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Aviation History
1963
1963 - 0483.PDF
FLIGHT International, 4 April 1963 461 SUPERSONIC NUTS AND BOLTS Engineering Details of Mach 2 Design ALMOST every lecture and symposium on supersonic trans ports has been aimed at defining the problems rather than suggesting how they may be solved. Now, in a paper given before the Society of Licensed Aircraft Engineers and Technolo gists, Mr M. A. Taunton, a design engineer with Bristol Aircraft, has given some facts about the way in which various engineering problems are being tackled in the design of the Anglo-French Mach 2 project. Mr Taunton dealt in detail with six aspects of the subject. Materials Under this heading the lecturer reviewed the effect of temperature on the strength of metallic structural materials. Fig 1 shows the kind of temperatures that occur at the leading edge and towards the rear of a supersonic aircraft at various Mach numbers. There are various parameters which may be used to define the structural efficiency of a material depending, of course, on the type of loading. By way of an example Mr Taunton pro duced Fig 2, based on the ratio of tensile strength to density, with the general conclusions that: (a) conventional high-strength alu minium alloys are very efficient up to 100 C, i.e., up to Mach 2.0; (b) that resistant aluminium alloys show reasonable efficiency up to 120°C, i.e., up to Mach 2.2; (c) titanium alloy or 12 per cent Cr steel is necessary at 280°C, i.e., Mach 3.0. Apart from the reduced static strength of aluminium alloys exposed to elevated temperatures while under sustained load the Fig I Variation of structure tempera tures with speed MACH NUMBER material will creep, resulting in a permanent deformation of the aircraft. As a basis for design it has generally been accepted that the maximum permissible permanent deformation should not be greater than that caused by a single application of proof load—i.e., that appropriate to a 0.1 per cent strain. With most heat-resistant materials this is not likely to impose a significant penalty, as fatigue considerations are likely to require similar stress limitations. Thermal Stresses One of the principal design problems of the supersonic transport aircraft is the stress induced by uneven heating of the structure; although present in subsonic aircraft, such heating is of negligible proportions. Critical temperature-gradients occur during acceleration and deceleration phases of flight. Fig 3 shows the typical flight plan of an SST. During the 6.5min in which the aircraft is accelerating from Mach 1.0 to Mach 2.2 the skin tem perature reaches 120QC, while the temperature at the centre of the spar web is only 20 C. In aluminium alloy this corresponds to a strain difference of 2.4 < 103—sufficient to impose tension on the spar web and compression on the skins. Allowing for the fact that the skin area is generally large in com parison with the web area, the majority of the elastic strain occurs in the spar webs, giving stresses only slightly below 24,0001b/sq in. The object of strain-relieving devices such as fluled webs and braced spars is to reduce the centre-line longitudinal stiffness of the web. Because such devices are not appropriate for highly loaded shear webs—for instance, where an undercarriage is attached to a spar —there is an induced fatigue problem, and the lecturer went on to describe the test equipment and procedure for applying repeated heating and cooling Cycles to a typical test panel. Temperature gradients are not peculiar to the wing. The tem perature of the outside skin of the fuselage rises rapidly during acceleration, while internal temperatures lag behind. Thus the shell expands but is restrained locally by the colder frames or bulkheads, resulting in significant bending stresses in skins and stringers. The lecturer produced a diagram showing the typical deflected shapes of a skin-and-stringer structure under these con ditions. The length of cylinder distortion varies with the longitudi nal bending stiffness of the stringers in such a way that the bending stresses do not vary significantly. Careful detail design is necessary to minimize stress concentrations in the areas of maximum thermal stress, and to minimize the radial restraint from bulkheads where possible. Glazing Problems High temperatures and steep temperature gradients are not the ideal environments for the glass and plastic materials currently used for airliner windscreens and cabin windows. Considerable research, however, has yielded small improvements in the thermal stress resistance of conventional materials and by careful detail design satisfactory solutions to the glazing prob lem seem possible. To provide acceptable vision for landing, without excessive drag in supersonic flight, the Concorde will have a two-part visor to shield the forward windscreen panels at supersonic speeds. The permanently exposed side panels of the screen are subject to fairly severe thermal conditions, but they do not have to resist bird impact. The main design criteria for the forward panels are: (a) to provide a fail-safe structure capable of supporting a cabin pressure differential of 10.7; (b) to resist the impact of a 41b bird at 500 m.p.h. (Vc at 8,000ft); (c) to provide adequate means for de-icing and de-misting in stand-off and ground conditions. ^ • ==\ HIGH ' STRENGTH ALUMINIUM " ALLOY —— " .TITANIUM ALLOY / ./ I2'4 Cr STEEL / ,18/8 STEEL / 1 >v j**~^—^_^y~—' —. \ HEAT RESISTANT \ N. ,,-ALUM INIUM ALLOV i i i Fig 2 (left) Elevated tem perature strength: weight ratio Fig -3 (right) A typical flight plan 1,600 1,200 aoo 400 - / M 2 2 J M-IO |rf tu>~T,ME- TRUE AIRSPEED | 1 / MINUTES A. , ALTITUDE / /> 97 ;IO = 2 2 M-ro ij QS3]\ - 40,000 100 200 300 TEMPERATURE ("c) 200 V 1,600 1,100 2000 DISTANCE (NAUTICAL MILES)
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