MATERIAL SELECTION FOR THE AIRCRAFT The aircraft industry is widely acknowledged as one the branches of engineering which makes the most stringent demands as far as materials are concerned. The very essence of aircraft construction lies in constructing an efficient
flying machine with the least weight, which makes an in‐depth knowledge of materials necessary. In addition, aircraft components are subject to a vast variety of structural and thermal loading conditions which necessitates the use of several classes of materials, including metals, nonmetals, plastics and composites. However, unlike most other mechanical engineering fields like civil engineering, the factor of safety involved in aerospace construction is very rarely more than 2, so the limiting properties required of the components closely match the properties of the materials commonly used. Design Criteria: There are criterias which govern material selection and airframe design: 1. Strength, 2. Stiffness and 3. Fatigue resistance. 4. Toughness 5. Effects of environmental heating 6. Ease of fabrication 7. Availability and consistency of of supply and 8. Cost. Members such as wing spars are designed primarily on the basis of
strength. In the engineer’s viewpoint, there are 2 important factors: • It does not matter if the spar partially deforms in the process of carrying load
• There is a certain amount of stress occurring in all flight conditions; it is necessary that the material selected is capable of withstanding all values of stress likely to be experienced by the member without total failure.
Stiffness‐based design is customary in components which nee d not necessarily fail to wreak havoc. A slight deformation can be catastrophic as in turbine blades, whose clearance with the casing is very minute. Even the slightest dimensional changes brought on by either aerodynamic or centripetal loads can cause the engine to come to a sudden halt. Fatigue is a phenomenon which occurs due to repeated loading of a member, which causes gradual failure. It is common in members subjected to cyclic or vibratory motion. Fatigue failure is one of the most unpredictable forms since in‐ service analysis of a member for fatigue damage is very difficult. Types of Loads: Loads on aircraft members can be broadly classified as static and dynamic. Static loads are those which are applied gradually such that the time effects of the application of loads are negligible. Dynamic or energy loads are those which are applied impulsively such that the energy with which the structure is loaded also needs to be considered. The weights of different components in the aircraft are static loads, while gust and landing loads on the aircraft are dynamic loads. Material Properties: In aircraft construction, weight is a parameter that the engineer tries to limit to a minimum, while strength is attempted to be maximized. Thus, a higher strength to weight ratio is one of the most important properties preferred by aerospace engineers. Strength is commonly determined by tensile testing of a specimen and plotting a graph between the stress applied and the strain developed
in it. A common stress‐strain curve obtained for metals is shown below: In the stress strain curve, the initial linear part corresponds to the elastic region, where loading and subsequent unloading causes the material to return to its
original state. When the curve fails to be linear by a strain of 0.001, the corresponding stress marks the proportional limit. Immediately after this, the yield stress is achieved. Loading of a member beyond its material elastic limit will result in a certain amount of permanent deformation. Finally, at the ultimate tensile stress, the material fails for good. The yield strength and ultimate strength of the material are crucial in strength‐ based designs, and the stresses developed in the members should always be within these limits.
The slope of the linear portion of the stress‐strain curve is called the Young’s Modulus of the material. Stiffness is defined as the ratio between the load applied and the corresponding elongation. Stiffness is related to Young’s modulus through volume. Hence, if it is necessary to have a stiff member, it is necessary to
select a material with a high Young’s modulus. Every time a material is loaded beyond its yield strength but below its ultimate strength, it undergoes a certain amount of plastic deformation and its properties including yield strength itself get modified. Plasticity coupled with repeated loading results in internal stress concentration which builds up with each loading. After certain number of loading cycles, cracks start propagating causing the material to eventually fail. Hence, the plastic region characteristics are critical in members prone to fatigue failure.
Fatigue is commonly analyzed using the stress intensity factor ‘k’. Creep is a highly temperature‐dependent phenomenon. It is one of the most common modes of failure in thermally highly‐stressed members, such as turbine blades and combustor linings. In order to analyze creep, knowledge of the thermal properties like thermal expansivity is required, in addition to carrying out experiments simulating simultaneous thermal and structural loading. Titanium alloys are very well known for their excellent creep‐resistant thermal properties. In addition to resisting static loads, fatigue and creep, energy loads must also be
resisted. The area under the linear portion of the stress‐strain curve is called the resilience modulus, while the area under the whole of the curve is called the toughness modulus. These moduli are indicators of the energy the material can absorb before failure, and are determined using standard tests like the Izod and Charpy tests.
Compression failure is likely in members that are short and thick. The compression characteristic curves are similar to the tensile characteristics, and have corresponding proportional and ultimate limits. Compressive limits have significance only related to bending stresses. All other compression related failures are due to phenomena such as buckling or crippling.
Common Aircraft Materials:
Aluminum‐based alloys such as duralumin are co nventionally preferred for the skin and most other structural components of the aircraft. This is because of its relatively high strength to weight ratio coupled with its extreme lightness. Nowadays, composites have come to replace aluminum largely due to their higher strength to weight ratio as well as other attractive properties such as lower overall weight and easier component tailoring. However, steel remains the standard material for crucial components like spars owing to its high strength and stiffness, predictable behavior under loading and reliability. Steel is avoided for other components due to its excessive weight.
Another useful class of metallic alloys is the titanium‐based ones, known for their superior thermal properties. High Mach number aircraft make extensive use of these alloys due to the high temperatures associated with aerodynamic heating. Steel: Advantages
1. High tensile strength 2. High Stiffness 3. Resistance to wear 4. Used where high tensile strength is required like in wing root attachments, fasteners etc. Disadvantages:
1. Extremely high specific gravity 2. Difficult to manufacture into finished components 3. Since weight is the most important consideration in aircraft structures, stee l is used only when tensile strength is of prime importance.
Titanium: Advantages:
1. Good fatigue strength 2. Good strength to weight ratio 3. Excellent temperature resistance 4. Good resistance to corrosion Disadvantages:
1. Relatively high density 2. High fabrication costs Aluminum alloys: Aluminum alloys are the most widely used in aircraft structures owing to their high strength to weight ratios. Airframe construction has depended for many years on the three groups of aluminum alloys: 1. The nickel free duralumin, 2. The derivatives of Y alloy
3. The aluminum-zinc-magnesium group. Some commonly used materials and their properties are listed below: MATERIAL
5Cr Mo V STEEL
Al 2024
Al 7075
Ti 6Al 4V
ULTIMATE TENSILE STRESS (MPa)
1654.7
530.896
228
900
YIELD TENSILE STRESS (MPa)
1378.9
455.05
103
830
YIELD COMPRESSIVE STRESS (MPa)
1516.8
525.1
-
-
YOUNG’S MODULUS (GPa)
206.8
73.1
71.4
114
SHEAR MODULUS (GPa)
75.84
21.48
26.95
44
ULTIMATE SHEAR STRENGTH (MPa)
999.7
206.84
-
-
DENSITY (kg/m3)
7790
2800
2810
4430
