9.pdf

DESIGNING THE AIRPLANE STRUCTURE FOR HIGH DURABILITY

Grigory .I. Nesterenko TsAGI, Zhukovsky Moscow Region, 140180,

Russia

Abstract The development of damage tolerance and fatigue Regulations in Russia is shown. The methods to provide long service goals are listed. Fail-safe and damage tolerance criteria are described. Improvements in fatigue and crack resistance performance are illustrated. Increase of service life due to the advance technology of primary structure elements joints is demonstrated. Stress values for modern aircraft structure are presented. Specifications of two structure types, integral stiffened and riveted, are given. Scopes of fatigue certification tests for new and aging airplanes are presented. The results of experiments performed to prove meeting the criteria of fatigue, crack growth time and residual strength of wing and fuselage structures are described. The developed method of analysis for residual strength of stiffened panels with use of R-curves is illustrated. The results of analysis of crack growth time based on a linear model as well as on crack growth retardation model are presented. The results of analytical-experimental research are presented regarding the following: residual strength of a structure with widespread fatigue damage (WFD), degradation of fatigue and damage tolerance performance of aging aircraft structures, corrosion. The possibility to provide high durability is proved by more than 40 years of service life of airplane-leaders in Russia.

Introduction The task to guarantee simultaneously reliability, long durability, minimum weight and cost effectiveness of transport airplanes is one of the most important problems in the contemporary aircraft building. The 50-year experience of creation and operating transport category airplanes in the USSR and Russia showed that to achieve long durability the design should be driven by three principles simultaneously1. Regular longitudinal joints of wing panels and fuselage joints must be designed meeting the safe-life principle. In the rest of airframe primary structure elements, the fail-safe and damage tolerance principles must be met simultaneously. Up to now, comprehensive data on fatigue, fail-safe and damage tolerance performance of airframes have been obtained in testing specimens, panels, and full-scale structures, the data being generalized in the present paper. The results of

analytical-experimental research performed in TsAGI together with Antonov, Ilyushin, Tupolev and Yakovlev Design Bureaus have also been generalized.

Improvement of Airworthiness Requirements Aircraft structural properties and reliability of operation depend on existing Airworthiness requirements. Requirements to safe-life, fail-safe and damage tolerance for civil aircraft in the USSR and Russia are presented in Table 1. In 1950-1970th only the safe-life concept was used to provide safety of long aircraft operation2. In 1976 the operational damage tolerance concept was introduced as an equitable alongside with safe-life. In the USSR and Russia practice the operational damage tolerance principle includes both damage tolerance itself and fail-safe principle. In 1994 Russian Aviation Regulations for Transport Category Airplanes (P 25.571) have been introduced, in which the principle of operational damage tolerance is set as the main one. According to the Regulations, recommendations for designers were developed to provide damage tolerance1,3 and fatigue strength4,5 of aircraft structures. Design service goals of airplanes in the USSR and Russia were provided on the basis of these recommendations (Fig.1). The main recommended criteria for providing operational damage tolerance at the stage of design development are shown in Figs. 2-41,3. Much attention was paid to provide operational damage tolerance of structures with regard to WFD. Fail-safe requirements for a wing in the case of WFD1 are presented in Fig. 5. With this damage, the wing structure must retain its strength at limit loads. To check probability of multi-site cracks, fatigue certification testing must cover at least three design goals1. It should be noted that in TsAGI classification, multiple site damage (MSD), multiple element damage (MED), and widespread fatigue damage (WFD) are united within a single term multiple site cracks. Multiple-site cracks are divided into two types: multiple site cracks in a single element (panel), and multiple site cracks in a structural cross-section which consists of several elements1.

1 American Institute of Aeronautics and Astronautics

AIAA/ICAS International Air and Space Symposium and Exposition: The Next 100 Y14-17 July 2003, Dayton, Ohio

AIAA 2003-2785

Copyright 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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Improved Material Properties One of the main activities to achieve long fatigue life of structures is to improve fatigue and crack resistance performance of aluminum alloys, which are the main materials in modern aircraft structures. These materials do have significant reserves for such an improvement. Aluminum alloys 1163T (Al-Cu) were developed and implemented in Soviet aviation industry in 1980th to replace D16T. Hot rolled plates and extruded panels of these new alloys have much longer fatigue life and higher fracture toughness than older D16T ones (Fig. 6). It should be noted that these new alloys have greater scattering of fatigue life among specimens from different melts. Fig. 6 depicts values for the best melts of 1163T, the values being the guideline for perfecting metallurgical technology.

Higher Fatigue Resistance of Joints Design development with regard to long fatigue life includes, as one of the most important tasks, rational design of structural elements and parts that makes stress concentration as minimum as possible. Fatigue quality of structural elements is associated with a coefficient of fatigue strength KF5, which is similar to stress intensity factor Keff. The structures are divided into the following categories against KF value: KF < 3 good structures; KF = 3 4 - satisfactory structures; KF > 4 unsatisfactory structures Fatigue performance of joints is driven by factors of structural design and manufacture technology. New types of rivets and riveting processes were developed to increase fatigue resistance of riveted joints6. Substantially higher fatigue performance in these joints is achieved by means of greater tightness due to plastic deformation. Increase of fatigue life of pressurized fuselage longitudinal lap joints by means of these advanced rivets is shown in Fig. 7 as an example.

