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Most common aircraft structures are typified by arrangements of thin, load bearing skins, frames and stiffeners, fabricated from lightweight, high strength materials of which aluminium alloys are the most widely used examples.
As a preliminary to the analysis of the basic aircraft structural forms presented in subsequent chapters we shall discuss general principles of stressed skin construction from the viewpoint of materials and the loading, function and fabrication of structural components.
1.1 Aircraft Materials
Several factors influence the selection of the structural material for an aircraft, but amongst this strength allied to lightness is probably the most important. Other properties having varying, though sometimes critical significance are stiffness, toughness, resistance to corrosion, fatigue and the effects of environmental heating, ease of fabrication, availability and consistency of supply and, not least important, cost.
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The main groups of materials used in aircraft construction have been wood, steel, aluminium alloys with, more recently, titanium alloys, and fibre-reinforced composites. In the field of engine design, titanium alloys are used in the early stages of a compressor while nickel-based alloys or steels are used for the hotter later stages. As we are concerned primarily with the materials involved in the construction of the airframe, discussion of materials used in engine manufacture falls outside the scope of this book. Before we consider the individual groups in detail it is interesting and instructive to examine briefly the history of aircraft materials from the birth of the industry, at the beginning of the 0th century, to the present day.
The first generation of conventional powered aircraft was constructed of wood and canvas. Spruce and birch were the most widely used timbers with tensile strengths of 70 N/mm and 100 N/mm respectively, specific gravities of 0.4 and 0.6 and Young’s moduli of 000 N/mm. Although these strength/weight ratios compare favourably with modern heat-treated aluminium alloys, natural wood had disadvantages. Changes in shape and dimensions resulted from moisture absorption and loss caused by changes in atmospheric humidity, while its structural properties exhibited the inconsistency common to natural products. Further, pronounced anisotropy caused by its grain structure gave a variation in the value of Young’s modulus in the ratio of 1501 depending on the direction of loading in relation to the grain. Associated effects were ratios of shear modulus and Poisson’s ratio of the order of 01 and 401 respectively.
The introduction of plywood and the development of synthetic resin adhesives brought improvements in the strength of spars and skins and enabled anisotropy to be eliminated or at worst controlled. However, the large amount of wood required for military aircraft construction, during the 114-18 war, revealed one of its most serious limitations. The most suitable forms were imported from overseas, requiring a large volume of shipping which was otherwise needed for the transport of food and troops. To avoid a similar critical situation arising at a future date the Air Ministry, in 14, prohibited the use of wood for the main load carrying parts of the structure. This decision obviously hastened the introduction of alternative metallic materials in airframe construction, although wood continued to make a significant contribution for many years. In fact, during the 1-45 war a particularly successful high performance aircraft, the de Havilland Mosquito, was built entirely of wood. It must be admitted, however, that the special circumstances of the time were the cause of this. There was a shortage of factories and skilled workers for metal fabrication, whereas the furniture industry was able to supply manpower and equipment. Moreover, wood could be adapted to rapid methods of construction and designers had acquired a substantial amount of experience in dealing with the problem of anisotropy in bulk timber. Furthermore, improvements in adhesives, for example the introduction of the Redux adhesives based on phenolformaldehyd thermosetting resin and the polyvinyl formal thermoplastic resin as a composite adhesive system, led to improved wood-wood, acceptable wood-metal and even metal-metal bonds.
Despite this relatively modern successful use of wood it became inevitable that its role as an important structural material should come to an end. The increased wing loadings and complex structural forms of present day turbojet aircraft cause high stress concentrations for which wood is not well adapted. Its anisotropy presents difficult problems for the designer while wooden aircraft require more maintenance than those constructed of metal. It is particularly unsuitable for use in tropical conditions where, as we have noted, large changes in humidity have serious effects on shape and dimensions. Attacks on the wood by termites is an additional problem.
The first practical all-metal aircraft was constructed in 115 by Junkers in Germany, of materials said to be ‘iron and steel’. Steel presented the advantages of a high modulus of elasticity, high proof stress and high tensile strength. Unfortunately these were accompanied by a high specific gravity, almost three times that of the aluminium alloys and about ten times that of plywood. Designers during the 10s were therefore forced to use steel in its thinnest forms, the usual preference being for a steel having a 0.1 per cent proof stress of 1000 N/mm. To ensure stability against buckling of the thin plate, intricate shapes for spar sections were devised; typical examples of these are shown in Fig. 7.1. Common gauges of the material were to 16 SWG (i.e. approximately 0.5 mm to 1.16 mm), with a composition of 0.5 per cent carbon, 1.5 per cent manganese steel to Specification DTD 17, a nickel chrome steel to Specification DTD 54A or a 1 per cent chromium steel to DTD 46A.
