The Air Force Manufacturing Technology Laser Peening Initiative

Probabilistic failure assessments for integrally bladed disks are system reliability problems where a failure in at least one blade constitutes a rotor system failure. Turbine engine fan and compressor blade life is dominated by High Cycle Fatigue (HCF) initiated either by pure HCF or Foreign Object Damage (FOD). To date performing an HCF life assessment for the entire rotor system has been too costly in analysis time to be practical. Although the substantial run-time has previously precluded a full-rotor probabilistic analysis, reduced order models make this process tractable as demonstrated in this work. The system model includes frequency prediction, modal stress variation, mistuning amplification, FOD effect, and random material capability. The model has many random variables which are most easily handled through simple random sampling.

Experimental study and fatigue life prediction on high cycle fatigue performance of laser-peened TC4 titanium alloy

PurposeThe purpose of this paper is to conduct a comparative study of the surface modifications induced by two different lasers on a 2050‐T8 aluminum alloy, with a specific consideration of residual stress and work‐hardening levels.Design/methodology/approachTwo lasers have been used for Laser shock peening (LSP) treatment in water‐confined regime: a Continuum Powerlite Plus laser, operating at 0.532 mm with 9 ns laser pulses, and near 1.5mm spot diameters; a new generation Gaia‐R Thales laser delivering 10 J‐10 ns impacts, with 4‐6mm homogeneous laser spots at 1.06 mm. Surface deformation, work‐hardening levels and residual stresses were analyzed for both LSP conditions. Residual stresses were compared with numerical simulations using a 3D finite element (FE) model, starting with the validation of surface deformations induced by a single laser impact.FindingsSimilar surface deformations and work‐hardening levels, but relatively lower residual stresses were obtained with the new large 4‐6 mm impact configuration. This was attributed to a reduced number of local cyclic loadings (2) compared with the small impact configuration (4). Additionally, more anisotropic stresses were obtained with small impacts. FE simulations using Johnson‐Cook's material' behavior were shown to simulate accurately surface deformations, but to overestimate maximum stress levels.Research limitations/implicationsThis work should provide LSP workers a better understanding of the possible benefits from the different LSP configurations currently co‐existing: using small (<2 mm) impacts at high‐cadency rates or large ones (>4‐5 mm). Moreover, experimental results and simulated data had never been presented on 2050‐T8 Al alloy.Originality/valueAn experimental (and numerical) comparison using two distinct laser sources for LSP, has never been presented before. This preliminary work should help LSP workers to choose adequate sources.

Understanding the role of monolayer graphene during long range shock strengthening of metal-graphene heterostructure

Experiment investigation of laser shock peening on TC6 titanium alloy to improve high cycle fatigue performance

Laser shock processing (LSP), also known as laser peening (LP), proves to be superior to conventional surface treatments such as shot peening, including deeper penetration of the residual stresses. The LSP treatment, which uses a very short pulse (ns) of intense (GW cm−2) laser beam to generate compressive residual stresses near the surface of the metallic samples, demonstrates a significant improvement of fatigue life and stress corrosion cracking resistance. In this paper, finite element analysis (FEA) combined with particle swarm optimization (PSO) method to predict the magnitude and distribution of optimized multiple LSP impacts on 304 stainless steel. The results of the simulation clearly show that optimized multiple LSP can mitigate residual stresses loss in the centre of the single impact zone and generate homogeneous compressive residual stresses at the surface. The results also reveal the optimized multiple LSP can lead to deeper penetration of the compressive residual stresses in the samples.

Chapter Three - Shock Processing

The purpose of this paper is to predict the compressive residual stress profile induced by the laser shock peening process and to optimize the surface conditions on a Ti-6Al-4V titanium super-alloy using the new technology based on the artificial neural networks. This study covers two principal cases: (i) The numerical analysis: the compressive residual stress distribution induced by the laser peening process is simulated based on the finite element method using ABAQUS software (ABAQUS/Explicit code). (ii) The mathematical modeling analysis: the using of the artificial neural network technique has been proposed to predict the residual stress profile induced by the laser shock processing in order to optimize the laser shock peening surface conditions. To train the artificial neural network, we use different numerical experiments measurements as training and test data. The best fitting training data set was obtained with four neurons in the hidden layers. After training and as seen from the mathematical experiments, the calculated residual stress and the damage state are obviously acceptable. The principal goal of this research paper is to predict the surface treatment state induced by the laser shock peening based on the artificial neural networks technique and on the numerical results without making multiple simulations that can take much more time.