Two Types of Structures: Integrally Stiffened and Riveted

Two design schools were formed in the Soviet Union. One of them develops riveted structures; another one deals with integrally stiffened (monolithic) structures made of extruded panels (Table 2). Opponents of integrally stiffened structures say that riveted ones have better fail-safe and damage tolerance performance due to dividing load carrying elements. They say that integrally stiffened structures made of extruded panels have lower corrosion resistance. Vise versa, opponents of rivets believe that much less number of holes is a great advantage of integrally stiffened structures since holes are stress concentrators and sites of crack initiation. Their analyses show that integrally stiffened structures have lower weight compared to riveted ones. For example, extruded panels with special tips for

transversal joints are used in AN-124 wing (Fig. 8), the tips substantially decreasing wing weight and providing easy inspection of wing transversal joints. Some issues of damage tolerance and fail-safe of integrally stiffened panels are studied in7. Comparison of damage tolerance and fail-safe characteristics of integrally stiffened and riveted structures has been performed in8. The results from these references show that fatigue and crack resistance performance of hot rolled plates used in riveted structures and that of extruded panels used in integrally stiffened structures are close to each other (Fig. 6). The 40-year experience of operating airplane structures made of integrally stiffened panels (Fig. 1, Table 2) confirms the possibility to provide protection against corrosion for extruded panels.

Full Scale Certification Tests In the Airworthiness requirements for civil airplanes (NLGS) in the USSR, results of full-scale fatigue and damage tolerance/fail-safe laboratory tests were regarded as the data of great importance. None of the aircraft types avoided full scale fatigue tests covering at least 3 design service goals (fatigue factor of safety). Meantime, for each aircraft type, several prototypes were tested, including those after some time of operation (Table 3). NDI methods for principal structural elements were advanced in these tests. Teardown inspections and subsequent flaw detections were performed to find small fatigue cracks. Methods of fatigue and damage tolerance analysis were updated on the basis of these results.

Stresses in Contemporary Structures Appropriate examination was performed, and stress levels in contemporary wide-body structures were determined (Table 4) to develop recommendations to provide safe life, fail-safe, and damage tolerance, as well as to set requirements to aluminum alloys properties. The wide-body aircraft structures are the most stressed ones compared to other types of structures. Thus, improvement of aluminum alloys is driven mostly by the requirements to materials for wide-body planes. Stress values under ultimate static loads and equivalent stresses eq

, *eq

eq

are presented in

Table 4. Equivalent skin stress equals maximum cyclic stress under the factor of cycle asymmetry R=0. Damage accumulated during one cycle of equals the damage over all the cycles during a standard flight. Values of depend on skin tensile loading. Value of

is the sum of and (the latter caused by loads carried by fasteners (rivets, bolts)). Fatigue equivalent stresses were determined by means of techniques from

eq

eq*eq eq

5,9 using Palmgren-Miner linear

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hypothesis of damage accumulation. Crack growth equivalent stresses take into account the effects of crack growth retardation and were determined in a dedicated experiment. Equivalent stresses in a wide-body wing structure have been determined for 7 to 8 hours flight duration.

Meeting Safe Life Requirements Safe life of aircraft structures is limited in most cases by fatigue in wing lower panel longitudinal joints and fuselage skin longitudinal lap joints, where hard to control multi-site cracks occur. Therefore, service goal of these longitudinal joints, and accordingly service goal of the whole airplane is defined using safe-life concept. Fatigue in longitudinal joints depends on fatigue properties of materials and joint manufacture technology. Full-scale experimental data on fatigue in wing lower panel longitudinal joints (Fig. 9) and fuselage skin longitudinal lap joints (Fig.10) have been generalized to determine aircraft service goals. Different aircraft types are marked in Figs. 9 and 10 with different point markers. Experimental points with arrows mean that no cracks occurred during given life time. The data on Boeing, McDonnel Douglas and Airbus Industrie planes are presented based on examining10-16. Russian aircraft structures presented in Figs. 9&10 are made of D16T. Minimum values of fatigue life vs. equivalent stress are determined approximately regarding lower boundary of experimental data. When making this minimum value estimation the following was taken into account: each experimental point was obtained from full scale tests of the structure with thousands of similar stress concentrators. One should also keep in mind that minimum fatigue life values for a structure made of advanced 1163T alloys will exceed those in Figs. 9&10 because of higher fatigue strength of 1163T as compared to D16T (Fig. 6). The data in Figs. 9&10 result in the following: service goals not less than 20000 flight for the wing and 30000 flights for the fuselage are provided under prescribed stress values in D-16T structures (Table 4).