In 10 Alfred Wilm, in Germany, accidentally discovered that an aluminium alloy containing .5 per cent copper, 0.5 per cent magnesium and silicon and iron as unintended impurities spontaneously hardened after quenching from about 480o C. The patent rights of this material were acquired by Durener Betallwerke who marketed the alloy under the name Duralumin. For half a century this alloy has been used in the wrought heat-treated, naturally aged condition possessing mechanical properties of 0.1 per cent proof stress not less than 0 N/mm, tensile strength not less than 0 N/mm and an elongation at fracture not less than 15 per cent. However, the improvements in these properties produced by artificial ageing at a raised temperature of, for example 175 o C were not exploited in the aircraft industry until about 14. Artificially aged duralumin has a 0.1 per cent proof stress of not less than 70 N/mm, a tensile strength not less than 460 N/mm and an elongation of 8 per cent.
In addition to the development of duralumin (first used as a main structural material by Junkers in 117) three other causes contributed to the replacement of steel by aluminium alloy. These were a better understanding of the process of heat treatment, the introduction of extrusions in a wide range of sections and the use of pure aluminium cladding to provide greater resistance to corrosion. By 18, three groups of aluminium alloys dominated the field of aircraft construction and, in fact, they retain their importance to the present day. The groups are separated by virtue of their chemical composition, to which they owe their capacity for strengthening under heat treatment.
The first group is contained under the general name duralumin having a typical composition of 4 per cent copper, 0.5 per cent magnesium, 0.5 per cent manganese, 0. per cent silicon, 0. per cent iron, with the remainder aluminium. The naturally aged version was covered by Air Ministry Specification DTD used in 14, while artificially aged duralumin came under Specification DTD 111 in 1. Typical properties of the two types have been quoted above although DTD 111 provided for slight reductions in 0.1 per cent proof stress and tensile strength.
The second group of aluminium alloys differs from duralumin chiefly by the introduction of 1 to per cent of nickel, a high content of magnesium and possible variations in the amounts of copper, silicon and iron. ‘Y’ alloy, the oldest member of the group, has a typical composition of 4 per cent copper, per cent nickel, 1.5 per cent magnesium, the remainder being aluminium and was covered by Specification DTA 58A issued in 17. Its most important property was its retention of strength at high temperatures, which meant that it was a particularly suitable material for aero engine pistons. Its use in airframe construction has been of a limited nature only. Research by Rolls-Royce and development by High Duty Alloys Ltd produced the ‘RR’ series of alloys. Based on Y alloy, the RR alloys had some of the nickel replaced by iron and the copper reduced. One of the earliest of the alloys RR 56, had approximately half of the per cent nickel replaced by iron, the copper content reduced from 4 to per cent, and was used for forgings and extrusions in aero engines and airframes. Specification DTD 10, issued in 10, listed minimum mechanical properties for RR56 of 0.1 per cent proof stress 10/Nmm, tensile strength of 00 N/mm and elongation of 10 per cent.
The third and latest group depends upon the inclusion of zinc and magnesium for their high strength. Covered by Specification DTD 6 issued in 17, these alloys had a nominal composition .5 per cent copper, 5 per cent zinc, per cent magnesium and up to 1 per cent nickel with mechanical properties, 0.1 per cent proof stress 510 N/mm, tensile strength 585 N/mm and an elongation of 8 per cent. In modern versions of this alloy nickel has been eliminated and provision made for the addition of chromium and further amounts of manganese.
Of the three basic structural materials described above, namely wood, steel and aluminium alloy, only wood is no longer of significance except in the form of laminates for non-structural bulkheads, floorings and furnishings. Most modern aircraft, for example Concorde, still rely on modified forms of the high strength aluminium alloys which were introduced during the early part of the 0th century. Steels are used where high strength, high stiffness and wear resistance are required. Other materials, such as titanium and fibre-reinforced composites first used about 150, are finding expanding uses in airframe construction. All these and some additional materials are now discussed in detail.
