Aircraft have changed enormously over the last century from the early Wright Flyer flown at Kittyhawk to the supersonic SR-71 Blackbird flown today. Of course the developments in aeronautical engineering can be broken down into separate divisions that have developed at different rates: a) the aerodynamics, b) power plant engineering, c) control, radios and navigation aids, d) airframe engineering (e.g. hydraulic/electrical systems, interior fittings etc.), and finally e) the structural design. For example, power plants have developed in two large steps separated by a series of sudden burst of ingenuity. In order to facilitate the first successful flight the Wright Brothers had to find a light yet powerful engine system. The next stride was the ingenious invention of the jet engine prior and during WWII by Sir Frank Whittle and Hans von Ohain. In between, the power output of piston engines “increased almost 200 times from 12 bhp to over 2000 bhp in just 40 years, with only a ten times increase in mass (3) “. As will be outlined in this article, the design of aerospace structures on the other hand has only made one fundamental stride forward, but this change was sufficient to change the complete design principle of modern aircraft. Today however, the strict environmental legislation and advent of the composite era may induce further leaps in structural design.
Fig. 1. A schematic drawing of the Wright Flyer (1)
Fig. 2. The modern supersonic SR-71 Blackbird (2)
1) Wire Braced Structures
If we look at the early design of aircraft such as the Wright Flyer in Figure 1 there can really be no misunderstanding of the construction style. The entire aircraft, including most notably the wings, forward and rear structures were all constructed from rectangular frames that were prevented from shearing (forming a parallelogram) or collapsing by diagonally stretched wire. There were two major innovative thoughts behind this design philosophy. Firstly, the idea that two parallel wings would facilitate a lighter yet stronger structure than a single wing, and secondly, that these two wings could be supported with two light wires rather than with a single, thicker wooden member. The structural advantage of the biplane construction is that the two wings, vertical struts and wires form a deep light beam, which is more resistant to bending and twisting than a single wing. Much like a composite sandwich beam it can be treated as two stiff outer skins for high bending rigidity connected by a lightweight “core” to provide resistance to shear and torsion.
Fig. 3. Cutaway drawing of the 1917 Sopwith Camel (3)
Fig. 4. Cutaway drawing of the 1935 Hawker Hurricane (3)
The biplane construction with wire bracing was the most notable feature of aircraft construction for much of the following years and paired nicely with lightweight materials such as bamboo and spruce (Figure 3). Wood is a composite of cellulose fibres embedded in a matrix of lignin and the early aeronautical engineers knew to take advantage of its high specific strength and stiffness. Strangely enough, after the era of metals we are now returning back to the composite roots of aircraft, albeit in a more advanced fashion. The biplane era lasted until the 1930s at which point metal was taking over as the prime aerospace material. Initially the design philosophy was not adapted to take full advantage of thin sheet metal manufacturing techniques such that wooden spars and struts were just replaced by thinner metal tubing. Consequently there remained a striking similarity in construction between a 1917 (Figure 3) and a 1931 (Figure 4) fighter. Even though some thin metal sheets were being used these components generally did not carry much load such that the main fuselage structure featured 4 horizontal longerons supported by vertical struts and wire bracing. This so called “Warren Girder” design can also be seen in some of earliest monoplane wing constructions such as the 1935 Hawker Hurricane. Aeronautical engineers were initially “unsure how to combine the new metal construction with a traditional fabric covering (3)” used on earlier aircraft. The onset of WWII meant that some safe and conservative design decisions were made to facilitate monoplane wings and the “Warren Girder” principle was directly copied to the internal framework of monoplane wings (Figure 5). These early designs were far from optimised and perfectly characterise the transition period between wire-frame structures and the semi-monocoque structures we use today.
Fig. 5. The Hawker Hurricane wing construction (3).
2) Semi-Monocoque Structures
The internal cross-bracing was initially acceptable for the early single or double seater aircraft, but would obviously not provide enough room for larger passenger aircrafts. To overcome this, inspiration was taken from the long tradition and expertise in boat building which had already been applied to construct the fuselages of early wooden flying boats. The highest standards of yacht construction at the time featured “bent wooden frames and double or triple skins…with a clear varnished finish…and presented a much more open and usable fuselage interior (3)”. The well-established boat building techniques were thus passed on to aircraft construction to produce newer aircraft with very smooth, aerodynamic profiles.
Fig. 6. Semi monocoque fuselage construction of an early wooden flying boat (4)
The major advantage of this type of construction is that the outer skin of the fuselage and wing no longer just define the shape and aerodynamic profile of the aircraft, but become an active load-carrying member of the structure as well. Thus, the structure becomes “multifunctional” and more efficient, unlike the braced fuselage which would be just as strong without the fabric covering the girders. As a consequence the whole structure is generally at a uniform and lower stress level, reducing stress concentrations and giving better fatigue life. Finally, as the majority of the material is located at the outer surface of the structure the second and polar moments of area, and therefore the bending and torsional rigidities are much increased. On the other hand, the thin-skinned construction means that compression and shear buckling become the most likely forms of failure. In order to increase the critical buckling loads the skins are stiffened by stringers and broken up into smaller sections by spars and ribs.
Fig. 7. Components of a semi monocoque wing (5)
Because the external skin is now a working part of the structure this type of construction became to be known as stressed skin or semi-monocoque, where monocoque means “shell in one piece” and “semi” is an english addition to describe the discrete discontinuities of internal stiffeners. The adoption of the semi-monocoque construction and a change from wood to metal naturally coincided since sheet metal production allowed a variety of thin skins to be easily manufactured quite cheaply, with better surface finish and superior material properties. Furthermore, metal construction was conducive to riveting which would overcome the adhesive problems of early wooden semi-monocoque aircraft such as the deHavilland Mosquito.