*eq

Meeting Fail-Safe Requirements

Meeting fail-safe requirements means that a structure with standardized flaws (Figs.2&3) retains static strength at limit stress lim equal to 67% of ultimate stress ult. A fuselage structure with longitudinal cracks must retain static strength at stress of 1.1 pr/t. Experimental data on residual strength of large panels and full scale structures (Figs. 11-13) were generalized to investigate conditions of meeting these requirements. The residual strength of riveted and integrally stiffened wing and fuselage D16T structures with a two-bay crack under the broken stringer is 220-240 MPa. The

residual strength of stiffened D16T structures is limited by strength of stringer material. Thus, allowable level of ultimate stress ult in these structures with regard to fail-safe concept must not exceed 330-360 MPa. Extruded stringers of D16T are mainly used in wings and fuselages of Russian airplanes. To achieve residual strength at ultimate stresses 380 MPa and more, stringers of high strength alloy B95 (7000 series) should be used. Residual strength of the pressurized fuselage with a longitudinal two-bay crack under the broken frame depends on skin failure criteria. Residual strength of fuselages at circumferential stress 115 MPa (Table 1) is provided in the following cases (Fig.13):

stoppers are installed under the frame, critical stress intensity factor in the skin material is Kapp = 135 MPa m .

no stoppers, but improved skin material with Kapp = 175 MPa m is used.

Meeting Damage Tolerance Requirements

When meeting damage tolerance requirements, visual inspections with intervals of 6000-7000 flights (approximately once per year) are regarded as the main means for crack monitoring. It is expected that skin cracks 50 mm long under the stiffener (stringer, frame) are reliably detected in these inspections. Starting from this initial length, a crack must not exceed two-bay length over the inspection interval. Scatter factor of 2 is assumed. Test data on crack growth time in full-scale structures have been generalized to investigate the possibility to meet these requirements (Figs. 14-16). It is concluded from these data that the formulated requirements are met for D16T structures of wing and fuselage (Table 4). There are reserves to increase frequency of inspections by 2-3 times.

Improved Methods of Residual Strength Analysis Linear fracture mechanics methods are used today for residual strength analysis. Most residual strength analyses of stiffened structures do not take into account stable additional growth of the skin crack under static loading. Neglecting this growth decreases analysis accuracy. Sometimes this neglecting causes uncertainty, i.e. which particular element skin or stringer - is a critical one with respect to residual strength. A method of residual strength analysis of stiffened structures with a two-bay skin crack under the broken stringer was developed in TsAGI on the basis of R-curves for skin material17 to specify the conditions of meeting fail-safe requirements. The method was checked by comparison of experimental and analytical values for residual strength of integrally stiffened and riveted wing and fuselage panels. The accuracy of residual strength analysis is 1-5%. Principles of the

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method are illustrated by analysis of residual strength of IL-86 fuselage with a two-bay skin crack under the broken stringer (Fig. 17).

Improved Analysis of Crack Growth Rate Accurate evaluation of the structural damage tolerance may be performed only when all the information about alternating load spectra is available, and interaction of loads is taken into account in crack growth rate analysis. Spectra of load factor increments for different aircraft types were examined in TsAGI, and typical loading program for heavy transport airplane wing was developed18 to solve the above problem. Analytical-experimental studies of skin crack growth rate in upper and lower wing surface of wide-body airplane were conducted using this program (Fig. 18)19. Effects of crack growth retardation and acceleration on the basis of modified Willenborg model20 were taken into account in the analysis. The reseach results presented in Fig. 18 show the following. Crack growth time values in the lower wing surface calculated on the basis of linear hypothesis (without taking into account interaction between loads of different amplitudes) will be conservative. Inspection intervals based on these values will be economically disadvantageous. Crack growth time calculated on the basis of linear hypothesis for the upper wing surface, is several times higher than test values. Therefore crack growth time analysis should be performed with the aid of appropriate retardation/acceleration models.

Reliable Operation of Aging Airplanes Safe operation of aging (long operated) airplanes is one of the most important problems in contemporary aviation. By the present time, airplanes of many types in Russia have worked out their design goals assigned at the period of their designing. Since it is impossible to replace them with newer ones, it is necessary to elongate service goals of old machines beyond design values (Fig 19). Safety of aging aircraft is provided by:

Analytical-experimental studies of damage tolerance and fail-safe;

Fatigue testing of structures taken after long operation;

Elongation of service goal for each individual airplane.

Service goal of each individual airplane is elongated every 1 or 2 years on the basis of special permissions (conclusions) approved by three organizations: design/manufacturing company, TsAGI, and State Research Institute of Civil Aviation (GosNIIGA). Three main scientific problems are solved while providing safe operation of aging aircraft:

damage tolerance/fail-safe of structures with multi site fatigue cracks;

degradation of crack resistance and fatigue strength performance during long operation;

initiation and growth time of corrosion damages. These results are also taken into account in designing of new aircraft for high durability. The experience of solving the above problems is described below.

