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Service Life of Airplane Structures

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Service Life of Airplane Structures ICAS 2002 CONGRESS SERVICE LIFE OF AIRPLANE STRUCTURES Prof. Grigory I. Nesterenko Central Aerohydrodynamic Institute (TsAGI), Russia Keywords: Fail safe, design goal, damage tolerance, fatigue Abstract stresses. The materials having improved strength, fatigue and crack resistance properties Test-analytical results of investigating into the ult fatigue, fail-safety and damage tolerance of were used in them. The stresses σ under the Russian aircraft are presented. Stresses in ultimate loads represent the maximum tolerable wide-body aircraft structures are given. stresses in static strength criteria. Equivalent skin Fatigue life curves for wing and fuselage stresses σequiv represent the maximum cycle structures are generated. Residual strength stresses having the stress ratio R=0. The data are presented for these structures having a structural damage per cycle, σequiv , is equal to skin crack under the broken stiffener. the structural damage for all the stress cycles Generalized curves for skin crack duration during one typical flight. The stresses σ*equiv are under the broken stiffener are presented. equal to the sum of stresses σ and additional stresses ∆σ under the loading of stiffeners 1 Introduction (rivets, bolts). The values of σequiv and σ*equiv in The design and service experience of the fatigue have been calculated by conservative transport aircraft for the recent 50 years has methods of Refs. [1, 2] using the linear damage shown that to ensure the aircraft reliability and accumulation hypothesis. The values of σequiv in efficiency it is required to provide in their crack growth with regard for their retardation structures the solution of three concepts have been outlined by the author, based on simultaneously, i.e. safe life, fail-safe and special test results. The stresses σequiv have been damage tolerance. The comprehensive data found for the wide-body aircraft having flight about these structural features have been duration of 7-8 hours. The crack growth has been collected by now in the course of testing inspected mainly by visual aids at annual coupons, panels and full-scale structures. This intervals (600-700 flights). It is supposed that presentation generalizes the information in such inspections detect reliably the skin cracks order to estimate tolerable stresses, fatigue 50 mm long under the broken stiffener. While performances, fail safety, damage tolerance of evaluating the stresses given in Table 1 the wide-body aircraft fabricated of modern analysis covered the stresses in the aircraft materials. The generalization covered the structures both for Russian aircraft and those results of the joint TsAGI activities with data published in Refs. [7-12]. Antonov, Ilyushin, Tupolev, Yakovlev companies. 3 Design goal The aircraft structural lifetime is limited by 2 Stress fatigue of longitudinal joints in the wing lower The stresses have been estimated in the wing surface panels and those in the fuselage skin and fuselage structures of contemporary wide- longitudinal joints. These joints are affected by body aircraft (Table 1). The tolerable maximum multiple site crack initiation at the end of the stresses have been found in the wide body fatigue life. Hence the lifetime for the aircraft structures as they have the highest longitudinal joints is estimated by safe-life 331.1 PROF. GRIGORY I. NESTERENKO concept. The fatigue life of the longitudinal residual strength of the lower wing surface [5] is joints depends on material properties and not ensured at the stresses taken from Table 1 assembly technology. Figs.1 and 2 present The residual strength in the structures of the minimum experimental fatigue lives of upper wing surface fabricated of high-strength longitudinal joints in the wing panels and in the alloys having the standardized damaged [13] and lim ult fuselage skin laps. These data are taken from stress value σ = 0.67 σ tens, defined from lifetime tests of full-scale structures. Different Table 1. It should be noted that in such structures aircraft models have different symbols for the a drastic crack length increase at the stresses of experimental points. The cracks 2-15 mm long 160-170 MPa is observed. have been detected in the longitudinal joints at In the structures of pressurized fuselages the mentioned lives. The relations of minimum having longitudinal two-bay skin cracks under lifetime values versus equivalent stresses the broken frame the residual strength is ensured σ*equiv have been determined from the Pr at the stresses σ = 1.15 (Table 1) in the experimental data (Figs.1 and 2). While t estimating the minimum lifetime values the fact following cases (Fig. 4): was taken into account that each experimental ● critical stress intensity factor in skin material is point in Figs.1 and 2 is the result of testing full- scale structures having thousands of similar Kapp = 135 MPa m and stoppers applied; stress concentrators in the