230 Principles of stressed skin construction Fig. 7.1 1 Wing ribs for the European Airbus (courtesy of British Aerospace). The different structural requirements of aircraft designed for differing operational roles lead to a variety of wing constructions. For instance, high-speed aircraft require relatively thin wing sections which support high wing loadings. To withstand the correspondingly high surface pressures and to obtain sufficient strength, much thicker skins are necessary. Wing panels are therefore frequently machined integrally with stringers from solid slabs of material, as are the wing ribs. Figure 7.11 shows wing ribs for the European Airbus in which web stiffeners, flanged lightness holes and skin attachment lugs have been integrally machined from solid. This integral method of construction involves no new design principles and has the advantages of combining a high grade of surface finish, free from irregularities, with a more efficient use of material since skin thicknesses are easily tapered to coincide with the spanwise decrease in bending stresses. An alternative form of construction is the sandwich panel, which comprises a light honeycomb or corrugated metal core sandwiched between two outer skins of the stress-bearing sheet (see Fig. 7.12). The primary function of the core is to stabilize the outer skins, although it may be stress-bearing as well. Sandwich panels are capable of developing high stresses, have smooth internal and external surfaces and require small numbers of supporting rings or frames. They also possess a high resistance to fatigue from jet efflux. The uses of this method of construction include lightweight ‘planks’ for cabin furniture, monolithic fairing shells generally having plastic facing skins, and the stiffening of flying control surfaces. Thus, for example, the ailerons 7.4 Fabrication of structural components 231 Typical flat panel edging methods Typical flat panel joints and corners Typical fastening methods Fig. 7.1 2 Sandwich panels (courtesy of Ciba-Geigy Plastics). 232 Principles of stressed skin construction and rudder of the British Aerospace Jaguar are fabricated from aluminium honey- comb, while fibreglass and aluminium faced honeycomb are used extensively in the wings and tail surfaces of the Boeing 747. Some problems, mainly disbonding and internal corrosion, have been encountered in service. The general principles relating to wing construction are applicable to fuselages, with the exception that integral construction is not used in fuselages for obvious reasons. Figures 7.7, 7.8 and 7.9 show that the same basic method of construction is employed in aircraft having widely differing roles. Generally, the fuselage frames that support large concentrated floor loads or loads from wing or tailplane attach- ment points are heavier than lightly loaded frames and require stiffening, with additional provision for transmitting the concentrated load into the frame and hence the skin. With the frames in position in the fuselage jig, stringers, passing through cut-outs, are riveted to the frame flanges. Before the skin is riveted to the frames and stringers, other subsidiary frames such as door and window frames are riveted or bolted in position. The areas of the fuselage in the regions of these cut-outs are reinforced by additional stringers, portions of frame and increased skin thickness, to react to the high shear flows and direct stresses developed. On completion, the various sub-assemblies are brought together for final assembly. Fuselage sections are usually bolted together through flanges around their periph- eries, while wings and the tailplane are attached to pick-up points on the relevant fuselage frames. Wing spars on low wing civil aircraft usually pass completely through the fuselage, simplifying wing design and the method of attachment. On smaller, military aircraft, engine installations frequently prevent this so that wing spars are attached directly to and terminate at the fuselage frame. Clearly, at these positions frame/stringer/skin structures require reinforcement. P.7.1 Review the historical development of the main materials of aircraft P.7.2 Contrast and describe the contributions of the aluminium alloys and steel P.7.3 Examine possible uses of new materials in future aircraft manufacture. P.7.4 Describe the main features of a stressed skin structure. Discuss the structural functions of the various components with particular reference either to the fuselage or to the wing of a medium sized transport aircraft. construction. to aircraft construction during the period 1945-70. Airworthiness and airframe loads The airworthiness of an aircraft is concerned with the standards of safety incorpo- rated in all aspects of its construction. These range from structural strength to the provision of certain safeguards in the event of crash landings, and include design requirements relating to aerodynamics, performance and electrical and hydraulic systems. The selection of minimum standards of safety is largely the concern of airworthiness authorities who prepare handbooks of official requirements. In the UK the relevant