Research on Widespread Fatigue Damage Studies related to the problem of widespread fatigue damage (WFD), which may grow from multi site damage (MSD) or multi element damage (MED), started in the USSR in 1972 after the accident of AN-10A passenger aircraft. Generalized results of this work are presented in17, 21-23. Tests results were used to define approximate criteria of residual strength for structures with MSD, and a method for analysis of residual strength of built-up structures with WFD has been developed (Fig. 20)17. Operation of Russian airplanes with MSD is not allowed today. Design of new airplane structures is managed to exclude any probability of WFD occurrence in operation.

Research on Degradation of Crack Resistance and

Fatigue Strength Performance Fatigue tests of long operated aircraft were performed in the USSR and are now performed in Russia. Stage-by-stage extension of service goal is made with regard to these test results24. Fatigue test results of airplanes taken from operation are compared with those for new airplanes tested under the same test program. Comparison of fatigue life of principal structural elements of new airplanes and those taken from operation22 is presented in Fig. 21. The comparison shows that structural fatigue strength decreases in the course of operation. The effect is the most significant for the elements loaded with static tensile stress on the ground. In 1999 an experiment in crack growth rate was performed on specimens cut out from long operated AN-12 fuselage skin and from long operated IL-18 wing skin. Skin material was D16T. Similar specimens were cut out from sheets taken from storehouse. Specimens from AN-12 fuselage skin were 1.2 mm thick; those from IL-18 wing skin were 5 mm thick. Crack growth rate was measured on the specimens in their initial state, and after heat treatment for re-aging (described in25). Temperature of heat treatment was 400C. Thin sheets from AN-12 did not demonstrate significant difference in crack growth time between specimens in the initial state and after treatment. Crack growth times for AN-12 skin and storehouse specimens are close to each other. Crack growth time for IL-18

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skin specimens is much less than for storehouse sheets. Heat treatment greatly increased crack growth time for IL-18 skin specimens, but it did not affect that for storehouse specimens (Fig. 22). This witnesses the degradation of crack growth rate properties in thicker wing skin specimens. It should be noted that D16T alloy contains much more traces of Si and Fe than improved material of higher purity 1163T. The issue on fatigue and crack resistance performance degradation is arguable and requires further investigation.

Prevention of Corrosion Damage Aluminum alloys, which are the main structural material in contemporary aircraft structures, must have good corrosion resistance. No cracks due to stress corrosion or intergranular corrosion during long operation should occur in these alloys. Therefore corrosion-resistant alloy B95ochT2 (Al-Zn) and weldable corrosion-resistant alloy AD37 (Al-Mg) were developed26. Alongside with new alloys development, appropriate methods of structure protection against corrosion are also created. The problem of safe operation with corrosion damages is being solved on the basis of experience. Residual structure strength is determined by analysis, in which corrosion damage is replaced with equivalent fatigue crack. For the case of corrosion damage, it is recommended to provide rated residual strength with standardized damages (Figs. 2&3) in the areas where corrosion damage is probable. Corrosion damage growth time is determined by analysis using the data from operation corrosion damage sizes detected and calendar operation time of the airplane where these damages have been detected. Some special TsAGI-developed techniques of mathematical statistics27 is used for this purpose. Fig. 23 shows the example of corrosion damage examination in IL-86 fuselage skin. Up to now service goals of 40 years have been achieved by transport airplanes in Russia based on the above principles of safe structure operation (Figs. 1&19). Extending service goals up to 50 years has been planned for some airplane types.

Conclusions The system to guarantee safe operation of civil transport aircraft has been formed for 45 years in the USSR and Russia and also included foreign experience. The system has proved its efficiency. Reliable operation of aging airplanes with service goal up to 40-50 years is

provided, and new aircraft are designed for high durability on the basis of this system. The high durability of aircraft structure with minimum weight is ensured by the following:

Improvement of Airworthiness Regulations in fatigue and damage tolerance.

Design development of structures with regard to safe life, fail-safe and damage tolerance simultaneously.

Design for maximum accessibility to principal structural elements for visual inspections.

Improvement of fatigue and crack resistance performance of structural materials.

Increase of fatigue strength by means of improving the structure and joint assembly technology.

Refining alternating load spectra. Improvement of crack growth rate and residual

strength analysis for substantiation of allowable stresses.

Proof of fatigue and damage tolerance analysis by full scale certification tests with the loads most close to actual operational spectra of alternating loads.

Design of structures to exclude widespread fatigue damage probability during aircraft long operation.

Use of structural materials not susceptible to degradation of properties in the course of long operation.

Use of corrosion-resistant materials and reliable protection of structures against corrosion.

References 1. Nesterenko, G.I. Damage tolerance of aircraft

structures, Inter-University Scientific Proceedings, Issue 2, Kiev Institute of Civil Aviation Engineers (KIIGA), Kiev, 1976, pp. 60-70 (in Russian).

2. Raikher, V.L., Dubinsky, V.S., Nesterenko, G.I. and Stuchalkin, Yu.A. The features of aircraft structure fatigue resistance certification and airworthiness maintenance in contemporary conditions, Test Facilities and Aircraft Certification International Symposium, Zhukovsky, Russia, August 22-25, 1995, pp. 233-245.

