The prime structural components within a semi monocoque aircraft structure comprise of but are not exclusive to frames, Ribs, Spars, Stringers and the skin. The following report will analyse and evaluate the various materials that are used within the aircraft structure with a comparison taken between aluminium alloys and composite materials with regards to the properties and characteristics of the required material to fill the operational performance of the aircraft. The loads and theatre that the aircraft will encounter and operate within will determine the material used for each structural component with drawbacks investigated for each material. The future of aircraft material usage will be investigated.
The primary task of the wing is to transfer the loads from the engine and lift forces generated by the air and transfer them to the fuselage structure. The wing cross-section takes the shape of an airfoil, which is a design based on the aerodynamic and operational requirement of the aircraft. The wing as a whole performs the combined function of a beam and a torsion member. It consists of axial members in stringers, bending members in spars and shear panels in the cover skin and webs of spars. The spar is a heavy beam running span wise to take transverse shear loads and span wise bending. It is usually composed of a thin shear panel (the web) with a dense flange or cap at the top and bottom to take bending. Wing ribs are planar structures capable of carrying in-plane loads. They are placed chord wise along the wing span. Ribs serve as load redistributors; ribs also hold the skin stringer to the aerodynamic shape of the wing and reduce the effective buckling length of the stringers.
In a typical rib construction the rib is supported by span wise spars. The cover skin of the wing together with the spar webs forms an efficient torsion member. For subsonic aircraft design, the skin structure is comparably thin and will almost within every design required to undergo post buckling. The thin skin can be assumed to make no contribution with regards the bending loads on the wing box structure, thus the bending moment is ultimately distributed between the stringers and spars.
A variance exists in cross sectional design within Aircraft with a two spar subsonic design concept; one type consists only of spars (the concentrated flange type) to take bending. The other type uses both spars and stringers to take bending. Within Supersonic aircraft design airfoils are relatively thin when a direct comparison is taken with the direct opposite design of subsonic airfoils. To with stand high surface air loads and to provide additional bending capability of the wing box structure, thicker skins are often necessary. In addition to increase structural efficiency, stiffeners can be manufactured by forging or machining as integral parts of the skin.
Fuselage Structure
The fuselage carries the payload, and is the main body to which all main aerodynamic structural parts are connected. It must be able to resist bending moments (caused by weight and lift from the tail), torsion loads (caused by fin and rudder) and cabin pressurization/depressurisation. The structural strength and stiffness of the fuselage must be high enough to withstand these loads, at the same time; the structural weight must be kept to a minimum.
semimonocoquefuselagedesign1.jpg
Fig 2-Semi monocoque Fuselage
Source-http://aviationglossary.com/aircraft-structure
In commercial passenger aircraft design, the vast majority of the fuselage is cylindrical in shape that narrow to a taper at the forward and aft sections. The semi monocoque construction consists of a stressed skin with added stringers to prevent buckling, attached to frames.
(www.wseas.us/e-library/conferences/2010/Corfu/.../HEAPFL-17.pdf)
The fuselage also has structural components perpendicular to the skin, that supports it and helps keep its shape. These structural supports are called frames if they are open/ring shaped or called bulkheads if they are closed. Disturbances in the perfect cylindrical shell, such as doors and windows, are called cut outs. They are usually unsuitable to carry many of the loads that are present on the surrounding structure. The direct load paths are interrupted and as a result the structure around the cut out must be reinforced to maintain the required strength. (aerostudents.com/files/aircraft Structures/aircraft Structures Full Version. pdf )
To provide shape and prevention from buckling when subjected to bending loads, incorporated within the fuselage are frames.. Stringers give a large increase in the stiffness of the skin under torsion and bending loads, with minimal increase in weight experienced overall within the fuselage structure. Frames and stringers make up the basic primary structure of the fuselage. Pressure bulkheads close the pressure cabin at both ends of the fuselage, and thus carry the loads imposed by pressurization.
Typical wing/Fuselage structure material
Aluminium Alloy
As previously investigated the loads on the main structural components of the wing and fuselage are varied and complex. Aluminium Alloys have been the primary material source within aircraft design, though in recent years some new alloys and composites have been applied, these materials will be investigated in the latter stages of the report. With its good strength to weight and cost ratio, aluminium is still used widely within the industry. The quantities of Aluminium required for aircraft structure is easily supplied as only silicon and oxygen elements are greater in number within the earth’s crust (www.aluminium.com/knowledge). It appears as a silvery white metal that has a strong resistance to corrosion and is malleable. If you analyze the contents of the stress strain graph below it is obvious to understand why Aluminium is the preferred option within metallurgy for Frames, Ribs, Spars, Stringers and the skin. The upper wing of a commercial aircraft is a compression dominated structure with Aluminium the material that meets this and the other requirements of damage tolerance and low cost.
