fig 42)

Fig 42 Oxidation of a niobium alloy C 103 RCS engine chamber after cyclic temperature testing The thermal

expansion mismatch between the protective silicide coating and the niobium substrate caused the coating to spall (a) Coating failure site and oxidation of the C 103 substrate after 81,700 s and 267,000 cycles at 650

to 815 °C (1200 to 1500 °F) (b) As oxidation of the substrate progresses the adjacent coating fails Here the coating is being lifted from the surface

Orbital Maneuvering System Nozzle Extension. The orbital maneuvering system provides the rocket propulsion for orbit insertion, translation, rendezvous, and deorbit of the space shuttle orbiter The conical OMS nozzle extension is approximately 1.3 m (50 in.) long and 1.2 m (46 in.) in diameter It is a welded sheet metal structure made of niobium alloy FS-85, typically 1.5 mm (0.060 in.) thick, and has an R512E silicide coating to protect it from oxidation at temperatures to 1360 °C (2480 °F)

A nozzle was removed from the Challenger vehicle when cracks adjacent to the weld bead were found (Fig 43) The

fracture face showed brittle quasi-cleavage with some grain-boundary fractures rather than ductile dimpling A hardness traverse indicated the fracture to be brittle Because welding was performed in a controlled chamber to maximum oxygen levels of 2 ppm, no weld contamination during manufacturing was suspected Inspection of the nozzle showed the coating had been breached and spalled in eyebrow-shaped areas as a result of the nozzle flexing under pressure A redesigned, stiffer nozzle was already available for replacement; the failed lightweight nozzle represented an earlier design

Fig 43 High temperature oxidation embrittlement of a niobium orbiter OMS rocket nozzle extension due to

mechanical damage to the silicide coating (a) Fracture of FS-85 alloy showing that hardness increases near the failed edge Hardness ranges from 67 HRC at location 4 to 46 HRC at location 8 145× (b) Fracture face

showing brittle combined quasi-cleavage and intergranular failure 90× (c) Enlargement of fracture face showing quasi-cleavage failure with no apparent ductility 295×

Auxiliary Power Unit Gas Generator Catalyst Bed. The auxiliary power unit of the space shuttle orbiter uses an iridium catalyst bed to decompose neat N2H4 The decomposed gases are then directed into a high-temperature turbine, which in turn drives hydraulic pumps for orbiter hydraulic pressure The gases resulting from the decomposition are nitrogen and hydrogen with a small amount of ammonia The latter gas will, at the temperatures of the catalyst (>930 °C,

or >1700 °F), cause nitriding of the Hastelloy alloy B used to house the catalyst bed Hastelloy alloy B was chosen for this application because of its relatively high resistance to nitriding Nevertheless, nitriding does occur, occasionally resulting in cracking of parts (Fig 44) Fortunately, the cracked parts are under low loads, failure is not critical, and no high-temperature coating is required

Fig 44 High-temperature nitriding of an orbiter APU gas generator catalyst bed housing The housing is made

of Hastelloy alloy B and is in service to 925 °C (1700 °F) in the presence of ammonia formed by hydrazine decomposition (a) Nitriding and crack originating from embrittled area 25× (b) Another area of housing showing nitriding and cracking 40×

The shuttle entry air data system nose cap is a modified reinforced carbon-carbon composite nose cap in which

14 pressure ports were added to provide air pressure distribution data throughout the flight entry profile Holes drilled in the composite nose cap permitted insertion of a niobium alloy C-103 plug that was connected on the internal side to thin-wall niobium tubing These tubes terminate at pressure sensors Concern was for the survival of the niobium plugs, because loss of a plug could result in ingestion of a stream of extremely hot plasma that could potentially damage the spacecraft Although the niobium was coated with a VH109 silicide coating (a proprietary mixture of chromium, titanium, silicon, and hafnium), there was concern that the silicide coatings might experience inadvertent and undetected damage during the manufacturing or launch cycle

A test program was run to determine whether a niobium port with damage through the coating to bare metal would survive a reentry temperature at the maximum oxygen partial pressure expected during the heating peak, thus ensuring fail-safe mission behavior (Fig 45) Flaws were made in the coating of a size readily visible by inspection using a chisel point indenter Each defect resulted in a coating spall area approximately 1 × 1.5 mm (0.040 × 0.060 in.) Penetration into the bare metal was approximately 0.25 × 0.5 × 0.75 mm (0.010 × 0.020 × 0.030 in.) deep

Fig 45 High-temperature oxidation of a C 103 niobium alloy nose cap plug in the orbiter SEADS program The

plug is coated with a silicide coating, which was purposely damaged for testing to verify flight safety (a) Three perforations (arrows) in the coating result in spalled areas of 1 × 1.5 mm (0.040 × 0.060 in.) each 2× (b) Enlargement showing base metal damage × × mm deep (0.010 × 0.020 × 0.030 in deep) 15× (c) Oxidation of niobium during one simulated entry cycle at a 1485-°C (2700-°F) peak temperature using 0.57- kPa (12-psi) oxygen partial pressure Craters A, B, and C grew in damaged areas to 6.3 mm diameter × 2.5

mm deep ( × 0.10 in.) 1.5× (d) Embrittlement of crater surface in section A is 0.2 mm (7 mils) deep Section C is undamaged base metal 50×

A maximum reentry cycle is expected to reach 1430 °C (2600 °F) The part shown in Fig 45 represents an overtest; it was heated to 1485 °C (2700 °F) Normally, reaction of niobium in air results in a canary yellow voluminous oxide, but flow characteristics caused the oxide to be blown free of the surface Craters grew at damaged coating areas during this cycle

to sizes of approximately 6.3 mm in diameter × 2.5 mm deep (0.25 × 0.10 in.), and some subsurface embrittlement was experienced Niobium, used in the highest-temperature location on the space shuttle orbiter, showed that it is capable of surviving at least one reentry regime even if it has local coating damage

Liquid-Metal Cracking

Liquid-metal cracking is cracking in a base metal that occurs in the presence of a liquid metal Often, parts that crack are under stress Intrusion of the liquid metal into grain boundaries often occurs, but in some cases, intrusion is difficult to detect The term liquid-metal embrittlement is often used because liquid-metal cracking often causes embrittlement

Liquid-metal cracking does not always lead to catastrophic failure of parts, as shown in the brazing example below, but it often totally destroys the structural integrity of the part The major structural materials on a manned space vehicle are primarily made of alloys of aluminum, steel, stainless steel, titanium, and nickel These are embrittled by liquid metals shown below:

Alloy Liquid metals causing embrittlement

Aluminum Mercury, indium, tin, zinc

Steel Tin, cadmium, zinc, lead, copper, lithium

Stainless steel Cadmium, aluminum, lead, copper

Titanium Cadmium, mercury

Nickel Zinc, cadmium, mercury

Most of these metals have a very low melting point The lowest, mercury, is often used in switches, instruments (thermometers, manometers), and vapor arc lamps Mercury is prohibited from use on the space shuttle orbiter to avoid contamination from spillage Cadmium, due to its sublimation in space, is again highly restricted and is never permitted to exceed 120 °C (250 °F) This is well below the liquid-metal embrittlement temperatures of the plated steel Zinc, except

as an alloy constituent, is generally not used on manned spacecraft

Many of the embrittling elements find their way into spacecraft usage in three families First, these elements are used in low-temperature soldersfor electrical and avionics applications These include lead-tin for electrically soldered connections and some indium-base solders for glass-to-metal seals Second, these embrittling elements are used as silver solders for higher-strength applications Most of these solders melt in the range of 605 to 800 °C (1125 to 1475 °F) and may contain copper, cadmium, zinc, lithium, and tin These have caused liquid-metal cracking, especially with incompatible metals under stress Particular attention is given to these applications They can be quite troublesome where torch brazing techniques are used, because temperature control depends upon the skill of the operator Third, many brazing alloys used on the space-craft for stainless steel plumbing systems contain copper Many years of brazing experience with automated equipment has proved this to be acceptable (see the following discussion)

Brazed Plumbing Joints. Radiographic inspection of three brazed manifold tube joints for the shuttle orbiter revealed that the 21-6-9 stainless steel tubing had cracked The tubing was 13 mm ( in.) in diameter and was brazed by automatic equipment with Nicoro 80 braze alloy (81.5Au-16.5Cu-2Ni) Metallurgical examination showed that all brazed joint cracks were completely filled with the braze alloy A review of all shuttle orbiter brazing up to this time indicated that eight of the 1165 joints ( 0.5%) had similar braze-filled cracks The cracks were both longitudinal and circumferential; they occurred in type 304L stainless steel, Inconel alloy 718, and 21-6-9 stainless steel in line sizes of 9.5, 13, and 19 mm ( , , and in.) (Fig 46) All cracks were completely under the braze union

