Aircraft Structures 1 2011 Part 7 potx

VI VI a 0 a P +- - ._ 1 Starboard all-moving tailplane 2 Tailplane composite construction 3 Tail radome 4 Military equlpment 5 Tail pitch control air valve 6 Yaw control air valves 7 Tail 'bullet' fairing 8 Reaction control system air ducting 9 Trim tab actuator 10 Rudder trim tab 11 Rudder composite construction 12 Rudder 13 Antenna 14 Fin tip aerial falrlng 15 Upper broad band communications antenna 16 Port tailplane 17 Graphite epoxy tailplane skin 18 Port side temperature probe 19 MAD compensator 20 Formation lighting strip 21 Fin construction 22 Fin attachment joint 23 Tailplane pivot seallng plate 24 Aerlals 25 Ventral fin 26 Tail bumper 27 Lower broad band communications antenna 28 Tailplane hydraulic jack 29 Heat exchanger air exhaust 30 Aft fuselage frames 31 Rudder hydraulic actuator 32 Avionics equlpment alr conditionlng plant 33 Avlonlcs equlpment racks 34 Heat exchanger ram air intake 35 Electrical system circuit breaker panels, port and starboard 38 Avionic equipment 37 Chaff and flare dispensers 38 Dlspenser electronic control units 39 Ventral airbrake 40 Alrbrake hydraulic Jack 41 Formation lighting strip 42 Avionics bay access door, port and starboard 43 Avionics equipment racks 44 Fuselage frame and stringer construction 45 Rear fuselage fuel tank 46 Main undercarriage wheel bay 47 Wing root fillet 48 Wlng sparlfuselage attachment joint 49 Water filler cap 50 Engine fire extinguisher bottle 51 Anti-colllsion light 52 Water tank 53 Flap hydraulic actuator 54 Flap hinge fitting 55 Nimonic fuselage heat shleld 56 Main undercarriage bay doors (closed after cycling of mainwheels) 57 Flap vane composite construction 58 Flap composite construction 59 Starboard slotted flap, lowered 60 Outrigger wheel fairing 61 Outrlgger leg doors 62 Starboard aileron 63 Aileron composite construction 64 Fuel jettison 65 Formation lighting panel 86 Roll control airvalve 67 Wing tip fairing 68 Starboard navigation light 69 Radar warnlng aerlal 70 Outboard pylon 71 Pylon attachment joint 72 Graphite epoxy composite wing construction 73 Aileron hydraulic actuator 74 Starboard outrigger wheel 75 BL755 6OO-lb (272-kg) cluster bomb (CBU) 76 Intermediate pylon 77 Reaction control air ducting 78 Alleron control rod 79 Outrlgger hydraulic retraction jack 80 Outrigger leg strut 81 Leg pivot fixing 82 Multi-spar wing construction 83 Leadingedge wlng fence 84 Outrigger pylon 85 Missile launch rail 86 AIM-9L Sldewlnder air-to-air mlsslle 87 External fuel tank, 300 US gal (1 135 I) 88 Inboard pylon 89 Aft retracting twin mainwheels 90 Inboard pylon attachment joint 91 Rear (hot stream) swivelling exhaust nozzle 92 Position of pressure refuelling connection on port side 93 Rear nozzle bearing 94 Centre fuselage flank fuel tank 95 Hydraulic reservoir 96 Nozzle bearing cooling air duct 97 Engine exhaust divider duct 98 Wing panel centre rib 99 Centre section integral fuel tank 100 Port wing Integral fuel tank 101 Flapvane 102 Port slotted flap, lowered 103 Outrigger wheel fairing 104 Port outrigger wheel 105 Torque scissor links 106 Port aileron 107 Aileron hydraullc actuator 108 Aileronlairvalve interconnection 109 Fuel jettison 110 Formation lighting panel 111 Port roll control air valve 112 Port navigation light 113 Radar warning aerial 114 Port wing reaction control air duct 115 Fuel pumps 116 Fuel system plplng 117 Port wlng leading-edge fence 118 Outboard pylon 119 BL755 cluster bombs (maximum load, seven) 120 Intermediate pylon 121 Port outrigger pylon 122 Missile launch rail 123 AlMBL Sidewinder air-to-air misslle 124 Port ieadlng-edge root extension (LERX) 125 Inboard pylon 126 Hydraulic pumps 127 APU intake 128 Gas turbine starterlauxiliary power unit (APU) 129 Alternator coollng air exhaust 130 APU exhaust 131 Engine fuel control unit 132 Engine bay venting ram air Intake 133 Rotary nozzle bearing 134 Nozzle fairing construction 135 Ammunition tank, 100 rounds 136 Cartridge case collector box 137 Ammunition feed chute 138 Fuel vent 139 Gun pack strake 140 Fuselage centrellne pylon 141 Zero scarf forward (fan air) nozzle 142 Ventral gun pack (two) 143 Aden 25-mm cannon 144 Engine drain mast 145 Hydraulic system ground connectors 146 Forward fuselage flank fuel tank 147 Engine electronic control unlla 148 Engine accessory equipment gearbox 149 Gearbox driven alternator 150 Rolls-Royce Pegasus 11 Mk 105 vectored thrust turbofan 151 Formation llghting strips 152 Engine oil tank 159 Bleed air soill duct 154 Air conditioning intake scoops 155 Cockpit air conditioning system heat exchanger 156 Engine compressorlfan face 157 Heat exchanger discharge to intake duct 158 Nose undercarriage