publications are Av.P.970 for military aircraft and British Civil Airworthiness Requirements (BCAR) for civil aircraft. The handbooks include operational requirements, minimum safety requirements, recommended practices and design data etc. In this chapter we shall concentrate on the structural aspects of airworthiness which depend chiefly on the strength and stiffness of the aircraft. Stiffness problems may be conveniently grouped under the heading aeroelasticity and are discussed in Chapter 13. Strength problems arise, as we have seen, from ground and air loads, and their magnitudes depend on the selection of manoeuvring and other conditions applicable to the operational requirements of a particular aircraft. The control of weight in aircraft design is of extreme importance. Increases in weight require stronger structures to support them, which in turn lead to further increases in weight and so on. Excesses of structural weight mean lesser amounts of payload, thereby affecting the economic viability of the aircraft. The aircraft designer is therefore constantly seeking to pare his aircraft’s weight to the minimum compatible with safety. However, to ensure general minimum standards of strength and safety, airworthiness regulations (Av.P.970 and BCAR) lay down several factors which the primary structure of the aircraft must satisfy. These are the limit load, which is the maximum load that the aircraft is expected to experience in normal operation, the proof load, which is the product of the limit load and the proof factor (1.0- 1.25), and the ultimate load, which is the product of the limit load and the ultimate factor (usually 1.5). The aircraft’s structure must withstand the proof load without detrimental distortion and should not fail until the ultimate load has been achieved. 234 Airworthiness and airframe loads nl (limit load) - Flight speed I Negative stall Fig. 8.1 Flight envelope. The proof and ultimate factors may be regarded as factors of safety and provide for various contingencies and uncertainties which are discussed in greater detail in Section 8.2. The basic strength and fight performance limits for a particular aircraft are selected by the airworthiness authorities and are contained in theflight envelope or Y-n diagram shown in Fig. 8.1. The curves OA and OF correspond to the stalled condition of the aircraft and are obtained from the well known aerodynamic relationship Lift = n w = f p v~sc~:~~ Thus, for speeds below VA (positive wing incidence) and VF (negative incidence) the maximum loads which can be applied to the aircraft are governed by CL,max. As the speed increases it is possible to apply the positive and negative limit loads, corresponding to nl and n3, without stalling the aircraft so that AC and FE represent maximum operational load factors for the aircraft. Above the design cruising speed V,, the cut-off lines CDI and D2E relieve the design cases to be covered since it is not expected that the limit loads will be applied at maximum speed. Values of nl, n2 and n3 are specified by the airworthiness authorities for particular aircraft; typical load factors laid down in BCAR are shown in Table 8.1. A particular flight envelope is applicable to one altitude only since CL,max is generally reduced with an increase of altitude, and the speed of sound decreases with altitude thereby reducing the critical Mach number and hence the design 8.2 Load factor determination 235 Table 8.1 Category Load factor n Normal Semi-aerobatic Aerobatic nl 2.1 + 24000/( W+ 10000) 4.5 6.0 n3 1 .o 1.8 3.0 n2 0.75nl but n2 < 2.0 3.1 4.5 diving speed V,. Flight envelopes are therefore drawn for a range of altitudes from sea level to the operational ceiling of the aircraft. Several problems require solutions before values for the various load factors in the flight envelope can be determined. The limit load, for example, may be produced by a specified manoeuvre or by an encounter with a particularly severe gust (gust cases and the associated gust envelope are discussed in Section 8.6). Clearly some knowledge of possible gust conditions is required to determine the limiting case. Furthermore, the fixing of the proof and ultimate factors also depends upon the degree of uncertainty of design, variations in structural strength, structural deteriora- tion etc. We shall now investigate some of these problems to see their comparative influence on load factor values. 