Fig 46 Liquid-metal cracking of brazed plumbing lines and unions in the shuttle orbiter due to overheating

Brazing alloy is Nicoro 80 (81.5Au-16.5Cu-2Ni) (a) Braze-filled crack (arrows) on Inconel alloy 718 tube end 5× (b) and (c) Intrusion of brazing alloy into 21-6-9 stainless steel (b) 130× (c) 45×

During the Apollo program, 30 tube ends of 5500 braze joints in type 304L stainless steel showed the same indication Extensive pressure testing of braze-filled cracked tube ends indicated that these tubes would burst in the tubular wall section rather than in the braze-filled crack under the union These joints were also found to be satisfactory in thermal shock and vibration tests Based upon this experience, tubes with braze-filled cracked ends, which remain totally under the braze union, are acceptable

The cause of the problem is thought to be liquid-metal attack The tube end may contain residual tensile stresses during its induction brazing or may contain small cracks that grow in contact with the braze metal The fresh surface of the crack is readily wet by the braze alloy Although the initial induction brazing heats the joint to approximately 980 °C (1800 °F), reheating a braze joint with the same induction brazing parameters increases the temperature because the braze alloy now conducts heat from the sleeve to the tube end more readily than in the initial braze, in which heat is transferred across the capillary gap to the tube by radiation It is believed that temperatures in tube ends may exceed 1095 °C (2000 °F) during reheat cycles used to reflow braze or debraze

Copper is known to be capable of causing liquid-metal embrittlement of stainless steels above 1095 °C (2000 °F) It is suspected that the high copper content of the braze alloy initiates the cracking when it occurs

The OMS rocket engine combustion chamber is a regeneratively cooled, axial flow type chamber in which the fuel passes through passages in the chamber wall The fuel acts as a coolant and at the same time is preheated before entering the firing chamber The chamber is made from type 304L stainless steel Its exterior is straddle milled to make longitudinal grooves, which are then closed into channels by electroless nickel plating

The testing program for the space shuttle orbiter OMS engine combustion chamber requires the evaluation of engine combustion instability This is simulated by igniting a small bomb in the combustion chamber during a firing sequence (The niobium nozzle extension, normally bolted to the engine in service, is not attached during this test.) Immediately after one such instability test, a portion of the chamber was found to be cracked (Fig 47)

Fig 47 Liquid-metal cracking of an orbiter OMS rocket chamber during testing The chamber is regeneratively

cooled and made from type 304L stainless steel Copper wiring broke loose and melted inside the chamber during instability testing, resulting in liquid copper embrittlement (a) Section of chamber showing a crack (arrow) in the fuel channel (b) Crack cross section 70× (c) Fracture face of crack showing intergranular appearance 355× (d) Enlargement of copper-filled crack in (b) 125× (e) Electron beam dot map of copper concentration in crack in (d) 125×

The cracked chamber and some bomb residue, scabbed onto the inner surface, were examined The scab material contained copper, copper oxide, and calcium carbonate The copper apparently came from copper wiring associated with the bomb The source of the calcium could not be determined Copper was found in crack areas and on crack faces and had penetrated the metal along grain boundaries

Reaction Control System Chamber. The reaction control system is used to control the roll, pitch, and yaw movements of the space shuttle orbiter and employs an engine using a silicide-coated niobium alloy C 103 chamber The hypergolic propellants N2O4 and hydrazine react within the chamber to provide engine thrust Normally, chambers are internally film cooled by the flow of the hydrazine fuel along chamber walls and operate well below 1315 °C (2400 °F) Under conditions of engine combustion instability, loss of the protective film may result in wall temperatures exceeding

1535 °C (2800 °F), and chamber burnthrough may take place To prevent any detrimental action to the surrounding spacecraft structure in the event of an instability and burnthrough, a wire sensing system on the chamber exterior will trigger shutoff of fuel and oxidizer valves

During a test of the wire sensing system, a reaction control chamber was purposely modified to induce instability, and significant cracking of the chamber resulted Cracking was associated with the melting of chromel-alumel thermocouple wires at temperatures slightly above 1425 °C (2600 °F) A test was set up in the laboratory on stressed coupons to simulate the suspected liquid-metal embrittlement failure

The test specimen measured 25 × 75 × 1.4 mm (1 × 3 × 0.056 in.), was coated with a silicide coating (75Si-2Cr-5Ti), and was subjected to elastic bending stresses Fine thermocouple wires were then placed on the niobium surface Figure 48(a) shows the extent of surface attack that occurred instantaneously as wires were melted Figures 48(b) and 48(c) show cross sections of liquid-metal embrittlement caused by melting the alumel (95Ni, Al, Si, Mn) and the chromel (90Ni-10Cr) thermocouples, respectively This example illustrates that even protective ceramic coatings can be breached during liquid-metal attack

Fig 48 Liquid-metal embrittlement of a niobium alloy C 103 test coupon intended to simulate embrittlement of

a C 103 shuttle RCS nozzle The coupon, with a silicide coating, was stressed, and fine chromel and alumel wires were melted on its surface The resulting coating damage is shown in (a) (b) Cross section showing damage caused by liquid alumel 100× (c) Cross section showing damage from liquid chromel 100×

Precipitation of Corrosion Products

To function properly, a valve for a manned spacecraft must be free of corrosion products Some spacecraft valves, in a full-open position, only move 0.25 mm (10 mils) off the valve seat Some have sealing lands only 125 m (5 mils) wide One of the cases of valve failures is the clogging that occurs from flow-decay products, that is, precipitation of dissolved metals as complex salts

Reaction Control System Quick Disconnect. After servicing of the forward reaction control system on the space shuttle orbiter in preparation for the second launch, a quick disconnect failed, spilling nitrogen tetroxide on the spacecraft

Columbia The N2O4 attacked the epoxy amine primer and resulted in the loosening or loss of about 100 tiles It also entered the spacecraft, exposing structure to corrosion and electrical wiring to possible damage Examination of the quick disconnect revealed deposition of complex iron nitrate compounds on its moving surfaces, causing jamming of the disconnect in an open position

Nitrogen tetroxide is commonly handled and stored in steel containers In the process, iron dissolves into the N2O4, and because it is present at concentrations of only a few parts per million, the integrity of the storage lines and containers are not jeopardized Under certain conditions of temperature and pressure, however, the iron will precipitate out as complex salts or form gels that cause valves to malfunction Figures 49(a) and 49(b) show the solubility of iron in N2O4 as a function of water content, temperature, and NO content At 20 °C (70 °F), the solubility of iron is 5.5 ppm when the NO content is 2.5% and water is 0.12% If the temperature were to cool to 10 °C (50 °F) after the spacecraft was loaded with this iron-saturated N2O4, only 2.6 ppm would remain soluble The rest could precipitate out A quick glance at the curves will show that under these conditions, a pickup of moisture by N2O4 (or loss) drastically decreases solubility, as does lowering the temperature and losing NO, a somewhat more volatile constituent

Fig 49 Solubility of iron in N2O4 as a function of water, temperature, and NO content (a) Solubility curves at 2.5% NO (b) Solubility curves at 1.25% NO

The failure of the quick disconnect delayed the launch of the second flight by more than a month for repair of the thermal protection system and evaluation and/or repair of other spacecraft problems created by the spillage These problems were minimized in future launches by the use of chillers and filters to process N2O4 prior to its loading Furthermore, samples are being taken at the spacecraft interface to verify that iron levels are below 2 ppm

The AC motor relief valves are used to control the flow of propellants (fuel and oxidizer) to the RCS engines Flow

decay products from nitrogen tetroxide caused an AC motor relief valve to fail to open at the specified pressure in an orbiter reaction control system test program The flow decay products are shown on the poppet in Fig 50

Fig 50 Failure of an orbiter RCS AC motor relief valve from flow-decay products in N2O4 testing (a) Front side

of poppet The round spots are deposits which caused the valve to malfunction (b) Back side of poppet showing similar surface deposits

Reaction Control System Helium Regulator Sensor Tubes. In the reaction control system of the space shuttle

orbiter, helium is used to expel propellants to the RCS rocket engines and vernier rocket engines The helium regulator valve, which controls this function, uses a sensor that has a fine orifice (0.18 mm, or 7 mils, in diameter) through which helium passes The pressure drop in this flow channel is a measure of the helium demand of the system Because the hole

is only about twice the diameter of a human hair, it is essential that no contamination, flow decay, or corrosion products restrict its flow

