Volume 13 - Corrosion Part 15 pptx

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 embrittl ement 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 titanium- cadmium. (a) Pin showing location of failure. Actual size. (b) Fail ure 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× 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 e mbrittlement 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 a t 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). We ar 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 st ainless 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 HNO 3 , 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 Hydrog en 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 pressur e. 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 sh owing 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 f ailure. 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 N 2 H 4 . 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 nitrid ing 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. [...]... steel HY -1 30 steel HY-140 steel 1095 spring steel 300-series stainless steels (unsensitized) 2 1-6 -9 stainless steel 20Cb stainless steel 20Cb-3 stainless steel A-286 stainless steel AM350 stainless steel AM355 stainless steel Almar 362 stainless steel Custom 455 stainless steel 1 5-5 PH stainless steel PH1 4-8 Mo stainless steel PH1 5-7 Mo stainless steel 1 7-7 PH stainless steel Nitronic 33 Wrought aluminum... high-pressure hydrogen at room temperature Alloy 250 maraging steel Type 410 stainless steel 1042 steel (quenched and tempered) 1 7-7 PH (TH1050) HP 9-4 -2 0 alloy steel H-11 high-strength steel Inconel alloy X-750 René 41 ED nickel 4140 steel Inconel alloy 718 MP35N Type 440C stainless steel Ti-6Al-4V (solution treated and aged) Monel alloy 400 D-979 stainless steel Nickel 270 CG27 stainless steel ASTM A 515, ... 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 Alloy Aluminum alloy 7079-T6 AM-355 stainless steel 1 7-7 PH stainless steel Aluminum alloy 7079-T6 Aluminum alloy 7075-T6 Aluminum... Haynes alloy 188 MP35N Ti-3Al-2.5V Ti-6Al-4V Ti-13V-11Cr-3Al Magnesium, M1A Magnesium, LA141 Magnesium, LAZ933 (a) (b) (c) T73 T73 T6 All All All All As cast As cast As cast As cast As cast All All All All All Annealed Annealed All All All All All All All 37 AT, HT(c) AT, HT(c) 37 90 40 37 10 37 37 40 40 50 50, Annealed Annealed All All All All All All All Stabilized All High-magnesium alloys 5456,... comparative stresscorrosion 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 Alloy Ferrous alloys Carbon steel (1000 series) Low-alloy steel ( 4130 , 4340, D6AC, etc.) Music wire (ASTM 228) HY-80 steel HY -1 30 steel HY-140 steel... 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 2 1-6 -9 stainless steel (b) 130 × (c) 45×... boosters Alloy Aluminum alloy 7079-T6 AM-355 stainless steel 1 7-7 PH stainless steel Aluminum alloy 7079-T6 Aluminum alloy 7075-T6 Aluminum alloy 7075-T6 Aluminum alloy 2024-T4 1 7-7 PH stainless steel 1 7-7 PH stainless steel Aluminum alloy 7178-T6 7079-T652 7079-T6 7079-T6 Material form Forging Bar Sheet Forging Plate Bar Bar Sheet Sheet Forging Forging Forging Forging Failure occurrence Prelaunch Prelaunch Prelaunch... 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 1 7-7 PH RH950 stainless steel Belleville washers is shown in Fig 59 The washers were... 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... PH1 3-8 Mo stainless steel 1 5-5 PH stainless steel 1 7-4 PH stainless steel Wrought aluminum alloys 2024 rod, bar, extrusion 2024 plate, extrusions 2124 plate 2048 plate 4032 7001 7049 7050 7075 7175 7475 7178 Cast aluminum alloys 319.0, A319.0 333.0, A333.0 Magnesium alloys AZ31B AK60A (a) Condition 1240 to 138 0 MPa (18 0-2 00 ksi) ultimate tensile strength 1240 to 138 0 MPa (18 0-2 00 ksi) ultimate tensile . 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),. through the tower leg bolts. The titanium tubing specified was Ti-6Al-4V with an 8 9- mm (3.5-in.) outside diameter and a 3.2-mm (0.125-in.) wall. The titanium was produced by a hot extrusion process. × 0.020 × 0.030 in. deep). 15 . (c) Oxidation of niobium during one simulated entry cycle at a 148 5- C (270 0- F) peak temperature using 0.5 7- kPa (1 2- psi) oxygen partial pressure. Craters A,