Aluminium Alloy Structures
The aerospace industry extensively uses three groups of aluminium alloys (i) nickel free duralumins, (ii) derivatives of Y alloy and (iii) aluminium-zinc-magnesium. Alloys from each group have been used extensively for airframes, skins and other stressed components, the choice of alloy being influenced by factors such as strength (proof and ultimate stress), ductility, ease of manufacture (e.g. in extrusion and forging), resistance to corrosion and amenability to protective treatment, fatigue strength, freedom from liability to sudden cracking due to internal stresses and resistance to fast crack propagation under load. Clearly, different types of aircraft have differing requirements. A military aircraft, for instance, having a relatively short life measured in hundreds of hours, does not call for the same degree of fatigue and corrosion resistance as a civil aircraft with a required life of 0,000 or more.
Unfortunately, as one particular property of aluminium alloys is improved, other desirable properties are sacrificed. For example, the extremely high static strength of the aluminium-zinc-magnesium alloys was accompanied for many years by a sudden liability to crack in an unloaded condition due to the retention of internal stresses in bars, forgings and sheet after heat treatment. Although variations in composition have eliminated this problem to a considerable extent other deficiencies showed themselves. Early Viscount service experience produced large numbers of stress-corrosion failures of forgings and extrusions. The problem became so serious that in 15 it was decided to replace as many aluminium-zinc-manganese components as possible with the aluminium-4 per cent copper Alloy L65 and to prohibit the use of forgings in zinc-bearing alloy in all future designs. However, improvements in the stress-corrosion resistance of the aluminium-zinc-manganese alloys have resulted in recent years from British, American and German research. Both British and American opinions agree on the benefits of including about 1 per cent copper but disagree on the inclusion of chromium and manganese, while in Germany the addition of silver has been found extremely beneficial. Improved control of casting techniques has brought further improvements in resistance to stress corrosion. The development of aluminium-zinc-manganese-copper alloys, called the 7000 series, has largely met the requirement for aluminium alloys possessing high strength, good fatigue crack growth resistance and adequate toughness. Further development will concentrate on the production of materials possessing higher specific properties, bringing benefits in relation to weight saving rather than increasing strength and stiffness.
The duralumin alloys possess a lower static strength than the above zinc-bearing alloys, but are preferred for portions of the structure where fatigue considerations are of primary importance such as the under-surfaces of wings where tensile fatigue loads predominate. Experience has shown that the naturally aged version of duralumin has important advantages over the fully heat-treated forms in fatigue endurance and resistance to crack-propagation. Furthermore, the inclusion of a higher percentage of magnesium was found, in America, to produce, in the naturally aged condition, mechanical properties between those of the normal naturally aged and artificially aged duralumin. This alloy, designated 04 (aluminium-copper alloys for the 000 series) has the nominal composition 4.5 per cent copper, 1.5 per cent magnesium, 0.6 per cent manganese, with the remainder aluminium, and appears to be a satisfactory compromise between the various important, but sometimes conflicting, mechanical properties.
Aluminium-magnesium-silicon alloys are that they are potentially cheaper than aluminium-copper alloys and, being weldable, are capable of reducing manufacturing costs and so have attracted renewed interest. Variants, i.e. ISO 601 alloy, also have improved property levels and, generally, possess a similar high fracture toughness and resistance to crack propagation as the 000 series alloys.
Frequently, a particular form of an alloy is developed for a particular aircraft. An outstanding example of such a development is the use of Hiduminium RR58 as the basis for the main structural material, designed CM001, for Concorde. Hiduminium RR58 is a complex aluminium-copper-magnesium-nickel-iron alloy developed during the 1-45 war specifically for the manufacture of forged components in gas turbine aero engines. The chemical composition of the version used in Concorde was decided on the basis of elevated temperature, creep, fatigue and tensile testing programmes and has the detailed specification of
%Cu %Mg %Si %Fe %Ni %Ti %A1
Minimum .5 1.5 0.18 0.0 1.0 -
Remainder
Maximum .70 1.65 0.5 1.0 1.0 0.0
Generally, CM001 is found to possess better overall strength/fatigue characteristics over a wide range of temperatures than any of the other possible aluminium alloys.