One would expect a helium system to be totally free of corrosion or precipitation products However, it must be recognized that molecular diffusion permits fuels and propellants to backstream into helium components Figure 51 shows

a result of contamination in the sensor passage as a combination of entrapped cleaning solution and reaction of this product with hydrazine The tube passage, on one of the shuttle flights, resulted in the sluggish response of the valve to the system demand Laboratory testing indicates that repeated wetting and drying with hydrazine vapors alone can cause plugging of this passage with decomposition products Experience has shown that the roughness of the walls of the hole probably contributed to iron dissolution and subsequent precipitation of complex iron hydrazine salts Fortunately, the valve can be checked out adequately prior to flight Studies were underway in 1986 to determine what design or processing changes could be made to minimize this problem

Fig 51 Failure of an orbiter RCS helium regulator valve due to contamination of the sensor orifice by hydrazine

salts (a) Sensor orifice assembly made from austenitic stainless steel (b) Orifice shown with screen removed Orifice is 0.18 mm (7 mils) in diameter 35 × (c) Deposits of hydrazine salts (arrow) 300× (d) Cross section through orifice showing salt deposits (e) and (f) Electron dot maps of orifice cross section The diagonal streak

is the orifice (e) Phosphorus (f) Sulfur Sulfur and phosphorus came from soap contaminants These reacted with hydrazine vapors to block orifice (d), (e), and (f) 145×

The APU injector tube of the space shuttle orbiter is made of Hastelloy alloy B It is the conduit for liquid hydrazine

to the catalyst bed in which the hydrazine decomposes to nitrogen, hydrogen, and some NH3 gases As the hydrazine approaches the catalyst bed, it is heated by the catalyst bed heater to about 650 °C (1200 °F) Metals dissolved in hydrazine, especially iron and molybdenum complex salts, precipitate on these walls when hydrazine vaporizes, as shown

in Fig 52 This precipitation is insufficient to restrict flow in this application If precipitation of this nature were to occur

on moving valve parts, it could result in valve leakage, sluggish response, or jamming

Fig 52 Encrustation of hydrazine salts in Hastelloy alloy B orbiter APU gas generator injector tubes (a) and (b)

Deposits occurring in sections of tubes above the 375 °C (700 °F) temperature range Both 415×

Atomic Oxygen in Low Earth Orbit

Atmospheres in low earth orbit principally consist of atoms and/or molecules of oxygen, nitrogen, argon, helium, and hydrogen (Fig 53a) Although the densities of these constituents are low at typical spacecraft altitudes of 300 to 500 km (185 to 310 miles), the orbital velocity of the spacecraft results in flux densities of 1013 to 1015 atoms/cm2 · s The interactions with the spacecraft produce drag effects that can limit orbital life in lower earth orbits as well as cause changes in the optical (thermal control) and physical properties of coatings and polymers by reflection, adsorption, or chemical reactions

Fig 53 Atmospheric composition (a) as a function of altitude Source: Ref 24 (b) Effect of solar (sunspot)

activity and atomic oxygen flux on thickness loss of Kapton- and Teflon-covered solar inertial facing surfaces (two sides exposed) Source: Ref 25

In the altitude range of approximately 200 to 650 km (125 to 400 miles), the atmosphere consists primarily of atomic oxygen The interaction of materials with atomic oxygen, which has impacting energy levels of 5 eV, can produce significant surface changes in some materials The effect of atomic oxygen depends on the fluence (or total flux of atoms per unit area), the impact angle, and the material impacted The greatest damage occurs to surfaces moving in the

direction of flight The flux level of atomic oxygen depends on solar activity and is related to the 11-year solar (sun spot) cycles, as shown in Fig 53(b)

Atomic oxygen degrades organic materials by breaking atomic bonds, by oxidation, and by causing loss of volatile species This degradation takes the form of surface erosion In the case of organic polymers, losses of 0.2 to 0.35 mm (8

to 14 mils) could be expected on a space station every solar cycle (11 years), depending on the orientation Silicone polymers and perfluorinated polymers are approximately 50 times more resistant to surface loss than organic polymers Only two metals show significant reaction silver and osmium Silver oxidizes readily, and expected surface losses, based

on Shuttle SPS-8 tests (Ref 26), could exceed 0.2 mm/yr (8 mils/yr) Osmium reacts to form OsO4, a volatile gas; this reaction also results in significant losses All other metals are nonreactive except for copper, which forms a superficial protective oxide film and undergoes essentially no changes in properties or dimensions

To prevent detrimental damage due to atomic oxygen, proper materials selection and the use of coatings are often acceptable approaches Some metals appear resistant in thicknesses as thin as 15 nm

Atomic Oxygen Degradation of Polymeric Insulation Films. The first shuttle flight resulted in surface roughening of the Kapton (polyimide) thermal control blankets used in the cargo bay areas The surfaces of the film, which were normally glossy and amber in color, changed to a flat, translucent yellow appearance that resulted from microscopic roughening of these surfaces Areas shielded by other parts showed no changes From an analysis of the data and environments, it was concluded that atomic oxygen was the responsible active species causing this surface reaction/erosion Subsequently, panels of test materials were exposed on racks in the shuttle cargo bay on the fifth and eighth flights to study the extent of degradation and to determine material reaction efficiencies

Samples of polyurethane paint before and after 40 h of exposure in the velocity direction on the eighth orbiter flight are shown in Fig 54(a) and 54(b) Samples of Kapton H polyimide film before and after 4 years of exposure on a satellite in low earth orbit are shown in Fig 54(c) and 54(d) Figure 54(e) shows the surface of a 0.05-mm (2-mil) thick sheet of FEP teflon coated with 100 nm of Inconel over 150 nm of silver and exposed to atomic oxygen and ultraviolet radiation in low earth orbit for 4 years The metal coating appears to be completely stripped from the surface

Fig 54 Atomic oxygen degradation of materials in low earth orbit (a) and (b) Polyurethane paint before (a)

and after (b) 40 h of low earth orbit exposure in the velocity direction on the eighth shuttle flight Both 14,000× (c) and (d) Sample of Kapton H (polyimide) film before (c) and after (d) approximately 4 years of exposure on a satellite in low earth orbit (e) Surface of a metallized Teflon film after about 4 years of exposure

on a satellite in low earth orbit The surface had been coated with 100 nm of vapor-deposited Inconel alloy over

150 nm of vapor-deposited silver Both metals appear to have been completely eroded away Courtesy of Materials and Process Laboratories, NASA George C Marshall Space Flight Center

Corrosion of Space Boosters and Space Satellites

Douglas B Franklin, George C Marshall Space Flight Center, National Aeronautics and Space Administration

The selection of corrosion prevention systems for space boosters and satellites requires the consideration of many different criteria related to the expected operating environments and various functional requirements Both boosters and satellites may be exposed to a seacoast environment for prolonged periods as well as to various propellants and operating fluids Satellites are also exposed to the high vacuum and solar radiation of space, and their external coating must provide controlled values for solar absorption and emittance to ensure proper temperature control Boosters, on the other hand, often have severe cryogenic and elevated temperature exposures

The need for lightweight structures, coupled with the low factors of safety used in component design, result in the use of materials selected more for strength than corrosion resistance This creates a high potential for stress-corrosion failure and makes the effects of surface corrosion very significant When the high reliability requirements for spacecraft are added to these considerations, it places a very demanding performance requirement on the corrosion protection system used No reuse was planned in earlier space booster programs, but for the space shuttle, economy demands the reuse of most of the major components Perhaps the most demanding environmental exposure is the solid rocket booster, which must survive periodic immersion in seawater for several days as well as damage to protective coatings resulting from towing and recovery operations (Fig 55) The boosters must be capable of up to 20 reuses

Fig 55 Ocean recovery of a shuttle solid rocket booster

Aluminum Alloys

The primary structural materials used for many of the propulsion components in the space shuttle transportation system (Fig 56) are high-strength aluminum alloys This includes aluminum alloy 2219 for welded structures (such as the propellant tankage) and aluminum alloys 7075 and 2024 for structures where mechanical joining methods can be used The protective system used on aluminum surfaces where exterior exposure is the primary concern consists of a chemical conversion coating (MIL-C-5541) to promote paint adhesion, followed by a chromate-inhibited epoxy primer 0.025 mm (1 mil) thick (Ref 27) and an epoxy topcoat 0.025 to 0.045 mm (1.0 to 1.8 mils) thick (Ref 28)