Fig. 8. Cutaway Drawing of the recently released A400M aircraft (6).
Figure 8 shows the typical construction of a modern aircraft. There have been numerous different structural arrangements over the past number of years but all generally feature some sort of vertical stiffener (ribs in the wings and rings in the fuselage) and longitudinal stiffener (called stringers). Over the years the main driver has been towards a) a reduction in the number of rivets by reverting to bonded assembly or ideally manufacturing separate components as a single piece and b) understanding the effects and growth of cracks under static and fatigue loading by building structures that can easily be inspected or have multiple redundancies (load paths). The design and manufacturing methods of semi-monocoque aircraft are now so automated that the development of a new aluminium, medium sized airliner “could be regarded as a routine exercise (1)”. However, the continuing legislative pressure to reduce weight and fuel consumption provides enough incentive for further development.
3) Sandwich Structures and Composite Materials
One of the major disadvantages of thin-skinned structures is their lack of rigidity under compressive loading which gives them a tendency to buckle. A sheet of paper nicely illustrates this point, since it is quite strong in tension but will provide no support under compression. One way of improving the rigidity of thin panels is by increasing the bending stiffness with the aid of external stiffeners, which at the same time break the structure up into smaller sections. The critical buckling load is a function of the square of the width of the plate over which the load is applied. Therefore skins can be made 4 times stronger in buckling by just cutting the width in half. As a wing bends upwards the main compressive loads act on the top skin along the length of the wing and therefore a large number of stringers are visible across the width.
Fig. 8. Buckling analysis of a stiffened wing panel. The stiffeners break the buckling mode shapes into smaller wavelengths that require higher energy to form compared to a single wave (7)
Another technique to provide more rigidity is sandwich construction. This generally features a very lightweight core, such as a honeycomb lattice or a foam, sandwiched between two thin yet stiff outer panels. Here the role of the sandwich core is to carry any shear loads and separate the two skins as far as possible. The second moment of area is a function of the cube of the depth and therefore the bending rigidity is greatly increased with this technique. Ideally, in this manner it would be possible to design an entire fuselage without any internal rings or stringers and the Beech Starship is an excellent example of a successful application. However, there are problems of forming honeycomb cores onto doubly curved shells since the material is susceptible to strong anticlastic curvature, forming a saddle shape when bent in one direction. Furthermore, there are problems with condensation and water ingress into the honeycomb cells and the ability to guarantee a good bond surface between the core and the outer skins. There is the possibility to use foam cores instead, but these tend to be heavier with lower mechanical properties. Perhaps the current trend is away from sandwich construction (10).
Fig. 9. A carbon fibre composite/honeycomb sandwich panel (9)
Fig. 10. The Beech Starship whose fuselage was design using sandwich construction with minimal internal bulkheads and ribs (8)
One of the major applications of honeycomb structures has been in combination with composite materials. Stiff carbon composite panels are the ideal candidate for the outer skins and the whole assembly can be co-cured together in an autoclave without having to perform any secondary bonding operations. Furthermore, the incredible specific strength and stiffness of carbon composites makes this combination an ultra lightweight yet resilient structure for aerospace applications. Indeed, we are now at the start of the “black” carbon age in commercial aircraft design. Apart from their excellent specific strength and stiffness properties composites exhibit the ability to tailor optimum mechanical properties by orientating the majority of plies in the direction of the load and allowing for less material waste during manufacture. As a result, the first generation of commercial aircraft that contain large proportions of composite parts, such as the Boeing 787 Dreamliner and Airbus A350 XWB, are planned to enter service throughout the next years.
Fig. 11. Considerable delamination leading to catastrophic failure (11)
Considerable effort has been made to mature composite technology in order to reduce manufacturing costs, guarantee reliably high quality laminates, understand the highly complex failure criteria and built hierarchical, multifunctional or self-healing structures. One of the major shortcomings is that the structural advantages of fibre-reinforced plastics must be viewed with respect to applications where the primary loads are aligned with the fibre direction. However, if a composite plate is subjected to significant out-of-plane stresses subsurface delaminations may develop between layers due to the weak through-thickness cohesive strength of the composite. These intralaminar delaminations are a significant problem as they are difficult to detect by visual inspection and may reduce the compressive strength of the laminate by up to 60%.
4) Novel Designs
With environmental legislation becoming ever so strict it is adamant that new concepts for lightweight and fuel efficient aircraft are found swiftly. Although the pressure on developing advanced composite materials is high it must be remembered that 100 years of innovation were required to reach the stage that large metal semi-monocoque structures could be manufactured in the 1940s and another 30 years to fully understand all failure criteria. Thus we may still require significant research and development before all current issues with composite materials are resolved. Apart from carbon fibre and other composites other researchers have been looking into completely redefining the shape of aircraft. Researchers at MIT have been developing the blended wing concept and NASA are exploring the technology of morphing or shape-changing aircraft, taking inspiration directly from nature.
Fig. 12. Illustration of the MIT Silent Aircraft concept (12).
Fig. 13. NASA morphing wing aircraft (13)
Whatever the final solution might look like the next 5o years in aerospace engineering will be incredibly innovative, ground-breaking and an exciting industry to be part of!