Aircraft structures for engineering students - part 5 pps

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.