Fig 56 Major propulsion components of the space shuttle transportation system

This system is used on the upper and lower skirt structures of the solid rocket boosters that are recovered from the ocean and reused With some modification, it also is used on the external tank (Fig 56) and provides a good base for bonding thermal insulation to provide protection from aerodynamic heating and heat from rocket exhaust plumes To obtain good paint adhesion, it has been found to be extremely important to maintain surface cleanliness between each processing step, particularly after application of the chemical conversion coating In addition, strict compliance with the recommendations

of the paint manufacturer for drying times and application procedures is required To ensure adequate quality control, the coatings are applied and tested to the requirements of MIL-F-18264

Where immersion in seawater and component reuse without refurbishment are required, all faying surfaces must be completely sealed (wet lay-up) with a polysulfide sealant meeting the requirements of MIL-S-8802 It is very important that there are no open gaps or voids not completely filled by the sealant, because seawater can be forced into these areas during recovery and will cause a serious corrosion problem Fasteners should also be installed with wet sealant, and fastener heads must be completely oversealed These procedures cannot be followed where electrical bonding is required For these areas, jumper cables are used with the contact surfaces bare The connection is then completely oversealed with polysulfide sealant

Another area requiring special consideration for protection is under the polyurethane foam insulation used on the external tank exterior to prevent excessive cryogenic propellant boil-off (liquid hydrogen and oxygen) during launch preparation and flight Because of the need for good cryogenic adhesive properties, the paint system previously described cannot be used It was also found that the bonding primer originally selected to promote adhesion between the tank surface and the spray foam would not adequately protect the aluminum alloy 2219 surface during extended storage, shipping, and launch preparations, particularly where multiple exposures to cryogenic temperatures are involved After laboratory tests, the material finally selected for this purpose consists of an epoxy bonding primer with an increased amount of strontium chromate pigment (Ref 29) This primer provides improved corrosion protection under the foam insulation without any significant degradation of cryogenic adhesive properties

Steel Alloys

One major area in which aluminum alloys are not used is the motor case of the solid rocket booster This application uses D6AC low-alloy steel heat treated to an ultimate tensile strength of 1340 to 1550 MPa (195 to 225 ksi) The paint system selected for corrosion protection is a sacrificial zinc-rich epoxy-polyamide primer 0.038 to 0.063 mm (1.5 to 2.5 mils) thick (Ref 30) and an epoxy-polyamide topcoat 0.038 to 0.063 mm (1.5 to 2.5 mils) thick (Ref 31) The steel surfaces are gritblasted to white metal (Steel Structures Painting Council Specification SSPC-SP6) prior to painting Although this system provides good protection to steel surfaces for prolonged periods of time, the paint is removed (by gritblasting) and reapplied after each flight because of magnetic-particle inspection requirements imposed on the motor case surfaces

There are a few areas in which paint cannot be used to provide corrosion protection These are the areas where the solid rocket booster motor case segments are joined and where the skirt segments join the motor case segments The material used to provide protection here is a heavy-duty calcium-base grease with special corrosion inhibitors for use in seawater (Ref 32) The grease is carefully applied to all bare areas during assembly and is removed and reapplied after each flight The grease protects the surfaces during storage and preflight operations and provides excellent protection for several days

of ocean exposure during recovery of the segments and until joint refurbishing operations can be initiated, which can be several weeks later The grease can also be applied diluted with trichloroethane or other solvents to provide protection to bare motor case segments during initial shipment by rail from the manufacturer as well as after the paint is stripped from used motor cases during refurbishment

Graphite/Epoxy Motor Case

A graphite/epoxy motor case that provides significant weight savings for the SRBs is under development Because of structural considerations and design requirements, D6AC steel adapter rings are used to join the motor case segments To prevent galvanic corrosion on the D6AC rings, the exterior surfaces of the graphite/epoxy are coated with the epoxy-polyamide topcoat discussed previously, and the machined joint area is sealed with a nonconductive sealant This minimizes the cathode surface area and reduces the galvanic current The D6AC rings are protected as described in the previous paragraphs

Electrical Cables

One area of the solid rocket booster that has been very difficult to protect is the electrical cables and connectors Because seawater contacting the ends of the stranded electrical wiring will quickly permeate through the wire and prohibit any reuse capability, electrical cables are enclosed in a watertight jacket of a polyether-based urethane plastic This requires bonding of the jacket to the connector at each end of the cable, as shown in Fig 57 In addition, the stainless steel connectors have a watertight O-ring seal to prevent the intrusion of seawater To provide additional protection against inadvertent leakage, the connector pins are coated with a film of heavy-duty calcium grease (Ref 32) prior to assembly The female connector sockets are designed so that the grease film is wiped off during insertion of the male pins and necessary electrical continuity through the connector is maintained

Fig 57 Schematic showing electrical cable with watertight jacket

Cathodic Protection

Carbon fiber reinforced phenolic ablators are used to line the solid rocket motor nozzles These materials appeared to aggravate corrosion in the solid rocket booster aft skirt during immersion in the ocean Sacrificial zinc anodes (MIL-A-18001) were therefore added to provide additional protection This includes zinc anodes for several individual aluminum components, the use of zinc for several nonstructural components, and the use of flame sprayed zinc on several aluminum components In addition, zinc anodes are attached at various locations by divers prior to towing the solid rocket boosters back to land for refurbishment These anodes have reduced significantly the galvanic attack of the aluminum surfaces in the aft skirt of the solid rocket booster

Other Alloys

Other alloys used for space booster systems include type 304 stainless steel, type 321 stainless steel (for welded components), Inconel alloy 718, titanium alloys Ti-6Al-4V and Ti-3Al-2.5V, and MP35N nickel-cobalt alloy Although these alloys are inherently corrosion resistant, special treatments are usually required to ensure that exposed surfaces are passivated to reduce possible pitting problems Surfaces exposed to seawater during solid rocket booster recovery are flushed with potable water (200 ppm chloride max), washed with a nonionic detergent (0.5% solution at 57 °C, or 135

°F), flushed again with potable water, and rinsed with deionized water until residual surface chlorides are below 50 ppm Components are refurbished as necessary after each flight to ensure that their integrity is not compromised

Control of Stress Corrosion

One other area that is carefully considered in design is the control of stress corrosion, particularly because of the widespread use of high-strength alloys that generally have poor resistance to stress corrosion Several stress-corrosion failures have occurred in earlier programs, many of which resulted in significant program impact

Table 4 lists these failures As shown in Table 4, most of the failures have occurred in high-strength aluminum alloys and

in the precipitation-hardening stainless steels in the seacoast environment There have also been instances in which unique environments were not adequately considered For example, stress-corrosion failure of a beryllium copper spring occurred because it was not recognized that a small amount of hydrazine could leak past an O-ring, thus exposing the spring to ammonia, a decomposition product of hydrazine

Table 4 Stress-corrosion failures in space boosters

Aluminum alloy 7079-T6 Forging Prelaunch LOX dome Saturn IB

AM-355 stainless steel Bar Prelaunch Flared tubing sleeve Saturn I

17-7PH stainless steel Sheet Prelaunch Wave spring Saturn IB

Aluminum alloy 7079-T6 Forging Manufacture Prevalve control support link Saturn V

Aluminum alloy 7075-T6 Plate Test Splice angle Saturn V

Aluminum alloy 7075-T6 Bar Assembly Prevalve control piston cylinder Saturn IB

Aluminum alloy 2024-T4 Bar Test Oxidizer check valve body Saturn IB

17-7PH stainless steel Sheet Test Actuator spring Saturn V

17-7PH stainless steel Sheet Test Prevalve Belleville spring Saturn V

Aluminum alloy 7178-T6 Forging Storage Upper E-beam Saturn IB

7079-T652 Forging Storage Rear spar Saturn IB

7079-T6 Forging Test Hold-down fitting Saturn IB

Figure 58 shows an actual stress-corrosion failure that occurred in an aluminum alloy 7079-T6 forging used as a main housing in a hydraulic actuator The failure occurred at the forging parting plane and was caused by excessive moisture in the hydraulic oil, residual stress from heat treating, and the interference fit of a small check valve that was press fit into the forging at the parting line The failure was found during acceptance testing of the part prior to installation

Fig 58 Stress-corrosion failure of 7079-T6 aluminum forging (a) Failed housing (b) Fractograph of failure

An example of stress-corrosion failure in 17-7PH RH950 stainless steel Belleville washers is shown in Fig 59 The washers were used as a Belleville spring (a stack of 54 washers) in a valve actuator Failure resulted from exposure to a humid atmosphere (water was trapped in the actuator housing) and the installation stresses caused by the preload on the spring stack

Fig 59 Stress-corrosion failure of 17-7PH RH950 stainless steel valve actuator Belleville spring (a) Failed

spring washers (b) SEM of fracture surface 2000×

The best method for controlling stress corrosion is to select, where possible, materials that are highly resistant to stress corrosion To this end, guidelines (MSFC-SPEC-522) have been prepared to aid the designer in the selection of materials for use in space booster and satellite systems Tables 5, 6, and 7 list alloys grouped to show their comparative stress-corrosion resistance when exposed to a seacoast environment The materials listed in Table 5 are considered resistant to stress corrosion in a seacoast atmosphere and can be used without restrictions

