Structural Failure Analysis and Prediction Methods for Aerospace Vehicles and Structures

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A new fatigue lifing approach has been developed by the DSTO for high performance metallic airframe structures. This approach is based on years of inspection and analysis of fatigue cracks in many airframe components and specimens, and is now an important additional method of determining fatigue lives for aircraft in the Royal Australian Air Force fleet. Like the earlier Damage Tolerance approach developed by the United States Air Force, the DSTO approach assumes that fatigue cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. The Damage Tolerance approach specifies a "standard" set of initial crack/flaw sizes based on Non-Destructive Inspection capabilities, and the important period of early (short) crack growth is estimated from back-extrapolation of long crack growth data. On the other hand, the DSTO approach uses initial crack/flaw sizes representative of small, fatigue-initiating discontinuities in the materials and structural components, and the crack growth lives are estimated from actual data for short-to-long cracks growing from these discontinuities. Furthermore, these data, particularly for lead cracks, are characterized by exponential crack growth behaviour. Owing to this general characteristic, the DSTO approach can use growth data for the lead crack to provide reasonable lower-bound estimates of component crack growth lives. Scatter factors based on engineering judgement are then applied to these estimates to determine the maximum allowable service life.

Acoustic fatigue is the major damage phenomenon induced by the high frequency lateral vibration of structural panels, such as that of an aircraft skin, under time varying pressure waves generated by engine and/or aerodynamic effects. For instance, acoustically-induced cracks have been discovered in the lower external surface of the nacelle skin and aft fuselage of the F/A-18 aircraft. In the case of the inlet nacelle overall sound pressure levels of the order of 172 dB have been recorded. Attempts to repair these cracks by applying standard bonded repairs developed for in-plane loads were made. However the cracks continued to grow at a similar rate as before the application of repairs. While the repair of cracked aircraft structures subjected to in-plane loads using bonded repairs has resulted in considerable aircraft life extension and hence cost savings, the use of bonded patches to repair panels with acoustically-induced cracks (acoustic fatigue) is only recent. In this chapter a generic design procedure is presented for the repair of aircraft panels containing acoustically induced cracks by incorporating the constrained layer damping technique. The analytical tools described in this chapter will enable the rapid design of damped repairs using closed form solutions that account for the effects of high frequency out-of-plane vibration on crack extension. A case study is also undertaken of a design of a repair to prevent acoustic fatigue cracking on the aft fuselage of the F/A-18.

This chapter gives a brief overview of the state-of-the-art and insight into the key aerothermal issues and challenges for aero-thermal-structural analysis and failure assessment of modern high-speed aerospace vehicles. Approximate engineering methods and computational methods to numerically simulate and understand the physics of high-speed flows surrounding the vehicle and hence, predict aerodynamic heating, the pressure and shear forces, and also the structural response to the aeroheating, aerodynamic and aeroacoustic loads are discussed in the context of the concerns of the aerospace vehicle designer and service life / structural failure analyst. The discussion is focused on those predictive methodologies that can be readily implemented, adequately robust and suitable for somewhat more complex vehicle designs and missions. Example applications, including a real-life application, are presented. Future directions, in particular the analytical / modelling challenges in the use of smart structures (e.g., SMAs for structural response / shape control), active cooling concepts and exotic composite materials for high-speed / performance aerospace vehicles are also discussed.

In aerospace vehicle design, structures always contain a range of geometric variations including various notches, holes and cutouts to achieve certain mechanical functions as well as weight savings. Those notches and holes can cause local stress concentrations and/or create hot spots for fatigue crack initiation and propagation. This chapter presents an analytical study of fatigue crack growth in aluminium alloy 7050-T7451 notched specimens under a fighter aircraft wing root bending moment spectrum. The crack growth data were measured by quantitative fractography for three groups of specimens with different stress concentration geometrical features. Under spectrum loading and for each spectrum peak stress level, a minimum of five specimens were tested. Based on the analysis of the measured spectrum crack growth data using linear elastic fracture mechanics, it was found that the concept of geometry factors formulated in the stress intensity factor could not collapse the crack growth rate data derived from each stress concentration feature, particularly near the small crack growth region. In order to investigate the possible reasons for this, three-dimensional elastic-plastic finite element analysis was used to determine notch plastic zone sizes for each stress concentration geometry. As a consequence, an alternative crack growth driving force which considered both notch elastic-plastic stress field and gross net-section stress field was used to interpret the fatigue crack growth data under spectrum loading. It was found that the predictions of crack growth under spectrum loading for different stress concentration factors at different peak load levels agreed reasonably well with the experimental results.