longitudinal joints. ● there are no stoppers, but Kapp ≈ 170 MPa m . It follows from Figs.1 and 2 that at the equivalent stresses σ*equiv given in Table 1 the n the structures of pressurized fuselages lifetime of 20000 flights is ensured in the having lateral two-bay skin cracks under the structures of contemporary wide-body broken stringer the residual strength is ensured at airplanes. the stresses σlim = 0.67 σult (Table 1) in those cases when the stringers are fabricated of high- 4 Fail safe strength materials of 7000 series [13]. The residual strength of the structure having D16 To ensure fail-safe concept it is required stringers is approximately 220 MPa (Fig. 5). that the structure has standardized damages in the form of two-bay skin crack under the broken stiffener to maintain the strength under 5 Damage tolerance the limit load of Plim = 0.67Pult. To evaluate this To ensure damage tolerance concept it is criterion, test-analytical results have been required that the crack growth duration from the generalized in terms of residual strength of the reliably detectable to the tolerable length is no wing and fuselage structures (Figs. 3 to 5). less than the aircraft structure inspection interval. Here the experimental data were used about the The safety factor for crack growth duration is residual strength of panels and full-scale assumed equal to 2. Visual inspections are structures published in Refs. [3-6]. The residual assumed principal. Inspection interval is 700 strength for the structure having such damages flights (once a year). The reliably detectable has been analyzed using the methods presented damage is assumed to be the skin crack having in Ref. [4]. The calculations utilized linear the length 2a = 50 mm under the broken stiffener fracture mechanics. (stringer, frame). The generalized average curves The residual strength of the lower wing of crack growth in the wing and fuselage surface structure having two-bay skin crack structures are presented in Figs. 6 to 8. These under the broken stringer is ensured at the limit figures also demonstrate the material properties stresses σlim = 0.67 σult (Fig. 3, Table 1). It of the structures under testing. should be noted that when there are skin cracks It follows from these curves that the under the broken caps of front or rear spar the damage tolerance concept is ensured in the wing 331.2 SERVICE LIFE OF AIRPLANE STRUCTURES and fuselage structures at the specified stresses [6] Vovnyanko A.G., Semenets A.I. Residual strength of and visual inspection intervals. integrally-stiffened structures of extruded panels of D16chT alloy and its modifications. – “Fiziko- khimicheskaya mekhanika materialov”, No. 2, 1983, 6 Conclusion Lvov, Ukraine, p.88-92 (in Russian). [7] Swift T. The application of fracture mechanics in the The values of tolerable stresses in the wing and development of the DC-10 fuselage. – AGARD-AG- fuselage structures of contemporary wide-body 176, v.5, 1974, p.227-287. aircraft have been determined. [8] Swift T. Design of redundant structures. – AGARD- LS-97, 1978, p.9.1 – 9.23. The relations of the wing and fuselage design [9] Goranson Ulf G. Damage tolerance. Facts and fiction. – 14th Plantema Memorial Lecture Presented at the goals for modern transports versus cyclic th stresses are defined. 17 Symposium of ICAF, Stockholm, Sweden, June 9, 1993. [10] Kevin R., Fowler and Watanabe Roy T. The residual strengths for the wing and Development of jet transport airframe test spectra. – fuselage structures with two-bay skin crack Boeing Commercial Airplanes, Seattle, Washington. under the broken stiffener have been outlined to [11] Gökgöl O. Crack free and cracked life of the estimate fail-safe parameters of these airplane pressurized cabin of the A-300B.- Calculation, tests structures. and design measurements to improve damage tolerance. – Aeronautical Journal, 1979, v.83, No. 817, p.1-15. The generalized average curves of skin crack [12] Spencer M.M. The 747 fatigue and fail safe test growth duration from visually detectable till program. – The American Society of Civil Engineers, tolerable lengths under broken stiffener are National Structural Engineering Meeting. San plotted for damage tolerance estimation. Francisco, California, 9-13/IV, 1973. [13] Nesterenko G.I. Ensuring damage tolerance of advanced wide-body aircraft structures. – Proceedings References of the 21st Symposium of ICAF, July 27-29, 2001, [1] Vorobyev A.Z., Leybov V.G., Olkin B.I., Stebenev Toulouse, France, v.II, p.965-977. V.N. Ensuring enhanced lifetime of wide-body aircraft. – In “Strength of aircraft structures”, Moscow, Mashinostroyeniye, 1982, p.122-151 (in Russian). [2] Vorobyev A.Z., Olkin B.I., Stebenev V.N., Rodchenko T.S. Fatigue strength of structural elements. - Moscow, Mashinostroyeniye, 1990, 240p (in Russian). [3] Nesterenko G.I., Selikhov A.F. Application of damage tolerance concept in developing wide-body aircraft.- In “Strength of aircraft structures”, Moscow, Mashinostroyeniye, 1982, p.151-189 (in Russian). [4] Nesterenko G.I.
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