Table 5 Alloys with high resistance to SCC

Ferrous alloys

Carbon steel (1000 series) Below 1240 MPa (180 ksi) ultimate tensile strength

Low-alloy steel (4130, 4340, D6AC, etc.) Below 1240 MPa (180 ksi) ultimate tensile strength

Music wire (ASTM 228) Cold drawn

HY-130 steel Quenched and tempered

HY-140 steel Quenched and tempered

1095 spring steel Quenched and tempered

300-series stainless steels (unsensitized) All

21-6-9 stainless steel All

20Cb stainless steel All

20Cb-3 stainless steel All

A-286 stainless steel All

AM350 stainless steel SCT1000 and above

AM355 stainless steel SCT1000 and above

Almar 362 stainless steel H1000 and above

Custom 455 stainless steel H1000 and above

15-5PH stainless steel H1000 and above

PH14-8 Mo stainless steel CH900 and SRH950 and above

Inconel alloy 600 Annealed

Inconel alloy 625 Annealed

Inconel alloy X-750 All

Ni-Span-C alloy 902 All

Magnesium, LA141 Stabilized

(a) High-magnesium alloys 5456, 5083, and 5086 should be used only in

controlled tempers (H111, H112, H116, H117, H323, H343) for resistance

to SCC and exfoliation Alloys with more than 3% Mg are not recommended for applications above 66 °C (150 °F)

(b) For copper alloys, numbers under "Condition" indicate the percentage of

cold work

(c) AT, annealed and precipitation hardened; HT, work hardened and

precipitation hardened

Table 6 Alloys with moderate resistance to SCC

Ferrous alloys

Carbon steel (1000 series) 1240 to 1380 MPa (180-200 ksi) ultimate tensile strength

Low-alloy steel (4130, 4340, D6AC, etc.) 1240 to 1380 MPa (180-200 ksi) ultimate tensile strength

Types 403, 410, 416, 431 stainless steels (a)

PH13-8Mo stainless steel All

15-5PH stainless steel Below H1000

17-4PH stainless steel All

Wrought aluminum alloys

2024 rod, bar, extrusion T6, T62

(a) Tempering between 370 and 595 °C (700 and 1100 °F) should be avoided

because resistance to SCC and corrosion is lowered

Table 7 Alloys with low resistance to SCC

Ferrous alloys

Carbon steel (1000 series) Above 1380 MPa (200 ksi) ultimate tensile strength

Low-alloy steel (4130, 4340, D6AC, etc.) Above 1380 MPa (200 ksi) ultimate tensile strength H-11 steel Above 1380 MPa (200 ksi) ultimate tensile strength

440C stainless steel All

18Ni maraging steel, 200 grade Aged at 480 °C (900 °F)

18Ni maraging steel, 250 grade Aged at 480 °C (900 °F)

18Ni maraging steel, 300 grade Aged at 480 °C (900 °F)

18Ni maraging steel, 350 grade Aged at 480 °C (900 °F)

AM350 stainless steel Below SCT1000

AM355 stainless steel Below SCT1000

Custom 455 stainless steel Below H1000

PH15-7 Mo stainless steel All except CH900

17-7PH stainless steel All except CH900

Wrought aluminum alloys

to stress corrosion This includes exposure conditions, protective treatments used, and sustained tensile stresses that may

be present It is extremely important to consider all sources of stress; not only operational stresses but also residual stresses and stresses that may be induced into the part during assembly must be considered Residual stresses can be the result of machining, forming, and heat-treating processes, while assembly stresses can result from improper fit-up tolerances, overtorquing, press fits, high-interference fasteners, and welding A more detailed discussion of these factors can be found in MSFC-SPEC-552

Special Environmental Effects

In addition to ordinary atmospheric environmental effects, corrosion control procedures for space boosters and satellites must take into account other special environmental factors Many of these are related to propellant compatibility

Oxygen. Most organic materials are not compatible with oxygen systems and can ignite under service conditions

Because of a lack of compatibility with liquid oxygen and stringent cleanliness requirements, organic coatings are not used on aluminum propellant tank (external tank) interiors Corrosion on these surfaces (aluminum alloy 2219-T87) is prevented by use of a chemical conversion coating (MIL-C-5541) In addition, rigorous drying procedures are required after conversion coating and tank cleaning (usually done in one continuous process), and the humidity inside the tanks is controlled to below 60% relative humidity during storage and purged to -12 °C (10 °F) dew point before shipping to the launch site

The primary metallic materials not considered acceptable in oxygen systems are tin, magnesium, and titanium alloys Their use may result in catastrophic impact ignition Tin is particularly reactive, and it has been found that copper alloys with a tin content as low as 2% can react when impacted in liquid oxygen All materials used in oxygen systems are selected to meet the requirements of NASA Handbook NHB 8060.1 at the temperature and pressure that will be encountered during use

Hydrogen. Propellant compatibility is also of concern in hydrogen systems, particularly at the high pressures up to 48 MPa (7000 psi) found in the space shuttle main engine (SSME) Table 8 shows the effect of exposure to high-pressure hydrogen on the notched strength ratio of several metal alloys One technique that has been used to protect alloys susceptible to hydrogen embrittlement is to copper plate the exposed surfaces A 0.13-mm (5-mil) thick coating of electroplated copper has been used for protecting Inconel alloy 718 in several SSME components Gold plating that is

0.13 mm (5 mils) thick has also been used for this purpose on Waspaloy alloy turbine disks For these coatings to be effective, very careful procedures are required to ensure that coating adhesion and integrity are of the highest quality

Table 8 Relative resistance to hydrogen embrittlement of various alloys in high-pressure hydrogen at room temperature

Pressure

factor, Kt MPa ksi

Ratio H2/He(a)

Type 410 stainless steel 8 69 10 0.22

1042 steel (quenched and tempered) 8 69 10 0.22

Type 440C stainless steel 8 69 10 0.50

Ti-6Al-4V (solution treated and aged) 8 69 10 0.58

Type 304 stainless steel 8 69 10 0.87

Type 321 stainless steel 8 34 5 0.87

Type 304N stainless steel 6.3 103 15 0.93

Type 310 stainless steel 8 69 10 0.93

Hydrazine and Nitrogen Tetroxide. Several other propellants are encountered in space booster and satellite systems that present unique compatibility problems The two most common are hydrazine and nitrogen tetroxide The major concern with hydrazine systems is not corrosive attack but the decomposition of hydrazine, which can result in failure due

to a large volumetric change For nitrogen tetroxide systems, most metallic materials are resistant when the N2O4 is dry However, because moisture can easily contaminate such systems (resulting in the formation of nitric acid), the primary materials of construction are those that also have high resistance to nitric acid, such as the aluminum alloys, stainless steels, and titanium alloys

Corrosion Control for Satellites

The criteria for selection of coatings for satellite systems are usually related to their thermal control properties and resistance to the effects of the space environment Properties can vary widely depending upon the specific requirements needed Figure 60 illustrates the variation in types of coatings and surface treatments that may be used

Fig 60 Chart showing the range in optical properties for several types of coatings and surfaces

Corrosion protection properties are usually of secondary importance Consequently, the environmental exposure conditions, particularly during manufacture and storage, must be carefully controlled to prevent corrosion and to prevent the deterioration of critical surfaces from contamination This means stringent controls on packaging of individual components, humidity control during component assembly, and environmental control during storage of completed assemblies Because most systems are assembled in a clean room to prevent surface contamination, keeping the relative

humidity below 70% should prevent most corrosion problems during the assembly process For storage purposes, particularly long-term storage, environmental conditions should be regulated so that the maximum relative humidity is below 60% and preferably below 50% Storage in uncontrolled environments is not permitted

References

1 Press Information, Space Shuttle Transportation System, Rockwell International, 1982

2 "The Effects of Silver on the Properties of Titanium," DMIC Technical Note, Defense Metals Information Center, Battelle Memorial Institute, 1965

3 "The Stress-Corrosion and Accelerated Crack-Propagation Behavior of Titanium and Titanium Alloys," DMIC Technical Note, Defense Metals Information Center, Battelle Memorial Institute, 1966

4 W.K Boyd and F.W Fink, "The Phenomenon of Hot-Salt Stress-Corrosion Cracking of Titanium Alloys," NASA CR-117, Battelle Memorial Institute