stress_strain_tension.gif
Fig 2-stress strain graph (metals) source- www.tms.org
Where lightness combined with a requirement to accommodate bending and compression loads particularly on spars and stringers then aluminium comes into its own as a material. On figure 3 we can see the bending load moment equation on a typical wing spar.
http://ciurpita.tripod.com/rc/notes/images/ibeam.gif
Fig 3 - Spar Structure- Source Anon
B H3 - b h3
M = -----------
6 H
To understand the bending strength of a spar it is first a necessity to ascertain the bending moment required. The bending strength is determined by the structure of the spar, its shape, and the materials used. Aluminium alloy provides great bending and shear load absorption and distribution as illustrated on figures 3 and 4.Determining the equations for the bending strength of the spar is not an easy and straight forward task. The problem is that the force is linear varying from zero at the neutral axis to its maximum at the surface. The b dimension represents the total empty space within the beam and may be split, as in the case of an I-beam. (www.seqair.com/Sequoia/FBL/BuilderLetter908EBook.pdf)
Aluminium alloys are a relatively light metal compared to steel with a good strength to weight ratio and very easily machined therefore it is the material of choice for formers, frame assemblies, and bulkheads. On figure 4 the strength to weight of aluminium is compared to other metal material.
metals3.jpg
Fig 4 -Strength to Density table.
Source- www-materials.eng.cam.ac.uk
These alloy structural members give cross sectional shape, rigidity, and strength to the fuselage. The shapes and sizes of these members vary considerably, depending on their function and position in the fuselage. Formers are secondary structures and are used primarily for skin attachments between the frames/bulkheads. Frame assemblies are the most abundant members in the fuselage and are visually present as strengthening devices. (http://www.freeed.net/sweethaven/Aviation)
As described earlier aluminium alloy is a very versatile metal and can be cast in any form. It can be rolled, stamped, drawn, spun, roll-formed, hammered and forged. The metal can be extruded into a variety of shapes and can be turned, milled, and bored in the machining process. (Www-materials.eng.cam.ac.uk)
Fuselage stringer extrusion is primarily manufactured from Aluminium alloy and is vital to modern pressurised aircraft. Extrusion can be formed into complex shapes with minimum material wastage encountered. The ability to place the material exactly where it is required (using a die) helps to overcome the fact that Aluminium alloy has a 1/3 of the modulus of elasticity of steel. Tension and compression loads are taken by the stringers assisted by the longerons and sometimes spars to prevent the fuselage skin from buckling and the wing from encountering excessive bending loads. Figure 5 below provides a generic schematic of stringer placement for one spar and skin assembly location.
744px-Stringer_cross_section_svg.png
Fig 5- spar, skin and stringer assembly-Source Anon
A crucial aspect of Aluminium alloy use in fuselage design is the ability of the material to be riveted, welded or resin bonded. Aluminium alloy needs no protective coating as certain machine finishes ensure a visually satisfying look, however it is often treated to ensure a more protective finish is guaranteed.
One of the great advantages of aluminium alloys is the ability to maintain and increase strength at low operational temperatures. As passenger aircraft regularly fly at high altitude thus low temperature operating environment (-50 degrees C) it is abundantly clear why commercial aircraft (Hirsch 2008) profit from this property. However during system operations that require high temperatures (200-250°C) aluminium alloys lose some of their strength and thus very rarely used in high temp environments (engine firewalls etc).
Aluminium alloys also have a strong resistance to corrosion which is a result of an oxide skin that forms as a result of reactions with the atmosphere. This is of very great benefit to aircraft skins where the elements take their toll on surface finish protection materials. This oxide skin protects aluminium from chemicals and acids, however alkaline substances can and will break through this defence to cause serious damage.
Composite Material
The development of composite materials has revolutionised the way aircraft are currently being designed and operate and within the following chapter we will analyse what properties within composites make this so.
Composite materials must be a combination of two or more components. These components can either be organic or inorganic with one taking the role as the matrix and the other as the reinforcement material.
The matrix basically bonds everything together within the composite and the reinforcing material does what it advertises (reinforces the composite)
They are numerous reinforcing materials currently in use within the aviation industry with the most common being fibreglass (light structural loads) and carbon fibre. The development of reinforcing materials has introduced boron, carbon and Aramid into the aircraft design arena and this is where the main structural components of future and current aircraft design have benefited.
Strength to weight ratio of composites is of great value within structural design particularly where this weight saving can allow operators to deliver more passengers utilising less fuel thus increasing profits. While aluminium is an excellent material for aircraft primary structural design, composites have now exceeded aluminium alloy in the most desired structural characteristics for aircraft design. On figure 6 we can see the benefits of composite materials compared to aluminium alloys with regards specific weight versus the specific modulus.
comp2.jpg
Fig 6 specific strength versus specific modulus.Source-www.machinedesign.com
To utilise carbon composites and achieve their optimum performance under main stress loads within aircraft structures, it is of the ultimate priority to position the carbon fibres in the direction of the stress loads.