8.2.1 Limit load An aircraft is subjected to a variety of loads during its operational life, the main classes of which are: manoeuvre loads, gust loads, undercarriage loads, cabin pressure loads, buffeting and induced vibrations. Of these, manoeuvre, undercarriage and cabin pressure loads are determined with reasonable simplicity since manoeuvre loads are controlled design cases, undercarriages are designed for given maximum descent rates and cabin pressures are specified. The remaining loads depend to a large extent on the atmospheric conditions encountered during flight. Estimates of the magnitudes of such loads are only possible therefore if in-flight data on these loads is available. It obviously requires a great number of hours of flying if the experi- mental data are to include possible extremes of atmospheric conditions. In practice, the amount of data required to establish the probable period of flight time before an aircraft encounters, say, a gust load of a given severity, is a great deal more than that available. It therefore becomes a problem in statistics to extrapolate the available data and calculate the probability of an aircraft being subjected to its proof or ultimate load during its operational life. The aim would be for a zero or negligible rate of occurrence of its ultimate load and an extremely low rate of occur- rence of its proof load. Having decided on an ultimate load, then the limit load may be fixed as defined in Section 8.1 although the value of the ultimate factor includes, as we have already noted, allowances for uncertainties in design, variation in structural strength and structural deterioration. 236 Airworthiness and airframe loads 8.2.2 Uncertainties in design and structural deterioration E Neither of these presents serious problems in modern aircraft construction and therefore do not require large factors of safety to minimize their effects. Modem methods of aircraft structural analysis are refined and, in any case, tests to determine actual failure loads are carried out on representative full scale components to verify design estimates. The problem of structural deterioration due to corrosion and wear may be largely eliminated by close inspection during service and the application of suitable protective treatments. 8.2.3 Variation in structural strength To minimize the effect of the variation in structural strength between two apparently identical components, strict controls are employed in the manufacture of materials and in the fabrication of the structure. Material control involves the observance of strict limits in chemical composition and close supervision of manufacturing methods such as machining, heat treatment, rolling etc. In addition, the inspection of samples by visual, radiographic and other means, and the carrying out of strength tests on specimens, enable below limit batches to be isolated and rejected. Thus, if a sample of a batch of material falls below a specified minimum strength then the batch is rejected. This means of course that an actual structure always comprises materials with properties equal to or better than those assumed for design purposes, an added but unallowed for ‘bonus’ in considering factors of safety. Similar precautions are applied to assembled structures with regard to dimension tolerances, quality of assembly, welding etc. Again, visual and other inspection methods are employed and, in certain cases, strength tests are carried out on sample structures. 8.2.4 Fatigue Although adequate precautions are taken to ensure that an aircraft’s structure possesses sufficient strength to withstand the most severe expected gust or manoeuvre load, there still remains the problem of fatigue. Practically all components of the aircraft’s structure are subjected to fluctuating loads which occur a great many times during the life of the aircraft. It has been known for many years that materials fail under fluctuating loads at much lower values of stress than their normal static failure stress. A graph of failure stress against number of repetitions of this stress has the typical form shown in Fig. 8.2. For some materials, such as mild steel, the curve (usually known as an S-N curve or diagram) is asymptotic to a certain minimum value, which means that the material has an actual infinite life stress. Curves for other materials, for example aluminium and its alloys, do not always appear to have asymptotic values so that these materials may not possess an inhite life stress. We shall discuss the implications of this a little later. 8.2 load factor determination 237 I I I I I I IO io2 io3 lo4 io5 io6 lo7 No. of repetitions Fig. 8.2 Typical form of S-N diagram. Prior to the mid-1940s little attention had been paid to fatigue considerations in the design of aircraft structures. It was felt that sufficient static strength would eliminate the possibility of fatigue failure. However, evidence began to accumulate that several aircraft crashes had been caused by fatigue failure. The seriousness of the situation was highlighted in the early 1950s by catastrophic fatigue failures of two Comet airliners. These were caused by the once-per-flight cabin pressurization cycle which produced circumferential and longitudinal stresses in the fuselage skin. Although these stresses were well below the allowable stresses for single cycle loading, stress concentrations occurred at the corners of the windows and around rivets which raised local stresses considerably above the general stress level. Repeated cycles of pressurization produced fatigue cracks which propagated disastrously, causing an explosion of the fuselage at high altitude. Several factors contributed to the emergence of fatigue as a major factor in design. For example, aircraft speeds and sizes increased, calling for higher wing and other loadings. Consequently, the effect of turbulence was