The effect of laser power density on the fatigue life of laser-shock-peened 7050 aluminium alloy

Laser shock peening (LSP) is an innovative surface treatment method that can result in significant improvement in the fatigue life of many metallic components. The process produces very little or no surface profile modification while producing a considerably deeper compressive residual stress layer than traditional shot peening operations. The work discussed here was designed to: (a) quantify the fatigue life improvement achieved by LSP in a typical high strength aircraft aluminium alloy and (b) identify any technological risks associated with its use. It is shown that when LSP conditions are optimal for the material and specimen configuration, a —three to four times increase in fatigue life over the as-machined specimens could be achieved for a representative fighter aircraft loading spectrum when applied at a representative load level. However, if the process parameters are not optimal for the material investigated here, fatigue lives of LSP treated specimens may be reduced instead of increased due to the occurrence of internal cracking. This paper details the effect of laser power density on fatigue life of 7050-T7451 aluminium alloy by experimental and numerical analysis.     

Typical fatigue-initiating discontinuities in metallic aircraft structures

A fatigue lifing framework using a lead crack concept, based on years of detailed inspection and analysis of fatigue cracks in many specimens and airframe components, has been developed by the DSTO for metallic primary airframe components. This framework is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force fleet. Like the original Damage Tolerance concept, developed by the United States Air Force, this framework assumes that fatigue cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial crack sizes and deriving early crack growth behaviour from back-extrapolation of growth data for long cracks, the framework uses data for real cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue cracks in typical metallic airframe structures. These discontinuities and the fatigue cracks that have grown from them are taken from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.     

A fatigue damage accumulation model for reliability analysis of engine components under combined cycle loadings

Rotor components of an aircraft engine in service are usually subjected to combined high and low cycle fatigue (CCF) loadings. In this work, combining with the load spectrum of CCF, a modified damage accumulation model for CCF life prediction of turbine blades is first put forward to take into account the effects of load consequence and load interaction caused by high‐cycle fatigue (HCF) loads and low‐cycle fatigue (LCF) loads under CCF loading conditions. The predicted results demonstrate that the proposed model presents a higher prediction accuracy than Miner, Manson‐Halford model does. Moreover, to evaluate the fatigue reliability of rotor components, reliability model with the failure mode of CCF is proposed on the basis of the stress‐strength interference method when considering the strength degeneration, and its results show that the reliability model with CCF is more suitable for aero‐engine components than that with the failure mode of single fatigue.     

A high-cycle fatigue apparatus at 20 kHz for low-cycle fatigue/high-cycle fatigue interaction testing

High-cycle fatigue (HCF) failures in aircraft engines are attributed to material damage states, created during processing or by in-service loading and environmental conditions, and then propagated to failure by HCF loading. The loading configuration experienced by aircraft engine turbine blades consists of an axial load caused by the centrifugal acceleration during rotation combined with the tensile and compressive loads caused by the natural vibrations of the blades themselves. To simulate these loading conditions a new testing apparatus was developed that is capable of providing interactive low-cycle fatigue/high-cycle fatigue (LCF/HCF) loading, in ratios (of magnitude and frequency) that give a realistic simulation of the actual flight loads experienced by engine components. This testing apparatus is based on a HCF cell operating at 20  kHz. The cell can also be integrated to a servo-hydraulic load frame, which provides a second fatigue cycle. The objective of this study was to demonstrate the capabilities of the new HCF apparatus via thermographic measurements and by performing LCF/HCF interaction tests.     

Improving the fretting performance of aero-engine tenon joint materials using surface strengthening

This paper summarises the regulation mechanism of the traditional and emerging surface strengthening treatments – mechanical shot peening (SP) and laser shock processing (LSP) treatment, respectively – on the fretting fatigue behaviour of the tenon joint materials of aero-engine cold-end parts, as surface strengthening treatment contributes to the improvement of service performance. From the point of view of the actual service environment of turbine components, this paper also expounds on the key problems faced by surface strengthening treatments in improving the mechanical properties of joint materials for aero-engine hot-end components. The application of LSP in the fretting fatigue life extension of turbine joint materials has strong feasibility and development prospects.     

