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Aviation History
1957
1957 - 1385.PDF
20 September 1957 475 f,g. 5. Location of titanium in X-3 aircraft (left) and in a typical DC-7 engine nacelle (Schapiro and LaBombard). The experience of titanium usage by the Douglas Aircraft Company, a pioneer in this field, was the subject of a significant paper presented on Tuesday, September 10, by Leo Schapiro. Co-author of the paper was Emerson LaBombard, and chairman of the conference session was Sir Sydney Camm, of Hawker Aircraft. _ , . . . Douglas first became interested m titanium as a constructionmaterial in 1948 (the paper began) when it undertook the task of designing and building the X-3 supersonic research aeroplane.The design incorporated a turbojet engine buried within the aft fuselage. The engine manufacturer was to furnish the tail-pipe,and the fuselage and wing structure would be conventionally of aluminium forward of the plane of the tail cone. Afr of thisstation, stainless steel would be the first-choice material unless the then-new lightweight metal titanium could be formed to theshapes required. Improved aeroplane performance would result from the lighter weight construction, but nothing was yet knownabout the workability or availability of titanium. On June 15, 1948, the first ingot of ductile titanium producedfor commercial sale was melted by the Remington Arms Company at Bridgeport, Connecticut. The material was unalloyed and wastermed "commercially pure." It was made available in four tempers: annealed, quarter-hard, half-hard and full-hard, andthe annealed and half-hard tempers were finally selected for production use.Laboratory tests were conducted to predict formability, and to evaluate shop-handling problems and expected dissimilar-metalscorrosion problems. Procurement for aeroplane construction began, limited only by a 35 lb maximum weight per piece becauseof the restricted ingot size being produced. Sheets were ordered up to 44in wide, up to 108in long and in a variety of gauges from0.016 to 0.250 in annealed temper and from 0.025 to 0.156 in the half-hard temper. For the 849 titanium parts for each of twoaeroplanes, the total number of sheets purchased did not all meet the company's procurement requirements. The failures and fabrication losses were surprisingly few con-sidering the complete lack of knowledge and experience available at the start of production. Practically none of the losses was dueto inherent characteristics of the material. They were almost entirely caused by correctable mill defects.Stretch forming, brake forming and hydro-press forming were all tried initially as room temperature operations. As the tendencyfor material defects to initiate cracks became a major factor in producing usable parts, heating was applied to all of these opera-tions, except stretch forming. The 849 parts per aeroplane, constituting a weight of 545 1b,became the internal shroud around the tail pipe, the fuselage skin and frames at the extreme aft end and the skin and frames of theempennage. The use of the titanium parts instead of stainless steel saved 395 lb. One material deficiency did arise after all detail parts werefabricated and assembly had started. A 0.035 gauge afterburner liner broke during shop handling, with indications of little duc-tility. Static tension tests with specimens cut from the part were satisfactory but bend tests were not. The microstructure was quitea surprise. The needle-shaped precipitates within the grains and the heavy constituent at the grain boundaries were the company'sfirst introduction to hydrides. This experience with finished parts constituting 65 per cent ofthe weight of purchased material was considered a satisfactory experience in the light of the material supplier's state of develop-ment, and the company next considered applying titanium to its then-new DC-7 commercial transport. Here it was planned to havethe firewall and skin and frames of die nacelle of titanium, to save weight over the stainless steel used on the DC-6. Only theannealed temper of commercially pure titanium was expected to be used, as the parts to be made would give minimum formingdifficulties in this temper. With the improvements effected by the material supplier, the procurement specification was modifiedto reflect the closer control of chemical composition then possible, but a less stringent bend quality requirement. Stretching and brake forming became room temperature opera-tions and parts breakage during shop handling became a new experience since they were no longer initiated by mill defects. InJanuary 1953 it was necessary to control still more shop breakage and low notched-strengdi (as the latter might affect serviceability)by establishing upper limits on the mechanical properties require- ments. These modifications were deemed desirable to retain thefavoured room temperature forming operations and to encourage the materials manufacturers to better control of dieir processingoperations. Sub-zero notched testing was avidly applied during this periodof shop-handling breakage and modification of procurement re- quirements. It was believed diat this test might become suitablefor distinguishing between brittle and ductile material for stretched parts. A considerable amount of such testing demonstrated thatthe sub-zero notched • strength of the material varied from 0.6 to 1.4 ratio (sub-zero notch strength divided by room-temperaturetension ultimate) and that a ratio of 1.0 might be an acceptable demarcation between acceptable and non-acceptable material. Widi the addition to the procurement specification of therequirement "freedom from hydride needles in the microstruc- ture," the company's relief was only short-lived. The loweredbend-ductility, decreased fatigue-strength and impaired ability to sustain a load in the presence of a notch with rapidly applied loadat low temperatures, were all associated with the presence of hydride needles, but all these properties were not at acceptablelevels when hydride needles were no longer discernible in the microstructure of newer materials. Consequent limitations in-cluded the use of a new "tear" test, and commercially pure titanium conforming to this requirement eliminated all shopbreakage. In die early consideration of titanium for a firewall, a 15 min2,000 deg F flame test was applied to the material without any loss in strength resulting. A more practical test of the ability ofthis material to resist fire occurred in December 1955 when a DC-7 operating out of Rome experienced a fire on its No. 3engine. The fire-resisting qualities were emphatically confirmed. In the four years that 450 lb of titanium per aircraft had beenflying on the early DC-7s (333 parts per aircraft), only nine parts which had failed in service had been reported to Douglas. All ofthese had been cracks in skins. In some cases, hydrogen in excess of the present requirements was associated with the service crack-ing. In one case no material deficiency was present and an engineering change to a heavier gauge was made. In its entirety,the titanium experience in the DC-7 programme was deemed sufficiently satisfactory to increase its usage on later models. TheModel DC-7C now had 554 titanium parts widi a flying weight of 800 lb.It was a logical step to continue using this structural material in the turbojet-powered DC-8 civil transport. Engine shroudingand pylons were expected to use 600 1b flying weight of annealed commercially pure titanium. Rip stoppers at frames and cut-outswere expected to use 330 lb flying weight of alloy sheet, and 15 lb 'HALF OF ARD MAIN ENTRANCE • JAMB DOUBLER JAMB DOUBLER ON ALL FOUR LOWER CARGO COMPARTMENT DOORS ENTIRE AFT MAIN ENTRANCE DOOR JAMB DOUBLER RIP STOPPERS ABOVEj FLOOR LINE ON FRAMES 1 IN THIS AREA UPPER HALF OF AFT SERVICE DOOR JAMB DOUBLER Fig. 6. Location of titanium in DC-8 aircraft (Schapiro and LaBombard). RIP STOPPERS BELOW &*?* LINE ON FRAMESIN THIS AREA ENTIRE FORWARD SERVICE DOOR JAMB DOUBLER RIP STOPPERS BELOW FLOOR LINE ON FRAMES IN THIS AREA
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