Aircraft Structural Design


Although the major focus of structural design in the early development of aircraft was on strength, now structural designers also deal with fail-safety, fatigue, corrosion, maintenance and inspectability, and producability.

Structural Concepts

Modern aircraft structures are designed using a semi-monocoque concept- a basic load-carrying shell reinforced by frames and longerons in the bodies, and a skin-stringer construction supported by spars and ribs in the surfaces.

Proper stress levels, a very complex problem in highly redundant structures, are calculated using versatile computer matrix methods to solve for detailed internal loads. Modern finite element models of aircraft components include tens-of-thousands of degrees-of-freedom and are used to determine the required skin thicknesses to avoid excessive stress levels, deflections, strains, or buckling. The goals of detailed design are to reduce or eliminate stress concentrations, residual stresses, fretting corrosion, hidden undetectable cracks, or single failure causing component failure. Open sections, such as Z or J sections, are used to permit inspection of stringers and avoid moisture accumulation.

Fail-safe design is achieved through material selection, proper stress levels, and multiple load path structural arrangements which maintain high strength in the presence of a crack or damage. Examples of the latter are:
a)Use of tear-stoppers
b)Spanwise wing and stabilizer skin splices

Analyses introduce cyclic loads from ground-air-ground cycle and from power spectral density descriptions of continuous turbulence. Component fatigue test results are fed into the program and the cumulative fatigue damage is calculated. Stress levels are adjusted to achieve required structural fatigue design life.

Design Life Criteria — Philosophy

Fatigue failure life of a structural member is usually defined as the time to initiate a crack which would tend to reduce the ultimate strength of the member.

Fatigue design life implies the average life to be expected under average aircraft utilization and loads environment. To this design life, application of a fatigue life scatter factor accounts for the typical variations from the average utilization, loading environments, and basic fatigue strength allowables. This leads to a safe-life period during which the probability of a structural crack occurring is very low. With fail-safe, inspectable design, the actual structural life is much greater.

The overall fatigue life of the aircraft is the time at which the repair of the structure is no longer economically feasible.

Scatter factors of 2 to 4 have been used to account for statistical variation in component fatigue tests and unknowns in loads. Load unknowns involve both methods of calculation and type of service actually experienced.

Primary structure for present transport aircraft is designed, based on average expected operational conditions and average fatigue test results, for 120,000 hrs. For the best current methods of design, a scatter factor of 2 is typically used, so that the expected crack-free structural life is 60,000 hrs, and the probability of attaining a crack-free structural life of 60,000 hrs is 94 percent as shown in the following figure and table.

s.f. = N / Np Probability of
Survival (%)
Np (Flight Hours)
(N = 120,000 hrs)
Np (Years)
(3,000 flight hrs / year)
2.0 94.0 60,000 20
2.5 97.5 48,000 16
3.0 98.8 40,000 13.3
3.5 99.3 34,300 11.4
4.0 99.54 30,000 10.0

With fail-safe design concepts, the usable structural life would be much greater, but in practice, each manufacturer has different goals regarding aircraft structural life.


Choice of materials emphasizes not only strength/weight ratio but also:

  • Fracture toughness
  • Crack propagation rate
  • Notch sensitivity
  • Stress corrosion resistance
  • Exfoliation corrosion resistance

Acoustic fatigue testing is important in affected portions of structure.

Doublers are used to reduce stress concentrations around splices, cut-outs, doors, windows, access panels, etc., and to serve as tear-stoppers at frames and longerons.

Generally DC-10 uses 2024-T3 aluminum for tension structure such as lower wing skins, pressure critical fuselage skins and minimum gage applications. This material has excellent fatigue strength, fracture toughness and notch sensitivity. 7075-T6 aluminum has the highest strength with acceptable toughness. It is used for strength critical structures such as fuselage floor beams, stabilizers and spar caps in control surfaces. It is also used for upper wing skins.

For those parts in which residual stresses could possibly be present, 7075-T73 material is used. 7075-T73 material has superior stress corrosion resistance and exfoliation corrosion resistance, and good fracture toughness. Typical applications are fittings that can have detrimental preloads induced during assembly or that are subjected to sustained operational loads. Thick-section forgings are 7075-T73, due to the possible residual stresses induced during heat treatment. The integral ends of 7075-T6 stringers and spar caps are overaged to T73 locally. This unique use of the T73 temper virtually eliminates possibility of stress corrosion cracking in critical joint areas.

Miscellaneous Numbers

Although the yield stress of 7075 or 2024 Aluminum is higher, a typical value for design stress at limit load is 54,000 psi. The density of aluminum is .101 lb / in3

Minimum usable material thickness is about 0.06 inches for high speed transport wings. This is set by lightning strike requirements. (Minimum skin gauge on other portions of the aircraft, such as the fuselage, is about 0.05 inches to permit countersinking for flush rivets.

On the Cessna Citation, a small high speed airplane, 0.04 inches is the minimum gauge on the inner portion of the wing, but 0.05 inches is preferred. Ribs may be as thin as 0.025 inches. Spar webs are about 0.06 inches at the tip.

For low speed aircraft where flush rivets are not a requirement and loads are low, minimum skin gauge is as low as 0.016 inches where little handling is likely, such as on outer wings and tail cones. Around fuel tanks (inboard wings) 0.03 inches is minimum. On light aircraft, the spar or spars carry almost all of the bending and shear loads. Wing skins are generally stiffened. Skins contribute to compression load only near the spars (which serve as stiffeners in a limited area). Lower skins do contribute to tension capability but the main function of the skin in these cases is to carry torsion loads and define the section shape.

In transport wings, skin thicknesses usually are large enough, when designed for bending, to handle torsion loads.

Fuel density is 6.7 lb/gallon.

Structural Optimization and Design

Structures are often analyzed using complex finite element analysis methods. These tools have evolved over the past decades to be the basis of most structural design tasks. A candidate structure is analyzed subject to the predicted loads and the finite element program predicts deflections, stresses, strains, and even buckling of the many elements. The designed can then resize components to reduce weight or prevent failure. In recent years, structural optimization has been combined with finite element analysis to determine component gauges that may minimize weight subject to a number of constraints. Such tools are becoming very useful and there are many examples of substantial weight reduction using these methods. Surprisingly, however, it appears that modern methods do not do a better job of predicting failure of the resulting designs, as shown by the figure below, constructed from recent Air Force data.

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