Vibration Modal Analysis of Compressor Blade based on ANSYS

Compressor is an important part of aero engine. In the environment of high temperature and high pressure,compressor blade will suffer from several physical and chemical processes,such as centrifugal force,aerodynamic force vibration and oxidation. These processes will lead compressor blade to fatigue fracture,and at the same time,make negative effects on the engine' s overall performance. Based on the software ANSYS15. 0,we made strength analysis and modal analysis of compressor blade in this paper. As a result,we got its natural frequencies,relevant modal parameters and vibration mode cloud pictures. After analyzing the influence that centrifugal force made on modal parameters,we predicted the expected damage of the blade. Eventually the analysis results will provide the basis for overall performance evaluation,structural crack detection,fatigue life estimation and strength calculation of aircraft engine compressor.     

Fatigue Fracture of a Compressor Disc of an Aeroengine

There was an accident to a single engine aircraft. From the eye and ear witness accounts, it was established that the accident occurred because of engine failure. After preliminary examination of the wreckage of the crashed engine, a few suspected components were identified for detailed laboratory investigation. The objective was to establish the primary failure in the engine. While majority of the engine components submitted for laboratory analysis showed secondary damages due to either crash impact forces or post-accident fire, the fracture pattern in one of the compressor disks was different from other components of the engine. Fractographic study revealed that the failure of the disk was by fatigue mechanism. Subsequent investigation showed that the fatigue fracture of the compressor disk was the first in the chain of events that led to the engine failure. This finding was further substantiated through fracture mechanics calculations.     

THE FATIGUE RESISTANCE OF PEENED 7050-T7451 ALUMINIUM ALLOY—REPAIR AND RE-TREATMENT OF A COMPONENT SURFACE

Abstract— Metal and glass-bead peening treatments, widely used throughout the aircraft industry to enhance the fatigue performance of many steels and titanium alloys, are now being routinely applied to high-strength aluminium-alloy components. This paper describes the effects of peening on the fatigue life of 7050 aluminium alloy material, which is representative of alloys used for many components in modern military aircraft. Using simulated service loading, two proposed peening/re-peening procedures were evaluated and compared with both the original peened surface and a simple hand-polished surface. The results show that optimisation of peening parameters to reduce surface damage can provide a substantial improvement in fatigue life over both the original peening treatment and the polished surface treatment, however, poor control of peening procedures, or unnecessary "overpeening" can lead to a relatively poor fatigue life. For re-peened surfaces, a procedure incorporating a polishing step, designed to repair any damage from the severe peening applied initially, gave the best fatigue performance. Results are discussed in relation to the stability of the residual surface stresses under fatigue loading, the surface roughness, and the number and types of defects introduced by the peening treatments.     

Fracture analysis of turbine disks and computational–experimental background of the operational decisions

In this paper, fatigue crack growth under operation conditions for rotating disks of aircraft gas turbine engines is analyzed. Initiation and growth of surface cracks for compressor disks made from two-phase titanium alloy has occurred in a disk and blade attachment. Damage accumulation and growth for turbine disks made from steel took place on the inner surface of hole in a hub of wheel. Suggested approach of simulation modeling is used for an analysis and prevention of operation failures of engine rotating components. In the approach described, finite-element models (FEMs) in two and three dimensions were applied to the study of stress–strain state and stress intensity factors for the basic configurations of compressor and turbine disks and their operational damage. Proposed design modifications and repair technologies to existing in-service aircraft gas-turbine engine rotating components are analyzed and substantiated on a static strength and fatigue life basis.     

Residual life prediction for healing fatigue damaged copper film by laser shock peening

A healing method for fatigue damage was studied by laser shock peening (LSP) with excimer laser for polycrystalline copper film. It is found that work hardening due to LSP could be responsible for the improvement of residual fatigue lives for the damaged and undamaged specimens by LSP, and the hardening degree for the damaged specimen by LSP is obviously higher than that for the undamaged specimen by LSP. In this paper, two basic mechanisms were identified. One is the dissipated energy enhancement mechanism, which improves the fatigue life caused by laser shock stress, and the other is the healing mechanism, which leads to a further improvement. Based on the two mechanisms, a residual fatigue life prediction method is proposed by the view of energy consumption before and after LSP. The predicted lives by the proposed method agree well with the experimental results.     