5 L.B Norwood, "Application of Beryllium on the Space Shuttle Orbiter," Paper presented at the 15th National SAMPE Conference, Cincinnati, OH, Society for the Advancement of Materials and Process Engineering, 1983

6 "Reactivity of Metals With Liquid and Gaseous Oxygen," DMIC Technical Note, Defense Metals Information Center, Battelle Memorial Institute, 1963

7 R.L Johnston, "Multidisciplinary Approach to the Design of High Pressure Oxygen Systems," Paper presented at Multidisciplinary Analysis and Optimization Symposium, 1984

8 A.C Bond, H.O Pohl, N.H Chaffee, W.W Guy, C.S Allton, R.L Johnston, W.L Castner, and J.S

Stradling, Design Guide for High Pressure Oxygen Systems, Reference Publication 113, National

Aeronautics and Space Administration, 1983

9 C.B Brownfield, "The Stress Corrosion of Titanium in Nitrogen Tetroxide, Methyl Alcohol and Other Fluids," SID 67-213, North American Aviation Inc., Space and Information Systems Division, 1967

10 "Stress Corrosion of Ti-6Al-4V in Liquid Nitrogen Tetroxide," DMIC Technical Note, Defense Metals Information Center, Battelle Memorial Institute, 1966

11 G.F Kappelt and E.J King, "Observations on the Stress Corrosion of 6Al-4V Titanium Alloy in Nitrogen Tetroxide," Paper presented at the AFML 50th Anniversary, Corrosion of Military and Aerospace Equipment Technical Conference, Denver, CO, Air Force Materials Laboratory, 1967

12 R.E Johnson, G.F Kappelt, and L.J Korb, "A Case History of Titanium Stress Corrosion in Nitrogen Tetroxide," Paper presented at the National Metals Conference, Chicago, IL, American Society for Metals,

15 F.K Lampson, The Marquardt Corporation, private communication, 1986

16 L.J Korb and R.E Johnson, "Stress Corrosion of Titanium Tanks in Methanol," Paper presented at the AFML Technical Conference on Corrosion of Military and Aerospace Equipment, Denver, CO, Air Force Materials Laboratory, 1967

17 R.E Johnson, "Apollo Experience Report The Problems of Stress Corrosion Cracking," NASA TN

D-7111, NASA Technical Note, National Aeronautics and Space Administration, 1973

18 W.B Lisagor, "Some Factors Affecting the Stress Corrosion Cracking of Ti-6Al-4V in Methanol," NASA

TN D-5557, Langley Research Center, 1969

19 R.L Johnston, R.E Johnson, G.M Ecord, and W.L Castner, "Stress Corrosion Cracking of Ti-6Al-4V in Methanol," NASA TN D-3868, National Aeronautics and Space Administration, 1967

20 F Mansfeld, The Effect of Water on Passivity and Pitting of Titanium in Solutions of Methanol and

Hydrogen Chloride, J Electrochem Soc., 1971

21 R.R Boyer and W.F Spurr, Characteristics of Sustained-Load Cracking and Hydrogen Effects in

Ti-6Al-4V, Metall Trans A, Vol 9A, Jan 1978

22 "Reaction of Titanium With Gaseous Hydrogen at Ambient Temperatures," DMIC Technical Note, Defense Metals Information Center, Battelle Memorial Institute, 1965

23 H.H Johnson and A.M Willner, Moisture and Stable Crack Growth in High Strength Steel, Appl Mater

Res., Vol 4, 1965, p 34

24 L.J Leger, J.T Visentine, and J.A Schliesing, "A Consideration of Atomic Oxygen Interactions With Space Station," Paper presented at the AIAA 23rd Aerospace Sciences Meeting, Reno, NV, American Institute of Aeronautics and Astronautics, Jan 1985

25 L.J Leger, J.T Visentine, and J.F Kuminecz, "Low Earth Orbit Atomic Oxygen Effects on Surfaces," Paper presented at the AIAA 22nd Aerospace Sciences Meeting, Reno, NV, American Institute of Aeronautics and Astronautics, Jan 1984

26 H.J Rockoff, "Materials System Requirements and Challenges on the Operational Space Station," Paper presented at the SME Space Tech Conference and Exposition, Anaheim, CA, Sept 1985

27 Product Bulletin 463-6-3, Sikkens Aerospace Finishes Division, Akzo Coatings America, Inc

28 Product Bulletin 400 Series, Sikkens Aerospace Finishes Division, Akzo Coatings America, Inc

29 Product Bulletin 515-346, Desoto, Inc

30 Technical Data Form 1042, Rust-Oleum Corporation

31 Technical Data Form 1047, Rust-Oleum Corporation

32 Product Bulletin Code 9760, Conoco, Inc

Corrosion in the Electronics Industry

Jack D Guttenplan, Rockwell International Corporation

Introduction

THE EFFECTS OF CORROSION on material degradation and component performance in the electronics industry have long been recognized Recent studies have shown that corrosion is becoming as even more significant factor in the reliability of electrical and electronic equipment This has occurred because of the following trends:

• Designs requiring high component density and faster signal processing, resulting in smaller components with closer spacings and thinner metallic sections

• New requirements for low-resistance, electrically stable grounding paths and electrical bonds to protect against stray electromagnetic radiation

• The exposure of electronics to more severe environments

The effects of environment and corrosion on the reliability of electronic equipment became very evident during the Vietnam War Military electronics were not designed to resist the effects of the hot, humid environment and the monsoon rains to which they were exposed, and high failure rates resulted Other examples of severe environments to which military electronic equipment are exposed include avionic systems in ship-based naval aircraft (Ref 1) and guidance systems on nuclear submarines (Ref 2)

The problem of corrosion of electronic equipment, however, is not unique to the military It exists throughout commercial and military applications worldwide and can occur in various severity levels of indoor and outdoor environments

Frequently, minute amounts of contaminants or corrosion products can cause serious degradation or complete failure Thin films of oxidation or corrosion products, invisible to the eye, can result in noise or resistance buildup in electrical

contacts and subsequent system failure Extremely low levels of moisture and corrosive contaminants have been known to cause corrosion problems with printed circuit boards, encapsulated integrated circuits, nichrome-film resistors, electrical connectors, other discrete devices, and a wide range of plated components

Studies have been initiated to determine both the nature of field environments and to develop laboratory test methods applicable to various classes of operating environments These studies have shown that parts per billion levels of selected pollutants are sufficient under proper conditions of temperature and humidity to accelerate corrosion reactions in electronic equipment (Ref 3) Corrosion can occur during manufacturing, storage, shipping, and service Moisture and such corrosive agents as chlorides, fluorides, hydrogen sulfide (H2S), sulfur dioxide (SO2), nitrogen compounds such as ammonia (NH3), and other airborne contaminants are the major culprits (Ref 4) The sources of these corrosives have been the subject of intensive investigation Some sources that have been identified include soldier flux residues, residual electroplating or other processing chemicals, sulfur from storage container materials, vaporized contaminants from adhesives, reactive substances in plastic materials and glass, environmental acid deposition, and a wide range of reactive airborne contaminants

Most of the same principles and mechanisms apply in the corrosion of electronic equipment as in, for example, the corrosion of a large structure However, the interaction among electrical, metallurgical, and environmental conditions, together with severe dimensional constraints, can lead to a unique set of corrosion problems for electronic systems These problems will be discussed briefly in the following sections

Corrosion Problems and Preventive Methods

Moisture Intrusion Into Black Boxes. Many manufacturers of electronic equipment have resorted to sealed black boxes to prevent corrosion However, moisture intrusion remains a major factor in electronic equipment failures The optimal way to exclude moisture is to make the box hermetic (airtight) by fusion Even if the box is hermetic, materials inside the box must be baked out to eliminate outgassing of moisture, and the box should be evacuated and pressurized with a dry inert gas This procedure will provide confidence of low relative humidity throughout the service life

However, the equipment for internal maintenance/testing and the increasing number of through-wall electrical input/output connections make fusion impractical in many cases Designers must select other approaches for sealing boxes Elastomers (rubber compounds) in the form of gaskets and O-rings are commonly used to seal lids and other points

of entry However, water vapor will eventually permeate through any elastomeric sealant and thus increase internal humidity Designers need to predict the long-term entry of water vapor through these seals Desiccants can postpone an increase in relative humidity, but the rate of moisture influx must still be estimated

A recent study provided data on the moisture vapor transmission rates for typical elastomeric sealants and illustrations on the use of these data in calculating the rates of moisture influx into sealed boxes (Ref 5, 6) It was concluded that a butyl rubber compound was the best sealant against water vapor permeation, with ethylene propylene rubber a less desirable alternative In an example using butyl rubber gaskets to seal a box, the calculations showed that, starting at 0% relative humidity inside and 90% relative humidity on the exterior, the time to reach 50% relative humidity at 40 °C (100 °F) inside the box was 7 years With a proper desiccant in the box, the time was extended to 20 years Using the same example, it can be calculated that the time for a silicone rubber gasket (without desiccant) would be 3 weeks and that the time for an ethylene propylene rubber gasket would be 1.9 years (see Ref 5 for the particular rubber compounds used.)