The work presented here is a compendium of theoretical results obtained by the authors between 2005 and 2009. Among these results are comprehensive analysis of the three-dimensional elastic stress and displacement fields near a tip of a through-the-thickness crack, generalization of the classical strip-yield model for plates having a finite thickness, and development of an analytical approach for calculating the plasticityinduced crack closure and crack growth rates at constant and variable amplitude loading. As an application of the developed approach, new predictive models of various non-linear fatigue crack growth phenomena in plates of finite thickness were developed. These include computational models of crack growth under small-scale yielding conditions and constant amplitude loading, growth of a fatigue crack emanating from a sharp notch, and crack growth retardation phenomenon following an overload cycle. All theoretical predictions were extensively compared with previous numerical and experimental studies demonstrating a great potential of the refined plate theory in the analysis of fracture and fatigue problems.

This chapter describes miniature stress sensor technology for monitoring the thermal stresses and ignition pressurization loads in solid rocket motors. The study was part of a larger international (TTCP) collaborative effort carried out from 1988 to 2002 to validate the instrumentation and analytical stress analysis and service life prediction methodologies for solid composite rocket motors, and thus establish improved, more reliable, cheaper and non-destructive capabilities for service life prediction and extension. Different motor configurations were used by the different countries in this collaborative program. The Australian effort, described in this chapter, used an end burning generic research motor (Pictor). The embedded transducers, in this end burning motor and in the different motor designs used by the other TTCP countries, were found to be stable in the temperature range used in the environmental testing program and gave consistent data during propellant cure, environmental testing and static firing of the motors. The rocket motor instrumentation and data reduction techniques were described. The data from the instrumented motors under various thermal storage loading conditions (multiple thermal cycling, shocking, accelerated ageing at elevated temperature) were used to validate the stresses and critical failure modes predicted by structural finite element modelling and a modified fracture mechanics approach for nonlinear viscoelastic materials. These studies verified the ability of the miniature bond stress sensors to detect cracking / damage in the propellant charge. The advancement in bond stress sensor technology was further used to investigate failure analysis of rocket motors under ignition pressurization conditions. Results from these studies demonstrated that the sensors are safe for static firing and could accurately measure pressure in different regions of the burning motor. The stress sensor data from this international collaborative program showed that the stress sensor technology could be used for real-time structural health monitoring of solid rocket motors to detect cracks and debonds in the propellant and to continuously monitor the extent of damage. The results from the instrumented Pictor motor verified and validated the thermal distributions, stress / strain states and regions of high propensity for crack propagation predicted by finite element modelling and fracture mechanics. The use of the instrumented motor data in probabilistic service life prediction methodologies and other NDE methods are also discussed.

This chapter gives an overview of the main uses and features of structural integrity assessment (SIA) of solid rockets. Details of the key elements and prediction techniques for analyzing the stress and strain response, failure criteria and the service environments in which solid rockets are required to operate are given. The development and use of these methods are essential to deal with the increasingly more stringent requirements in performance, safety, reliability and cost. The role and extent to which SIA is employed in design and development, service life prediction and vulnerability / safety assessment of solid rockets are discussed. The development of appropriate nonlinear constitutive models (temperature and strain-rate dependent and incorporates damage) and their implementation in finite element codes, suitable failure criteria (e.g. based on facture mechanics), and service life prediction methodologies are discussed, and an example solution procedure is provided for a real-life rocket. A method for determining high strain-rate mechanical properties, using the Hopkinson Bar technique, to assess the structural response of rocket component materials to impact loading conditions is described.

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