magnified and the magnitudes of the fluctuating loads became larger. In civil aviation, airliners had a greater utiliza- tion and a longer operational life. The new ‘zinc rich’ alloys, used for their high static strength properties, did not show a proportional improvement in fatigue strength, exhibited high crack propagation rates and were extremely notch sensitive. Despite the fact that the causes of fatigue were reasonably clear at that time its elim- ination as a threat to aircraft safety was a different matter. The fatigue problem has two major facets: the prediction of the fatigue strength of a structure and a knowledge of the loads causing fatigue. Information was lacking on both counts. The Royal Aircraft Establishment (RAE) and the aircraft industry therefore embarked on an extensive test programme to determine the behaviour of complete components, joints and other detail parts under fluctuating loads. These included fatigue testing by the RAE of some 50 Meteor 4 tailplanes at a range of temperatures, plus research, also by the RAE, into the fatigue behaviour of joints and connections. Further work was undertaken by some universities and by the industry itself into the effects of stress concentrations. In conjunction with their fatigue strength testing, the RAE initiated research to develop a suitable instrument for counting and recording gust loads over long periods 238 Airworthiness and airframe loads of time. Such an instrument was developed by J. Taylor in 1950 and was designed so that the response fell off rapidly above 10 Hz. Crossings of g thresholds from 0.2g to 1.8g at 0.lg intervals were recorded (note that steady level flight is 1g flight) during experimental flying at the RAE on three different aircraft over 28 000 km, and the best techniques for extracting information from the data established. Civil airlines cooperated by carrying the instruments on their regular air services for a number of years. Eight different types of aircraft were equipped so that by 1961 records had been obtained for regions including Europe, the Atlantic, Africa, India and the Far East, representing 19 000 hours and 8 million km of flying. Atmospheric turbulence and the cabin pressurization cycle are only two of the many fluctuating loads which cause fatigue damage in aircraft. On the ground the wing is supported on the undercarriage and experiences tensile stresses in its upper surfaces and compressive stresses in its lower surfaces. In flight these stresses are reversed as aerodynamic lift supports the wing. Also, the impact of landing and ground manoeuvring on imperfect surfaces cause stress fluctuations while, during landing and take-off, flaps are lowered and raised, producing additional load cycles in the flap support structure. Engine pylons are subjected to fatigue loading from thrust variations in take-off and landing and also to inertia loads produced by lateral gusts on the complete aircraft. A more detailed investigation of fatigue and its associated problems is presented in Section 8.7 after the consideration of basic manoeuvre and gust loads. The maximum loads on the components of an aircraft’s structure generally occur when the aircraft is undergoing some form of acceleration or deceleration, such as in landings, take-offs and manoeuvres within the flight and gust envelopes. Thus, before a structural component can be designed, the inertia loads corresponding to these accelerations and decelerations must be calculated. For these purposes we shall suppose that an aircraft is a rigid body and represent it by a rigid mass, 111, as shown in Fig. 8.3. We shall also, at this stage, consider motion in the plane of the mass which would correspond to pitching of the aircraft without roll or yaw. We shall also suppose that the centre of gravity (CG) of the mass has coordinates 2, 3 referred to x and y axes having an arbitrary origin 0; the mass is rotating about an axis through 0 perpendicular to the +XJ’ plane with a constant angular velocity w. The acceleration of any point, a distance r from 0, is w2r and is directed towards 0. Thus, the inertia force acting on the element, bm, is w’rSm in a direction opposite to the acceleration, as shown in Fig. 8.3. The components of this inertia force, parallel to the x and y axes, are w2rSm cos 6 and w2rSn? sin 6 respectively, or, in terms of .Y and J’, w2xSm and w2ySm. The resultant inertia forces, F, and F,., are then given by F, = w xdm = S’ F,. = w ydm = wL ydm s? ’J’ 8.3 Aircraft inertia loads 239 0 CG (F, 8) LkkJ Fig. 8.3 Inertia forces on a rigid mass having a constant angular velocity. in which we note that the angular velocity u is constant and may therefore be taken outside the integral sign. In the above expressions J x drn and J y dm are the moments of the mass, nz, about the y and x axes respectively, so that Fy = w2m (8.1) F,, = JJin (8.2) and If the CG lies on the x axis, J = 0 and F,, = 0. Similarly, if the CG lies on the y axis, Fy = 0. Clearly, if 0 coincides with the CG, X = J = 0 and F, = F,. = 0. Suppose now that the rigid body is subjected to an angular acceleration (or deceleration) Q! in addition to the constant angular velocity, w, as shown in Fig. 8.4. An additional inertia force, curSrn, acts on the element Srn in a direction perpendicular to r and in the opposite sense to the angular acceleration. This inertia force has components ar6m cos e and tur6nt sin 8, i.e. axbin and aySi71, in the y and x directions respectively. Thus, the resultant inertia forces, Fy and F', are given by Fy= aydrn=cr ydm JS Fig. 8.4 Inertia forces on a rigid mass subjected to an angular acceleration. [...]