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Aviation History
1953
1953 - 0935.PDF
FLIGHT, 17 July 1953 01 Light Alloys and Production Problems Two Further Papers from the A.F.I.T.A. Congress SIX papers by British aeronautical engineers were pre sented at the international congress organized by PAssociation Francaise des Ingenieurs et Techniciens de l'Aeronautique (A.F.I.T.A.) on the occasion of the Paris Aero Show. The contribution by Mr. Petter of Follands was summarized in our issue of July 3rd. We here present slightly condensed versions of two further papers, High- strength Light Alloys, by Mr. A. Black of Vickers-Super- marine's research department, and The Transition from Prototype to Production in Aircraft Manufacture, by Mr. W. S. Hollyhock of the production department of Hawker Aircraft, Ltd. Of the other three papers, two had already been given in England before the Royal Aeronautical Society, and we hope to give an account of the remaining one, Some Problems in the Design of High-speed Aircraft and Theoretical Methods of Solution, by Mr. B. A. Hunn of Hawkers, in a forthcoming issue. HIGH-STRENGTH LIGHT ALLOYS The object of Mr. Black's paper was to discuss various limita tions and manufacturing difficulties which had arisen in the application of such aluminium alloys to British aircraft construc tion, and to indicate methods which had proved successful in overcoming these difficulties. The alloys considered were those of specifications D.T.D. 363A and D.T.D. 683. In the introduc tory tables (I and II, right), the properties of the older alloy. B.S.S. L.65, were included for comparison. Considered in turn by the lecturer were certain fundamental properties. The first was that of ductility. In common with the majority of metallic alloys, Mr. Black stated, the increased strength of the materials was accompanied by a lowering of ductility as measured by the elongation obtained from a tensile test. A low elongation value was undoubtedly undesirable in an aircraft material, because only small amounts of permanent distortion could take place before a fracture occurred, and large amounts of cold work could not be withstood without fracture. The specification values for the alloys considered indicated a ratio of proof stress to maximum stress of approximately 85 per cent, but in fact this value commonly exceeded 90 per cent. In these circumstances, the range of stress over which no permanent distortion occurred was greatly increased, and therefore there was a decreased range of strain over which permanent distortion with out failure could take place. The new materials showed large increases in tensile strength, but unfortunately no improvement in fatigue strength. In the stressing of aircraft structures, the fatigue strength of the material used was rarely considered, and it must be assumed that the comparative absence of fatigue failures was due to the factors of safety which were applied to the calculations. Directional effects were next considered by the lecturer, who said that the ductility of the alloys varied greatly with the direction of grain. The reduced values in the transverse direction might, in fact, be lower than could be considered acceptable for an aircraft structural material, the effect being extremely marked in extru sions and also in large forgings with widely differing sections. A decrease in tensile strength across the grain might also be present, and typical figures for D.T.D. 683 forgings showed well the extent of these directional effects (Table III). It had been noted that the alloys had a reduced capacity to absorb permanent distortion, a low ductility and a low ratio of fatigue strength to tensile strength. These properties were com bined with a high normal stress level in use, and the result was a greatly increased sensitivity to stress concentrations resulting from bad design or surface notches. It was, therefore, essential in design to minimize stress concentration. Internal stress was produced both by cold bending and by rapid quenching. The magnitude of such stress depended on the relationship between the elastic limit and the maximum stress. If this ratio was high, as was the case with these alloys, residual internal stress would reach a high level, and was very undesirable. Turning to the question of methods of overcoming manufac turing difficulties, the speaker dealt first with bending and mani pulation. Cold bending had for many years been the most con venient method of producing required shapes or of correcting unwanted distortion. The advent of the high-strength alloys had introduced considerable difficulties, and the alteration of estab lished techniques was necessary. In particular, an assessment of TABLE I: CHEMICAL COMPOSITION OF ALLOYS Element Copper Magnesium Silicon Iron Manganese Zinc Titanium Chromium Aluminium Max. 4.8 0.85 0.9 1.0 1.2 0.3 L65 Min. 3.8 0.55 0.6 0.4 Pe Typ. 4.25 0.6 0.75 0.4 rcentage Compositi D.T.D. 363A Max. 3.0 4.0 0.6 0.6 1.0 8.5 0.3 1.0 Min. 4.0 Typ. 1.8 2.0 0.3 7.0 0.13 the remainder on D.T.D. 683 Max. 1.5 3.5 0.5 0.5 4.0 6.5 0.3 0.5 Min. 2.0 0.25 4.5 Typ. 1.3 2.5 0.25 5.75 0.13 TABLE II: PHYSICAL PROPERTIES Properties:— Specific gravity Weight, Ib/cu in ... Specific heat 0-100 deg C. calories/8 Coefficient of expansion per degree,.,. 20-300 deg C Thermal conductivity, C.G.S. units Specific resistance, microhms'cc Heat treatment:— Nominal solution treatment temp.... deg. C Ageing temperature, deg C Ageing time, hours L.65 2.79 0.101 0.21 25 x 10-« 0.27 4.5 505 170 10 D.T.D. 363A 2.85 0.103 0.24 23 x 10-» 0.32 5.0 460 135 8 D.T.D. 683 2.82 0.102 0.24 23 x 10-<> 0.32 5.0 465 135 12 Note: The figures given are the normal recommended temperatures for solution treatment and precipitation, and times for precipitation. They may be varied by the material manufacturer or by the aircraft constructor to suit special circumstances the permissible degree of bending was needed. In attempting to assess the permissible amount of bend on a given section, the elastic bending formula had been used by various firms in an attempt to forecast the resulting elongation. The formula was ioo zT+i where T = the ratio of the radius of bend to the depth of section. Such a formula did not take into account the angle through which the material was bent, and applied only within the elastic limit. From the results of a small number of tests carried out at the lecturer's firm on the bending of the 8 per cent zinc alloy, it had been possible to evolve a formula giving the relationship between the residual elongation and the various bend factors which arose. This formula was KOVD e 2T+1 where e = residual elongation per cent D = depth of rectangular section T = r:D r = radius of bend block 6 = angle of bend K = a constant for the material in a given heat treatment condition. Approximate values for K for the 8 per cent zinc alloy were :— annealed condition, 1.10; freshly quenched, 1.50; fully heat treated, 2.05. In view of the small number of tests from which this information was»obtained, it was put forward simply as a basis on which further tests might be carried out. It had been seen that the permissible degree of cold bending must be assessed by considering the properties of the material in the plastic range with particular reference to the angle of bend, radius of bend, and size and condition of material. Such cold bending would inevitably result in an internal stress—possibly of an extremely high order—unless subsequent full heat-treatment was carried out. Cold bending was normally carried out by means of hydraulic or fly-presses, using three-point loading. The use of TABLE III: TRANSVERSE EFFECTS Property Maximum stress, tons/sq in 0.1 per cent proof stress, tons/sq in Elongation per cent Specification value 32 27 7.0 Transverse value 28 25 2.5
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