The latest aluminium alloys to find general use in the aerospace industry are the aluminium-lithium alloys. Of these, the aluminium-lithium-copper-manganese alloy, 800, developed in the UK, is extensively used in the main fuselage structure of GKN Westland Helicopters’ most recent design EH101; it has also been qualified for Eurofighter 000 (now named the Typhoon) but has yet to be embodied. In the USA the aluminium-lithium-copper-manganese alloy, 05 has been used in the fuselage frames of the F16 as a replacement for 14, resulting in a fivefold increase in fatigue life and a reduction in weight. Aluminium-lithium alloys can be successfully welded, possess a high fracture toughness and exhibit a high resistance to crack propagation.
Steels
The use of steel for the manufacture of thin-walled, box-section spars in the 10s has been described previously in this section. Clearly, its high specific gravity prevented its widespread use in aircraft construction, but it has retained some value as a material for castings for small components demanding high tensile strengths, high stiffness and high resistance to wear. Such components include undercarriage pivot brackets, wing-root attachments, fasteners and tracks.
Although the attainment of high and ultra-high tensile strengths presents no difficulty with steel, it is found that other properties are sacrificed and that it is difficult to manufacture into finished components. To overcome some of these difficulties types of steel known as maraging steels were developed in 161, from which carbon is either eliminated entirely or present only in very small amounts. Carbon, while producing the necessary hardening of conventional high tensile steels, causes brittleness and distortion; the latter is not easily rectifiable as machining is difficult and cold forming impracticable. Welded fabrication is also almost impossible or very expensive. The hardening of maraging steels is achieved by the addition of other elements such as nickel, cobalt and molybdenum. A typical maraging steel would have these elements present in the proportions nickel 17-1 per cent, cobalt 8- per cent, molybdenum -.5 per cent, with titanium 0.15-0.5 per cent. The carbon content would be a maximum of 0.0 per cent, with traces of manganese, silicon, sulphur, phosphorus, aluminium, boron, calcium and zirconium. Its 0. per cent proof stress would be nominally 1400 N/mm and its modulus of elasticity 1800 000 N/mm.
The main advantages of maraging steels over conventional low alloy steels are higher fracture toughness and notched strength, simpler heat treatment, much lower volume change and distortion during hardening, very much simpler to weld, easier to machine and better resistance to stress corrosion/hydrogen embrittlement.
On the other hand, the material cost of maraging steels is three or more times greater than the cost of conventional steels, although this may be more than offset by the increased cost of fabricating a complex component from the latter steel.
Maraging steels have been used in aircraft arrester hooks, rocket motor cases, helicopter undercarriages, gears, ejector seats and various structural forgings.
In addition to the above, steel in its stainless form has found applications primarily in the construction of super- and hypersonic experimental and research aircraft, where temperature effects are considerable. Stainless steel formed the primary structural material in the Bristol 188, built to investigate kinetic heating effects, and also in the American rocket aircraft, the X-15, capable of speeds of the order of Mach 5-6.
Titanium
The use of titanium alloys increased significantly in the 180s, particularly in the construction of combat aircraft as opposed to transport aircraft. This increase has continued in the 10s to the stage where, for combat aircraft, the percentage of titanium alloy as a fraction of structural weight is of the same order as that of aluminium alloy. Titanium alloys possess high specific properties, have a good fatigue strength/tensile strength ratio with a distinct fatigue limit, and some retain considerable strength at temperatures up to 400o-500 o C. Generally, there is also a good resistance to corrosion and corrosion fatigue although properties are adversely affected by exposure to temperature and stress in a salt environment. The latter poses particular problems in the engines of carrier-operated aircraft. Further disadvantages are a relatively high density so that weight penalties are imposed if the alloy is extensively used, coupled with high primary and high fabrication costs, approximately seven times those of aluminium and steel.
In spite of this, titanium alloys were used in the airframe and engines of Concorde, while the Tornado wing carry-through box is fabricated from a weldable medium strength titanium alloy. Titanium alloys are also used extensively in the F15 and F American fighter aircraft and are incorporated in the tail assembly of the Boeing 777 civil airliner. Other uses include forged components such as flap and slat tracks and undercarriage parts.
New fabrication processes (e.g. superplastic forming combined with diffusion bonding) enable large and complex components to be produced, resulting in a reduction in production man-hours and weight. Typical savings are 0 per cent in man-hours, 0 per cent in weight and 50 per cent in cost compared with conventional riveted titanium structures. It is predicted that the number of titanium components fabricated in this way for aircraft will increase significantly and include items such as access doors, sheet for areas of hot gas impingement etc.