... parallel to the axis of the fuselage N -T + mlacos 10" - 4 .5 sin 10" = 0 1.e N - 137.1 + 13 .5~ 0~10 ~-4 .5sin1O0=O 4 .5 kN Fig 8.6 Shear and axial loads at the section AA of the aircraft of Example 8.1 242 Airworthiness and airframe loads whence N = 124.6kN Now resolving forces perpendicular to the axis of the fuselage S - rnlusin 10" - 4 .5~ 0s = 0 10" i.e S - 13 .5 sin lo" - 4 .5 cos 10" = 0 so that S = 6.8kN... inertia force Note that the actual normal acceleration in this particular case is (n - 1)g For vertical equilibrium of the aircraft, we have, referring to Fig 8.9 where the aircraft is shown at the lowest point of the pull-out L + P+ Tsiny - nW =0 (8.12) For horizontal equilibrium T COSY +fw - D = 0 (8.13) and for pitching moment equilibrium about the aircraft' s centre of gravity La - Db - Tc - Mo - P... Fig 8.7; pn is the mass of the aircraft and a, and a,, its accelerations in the horizontal and vertical directions respectively Then, resolving forces horizontally ma, - 400 = 0 whence ma, = 400 kN Now resolving forces vertically ma, + 250 - 1200 = 0 which gives ma, = 950 kN Then 950 - 950 a, = - 3.8g m 250 /g Now taking moments about the CG I C G- 1200 x 1.0 - 400 x 2 .5 = 0 ~ from which I c ~ = 2200... Therefore, we shall take CL = 1.099 From Fig 8.10(a) CD = 0.08 75 The values of lift, tail load, drag and forward inertia force then follow: Lift L = ipV2SCL= 4 x 1.223 x 602 x 14 .5 x 1.099 = 350 00N Tailload P = n W - L = 4 5 ~ 8 0 0 0 - 3 5 0 0 0 = Drag D = i p V 2 S C D = l000N i x 1.223 x 602 x 14 .5 x 0.08 75 = 2790N Forward inertia force fW =D (from Eq (8.13)) = 2790 N In Section 8.4 we determined aircraft. .. aircraft before it is brought to rest if the touchdown speed is 25 m/s The aircraft is subjected to a horizontal inertia force ma where m is the mass of the aircraft and a its deceleration Thus, resolving forces horizontally T cos IO" - ma = 0 8.3 Aircraft inertia loads 241 A , \, " Wheel reaction R / Arrester hook Fig 8 .5 Forces on the aircraft of Example 8.1 i.e which gives T = 137.1kN Now resolving forces... aircraft is in a steady, unaccelerated, level fight condition Thus for vertical equilibrium L+P- w=o (8.7) for horizontal equilibrium T-D=O and taking moments about the aircraft s centre of gravity in the plane of symmetry La - Db - Tc - Mo -PI = 0 (8.9) For a given aircraft weight, speed and altitude, Eqs (8.7), (8.8) and (8.9) may be solved for the unknown lift, drag and tail loads However, other parameters... the yawing inertia of the aircraft 8.6.3 Gust envelope - _ I _1 "- , = ^~ I I I - ~ l .-= "-_ - _ _.l .-_ *." _ Airworthiness requirements usually specify that gust loads shall be calculated at certain combinations of gust and flight speed The equations for gust load factor in the above analysis show that n is proportional to aircraft speed for a given gust velocity Therefore, we may plot a gust envelope... introduce a scatter factor K,, ( > 1 ) to allow for this; Eq (8.47) then becomes s- (1 + C/m) + C / a ) Sa,m a - Kn(l (8.48) K, varies with the number of test results available and for a coefficient of variation of 0.1, K,, = 1. 45 for six specimens, K,, = 1.4 45 for 10 specimens, K,, = 1.44 for 20 specimens and for 100 specimens or more K,, = 1.43 For typical S-N curves a scatter factor of 1.43 is equivalent... symmetric manoeuvre l*-llll_ -s._~ _-~ _YI _I_Y _-_ -_ -* I,_I_Y_LIY.I-Ylli In a rapid pull-out from a dive a downward load is applied to the tailplane, causing the aircraft to pitch nose upwards The downward load is achieved by a backward movement of the control column, thereby applying negative incidence to the elevators, 246 Airworthiness and airframe loads Fig 8.9 Aircraft loads in a pull-out from a dive... Section 8.4) (8. 15) At the lowest point of the pull-out, e = 0, and V2 n =-+ I gR V Fig 8.11 Aircraft loads and acceleration during a steady pull-out (8.16) 250 Airworthiness and airframe loads We see from either Eq (8. 15) or Eq (8.16) that the smaller the radius of the flight path, that is the more severe the pull-out, the greater the value of n It is quite possible therefore for a severe pull-out to overstress . Thus 4 .5 g mla =-3 g= 13.5kN Resolving forces parallel to the axis of the fuselage N - T + mlacos 10" - 4 .5 sin 10" = 0 N- 137.1 + 13 .5~ 0~10 ~-4 .5sin1O0=O 1.e. 4 .5 kN. resolving forces vertically ma, + 250 - 1200 = 0 which gives ma, = 950 kN Then - 3.8g 950 - 950 m 250 /g a, = - Now taking moments about the CG ICG~ - 1200 x 1.0 - 400. 1.223 x 602 x 14 .5 x 1.099 = 350 00N Tailload P=nW-L=4 .5~ 800 0-3 50 00= l000N Drag D = ipV2SCD = i x 1.223 x 602 x 14 .5 x 0.08 75 = 2790N Forward inertia force fW = D (from