Investigation of the fatigue life of pre- and post-drilling hole in dog-bone specimen subjected to laser shot peening

The influence of processing sequence of laser shot peening (LSP) on the fatigue properties of fastener hole was investigated with finite element method and experiments. The results show that different processing sequences lead to different residual stress distributions and different fatigue lives. The compressive residual stresses (CRS) are squeezed into two-sided surface layers of fastener hole by two sided laser shot peening, and the ellipse CRS fields are found on both sided surfaces of sample. However, when the pre-drilling hole in dog-bone specimen is subjected to LSP, the tensile stresses appear at its mid-thickness region, while the CRS distribute in the entire thickness region of the post-drilling hole after LSP. The fatigue crack initiation of specimens treated by LSP stems from the subsurface layer of hole edge. The fatigue striation spacing of specimen with post-drilling hole after LSP is narrower in comparison with that of case with pre-drilling hole before LSP. The fatigue life of post-drilling hole is longer than that of the pre-drilling hole.     

Laser shock processing of aluminium alloys. Application to high cycle fatigue behaviour

Subjecting target metallic samples to a very short pulse (about 20 ns) of intense (GW cm⁻²) laser light generates, through a surface plasma, a high-pressure stress wave propagating to the first millimetre in depth, which is commonly called laser shock processing (LSP). The purpose of this work was to evaluate the role of this novel process on the cyclic properties of A356, Al12Si and 7075 aluminium alloys. Major contributors to the fatigue performance improvements were investigated in order to determine the optimum shock conditions. These were mainly compressive residual stress (RS) levels for which a large range of incident shock conditions was performed. We showed that stress levels were very sensitive to the laser fluence and the number of local impacts, and experimental RS measurements were found to be in good agreement with analytical modelling results. In comparison, a conventional shot peening (SP) treatment was found to lead to higher surface hardening and RS levels, but with a very detrimental roughening not observed after LSP. High cycle (10⁷) fatigue tests carried out on laser- processed, shot-peened and untreated notched samples illustrated the efficiency of LSP as a new, promising method to improve the fatigue limits σ_D of structures, especially in comparison with enhancements displayed by SP (+22% vs. +10%). According to crack detection electric measurements, fatigue performance improvements with LSP mainly occurred during the crack initiation stage.     

PROTECTIVE COATINGS FOR AERO ENGINE HOT SECTION COMPONENTS

Aluminide coatings have been widely used in the aircraft industries for the protection of gas turbine engine hot section components against oxidation and/or hot corrosion. This paper considers modes of coating degradation under conditions of cyclic oxidation, hot corrosion and corrosion-erosion interactions during service, as well as the effects of interdiffusion between coating and substrate alloys either during service or coating application. It also discusses means of improving existing coatings as well as advanced coating systems currently under development. In assessing coating performance, consideration must be given to the influence that coatings may have on substrate properties such as mechanical strength, resistance to creep and resistance to mechanical and thermal fatigue. Finally, it is stressed that proven performance for a given coating/substrate combination is no guarantee that no deleterious reaction will occur when the same coating is used with a different substrate alloy. Therefore, coating substitution requires requalification.     

飞机发动机尾喷口调节片断口的电子显微分析

利用扫描电子显微镜及附带的能谱分析仪对飞机发动机尾喷口调节片断口进行了微观分析研究。结果表明:调节片边缘的损伤特征以磨损为主,而中部则以接触疲劳为主,疲劳裂纹易在微动区产生。根据上述结果讨论了促使疲劳裂纹萌生的因素,即循环应力和摩擦力引起材料表层塑性变形,以及磨损破坏了材料表面的完整性,造成裂纹尖端应力集中效应。     

Failure analysis of UHPFRC panels subjected to aircraft engine model impact

Using the finite element approach, this paper evaluates the punching resistance of ultra-high performance fiber reinforced concrete (UHPFRC) panels subjected to the impact of an aircraft engine. The models are analyzed using LS-DYNA, a commercially available software program. The structural components of the UHPFRC panels, aircraft engine model, and their contacts are fully modeled. Included in the analysis is material nonlinearity, which considers damage and failure. The analysis results are then verified with the test results. A parametric study with varying fiber contents is carried out to investigate the punching behavior of the UHPFRC panels under aircraft engine impact. The penetration depth, residual velocity of the aircraft engine, scabbing area, and failure mode of various UHPFRC panels are examined. Punching resistance capacities of reinforced concrete (RC) and steel plated concrete (SC) panels are also investigated in this study.     