3. Nesterenko, G.I., Selikhov, A.F., Using damage tolerance & fail-safe approach in design of wide-body airplanes, Aircraft Structure Strength, Moscow, Mashinostroenye 1982, pp. 151-189 (in Russian).

4. Selikhov, A.F. Major tasks and specific features of ensuring the strength of wide-body passenger

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airplanes, Aircraft Structure Strength, Moscow, Mashinostroenye , 1982, pp. 7-45 (in Russian).

5. Vorobyev, A.Z., Leibov, V.G., Olkin, B.I., Stebenev, V.N., Achievivg higher service goals of wide-body airplanes, Aircraft Structure Strength, Moscow, Mashinostroenye, 1982, pp.122-151 (in Russian).

6. Stebenev, V.N., Komarov, V.I., Gorshkov, S.V., Manyukov, V.I., Damage tolerance of structures with rivets of new types, Proceedings of Scientific Conference on Durability and Fatigue of Aviation Structures, TsAGI, 1978, pp. 115-118 (in Russian).

7. Vovnyanko, A.G., Semenets, A.I., Residual strength of built-up integrally stiffened structures made of extruded panels of D16chT and its modifications, Physiko-Chimicheskaya Mechanica Materialov, 1983, N2, Lvov, the Ukraine, pp.88-92 (in Russian).

8. Nesterenko, Grigory I. Damage tolerance of integrally stiffened and riveted stiffened structures, Proceedings of the 20th Symposium of the international Committee on Aeronautical Fatigue, 14-16 July, 1999, Bellevue, WA, USA, Volume II, pp. 873-894.

9. Vorobyev, A.Z., Olkin, B.I., Stebenev, V.N., Rodchenko, T.S., Fatigue strength of structural elements, Moscow, Mashinostroenye, 1990, 240 pp. (in Russian).

10. Swift, T. The application of fracture mechanics in the development of the DC-10 fuselage, AGARD-AG-176, V.5, 1974, pp.227-287.

11. Swift, T. Design of redundant structures, AGARD-LS-97, 1978, pp.9.1 9.23.

12. Goranson, Ulf G. Damage tolerance. Facts and fiction. 14th Plantema Memorial Lecture Presented at the 17th Symposium of ICAF, Stockholm, Sweden, June 9, 1993, 53p.

13. Fowler, Kevin, R., and Watanabe, Roy T. Development of jet transport airframe test spectra, Boeing Commercial Airplanes, Seattle, WA, USA, May 1989, 16 p.

14. Gkgl, O. Crack free and cracked life of the pressurized cabin of the A300B. Calculation , tests and design measurements to improve damage tolerance, Aeronautical Journal, 1979, V.83, No 817, pp. 1-15.

15. Spencer, M.M. The 747 fatigue and fail-safe test program, The American Society of Civil Engineers, National Structural Engineering Meeting, San Francisco, CA, 9-13/IV, 1973.

16. Boeing Structural Design and Technology Improvements. Airliner, Boeing, April-June 1996.

17. Nesterenko, Grigory I., Nesterenko Boris G., Residual strength analysis of the stiffened

structures with WFD, MSD and single crack, Proceedings of the 6th Joined FAA/DoD/NASA Conference on Aging Aircraft. San Francisco, CA, USA, September 16-19, 2002, CD-ROM Proceedings.

18. Basov, V.N., Nesterenko G.I., Strizhius V.Ye., Standardized program of heavy transport wing loading, Trudy TsAGI, Issue 2642, Moscow, 2001, pp.26-34 (in Russian).

19. Nesterenko, B.G. Analytically-experimental study of damage tolerance of aircraft structures, Proceedings of the 23rd International Congress of Aeronautical Sciences ICAS 2002, Toronto, Canada, 8 to 13 September, 2002, CD-ROM Proceedings.

20. Gallagher, J.P., Miedlar, P.C., Cross, C.W., Papp M.L. Cracks 93 system users manual, University of Dayton Research Institute, UDR-TR-93-107, 1993.

21. Nesterenko, G.I. Multiple site fatigue damages of aircraft structures, AGARD Conference Proceedings 568 (AGARD-CP-568) Widespread Fatigue Damage in Military Aircraft. Rotterdam, the Netherlands, 10-11 May 1995, pp.11-111-8.

22. Nesterenko, G.I. Fatigue and damage tolerance of aging aircraft structures, Proceedings of the FAA-NASA Symposium on the Continued Airworthiness of Aircraft Structures. Atlanta, GA, August 28-30, 1996, pp. 279-300.

23. Nesterenko, G.I. Fatigue and damage tolerance of aging aircraft structures, Proceedings of the 19th Symposium of the International Committee on Aeronautical Fatigue. June 18-20, 1997, Edinburgh, Scotland, pp. 731-742.

