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Aluminium and its alloys have already been discussed in Section 6.1. It seems certain that these will remain as the major materials of aircraft con-struction for many years to come, although, in percentage terms, titanium may, and reinforced plastics will, make progressively larger inroads. There is also, at present, a limitation on the use of light alloy on supersonic air-craft because at high speed the skin of the airair-craft can approach tempera-tures which are detrimental to the strength of the material. Concorde was designed as a Mach 2.2 aeroplane because higher speeds would have dic-tated the use of a material other than aluminium alloy. Even then a special alloy was developed particularly for the aircraft with the object of obtain-ing an improvement in the strength and fatigue characteristics over a wide range of temperatures.

In the notes on steel above we said that steel was graded or denomi-nated by its carbon content, and then by the quantity of other elements in its composition. A similar system applies to light alloy, which falls into three groups. The original formulation was patented and marketed under the trade name of Duralumin. For many years the name ‘Dural’ was freely applied to all varieties of light alloy (at least in Europe) in much the same way that the name ‘Hoover’ was applied to all makes of vacuum cleaner.

Dural is a copper-bearing alloy which was reputedly discovered by acci-dent and possessed the rather strange characteristic that if it was heated

(to about 480°C) and suddenly quenched in water it became soft and easily bent or formed, but then the formed shapes, over a period of a few hours, regained their earlier strength and hardness naturally and without any further attention. This remarkable property, which is entirely differ-ent from the work-hardening characteristic of copper and brass (although most aluminium alloys also work harden), has influenced the techniques, and hence the design philosophy, of aircraft manufacture for 80 years. It was later found that a further heat treatment, which was called artificial ageing (at about 170°C), made an additional improvement in the strength.

For comparison, the 0.1% proof stress of naturally aged Dural was about 33 000 lbf/in² (230 N/mm²) and modern derivatives still have much the same strength. The same alloy artificially aged would have a 0.1% proof stress of about 51 000 lbf/in²(350 N/mm²).

The artificially aged copper-bearing alloy was well established by the early 1930s, with the second group of alloys already making an appear-ance. This group differed from the Duralumin group by the addition of nickel and a higher magnesium content. Not widely used for structures, this group was mainly employed in aero engines and as forgings.

The third group, which came into use in the middle 1940s, used zinc and magnesium as alloying elements. The 0.1% proof stress of a typical early member of this group was 74 000 lbf/in²(510 N/mm²), which represented an enormous increase over the previous families of alloys. Unfortunately, possibly because wartime pressures had influenced the development pro-gramme, the desirable properties were accompanied by some highly unde-sirable tendencies. Fatigue cracking (see Section 6.1) accompanied by higher notch sensitivity (Section 9.3) led to some unpleasant failures. A new phenomenon called stress corrosion, initiated by unrelieved internal stresses caused by heat treatment, led to cracking which was very difficult to detect. Many designers considered that the disadvantages outweighed the benefits, but recently improved versions have brought this group of alloys into more general use.

Table 6.2 lists the mechanical properties of some of the widely used alloys. The designation used is nominally international but most familiar in America. The current equivalent British specifications can easily be checked in the brochures published by the materials producers. Some notes on the various methods of numbering specifications are given in Section 6.6.

6.2.4 Titanium

Titanium is effectively the last on the list of metals used in aircraft struc-tures, although magnesium was used in the past and beryllium may be used in the future. One of the most plentiful metals on Earth, it is diffi-cult to refine and therefore expensive. Its weight is between that of steel and light alloy, but it is strong and corrosion resistant even at high tem-peratures. These characteristics have led to its increasing use for spe-cialised pieces of structure, such as fairings near jet engine efflux or intake

Understanding Aircraft Structures Permissible tensile stress

Country of Elongation Brinell hardness

origin Ref. No. 0.2% proof ultimate (UTS) at break (%) No. Remarks

USA 2014–T4 L.164 is sheet material

UK L.164 35 000 56 000 14 105 L.102 is a bar

UK L.102

}

USA 2014–T6 50 000 61 000 8 135 Note the effect of heat treatment on

UK L.165

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stress increases 43%

USA 2024–T4 39 000 59 000 15 120

UK L.109

}

USA QQ–A–250/11

UK L.111 37 000 43 000 8 L.111 is a bar This is weldable alloy.

UK L.113 L.113 is a sheet

}

UK L.114 L.114 is a tube welding

USA 7075–T6 62 000 71 000 8 150 Aluminium–zinc alloy

UK L.88

}

(1) The properties are approximated to illustrate the grouping. Figures must not be used for design and reference must be made to the specifications.

(2) Stress is quoted in lbf/in.². To convert to N/mm²divide the figure given by 145.0 (see also note 5).

(3) Young’s Modulus for light alloy is 9.5–10.7 ¥ 10⁶lbf/in.². (4) Poisson’s ratio for light alloy is 0.3.

(5) Stress may be quoted in units called ‘kips’. 1 kilo pound = 1000 lb.

(6) The suffixes –T4 and –T6 signify that the material can receive two different heat treatments. The effect is that UTS increases 9% and proof stress increases 43%. L.164 and L.165 are the same basic alloy but in the UK system the different heat treatments give the two specifications separate reference numbers.

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edges where stainless steel may have been used previously. On the Boeing 727 designed in the early 1960s, less than 2% of the structure weight was titanium, but by the late 1960s the Boeing 747 structure was almost 10%

titanium by weight and this trend was almost certainly exceeded by mili-tary aircraft. In recent years the move towards composite structures has rather reversed the trend towards titanium. Bearing in mind the high initial tooling cost of composites and the comparatively low repair cost of metal structures, the pendulum may yet swing back towards titanium.

Returning to the Boeing 747, it is interesting to speculate on how much the designers were driven, against their will, to use titanium because weight in the final design exceeded that of the project calculations. If they were forced to use more of the expensive titanium than they would have wished, then the price of the aircraft must have been higher than expected;

but in the event it is a magnificent money-making machine and supports the argument that the cost of raw material should not necessarily be a deterrent to its use. (See also Section 6.4.)

As with other metals, titanium is used as an alloy, in fact as two fami-lies of alloys, one with aluminium and one with manganese. These give a good range of available properties (see Table 6.3) which should be com-pared with those in Tables 6.1 and 6.2.