To investigate the effect of power density on the residual stress distribution on the surface of 17-7 PH stainless steel welded joints during laser shock process, we utilized ABAQUS finite element software for analyzing its distribution under different laser shock power densities. The material was tested for residual stress using an X-ray stress diffractometer to verify the accuracy of the simulation results and lay the foundation for finding the optimal process parameters. The simulation results show that when the laser power density is 2.79 GW/cm2, the maximum residual compressive stress value on the surface of the sample after laser shock peening is −37.2 MPa; When the laser power density is 2.79 GW/cm2, the maximum residual compressive stress value on the surface of the sample is −94.1 MPa; When the power density reaches 5.17 GW/cm2, the maximum residual stress value on the surface of the laser shock peening sample is −144.8 MPa, and residual stress cavity is formed. The experimental results show that when the laser power density is 2.79 GW/cm2, the maximum residual compressive stress on the surface of the sample is −37.3 MPa; When the power density is 3.98 GW/cm2, the uniformity of compressive stress distribution on the surface of the sample is optimal, with a maximum residual compressive stress value of −99.0 MPa; When the power density is 5.17 GW/cm2, the maximum residual compressive stress value is −146.1 MPa, and residual stress cavity is formed on the material surface. The simulation results are close to the experimental results, and the model error ranges from 0.2% to 4.9%. The simulation results are accurate and reliable.

A compressor blade failure was experienced at the 69 MW gas turbine of a combined cycle (C.C.) unit after four years operation since the last overhaul (January 2005). The unit accumulated 27,000 service hours and 97 start-ups since the last overhaul. This unit consists of four gas turbine stages and 19 compressor stages and operates at 3600 rpm. In 2006, the unit was equipped with a fogging system at the compressor air inlet duct to increment unit power output during high ambient temperature days (hot days). These fog water nozzles were installed upstream of the compressor inlet air filter without any water filter/catcher before the water spray nozzles. Three unit failure events occurred at small periods, which caused forced outage. The first failure occurred in December 2008, a second event in March 2009 and the third event in May 2009. Visual examination carried out after the first failure event indicated that the compressor vanes (diaphragms) had cracks in their airfoils initiating at blade tenons welded to the diaphragm outer shroud at stages 3, 8, 9, 10 and 11. Also, many stationary vanes and moving blades at each stage of the compressor showed foreign object damage (FOD) and fractures at the airfoil. Visual examination performed for the second failure event after 60 unit operation hours indicated that many compressor vanes (diaphragms) and moving blades had FOD at the airfoil. This was attributed to fractures of the fogging system water spray nozzle, which were then induced to the compressor flow path channel at high velocity causing the above-mentioned damage. Visual examination completed upon the third failure event after two unit startup attempts indicated damage of compressor stationary vanes and moving blades principally at stages 12 to 16, and also stages 17 to 19. The damage consisted of airfoil fracture in stationary vanes and moving blades, FOD, moving blade tip rubbing, and bending of stationary vanes, moving blades and diaphragm shrouds. A laboratory evaluation of stationary vane tenon fracture indicated a high cycle fatigue (HCF) failure mechanism, and crack initiation was accelerated by corrosion picks on blade surfaces due to high humidity air generated by the fogging system. Stationary vane damage was caused by a rotating stall phenomenon, which generates vibratory stresses in stationary vanes and moving blades during unit start-ups. During the third failure event, stationary vane HCF damage was highly accelerated due to pre-existent partial fractures in tenons generated during previous failure events, which had not been detected by non-destructive tests. Stationary vane and moving blade failure was also influenced by high tenon brittleness in stationary vanes and moving blades generated during manufacture by welding (diaphragms) and repair welding (moving blades) without adequate post-weld heat treatment (stress relieving). A compressor stationary vane and moving blade failure evaluation was completed. This investigation included cracked blade metallographic analysis, unit operation parameter analysis, history-of-events analysis, and crack initiation and propagation analysis. This paper provides an overview of the compressor failure investigation, which led to identification of the HCF failure mechanism generated by rotating stall during unit start-ups, highly accelerated by corrosion generated by the fogging system and influenced by high stationary vane and moving blade brittleness as the primary contribution to the observed failure.