An additional conclusion of this study was that in order to have high confidence that corrosion will not be a problem inside a sealed black box the relative humidity should be maintained at 40% or less at room temperature (Ref 6) Also, to prevent condensation, care must be taken to keep the internal temperature of the box from dropping below the dew point This study did not address gross leakage of moisture due to improperly designed or sealed seals and defective materials or the incompatibility of sealant materials with environmental fluids

One of the standard methods of cooling an electronic enclosure is the use of force ventilation, with air generally drawn from the atmosphere surrounding the equipment enclosure In sites with aggressive atmospheres, this type of cooling will greatly accelerate corrosion because the circulating contaminated air comes in intimate contact with sensitive electronics Unless the outside environment is benign, the introduction of outside air into an electronic equipment cabinet should be eliminated or limited to the lower rate possible (Ref 7)

For equipment that is neither sealed nor pressurized, corrosion protection should be designed for worst-case conditions on the assumption that moisture will get in Corrosion can be minimized by:

• Providing low point drains so that water cannot collect

• Encapsulating components so that moisture cannot reach them

• Conformal coating printed circuit boards for protection against moisture and contaminants

• Mounting printed circuit boards vertically and well above the bottom of the housing to prevent moisture and debris from collecting on the boards

• Locating edge connectors on the vertical sides, not the bottom, of the printed circuit board so that moisture and debris will not collect in the mated connectors and degrade the contacts

• Locating feed-through connectors on the sides of the box, not on the bottom, for similar reasons

• Designing cabling to lead down and away from connectors and providing drip loops where possible

• Avoiding hygroscopic materials that will hold or wick up moisture

• Avoiding materials that emit corrosive vapors

• Using a volatile corrosion inhibitor that is carefully selected to protect those metals of concern in a particular application

Dissimilar-Metal Corrosion. Electronic design is unique in the wide variety of metals used because of particular physical and electrical properties Some of the more common metals and their uses in an electronic system are given in Table 1 These metals are combined to form a myriad of dissimilar-metal couples in electronic equipment In the presence

of moisture (an electrolyte), destructive galvanic corrosion can take place (see the section "Galvanic Corrosion" of the article "General Corrosion" in this Volume)

Table 1 Metals and alloys commonly used in electronic systems

Gold Electrical connector contacts, printed circuit board edge connectors, leaf-type relays, miniature coaxial connectors,

semiconductor leads, and microminiature and hybrid circuits

Silver Protective coating on relay contacts, wave guide interiors, wire, high-frequency cavities, EMI/EMP shields, and

EMI gaskets

Magnesium alloys Radar antenna dishes and lightweight structures, such as chassis, supports, and frames

Iron, steel, and

ferrous alloys

Component leads, magnetic shields, magnetic coatings on memory disks, transformers, brackets, racks, hermetic electrical connector shells, and fastener hardware

Aluminum alloys Equipment housings, chassis, mounting racks, supports, frames, electrical connector shells, and printed circuit

board heat sinks

Copper and copper

alloys

Wire, printed circuit board circuitry and heat sinks, component leads, terminals, bus bars, nuts and bolts, and radio frequently gaskets

Cadmium plating Sacrificial protective coating on ferrous fastener hardware and on electrical connectors

Nickel plating Barrier-type layer between copper and gold in electrical contacts, for corrosion protection on electrical connectors,

printed circuit board heat sinks, electrical bonds in EMI applications, and for compatibility in dissimilar-metal junctions

Tin plating For corrosion protection, solderability, and compatibility between dissimilar metals, on electrical connectors, radio

frequency shields, filters, small enclosures, component leads, and automatic switching devices

Solder and solder

plating

For joining, solderability, and corrosion protection

Beryllium Inertial guidance instruments

Source: Ref 8

The principles of galvanic corrosion are discussed elsewhere in this Volume, and will not be dealt with in this article To minimize the effects of galvanic corrosion in electronic equipment, the joining of dissimilar metals as defined in MILSTD-889 should be avoided wherever possible Where similar metals cannot be used because of design requirements, one or more of the following steps should be taken:

• Design the couple so that the area of the more noble metal (cathode) is appreciably smaller than the area

of the more active metal (anode) Decrease the cathode area by painting or coating

• Plate the cathode (and/or anode) with a compatible metal

• Interpose a compatible metallic washer or gasket between the dissimilar metals

• Interpose a nonabsorbing, insulative washer, gasket, or coupling between the joined metals

• Paint the faying surfaces prior to joining

• Seal the interfaces to preclude entrance of moisture

• Place the electronics in a hermetically sealed enclosure and pressurize with an inert gas If impossible to hermetically seal, used elastomeric seals and maintain relative humidity below 40% at room temperature (see the section "Moisture Intrusion Into Black Boxes" in this article)

• Where electrical bonding is required, follow the guidelines discussed in the section "Electromagnetic Interference (EMI)" of this article

The prolific use of noble (more cathodic) metal platings on anodic substrates as corrosion barriers for electronic components is a related problem Where the plating is excessively thin (porous), is cracked due to flexure or differential thermal expansion, or suffers mechanical damage, then the anodic base metal is exposed In the presence of moisture, accelerated galvanic corrosion will occur This corrosion is particularly severe because of the unfavorable area relationship (large cathode-to-anode area ratio) Care must be taken to ensure that minimum thicknesses of plating necessary to eliminate porosity are specified on engineering drawings and are enforced by quality control testing Reflowing of tin and solder coatings helps to eliminate porosity Where flexure of the plating is anticipated, for example,

on component leads, a low-stress ductile plating should be used In this example, a solder plate would be preferred to the more brittle electroless nickel plate

Electromagnetic Interference (EMI). Even the most reliable electronic circuits and components are susceptible to malfunction due to interference from natural and man-made electromagnetic emissions Filtering, shielding, and grounding are three ways to minimize these effects, known as electromagnetic interference, and to keep the equipment itself from being a source of interference to other electronics (Ref 9)

Shielding and grounding involve surrounding the electronics with a conductive shield or envelope grounded back to the main structural section or airframe This requirement has resulted in an increasing number of electrically bonded interfaces and grounding paths in which low direct current (dc) resistance must be maintained Extensive use of the light (active) metals for housings and chassis in electronic systems, together with the inability to use insulating-type coatings at electrically bonded interfaces, has compounded the difficulties of corrosion control Normal corrosion rates, accelerated

by galvanic action where dissimilar metals must be bonded, can lead to an increase in electrical resistance due to oxide and corrosion product formation This will result in the eventual loss of EMI protection

Protective treatments for electrically bonded interfaces must satisfy the conflicting requirements of corrosion control while maintaining good electrical continuity Several methods are helpful in meeting these requirements, including plated metals, chemical conversion films, metal-to-metal contacts sealed with an organic moisture barrier, water-displacing ultrathin-film corrosion-preventive compounds as specified in MIL-C-81309, and combinations of the above

The use of MIL-C-5541, Class 3, chemical conversion films for the protection of aluminum in electrically bonded interfaces has been investigated for aerospace electronic equipment (Ref 10) These Class 3 films combine good corrosion resistance (that is, they must pass 168 h of exposure to 5% salt spray test per ASTM B 117) with low electrical resistance,

a specified in MIL-C-81706

The effects of temperature, mechanical action, and aging on Class 3 chemical film properties have also been evaluated (Ref 10) All three effects are important because they could influence the reliability of EMI protection High temperatures are known to degrade the corrosion resistance of chemical films Scratches or abrasion could expose base metal and degrade corrosion protection Finally, long storage and service life requirements of 10 years or more for military hardware, much of which is without inspection, make the effects of aging an important consideration Electrical resistance must remain stable to meet the requirements for radio frequency junctions specified in MIL-B-5087, "Bonding, Electrical, and Lightning Protection for Aerospace Systems." The requirement is a maximum dc resistance of 0.0025 /junction Three principal conclusions were reached in this investigation

First, Class 3 chemical films of MIL-C-5541 will withstand temperatures to 65 °C (150 °F) for extended periods of time without loss of properties The maximum exposure without serious degradation is a few hours at 95 °C (200 °F) Thermal treatments or processing that exceed this exposure cannot be tolerated