hydraulic retraction iack 159 intake blow-in doors 160 Englne bay ventlng alr scoop 161 Cannon muzzle fairing 162 Lift augmentation retractable cross-dam 163 164 165 166 167 168 169 170 171 172 173 174 175 176 177 176 179 180 181 182 183 184 185 186 187 188 189 190 191 192 193 194 195 196 197 1% 199 200 201 202 203 204 205 206 207 208 209 210 21 1 212 213 214 Cross-dam hydraulic jack Nosewheel Nosewheel forks Landlngltaxying lamp Retractable boarding step Nosewheel doors (closed after cycling of undercarrlage) Nosewheel door jack Boundary layer bleed air duct Nose undercarriage wheel bay Kick-in boarding steps Cockplt rear pressure bulkhead Starboard side console panel Martin-Baker Type 12 ejectlon seat Safety harness Election seat headrest Port engine air intake Probe hydraulic jack Retractable in-flight refuelling probe (bolt-on Cockplt canopy cover Miniature detonating cord (MDC) canopy breaker Canopy frame Englne throttle and nozzle angle control levers Pilot's head-up display Instrument panel Moving map display Control column Central warning system Dane1 Cockplt pressure floor Underfloor control runs Formation lighting strips Aileron trim actuator Rudder pedals Cockpit section composite construction Instrument panel shroud One-plece wrap-around windscreen panel Ram air intake (cockpit fresh alr) Front pressure bulkhead Incidence vane Air data computer Pitot tube Lower IFF aerial Nose pitch control air valve Pltch trlm control actuator Electrical system equipment Yaw vane Upper IFF aerial Avionic equipment ARBS heat exchanger MlRLS sensors Hughes Angle Rate Bombing System (ARBS) Composite constructlon nose cone ARBS glazed aperture pack) 228 Principles of stressed skin construction Fig. 7.9 British Aerospace 146 (courtesy of British Aerospace). modern aircraft, coupled with a drop in the structural percentage of the total weight from 30-40 per cent to 22-25 per cent, gives some indication of the improvements in materials and structural design. For purposes of construction, aircraft are divided into a number of sub-assemblies. These are built in specially designed jigs, possibly in different parts of the factory or even different factories, before being forwarded to the final assembly shop. A typical breakdown into sub-assemblies of a medium-sized civil aircraft is shown in Fig. 7.10. Each sub-assembly relies on numerous minor assemblies such as spar webs, ribs, frames, and these, in turn, are supplied with individual components from the detail workshop. Although the wings (and tailsurfaces) of fixed wing aircraft generally consist of spars, ribs, skin and stringers, methods of fabrication and assembly differ. The wing of the aircraft of Fig. 7.7 relies on fabrication techniques that have been employed for many years. In this form of construction the spars comprise thin aluminium alloy webs and flanges, the latter being extruded or machined and are 7.4 Fabrication of structural components 229 Vertical tail -!k Rudder Rear fuselage Centre fuselage Mainplane or centre section Mose Outer wing Wing tip Fig. 7.10 Typical sub-assembly breakdown. bolted or riveted to the web. The ribs are formed in three parts from sheet metal by large presses and rubber dies and have flanges round their edges so that they can be riveted to the skin and spar webs; cut-outs around their edges allow the passage of spanwise stringers. Holes are cut in the ribs at positions of low stress for lightness and to accommodate control runs, fuel and electrical systems. Finally, the skin is riveted to the rib flanges and longitudinal stiffeners. Where the curvature of the skin is large, for example at the leading edge, the aluminium alloy sheets are passed through ‘rolls’ to pre-form them to the correct shape. A further, aerodynamic, requirement is that forward chordwise sections of the wing should be as smooth as possible to delay transition from laminar to turbulent flow. Thus, countersunk rivets are used in these positions as opposed to dome-headed rivets nearer the trailing edge. The wing is attached to the fuselage through reinforced fuselage frames, frequently by bolts. In some aircraft the wing spars are continuous through the fuselage depend- ing on the demands of space. In a high wing aircraft (Fig. 7.7) deep spars passing througn the fuselage would cause obstruction problems. In this case a short third spar provides an additional attachment point. The ideal arrangement is obviously where continuity of the structure is maintained over the entire surface of the wing. In most practical cases this is impossible since cut-outs in the wing surface are required for retracting undercarriages, bomb and gun bays, inspection panels etc. The last are usually located on the under surface of the wing and are fastened to stiffeners and rib flanges by screws, enabling them to resist direct and shear loads. Doors covering undercarriage wells and weapon bays are incapable of resisting wing stresses so that provision must be made for transferring the loads from skin, flanges and shear webs around the cut-out. This may be achieved by inserting strong bulkheads or increasing the spar flange areas, although, no matter the method employed, increased cost and weight result. 