The paper presents three algorithms for the measurement of gasturbine engine (GTE) compressor blades airfoil (suction side, pressure side, leading and trailing edges) geometry using coordinate measuring machines. The first algorithm is a single scan of the heightwise blade section using a test point. The second algorithm is based on the pre-alignment of the nominal profile with the measured actual profile. An iterative closest-point algorithm is used to carry out the procedure of the best possible alignment of the measured ptofile and the nominal one. In the third algorithm, the coordinates of the measured points of the edges are calculated using the spline defined by the coordinates of the test point (TP) center. The research of errors of measuring the geometry of the elements using the algorithms mentioned above was carried out. To perform the research a model was developed simulating the contact of the TP and the surface of the part being measured. The model makes it possible to calculate the coordinates of the contact point, the center coordinates and the coordinates of the MT of the measured points. The search for the coordinates of the contact was performed using the method of sequential quadratic programming. A procedure was also used that makes it possible to simulate the deviation of the profile and surface position from the nominal. In the course of the simulation dependences for determining the average, lower and upper boundaries of the measurement error distribution for any airfoil point being measured were obtained.

Research on a different method to reach the saturate limit of titanium aluminide alloy surface mechanical and fatigue properties by laser shock process

Welded joints made of 6061-T6 Al alloy were studied to evaluate warm laser shock peening (WLSP) and laser shock peening (LSP) processes. The estimation model of laser-induced surface residual stress was examined by means of experiments and numerical analysis. The high-cycle fatigue lives of welded joint specimens treated with WLSP and LSP were estimated by conducting tensile fatigue tests. The fatigue fracture mechanisms of these specimens are studied by surface integrity and fracture surface tests. Experimental results and analysis indicated that the fatigue life of the specimens processed by WLSP was higher than that with LSP. The large increase in fatigue life appeared to be the result of the larger residual stress, more uniform microstructure refinement and the lower surface roughness of the WLSP specimens.

Following foreign object damage (FOD), a decision to repair components using novel additive manufacturing (AM) technologies has good potential to enable cost‐effective and efficient solutions for aircraft gas turbine engine maintenance. To implement any new technology in the gas turbine remanufacturing world, the performance of the repair must be developed and understood through careful consideration of the impact of service life‐limiting factors on the structural integrity of the component. In modern gas turbine engines, high cycle fatigue (HCF) is one of the greatest causes of component failure. However, conventional uniaxial fatigue data is inadequate in representing the predominant HCF failure mode of gas turbine components that is caused by vibration. In this study, the vibratory fatigue behavior of Ti6Al4V deposited using wire‐fed electron beam additive manufacturing (EBAM) was examined with the motivation of developing an advanced repair solution for fatigue critical cold‐section parts, such as blades and vanes, in gas turbine engine applications. High cycle fatigue data, generated using a combination of step‐testing procedure and vibration (resonance) fatigue testing, was analyzed through Dixon–Mood statistics to calculate the endurance limits and standard deviations of the EBAM and wrought Ti6Al4V materials. Also plots of stress (S) against the number of cycles to failure (N) were obtained for both materials. The average fatigue endurance limit of the EBAM Ti6Al4V was determined to be greater than the wrought counterpart. But the lower limit (95% reliability) of 426 MPa for the EBAM Ti6Al4V was lower than the value of 497 MPa determined for wrought Ti6Al4V and was attributed to the slightly higher data scatter–as reflected by the higher standard deviation–of the former material.

Fretting fatigue of laser shock peened Ti–6Al–4V