Computational and Experimental Study of Fatigue Crack Growth Duration in a Stiffened Fuselage Panel

Aeronautical and Space-Rocket Engineering


Аuthors

Kim Y. L.*, Pegov D. M.**, Sviridov A. A.***

Central Aerohydrodynamic Institute named after N.E. Zhukovsky (TsAGI), 1, Zhukovsky str., Zhukovsky, Moscow Region, 140180, Russia

*e-mail: julia.kim2202@gmail.com
**e-mail: pegov.dm@phystech.edu
***e-mail: sviridov98@yandex.ru

Abstract

Ensuring operational survivability and safety of modern aircraft is a paramount challenge in aerospace engineering. Historical experience, including the catastrophic failures of the An-10A and Boeing-737 aircraft, has dictated a global shift from the “safe-life” concept to the principles of “fail-safe characteristic” and “damage acceptability”. In the context of the design service lives increasing and transition to the condition-based maintenance, the development of reliable computational and experimental methodologies for the fatigue crack growth (FCG) predicting in primary airframe structures, such as stiffened fuselage panels, is of critical importance. The article presents a comprehensive technique for the duration of through-thickness fatigue crack growth computing in a stiffened fuselage skin, explicitly accounting for the skin bulging effect, internal cabin pressure, and the presence of bonded crack stoppers.
The classical Forman equation is being employed to describe the fracture kinetics. This model is selected due to its effective description of the crack behavior in the high stress intensity factor (SIF) area and accounts for the load cycle asymmetry. Computations utilize the fracture toughness characteristics of the domestic 1163AT aluminum alloy.
The key challenge in computing the crack growth in stiffened panels is accurate SIF determination, considering actual geometry of the structure. Application of analytical solutions for an infinite plate leads to significant errors. In the presented work, the SIF is being determined through a combined correction factor that integrates the geometric impact of the stiffeners and the bulging effect of the curved shell under internal pressure. To create a generalized correction applicable to a wide range of panel configurations, the stiffener impact is being approximated by a polynomial function based on parametric data.
Furthermore, the study introduces additional corrective coefficients, kred and kstop, directly into the SIF formulation. The coefficient kred quantitatively assesses the impact of the adhesive joint between the skin and the central frame doubler. Based on experimental observations, it is assumed that adhesive failure occurs at the crack half-length of 200 mm., Until this moment the back-up restricts the crack opening; after failure, the crack growth rate increases substantially. The coefficient kstop models the crack growth deceleration caused by the edge frame stoppers as the crack tip approaches them.
The proposed computational method was validated against the experimental data obtained for various damage configurations in stiffened panels. The calculations were being performed for the two typical scenarios: a crack located over an intact frame and a crack located over a failed frame. The results demonstrate acceptable convergence with the experimental data. Notably, for the case of a crack over an intact frame, the deviation between the calculated and experimental data does not exceed 1.5–4%, confirming accurate determination of the effective SIF value. The study concludes that the proposed method allows for the crack growth duration assessment in stiffened fuselage panels with acceptable accuracy.

Keywords:

stiffened panel, pressurized fuselage, crack growth life, adhesive joint, crack stopper, skin buckling