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. [...]... P.Therefore, from Eq (8 .12 ), since T = 0, we have LznW 0.20 20 0 .15 1 5 0.IO 1 0 0.05 5 0 0 (b) Fig 8 .10 (a) C ,, cy, CMgG - C curves for Example 8.3; (b) geometry of Example 8.3 , (i> 248 Airworthiness and airframe loads Hence L cL= N ipV2S i 4.5 x 8000 = 1. 113 x 1. 223 x 602 x 14 .5 From Fig 8 .10 (a), a = 13 .75 " and C M , c G = 0. 075 The tail arm I, from Fig 8 .10 (b), is 1 =4 .18 cos(a-2)+0.31sin(a-2) (ii) Substituting... mla=-3g= 13 .5kN g Resolving forces parallel to the axis of the fuselage N -T + mlacos 10 " - 4.5 sin 10 " = 0 1. e N - 13 7 .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 = 12 4.6kN Now resolving forces perpendicular to the axis of the fuselage S - rnlusin 10 " - 4.5~0s = 0 10 " i.e S - 13 .5 sin... non-aerobatic aircraft for which nl = 2.5, w = 2400N/m2 and aCL/acw = 5.0/ rad Taking F = 0. 71 5 we have, from Eq (8.33) n=l+ ! j x 1. 223Vc x 5.0 x 0. 71 5 x 15 .25 2400 + giving n = 1 0. 013 9Vc, where the cruising speed Vc is expressed as an equivalent airspeed For the gust case to be critical 1 + 0. 013 9Vc > 2.5 or Vc > 10 8m/s Thus, for civil aircraft of this type having cruising speeds in excess of 10 8m/s,... 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 2 41 A , \, " Wheel reaction R / Arrester hook Fig 8.5 Forces on the aircraft of Example 8 .1 i.e which gives T = 13 7. 1kN Now resolving forces vertically R - W-TsinlO"=O i.e R = 45 + 13 1.1sin 10 " = 68.8... CL = 1. 099 From Fig 8 .10 (a) CD = 0.0 875 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 = 35000N 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 0.0 875 = 279 0N Forward inertia force fW =D (from Eq (8 .13 )) = 279 0 N In Section 8.4 we determined aircraft loads corresponding... gives 1 = 4 .12 3m In Eq (8 .14 ) the terms La - Db - Mo are equivalent to the aircraft pitching moment MCGabout its centre of gravity Thus, Eq (8 .14 ) may be written M C G - P1 0 or PI = i p v 2 s c c M M , C G (iii) where c = wing mean chord Substituting P from Eq (iii) into Eq (8 .12 ) we have =nW or dividing through by 4pV2S We now obtain a more accurate value for CL from Eq (iv) 1. 35 C = 1. 113 -L x 0. 075 ... forces acting on the aircraft are those shown in Fig 8 .12 The horizontal component of the lift vector in this case provides the force necessary to produce the centripetal acceleration of the aircraft towards the centre of the turn Thus wv2 Lsin4 =- gR (8 . 17 ) and for vertical equilibrium Leos+= w (8 .18 ) L = Wsec4 (8 .19 ) or Fig 8 .12 Correctly banked turn 8.6 Gust loads 2 51 From Eq (8 .19 ) we see that the... = 1. 088 4 .12 3 giving a = 13 .3" and C M , c G = 0. 073 Substituting this value of a into Eq (ii) gives a second approximation for I , namely 1 = 4 .16 1 m Equation (iv) now gives a third approximation for CL,i.e CL = 1. 099 Since the three calculated values of CL are all extremely close further approximations will not give values of CL very much different to those above Therefore, we shall take CL = 1. 099... 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 = 0 I (8 .14 ) Equation (8 .14 ) contains no terms... and CM,CG a light aircraft are shown in Fig 8 .10 (a) The aircraft , , for weight is 8000 N, its wing area 14 .5m2 and its mean chord 1. 35m Determine the lift, drag, tail load and forward inertia force for a symmetricmanoeuvre corresponding to n = 4.5 and a speed of 60 m/s Assume that engine-off conditions apply and that the air density is 1. 223kg/m2 Figure 8 .10 (b) shows the relevant aircraft dimensions . cross-dam 16 3 16 4 16 5 16 6 16 7 16 8 16 9 17 0 17 1 17 2 17 3 17 4 17 5 17 6 17 7 17 6 17 9 18 0 18 1 18 2 18 3 18 4 18 5 18 6 18 7 18 8 18 9 19 0 19 1 19 2 19 3 19 4 19 5 19 6 19 7 1% 19 9 200. lighting panel 11 1 Port roll control air valve 11 2 Port navigation light 11 3 Radar warning aerial 11 4 Port wing reaction control air duct 11 5 Fuel pumps 11 6 Fuel system plplng 11 7 Port. 11 8 Outboard pylon 11 9 BL755 cluster bombs (maximum load, seven) 12 0 Intermediate pylon 12 1 Port outrigger pylon 12 2 Missile launch rail 12 3 AlMBL Sidewinder air-to-air misslle 12 4