An example of the fibre positioning within a primary structure would be on the wing spar and ribs within the wing. The fibres are positioned at 45 degree angles within the spars and ribs to accommodate the shear loads experienced. Figure 7 below shows the shear forces experienced on the wing spar with the torsion, compression and shear forces illustrated.
Fig 7 wing rib with integrated spar
Source-INGENIA ISSUE 36 SEPTEMBER 2008
Carbon fibre composite material within fuselage structural components is becoming more common due to the ease of manufacture of such components and the strength to weight ratio when assembled in modular form. The reduction in maintenance requirements and thus a reduction in operational cost have ensured that composite fuselages are of the ultimate goal in large passenger aircraft design.
The structural load advantages of carbon composite fuselage structures vastly out way the traditional aluminium alloy design. The notable high strength of fatigue and specific strength coupled with a corrosion resistance currently unmatched in aircraft design ensure a formidable structural material
composite_barrel_lg.jpg
Fig 9 Carbon composite barrel
Source-www.boeingblogs.com
Carbon composite fuselages usually integrate the skin and stringers into one mould utilising heat and pressure (Fig 9). The advantage of this process allows the fibres to be positioned at such angles and when combined with various layers within the resin it produces a composite that sustains specific loads from the aircraft. Tension and compression loads encountered normally through the stringers and skin can then be sustained and distributed with greater accuracy than the typical aluminium alloy design. Figure 10 highlights the areas that impact damage can occur on the fuselage and potential causes of such damage.
Fig 10 Impact scenarios for fuselage structure. Source-Anon
One of the disadvantages of manufacturing carbon composite modular fuselages is the identification of impact damage and delamination sustained during a typical operation. Unlike aluminium alloys which exhibit that great quality of placicity, Carbon fibre epoxy composites do not. This presents a serious challenge for structural engineers as the smallest of operational impacts ultimately result in a failure within the composite laminate. The diagnosis of the extent of this damage is not easily performed by a typical and cost effective manual visual inspection. Non destructive inspection techniques will be required resulting in a greater time on the ground for the aircraft resulting in a significant financial cost implication.
Conclusion
The design of commercial aircraft will for the foreseeable future still mainly rely on aluminium alloys as the primary material for manufacture. The cost of the initial set up of composite machinery requirements and training needs for staff will have a large bearing on this outcome. Although the cost of the raw material involved with manufacturing and repairing composites has declined over the last decade it is still not as abundant as and therefore more expensive than aluminium alloy. Both materials produce excellent results for all load factors experienced on the aircraft and its primary structure as discussed previously. Although carbon fibre composites have a great strength to low weight, composites have not been a miracle solution for aircraft structures. Composites are extremely difficult to inspect for damage requiring special and expensive techniques with the majority suspect to moisture absorption unlike aluminium alloy. Aluminium alloys are far simpler to repair and have the robustness required to resist impact fatigue damage encountered during every day operations. Confidence in material usage is a massive factor in dictating the way aircraft are designed, composites are still a relatively new material and until the mystery of the mechanics are widely adopted and understood it will remain underused.
Future Materials
The future of both aluminium alloys and composite structures are extremely bright with low density aluminium Lithium alloy and ceramic matrix composites leading the way. Aluminium Lithium alloy has been primarily created for major structural components within aircraft design. It has a greater strength to weight ratio than aluminium alloy and less prone to flaws during the manufacturing process. Although this would ensure a great cost saving to aircraft operators with regards fuel and passenger numbers, it does have downsides. They are two main concerns for the advancement of Al-Li alloy structural design. The manufacturing processes are problematic due to the risk of explosion when molten AL-Li comes into contact with water. Various factors will determine the size of explosion (depth of water and diameter/velocity of molten stream). The cost of Al-Li alloys are also up to five times the cost of the typical aluminium alloys in use, thus limiting usage within the design market. (http://ezinearticles.com/?Introduction-to-Aluminum-Lithium-Alloys&id=950764)
The advancement within ceramic matrix composites is a massive step forwards with regards structure material usage. CMC’s provide great properties for aviation design development. The material is lightweight and has high temperature strength, increased fracture toughness, extremely damage tolerant and resistance to thermal shock. Further research is currently being carried out at Boeing and Augusta Westland on the potential of this composite. The downside to CMC’s is the same as standard carbon fibre composites with impact damage very difficult to identify using standard visual inspection techniques.
Both of the materials give a glimpse of the future with their introduction into aviation structural design a certainty over the coming decades