References

  1. Methods for determining compliance (MOS) with aviation regulation AP 25.571 “Ensuring structural safety under conditions of strength during long-term operation”, 1996. 29 p. (In Russ.).
  2.  Dotsenko AM, Vorob'ev AZ. The development of fatigue cracks. Journal Factory Laboratory. 1970(6):714-717. (In Russ.).
  3.  Nesterenko BG, Nesterenko GI. Ging aircraft structures life time. Civil Aviation High Technologies (Nauchnyi Vestnik MGTU GA). 2014(199):11-22. (In Russ.).
  4.  Paris P, Erdogan F. A critical analysis of crack propagation laws. Journal of Basic Engineering, Transactions of the American Society of Mechanical Engineers. 1963;85(4):528-534. DOI: 10.1115/1.3656900 
  5.  Walker K. The effect of stress ratio during crack propagation and fatigue for 2024-T3 and 7055-T6 aluminium. In: Rosenfeld MS. Effects of Environment and Complex Load History on Fatigue Life: Special Technical Publication 462. Philadelphia: American Society for Testing and Materials; 1970. p. 1-14.  DOI: 10.1520/STP32032S
  6.  Forman RG, Kearney VE, Engle RM. Numerical Analysis of Crack Propagation in a Cyclic-Loaded Structure. Journal Basic Engineering. 1067;89(3):459-463. DOI: 10.1115/1.3609637
  7.  Colliprist JE. An Experimentalist’s View of the Surface Flaw Problem. In: Swedlow JL. (ed) The Surface Crack: Physical Problems and Computational Solutions. ASME Publ.; 1972. p. 43-62.
  8.  Forman RG, Mettu SR. Behavior of Surface and Corner Cracks Subjected to Tensile and Bending Loads in Ti-6Al-4V Alloy. NASA Technical Reports Server (NTRS), 1990.  68 p.
  9.  Nikonov VV, Shapkhin VS. Using models of Paris-Ardogan for prediction of aircraft structures failure. Nauchnyi vestnik MGTU GA. 2010(161):15–23. (In Russ.).
  10.  Kim AS. Investigation of crack growth propagation in constructions of russian aircrafts by using NASGRO equation. Nauchnyi vestnik GosNII GA. 2016(12):103-112. (In Russ.).
  11.  Harter JA. AFGROW user guide and technical manual. Wright-Patterson AFB: Air Force Research Laboratory, 1999. 191 p. DOI: 10.1111/j.1460-2695.2010.01476.x
  12.  Chakherlou TN, Mirzajanzadeh M, Saeedi Kh. Fatigue crack growth and life prediction of a single fitted hole plate. Fatigue & Fracture of Engineering Materials & structures. 2010;33(10):633–644.
  13.  Kim AS, Borisov MP. Investigation of the effect of bulging of the fuselage skin on the growth rate of a longitudinal crack under a whole/destroyed frame and between frames. Sbornik statei nauchno-tekhnicheskoi konferentsii “Prochnost' konstruktsii letatel'nykh apparato”. Ser. “Trudy TSAGI Issue 2782. Zhukovskii: TSAGI; 2018. p. 45-48. (In Russ.).
  14.  Sysoyeva IV. The calculations of stress intensity factors for typical aviation structures with cracks. Trudy MAI. 2011(45). (In Russ.). URL: https://trudymai.ru/eng/published.php?ID=25559&PAGEN_2=2
  15.  Dement'ev AD. Calculation of stress intensity coefficients at the tip of a through crack using a specialized hybrid finite element. Uchenye zapiski TSAGI. 1987;18(6):84-91. (In Russ.).
  16.  Leontiev MK., Lanevskiy TM. Calculation of crack stress intensity factor in high pressure turbine disk. Aerospace MAI Journal. 2012;19(3):121-127. (In Russ.).
  17.  Tumanov NV. Physical and mechanical aspects of stable fatigue crack growth. Aerospace MAI Journal. 2011;18(2):132-136. (In Russ.).
  18.  Kartamyshev AI, Yablonskii IS. (eds) Design characteristics of structural aircraft materials: Handbook. Issue 1. Zhukovsky: TsAGI; 1987. 215 p.
  19.  Rooke DP, Cartwright DJ. Compendium of Stress Intensity Factors. London: Her Majesty’s Stationery Office; 1976. 330 p.
  20.  Chen D. Bulging of fatigue cracks in a pressurized aircraft fuselage. Doctoral Thesis. Delft University of Technology, Faculty of Aerospace Engineering, 1991.
  21.  Kim AS. Methods for calculating the residual strength and duration of crack growth in the wing and fuselage skin. PhD thesis. Zhukovsky: TSAGI; 2022. 173 p. (In Russ.).

mai.ru — informational site of MAI

Copyright © 1994-2026 by MAI