24. Dubinsky, V.S., Nesterenko, G.I., Raikher, V.L., Stuchalkin, Yu.A. Continued airworthiness of aircraft structures certified for fatigue, Trudy TsAGI, Issue 2631, Moscow, 1998, pp.73-75 (in Russian).

25. Schmidt, Hans-Jrgen Damage tolerance philosophy, methods and experiments applied to modern aircraft large transport aircraft structure for compliance with applicable FAA/JAA regulations, Ph.D. Thesis. A.A. Blagonravov Institute of Machine Studies of the Russian Academy of Sciences, Moscow, 2002.

26. Fridlyander, I.N., Aluminum alloys in flying vehicles in 19702000 and 20012015, Journal Metallovedenie i Termicheskaya Obrabotka Metallov (Metal Studies and Metal Heat Treatment), N16, 2001, Moscow (in Russian).

27. Senik, V.Ya., Examination of fatigue crack growth in aircraft structural elements based on operation data, Trudy TsAGI, Issue 1671, Moscow, 1975, pp.17-27 (in Russian).

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Table 1. Fatigue and damage tolerance requirements for the USSR and Russian passenger airframes

Civil aircraft airworthiness requirements (NLGS)

Publication date Service life ensuring concept

NLGS-1 1967 Safe life NLGS-2 1974 Safe life NLGS-2

Amendment 2 to part 4 1976 Safe life or fail-safe and damage tolerance

simultaneously

NLGS-3 1984 Safe life or fail-safe and damage tolerance simultaneously

Aviation regulations AP-25 1994 Fail-safe and damage tolerance simultaneously Methods of compliance assessment AP-25

MOC25.571 (advisory circular) 1996 Fail-safe and damage tolerance

simultaneously

Table 2. Stiffened wing structures Riveted Integrally stiffened Combined (center wing integrally

stiffened, outer wing riveted)

AN-10, AN-12, TU-104, TU-134, TU-154, TU-204,

IL-18, IL-96, YAK-40, YAK-42

AN-22, AN-124, AN-225, IL-62, IL-76, IL-86

AN-24, AN-26, AN-30, AN-32, AN-70, AN-72, AN-74, AN-140

Table 3. Fatigue and damage tolerance test volumes for aircraft structures New structures Structures after operation

Aircraft Full-scale aircraft

Wing Fuselage Full-scale aircraft

Wing Fuselage

AN-10 3 2 2 AN-12 2 4 AN-22 1 AN-24 1 1 3 3 2 AN-124 1 IL-18 1 3 1 3 IL-62 2 1 1 IL-76 1 2 2 IL-86 1 2 1 IL-96 1 1

TU-104 1 5 1 2 2 TU-114 1 1 1 TU-124 1 1 TU-134 1 2 2 1 1 TU-144 3 TU-154 4 1 1 TU-204 2 YAK-40 1 1 1 YAK-42 2 1

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Table 4. Stresses in the structures of wide body airplanes

Lower wing surface Upper wing surface Fuselage

Equivalent stresses, R=0

Equivalent stresses, R=0

Ulti

mat

e te

nsile

st

ress

es,

ult ,

MPa

eq

, MPa

*

eq, M

Pa

Ulti

mat

e c

ompr

es-

sive

stre

sses

ult C

OM

P, M

Pa

Ulti

mat

e te

nsile

st

ress

es

ul

t TEN

S, M

Pa

eq, M

Pa

*

eq, M

Pa

Ulti

mat

e lo

ngitu

-di

nal t

ensi

le s

tress

es

ul

t , M

Pa

Equi

vale

nt l

ongi

tu-

dina

l ten

sile

stre

sses

eq,

MPa

Hoo

p st

ress

es

Pr/t

, MPa

Fatigue

170

Fatigue

190

Fatigue

132

Fatigue

145

Fatigue

130

380

Crack growth

130

490

245

Crack growth

60

360

Crack growth

110

105

Years

1950

1955

1960

1965

1970

1975

1980

1985

1990

1995

2000

2005

0 25000 50000 75000

Service life, flight hours

IL-76

TU-334

AN- 140

IL -114

AN-74

AN-32

AN-26

AN-12

IL -96TU-204

IL-18AN-10

TU-104

TU-114 TU-124AN-24

TU-144YaK-40

TU-134

IL-62 TU-154

AN-124YaK-42

IL-86

Figure 1. Research in design goal enhancement for civil aircraft

Fatigue & crack resistance under random load spectra

Aging aircraft service lifeSoftware to calculate fatigue and damage tolerance

Degradation of material properties

Acoustic strength testing

Service life of mechanismsMulti-channel systems of quasi-random loading

Requirements to material fatigue & crack resistancePhysical methods of fatigue study

Designing for service lifeComposite materials, crack resistance

Non-destructive inspection methodsStructural damage tolerance

Analytical life estimates

Probabilistic methods of service life evaluationFatigue tests of full-scale structures

Fatigue under vibration loadsFatigue under repeated-static load

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L=2b L=2b

b

H

L=b

Simultaneous initiation of one crack per panel;crack length underneath broken stringer is equal to twobay length

Spar cap broken; spar web crack length =1/3 of webheight; skin crack is equal to one bay length.Spar web broken

One panel broken

1/3H

Figure 2. Standardized wing damages

b1

b2

L=2b =350-400 mm2

L=2b =1000 mm1

L=150 mm

L=500-1000 mm

Figure 3. Standardized fuselage damages

Cutout-initiated crack 150 mm long; skin and edge shape damaged. Skin crack in pressure bulkhead.