Plastics
Plain plastic materials are heavier than wood although of comparable strength. However, their specific gravities are less than half those of the aluminium alloys so that they find uses as windows or lightly stressed parts whose dimensions are established by handling requirements rather than strength. They are also particularly useful as electrical insulators.
Glass
The majority of modern aircraft have cabins pressurized for flight at high altitudes. Windscreens and windows are therefore subjected to loads normal to their midplanes. Glass is frequently the material employed for this purpose in the form of plain or laminated plate or heat-strengthened plate. The types of plate glass used in aircraft have a modulus of elasticity between 70 000 N/mm with a modulus of rupture in bending of 45 N/mm. Heat strengthened plate has a modulus of rupture of about four and a half times this figure.
Fibre Composite Materials/Structures
Composite materials consist of strong fibres such as boron, glass or carbon set in an epoxy or a thermoplastic resin, which is mechanically and chemically protective. The fibres may be continuous or discontinuous but possess a strength very much greater than that of the same bulk materials. For example, carbon fibres have a tensile strength of the order of 400 N/mm and a modulus of elasticity of 400 000 N/mm.
A sheet of fibre-reinforced material is anisotropic, that is, its properties depend on the direction of the fibres. Generally, therefore, in structural form two or more sheets are sandwiched together to form a lay-up so that the fibre directions match those of the major loads.
In the early stages of the development of composite materials glass fibres were used in a matrix of epoxy resin. This glass reinforced plastic (GRP) was used for radomes and helicopter blades but found limited use in components of fixed wing aircraft due to its low stiffness. In the 160s, new fibrous reinforcements were introduced; Kevlar, for example, is an aramid material with the same strength as glass but is stiffer. Kevlar composites are tough but poor in compression and difficult to machine, so they were used in secondary structures. Another composite, using boron fibre and developed in the USA, was the first to possess sufficient strength and stiffness for primary structures.
These composites have now been replaced by carbon fibre reinforced plastics (CFRP), which have similar properties to boron composites but are very much cheaper. Typically, CFRP has a modulus of the order of three times that of GRP, one and a half times that of a Kevlar composite and twice that of aluminium alloy. Its strength is three times that of aluminium alloy, approximately the same as that of GRP, and slightly less than that of Kevlar composites. CFRP does, however, suffer from some disadvantages. It is a brittle material and therefore does not yield plastically in regions of high stress concentration. Its strength is reduced by impact damage which may not be visible and the epoxy resin matrices can absorb moisture over a long period which reduces its matrix dependent properties, such as its compressive strength; this effect increases with increase of temperature. Further, the properties of CFRP are subject to more random variation than those of metals. All these factors must be allowed for in design. On the other hand, the stiffness of CFRP is much less affected than its strength by the above and it is less prone to fatigue damage than metals. It is estimated that replacing 40% of an aluminium alloy structure by CFRP would result in a 1% saving in total structural weight.
Fig 1. Crossection of a helicopter rotor blade
CFRP is included in the wing, tailplane and forward fuselage of the latest Harrier development, is used in the Tornado taileron and has been used to construct a complete Jaguar wing and engine bay door for testing purposes. The use of CFRP in the fabrication of helicopter blades has led to significant increases in their service life, where fatigue resistance rather than stiffness is of primary importance. Figure 7. shows the structural complexity of a Sea King helicopter rotor blade which incorporates CFRP, GRP, stainless steel, a honeycomb core and foam filling. An additional advantage of the use of composites for helicopter rotor blades is that the moulding techniques employed allow variations of cross-section along the span, resulting in substantial aerodynamic benefits. This approach is being employed in the fabrication of the main rotor blades of the GKN Westland Helicopter EH 101.
A composite (fibreglass and aluminium) is used in the tail assembly of the Boeing 777 while the leading edge of the Airbus A10-00/A0 fin assembly is of conventional reinforced glassfibre construction, reinforced at the nose to withstand bird strikes. A complete composite airframe was produced for the Beechcraft Starship turboprop executive aircraft which, however, was not a commercial success due to its canard configuration causing drag and weight penalties.
The development of composite materials is continuing with research into the removal of strength-reducing flaws and local imperfections from carbon fibres. Other matrices such as polyetheretherketone, which absorbs much less moisture than epoxy resin, has an indefinite shelf life and performs well under impact, are being developed; fabrication, however, requires much higher temperatures. Metal matrix composites such as graphite-aluminium and boron-aluminium are light-weight and retain their strength at higher temperatures than aluminium alloys, but are expensive to produce.