Foreign object damage and fatigue crack threshold: Cracking outside shallow indents

Foreign Object Damage (FOD) usually happens when objects are ingested into jet engines powering military or civil aircraft. Under extreme conditions, FOD can lead to severe structural damage. More commonly it produces local impacted sites of the fan and compressor airfoils, lowering fatigue life of these components. FOD is a prime cause for maintenance and repair in aircraft engines. In this paper, a framework for analyzing FOD and its effect on fatigue cracking is established. A finite element analysis is used to identify three relevant regimes of FOD related to the depth of penetration into the substrate, and to determine the residual stresses. Most of the emphasis in this paper focuses on fatigue cracks emerging from shallow indentations, which are generally expected to be of most practical concern. Full three-dimensional finite element solutions are obtained for semi-circular surface cracks emerging from specific locations at the indentation revealing the influence of the residual stress on the stress intensity factor distribution. For shallow indents, a relatively simple dimensionless formula for the relation between the residual stress intensity factor, the crack size, and the indentation width are developed. These results, together with results for the intensity factor variations due to cyclic loading, have been used to address the question: To what extent do the residual stresses caused by the FOD reduce the critical crack size associated with threshold fatigue crack growth? Formulas for the critical crack size are obtained. Specific results are presented for the blade alloy, Ti-6Al-4V, revealing that FOD can reduce the critical crack size by as much as 60%.     

Investigation of In-Flight Shutdown of Turboprop Engine Due to Electrical Discharge Damage

A turboprop training aircraft experienced an in-flight shutdown failure with complete seizure of its propeller. Disassembly of the mishap engine revealed that many of the engine components were severely damaged. The laboratory investigation of the failed engine components determined that mechanical failure of the driveshaft bearing in the gearbox was the principal contributing factor that led to in-flight complete seizure of the propeller shaft. Microscopic examination of the failed bearing remnants found electrical arc-induced pittings which played a role as crack initiation sites resulting in premature rolling contact fatigue cracking during continued engine operation. The investigation established clear evidence of electrical discharge damage (EDD) on engine components connecting from the starter-generator to the failed input driveshaft bearing. The evidence of EDD observed in multiple elements located along the electrical current path and the residual magnetism measurement suggested that the starter-generator is highly associated with the source of the EDD.     

A note on fatal aircraft accidents involving metal fatigue

A world-wide survey of aircraft accidents was done to determine the extent to which metal fatigue is still a problem for aircraft. A total of 306 fatal accidents since 1934 were identified as having metal fatigue as a related cause, and these accidents resulted in 1803 fatalities. The accidents cover civil and, to a limited extent, military aircraft. Wing failure and engine failure were the most frequent cause of fixed-wing accidents, while for helicopters, failures of the main and tail rotors were the most common problem. Currently, about 18 fatal accidents per year are attributable to metal fatigue.     

Investigations on fretting fatigue in aircraft engine compressor blade

An investigation of several cracked blade tangs in the military aircraft engine compressor was conducted to identify the root cause of the failure. These cracks were found during the scheduled maintenance with fluorescent penetration inspection. The engine compressor blade made of Ti–6Al–4V is attached to compressor rotor by means of inserting retaining pin through rotor and blade tang. By analyzing the fracture surface of the failed blade tang, it is found that the crack in the blade tang was initiated by fretting fatigue and propagated under low cycle fatigue. Stress analysis of the blade using a non-linear finite element method is coincident with the results of fractography. The clearance between retaining pin and tang hole caused small amplitude of sliding motion leading to fretting wear during engine operation. Consequently, the damaged area due to fretting wear acts as a stress raiser inside tang hole and contributes to accelerate fretting fatigue.     

航空发动机用树脂基复合材料应用进展与发展趋势EI北大核心CSCD

树脂基复合材料具有比强度和比模量高、疲劳性能和耐腐蚀性能好等优点,已经成为航空发动机冷端部件的应用和发展趋势。国外航空发动机用树脂基复合材料研究起步较早,已经在多型发动机的风扇叶片、风扇机匣、外涵机匣、短舱等部件得到成熟应用,并朝着结构形式更优、材料性能更好、制造成本更低、自动化程度更高的方向发展。国内树脂基复合材料发展基础良好,但与国外相比在发动机上应用比例不高,需要进一步提升设计、材料、制造、实验技术水平及工程化能力。本文重点论述国外航空发动机复合材料构件的结构、材料和工艺发展现状,分析发展趋势,从建立航空发动机用复合材料体系、加强应用研究和设计牵引、推进预研成果转化和自动化技术应用等方面提出相关建议。