Transverse skin crack length = two bay lengths; stringer broken. Longitudinal skin crack length = two bay lengths frame broken.

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0,1service life

2, mm

, flights

service life

2575

5 0,5

Lstandexternal in-serviceinspection

non-destructiveinspection at factory

non-inspectedelements

0,25service life

a) The case

b) The case

Figure 5. Example of reg

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Figure 4. Requirements to crack growth time of wide spread fatigue damage of wing panel

of wide spread fatigue damage of the wing cross section

ulated damages of wing structure for the case of wide spread

ican Institute of Aeronautics and Astronautics 10

c) Fracture tougness

0 2 0 4 0 6 0 8 0 1 0 0 1 2 0 1 4 0 1 6 0 1 8 0 2 0 0

m P,K app

5

8

cyclemm ,

dNda 310

Hzf,R

,mPK32010

31==

=

5

8

Hzf,R,Pgrossmax32010

133==

=

a) Fatigue

Common material Improved material

Wing

Fuselage

0 1 0 0 2 0 0 3 0 0 4 0 0 5 0 0 6 0 0

0 100 200 300 400 500 600 N103, cycles

Sheet D16ATV t= 1,5 mm

Sheet 1163ATV t= 1,5 mm

Sheet D16ATV t= 5,0 mm

Extruded panelD16 t=5,0 mm

Extruded panel1163 t=5,0 mm

Plate 1163t= 5,0 mm

1

2

3

4

6

7

b) Crack growth rate

2

Common material

Improved material

Wing

Fuselage

0 1 2 3 4 5 6

0 1 2 3 4 5 6

Sheet D16ATV t= 1,5 mm

Sheet 1163ATV t= 1,5 mm

Sheet D16ATV t= 5,0 mm

Extruded panelD16 t=8,0 mmExtruded panel

1163 t= 8,0 mmPlate 1163

t= 8,0 mm1

2

3

4

6

7

c) Fracture tougness

Fuselage

0 20 40 60 80 100 120 140 160 180 200

Common material Improved material

Wing

1

2

3

4

5

6

7

8Sheet D16ATV t= 1,5 mm

Sheet 1163ATV t= 1,5 mm

Sheet D16ATV t= 5,0 mm

Extruded panel D16 t=8,0 mm

Extruded panel1163 t= 8,0 mm

Plate 1163 t= 8,0 mm

Figure 6. Fatigue and crack resistance of common and improved Al-alloys

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0 50000 100000 150000 200000N, cycles

Sheet D16V

Fracture

Fracture.R,P

tPr 0110 ==

Figure 7. Effect of joint assembly technology on fatigue resistance in longitudinal lap joints of fuselage skin

AN-124, integrally stiffened extruded panels IL-96-300, riveted panels

Panel joint Panel joint

Figure 8. Wing lower panel layouts

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104 105

B-747

AN-124AN-22

IL-86

YAK-42

IL-76M

IL-96

B-767 B-777

TU-154

IL-86

TU-204 AN-24

IL-62 d

W

Extruded panelD16 chT

K t netto = 2.6

61

=Wd

Tmin

T, flights

*eq , MPa

Figure 9. Fatigue of the longitudinal panel joints of the lower wing surface. Full-scale structure tests

Service goal TS.G. Tmin Service goal TS.G. Tmin

Pr/t, MPa

2 4 6 8 2 4 6104 105 40

60

80

100

120

140

B-747-100SR B-747-400

IL-96B-737

A-300BIL-76

IL-86 TU-154

IL-86

YAK-42TU-134

AN-24

IL-18TU-124TU-104

TU-204IL-62

TU-114

6 8

YAK-40

T, pressurizations

d

W

Extruded panelD16 chT

K t netto = 2.6

6

1

w

d=

Tmin

Service goal TS.G. Tmin

Figure 10. Fatigue of the longitudinal skin joints of the pressurized fuselages. Full-scale structure tests

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F

0

1

3

, P

3,

b=130 mm

t=8 mm2a

1, 2

300

200

280

260

240

220

180

D16chT material. F / F =0.9-1.0str skin

Figure 11. Residual str

, P

1, 2, 3, 4, 5

tott

tott

Figure 12. Re

D16ch=2

D16=

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F/D16chT material,

= 0.9 1.0 skinstr

0.5 1.0 1.5 2.0 2.5

2

4

2 b

4 b=130 mm

2at=8 mm

ength of integrally stiffened and riveted wing panels made of improved alloys

1 2

34

5

7

6

6, 7

skinstr FF /

skinstr FF /

sidual strength of integrally stiffened and riveted fuselage panels

T material, b=130 mm, .2 mm, = 0.35

chT material, b=130 mm, 3.2 mm, = 0.25

merican Institute of Aeronautics and Astronautics 14

= Pr/t, MPa

0

20

40

60

80

100

120

140

0 250 500 750 1 000 2a0, mm

m

2a

Skin D16ATV, KAPP = 135 MPa Frame D16T , TU = 460 MPa

compound frame with stopper

compound frame

single-element frame

unstiffened cylinder

Figure 13. Residual strength of the pressurized fuselage having longitudinal crack in skin under the broken frame