Second, scratches up to 1.6 mm ( in.) wide and mechanical abrasion where the chemical film is not completely removed pose no threat Apparently, hexavalent chromium (Cr6+) leaching slowly out of the film acts as a corrosion inhibitor to protect bare aluminum

Finally, preliminary aging results indicate that the properties of chemical films remain stable at both low (20% relative humidity) and high (cycling temperature humidity per MIL-STD-202, Method 106) humidities after 1 year and 2 years of exposure The low-humidity test represented exposure inside a sealed black box, but 10 days of exposure to the high humidity is a standard accelerated corrosion test for electronic systems

These results indicate that MIL-C-5541, Class 3, chemical films are an excellent choice for the protection of aluminum in electrical bonding applications Galvanic effects are nonexistent, as they are with noble metal platings; the chemical films are easily repaired; and the properties remain stable Where aluminum is joined to more noble metals (for example, nickel, stainless steel, or silver), sealing of the interface with an organic sealant will be required

A recent trend in commercial electronic system is the movement away from metal and toward plastic enclosures (Ref 11) Metals are excellent shielding materials, while plastics are transparent to electromagnetic signals Two approaches are being used to impart metal shielding properties to plastics One is to incorporate conductive fillers in molded plastics, and the other approach is to apply metallized surface coatings to the interior surfaces of plastic enclosures Both techniques have some major problems that are currently being addressed

System-Generated Electromagnetic Pulse. A problem related to EMI protection is protection against generated electromagnetic pulse This is a new requirement for military hardware Secondary electrons emitted from an irradiated metal and accelerated out through the double layer at the metal surface at high velocity are capable of harming electronic circuits or components that they might contact To prevent this, a gas fill should be used inside of a sealed

system-black box that will slow down or absorb electrons, and a low atomic number (low-Z) coating must be applied to the inside surfaces of the box to reduce the electron emission efficiency of those surfaces The low-Z coating is defined as a coating containing less than 1% by mass of elements of atomic number (Z ) greater than 9 The thickness and the coverage requirements for the low-Z coating will vary, depending on the atmosphere/lack of atmosphere within the box In many cases, a clear epoxy polyamide coating 0.025 to 0.1 mm (1 to 4 mils) thick has been used for this purpose The low-Z

coating can substitute for a corrosion-preventive coating on interior surfaces

Flux Residues. The main functions of a soldering flux are to remove oxides, tarnish films, and other impurities from the surfaces of metals being soldered and to exclude atmosphere from the surfaces faces during soldering to prevent the formation of new oxides A soldering flux should also lower the surface tension of the molten solder, allowing the solder

to flow readily and adhere to more of the base metal surfaces (Ref 12)

In the case of liquid-type soldering fluxes, many organic and inorganic materials provide excellent fluxing action, but most of these materials leave extremely corrosive residues after cooling (Ref 13) Corrosive flux residues that attack and consume the solder alloy and base metals can weaken or embrittle the soldered connections They can also increase the electrical resistance of the connections or open them entirely On the other hand, conductive corrosive residues can pick

up moisture from the atmosphere and can lower insulation resistance and form conductive paths, causing current leakage

or short circuits Nonconductive corrosion products, especially those carried by fumes, can cause damage by building up

on electrical contact surfaces

The fluxes used for electronic soldering can be divided into two categories: the solvent-soluble, or rosin types and the water-soluble types In general, the water-soluble fluxes and their residues are much more active than the rosin fluxes and residues, but mild water-soluble fluxes and highly activated rosin fluxes are available Inorganic zinc chloride-hydrochloric acid (ZnCl2-HCl) type fluxes are included in the water-soluble category, but their highly corrosive nature and the difficulty in removing the residues preclude their use in electronic soldering

Water white rosin, a widely used soldering flux material, leaves essentially noncorrosive, nonconductive, and nonhygroscopic residues Because rosin is a very weak organic acid, its practical ability to clean a metal oxide surface is limited To increase the ability of rosin fluxes to clean oxide surfaces, activating agents are added to produce rosin mildly activated (RMA) and rosin activated (RA) fluxes For RMA fluxes, and activator may be any of a number of amines, organic acids, amides, or halogen-containing materials (Ref 14) Because RMA fluxes must be noncorrosive, as specified

in military specification MIL-F-14256, only a small, limited amount of these materials may be added The activating agents for RA fluxes are usually amine-neutralized HCl or occasionally a halogen-substituted organic material These proprietary materials are water soluble and are the primary active constituents of the organic chloride type fluxes When

heated, these activators decompose and liberate HCl This liberated acid, unlike rosin, not only reduces the metal oxide but readily attacks the cleaned metal surface as well as the surface of the solder

The organic chloride type fluxes are the most active fluxes used in soldering electronic assemblies The active components are usually amine hydrochlorides (or hydrobromides) combined with water-soluble organic acids Glutamic acid hydrochloride is an example Fast and efficient removal of the residues is necessary to prevent initiation of corrosion

Water-soluble chloride-free fluxes have gained some popularity The active components of these fluxes are water-soluble organic acids The residues are similar to those of the organic chloride type fluxes, but they eliminate the possibility of chloride ion (Cl-) entering the corrosion cycle If left on the board too long, they may discolor the solder surface A water rinse is usually sufficient for flux removal If polymerization occurs during soldering, alkaline cleaners may facilitate removal

In a study performed for the U.S Army Electronics Command, it was determined that the single largest factor causing corrosion of printed circuit board assemblies was flux residues (Ref 15) To prevent this form of corrosive attack, the lowest acid content flux possible should be used for soldering (Ref 8) In addition, cleaning processes should be employed that are tailored to the type of flux used and will completely remove all residues

Component Lead Materials/Finishes. High reliability in the operation of moisture-sensitive microelectronic devices is often achieved by packaging these devices in hermetically sealed containers This generally limits the electrical feed-through leads to low coefficient of expansion materials such as Kovar, Dumet, and Alloy 42 because of the need to form a seal with glass These low-expansion materials, however, rate very poorly for solderability They are finished, therefore, with a metal that is selected for its solderability and compatibility with internal wire bonding and die-bonding operations The commonly used lead materials are difficult to plate, and many solderable finishes cannot be plated directly onto these materials Gold is predominantly used because it can be plated directly on Kovar, has good solderability, is corrosion resistant, and is compatible with bonding operations Tin and tin alloy electroplated finishes are also used in many cases for solderability and corrosion control

Although gold itself is quite inert, severe galvanic actin can occur on gold-plated Kovar if the coating is not pore free (Ref 16) A potential difference of approximately 0.6 V can be generated by the dissimilar-metal couple of gold to Kovar (Ref 17), and the area relationships are poor (large cathode-to-anode area ratio), resulting in accelerated attack of the Kovar base material

Tin or tin-lead solder coatings, on the other hand, are preferable for corrosion control because they are close to Kovar in the electromotive series and because the potential difference is low (0.02 V) (Ref 17) Quite often, gold-plated component leads are pretinned, that is, immersed in the solder pot to give a solder coating Studies have shown that gold coatings as thick as 7.5 m (300 in.) completely dissolve in molten solder to give a solder-Kovar bond On most metals, 2.5 m (100 in.) of gold is sufficient for solderability Greater thicknesses are of little value because the gold simply dissolves

in the molten solder The pretinning procedure effectively eliminates the galvanic-corrosion problem

Other corrosion-sensitive areas for component leads are the lead-glass interface at the lead egress from the container and lead bends, which are susceptible to stress-corrosion cracking When the glass is fused during the sealing process, a meniscus is formed, resulting in a thin coating of glass extending out on the lead material The lead is then gold plated up

to this meniscus The thin coating of glass can be cracked or broken away during plating and subsequent handling, packaging, lead bending, and soldering, leaving bare lead material This site is very susceptible to corrosive attack because of the galvanic couple formed between the more noble material finish and the base metal

Stress-corrosion failures of leads have been observed for both transistor cans and integrated circuit packages when stress and moisture were combined (Ref 16, 18) Tensile stresses arise from lead bending during installation or even by differential thermal expansion in a rigid mounting because of the low expansion of Kovar alloy and the high expansion of the rigid circuit board The practice of pretinning and organic conformal coating has been an effective solution to both of the above corrosion-sensitive areas

Special care must be taken in the pretinning of leads on nonhermetic devices, particularly nichrome-film resistors Thermal stress from the solder-dipping process can damage epoxy-to-metal end seals and allow the entry of fluxes, chlorinated solvents, and moisture The nichrome film is susceptible to corrosion in the presence of moisture and a chloride contaminant; corrosion (electrolysis) can become rapid if a bias is imposed The temperature of the solder, the residence time in the solder, and the proximity of molten solder to the body of the resistor are critical parameters that