Generally, the use of composites in aircraft construction appears to have reached a plateau, particularly in civil subsonic aircraft where the fraction of the structure comprising composites is approximately 15%. This is due largely to the greater cost of manufacturing composites compared with aluminium alloy structures since composites require hand crafting of the materials and manual construction processes. These increased costs are particularly important in civil aircraft construction and are becoming increasingly important in military aircraft.
It is possible to lower costs by
• Innovative design concepts, which consider producibility
• Lowering the part counts
• The use of fewer fasteners
• The use of automation methods to cutdown manufacturing costs
The use of composites presents the engineer with a range of new freedoms
᠋ Ability to engineer the material as well as the structure
᠋ Superior structural properties
᠋ Ability to assemble and form in a soft condition and bake to harden.
With it comes a range of new problems
᠋ New stress methods (anisotropic or orthotropic) needed
᠋ Computer analysis programs requires
᠋ Designing around material weaknesses
᠋ No reservoir of service experience
Advantages of Composites over metals
We can summarise the advantages and disadvantages of composites as follows
᠋ Light weight
᠋ Resistance to corrosion
᠋ High resistance to fatigue damage
᠋ Reduced machining
᠋ Tapered sections and compound contours easily accomplished
᠋ Can orientate fibers in direction of strength/stiffness needed
᠋ Reduced number of assemblies and reduced fastener count when cocure or consolidation is used
᠋ Absorb radar microwaves (stealth capability)
᠋ Thermal expansion close to zero reduces thermal problems in outer space applications.
Disadvantages of Composites over Metals
᠋ Material is expensive
᠋ Lack of established design allowables
᠋ Corrosion problems can result from improper coupling with metals, especially when carbon or graphite is used (sealing is essential)
᠋ Degradation of structural properties under temperature extremes and wet conditions
᠋ Poor energy absorption and impact damage.
᠋ May require lightning strike protection
᠋ Expensive and complicated inspection methods
᠋ Reliable detection of substandard bonds is difficult
᠋ Defects can be known to exist but precise location cannot be determined.
HISTORY OF SERVICE PROBLEMS IN U.S. AIR FORCE AIRCRAFT AND ITS INFLUENCE ON DESIGN AND MATERIAL SELECTION
Corrosion and cracking related problems are the main drivers in aircraft design and maintenance. The distribution and magnitude of service cracking problems in Air Force aircraft is shown below. There are a total of 1,4 major and minor cracking problems recorded on 1 types of military aircraft.
During a 1 month period, in one study, 16 major cracking/failure incidents were reported. The majority of these were fatigue initiated with corrosion fatigue second followed by stress corrosion. In another study, out of 64 major cracking incidents reported, the majority were stress corrosion followed by corrosion fatigue and fatigue about equal. It is noted on both of these charts on that some failures were attributed to overload. This is rare in commercial transport history.
The distribution of cracking and failure origins is shown below. The majority of failures were due to poor quality where cracks initiated at holes. Material flaws, defects and scratches were second in magnitude with poor design details following this in magnitude.
This magnitude of cracking incidents also contributed to an Air Force decision to change the design philosophy of their structures to that of damage tolerance. Prior to this time, the main philosophy had been a safe life approach where the design was based on a full-scale fatigue test to 4 lifetimes.
Civil Aircraft industry also moved to damage tolerant design philosophy. This was also influenced by early Comet failures, see below. In this instance the investigation of the Comet failure on Yoke Peter on the structure recovered near Elba confirmed that the primary cause of failure was pressure cabin failure due to fatigue. The origin in this case was at the corner of the Automatic Direction Finding (ADF) windows on the top centre line of the cabin.
Yoke Uncle was repaired and the fuselage skin was strain gauged near the window corners. The peak stresses measured were 4,100 PSI for 8.5 PSI cabin pressure plus 650 PSI for 1g flight and 150 PSI for a 10 ft/sec gust making a total of 45.700 PSI. The material was DTD 546 having an ultimate strength of 65,000 PSI. Therefore, the 1P + 1g stress was 70% of the Ultimate Material Strength.
Thus, the cause of the failures was determined to be fatigue due to high stresses at the corner of windows in the pressure cabin. This investigation resulted in considerable attention to detail design in all future pressure cabin designs and resulted in full scale fatigue tests becoming recognised as being required.
Corrosion
Corrosion is also one of the main drivers in aerospace design and maintenance considerations.
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