0

100

200

300

400

0 1 000 2 000 3 000 4 000T, flights

2a, mm

m

eq = 130 MPa; D16chT alloy ; da/dN =0.0025 mm/cycle for K = 31 MPa ;Fstr/Fskin=0.9-1.0

Figure 14. Crack growth duration in the skin under the broken stringer in the lower wing surface

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0

100

200

300

400

500

600

700

0 1000 2000 3000 4000

2a, mm

m

2a

Pr/t = 110 MPa; skin D16ATV alloy da /dN = 0.002 mm/cycle for K=31 MPa

T, flights

Figure 15. Longitudinal crack growth duration in the fuselage skin

0

100

200

300

0 1000 2000 3000

2a, mm

meq = 110 MPa; skin D16ATV alloy da/dN = 0.002 for K = 31 MPa

2

T, flights

Figure 16. Transversal crack growth duration in the skin under the broken stringer in the fuselage

American Institute of Aeronautics and Astronautics

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mMPaK,

250 300 350

0

20

40

60

80

100

120

140

160

180

200 CaK eff= =265MPa

2

1

bt

KR-curve 2

Mb

250 300 350

120

160

200

240

280

320

=

TU

MPa,

12 Analsis

Figure 17. Residual strength analysis of

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aterial D16AT, Fstr/Fskin = 0.35 =170 mm; t =2.2 mm 2080400 450 500 2aeff , mm

skin

400 450 5002aeff , mm

400=

str

1 2 =220 MPa experimental data on wide body airplane IL-86 panel test

R - curve

stiffened structures using R-curves for skin material

nautics and Astronautics 7

a) Lower wing surface, 2324-T39 alloy

0

20

40

60

80

0 5000 10000 15000 20000 25000 30000 35000 40000N, flights

Cra

ck le

ngth

2a

, mm

-0.2

0.4

0

Linear model analysis

Test results

Modified Willenborg model analysis

b) Upper wing surface, 7075-T77 alloy

0

20

40

60

80

0 100000 200000 300000 400000 500000 600000N, flights

Cra

ck le

ngth

2a,

mm

0

-0.2

-0.4

-0.6

0.2

Modified Willenborg model analysis

Linear model analysis

Test results

Figure 18. Crack growth rate analysis in wide-body aircraft wing under typical loading spectrum

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0

0,5

1

1,5

2

2,5

3

YAK-40 IL-62 TU-134 TU-154B AN-12 AN-24

Serv

ice

goal

(life

tim

e) o

f lea

der-

airc

raft

/Des

ign

goa

l

flight Flight hours Years

Figure 19. Ensuring service goal of long operated civil aircraft

Wing cross section of the airplane AN-10A # 11222

Wing cross section of the airplane AN-10A # 11202rear spar

str. 8 str. 4 str. 1

front panel

Stringers #8,#7,#6,#5 broken

00

0.2 0.4 0.6 0.8 1.0

0.2

0.4

0.6

0.8

1.0

#11222

Stringers #4,#3,#2,#1 broken

F /Fdam 0

P /Pfrac 0

Figure 20. Comparison of the analytical and test values of the residual strength of the AN-10A airplane wing with widespread fatigue damages (WFD)

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0,2

0,4

0,6

0,8

1

AN-24wing

TU-134Awing

TU-104body

AN-12wing

YAK-40wing

IL-18wing

IL-62wing

G

NNN oper

0

+=G

N0 - new structure life; Noper operational lifetime; N life after operation

New structures

Figure 21. Relative fatigue strength of aircraft structures before (light plus dark zones) and after (dark zones) service

30

40

50

60

70

80

0 1000 2000 3000 4000 5000N, cycles

D16ATB, max =133 MPa, R=0.01, f=1 Hz

Store-house sheet

Store-house sheet after heat treatment

2, mm

IL-18 wing skin after heat treatment

IL-18 wing skin

Figure 22. Effect of heat treatment (annealing) on crack growth rates in old (being operated) and new (from store-house) wing skin sheets

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0

0,2

0,4

0,6

0,8

1

1,2

1,4

1,6

1,8

2

4 5 6 7

Cor

rosi

on d

epth

, mm

probability p=0.001 p=0.05 p=0.5

1

Figure 23. Corrosion damag

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2

8 9 10 11 12 13 14Service life, year

e growth rate analysis in IL-86 fuselage skin

eronautics and Astronautics 21