High-pressure Compressor Blades Research Articles

Numerical Simulation of Residual Stresses and Structural Vibration of Welding Repaired Blades

Abstract Taking a typical aero-engine high-pressure compressor blade as the research object, the Gauss heat source was used to numerically simulate the repair process of blade surfacing welding. The temperature data and stress data of the blade were obtained when the heat source velocity was 1 mm/s, 2 mm/s, and 4 mm/s, respectively. The results showed that both the temperature and stress values decreased with the heat source velocity increasing. Then, the residual stress of the repaired blade is introduced into the modal analysis module as prestress, and the Campbell diagram is drawn to analyze the natural frequency and speed margin of the blade respectively. The results show that as the heat source speed increases, the frequencies of each order of the blade decrease. The speed margin of each resonance point is greater than 10%, and the modal dynamic stress corresponding to the resonance speed can be used as the primary criterion for blade failure.

Unsteady Flows and Component Interaction in Turbomachinery

Unsteady component interaction represents a crucial topic in turbomachinery design and analysis. Combustor/turbine interaction is one of the most widely studied topics both using experimental and numerical methods due to the risk of failure of high-pressure turbine blades by unexpected deviation of hot flow trajectory and local heat transfer characteristics. Compressor/combustor interaction is also of interest since it has been demonstrated that, under certain conditions, a non-uniform flow field feeds the primary zone of the combustor where the high-pressure compressor blade passing frequency can be clearly individuated. At the integral scale, the relative motion between vanes and blades in compressor and turbine stages governs the aerothermal performance of the gas turbine, especially in the presence of shocks. At the inertial scale, high turbulence levels generated in the combustion chamber govern wall heat transfer in the high-pressure turbine stage, and wakes generated by low-pressure turbine vanes interact with separation bubbles at low-Reynolds conditions by suppressing them. The necessity to correctly analyze these phenomena obliges the scientific community, the industry, and public funding bodies to cooperate and continuously build new test rigs equipped with highly accurate instrumentation to account for real machine effects. In computational fluid dynamics, researchers developed fast and reliable methods to analyze unsteady blade-row interaction in the case of uneven blade count conditions as well as component interaction by using different closures for turbulence in each domain using high-performance computing. This research effort results in countless publications that contribute to unveiling the actual behavior of turbomachinery flow. However, the great number of publications also results in fragmented information that risks being useless in a practical situation. Therefore, it is useful to collect the most relevant outcomes and derive general conclusions that may help the design of next-gen turbomachines. In fact, the necessity to meet the emission limits defined by the Paris agreement in 2015 obliges the turbomachinery community to consider revolutionary cycles in which component interaction plays a crucial role. In the present paper, the authors try to summarize almost 40 years of experimental and numerical research in the component interaction field, aiming at both providing a comprehensive overview and defining the most relevant conclusions obtained in this demanding research field.

A Deep-Learning-Based Approach for Aircraft Engine Defect Detection

Borescope inspection is a labour-intensive process used to find defects in aircraft engines that contain areas not visible during a general visual inspection. The outcome of the process largely depends on the judgment of the maintenance professionals who perform it. This research develops a novel deep learning framework for automated borescope inspection. In the framework, a customised U-Net architecture is developed to detect the defects on high-pressure compressor blades. Since motion blur is introduced in some images while the blades are rotated during the inspection, a hybrid motion deblurring method for image sharpening and denoising is applied to remove the effect based on classic computer vision techniques in combination with a customised GAN model. The framework also addresses the data imbalance, small size of the defects and data availability issues in part by testing different loss functions and generating synthetic images using a customised generative adversarial net (GAN) model, respectively. The results obtained from the implementation of the deep learning framework achieve precisions and recalls of over 90%. The hybrid model for motion deblurring results in a 10× improvement in image quality. However, the framework only achieves modest success with particular loss functions for very small sizes of defects. The future study will focus on very small defects detection and extend the deep learning framework to general borescope inspection.

Experimental and numerical investigation into the effect of surface roughness on particle rebound

Erosion damage and particle deposition are crucial wear phenomena in gas turbine engines. As a result, compressor efficiency decreases, stability margin reduces, and maintenance cost increases. Hence, predicting these phenomena in an accurate manner is of paramount importance for a cost-efficient, safe, and sustainable operation. Erosion and particle deposition in the annulus are affected by particle transportation in the fluid and particle-wall interaction. The latter involves the particle impact, the potential damage of the surface and/or the particle, and the particle rebound. Particle rebounds are statistical in nature due to the target surface roughness, the variability in particle sizes, and superimposed effects caused by particle shapes as well as particle rotation and particle break-up during contact. Multiple studies investigated the statistics of particle rebound, providing empirical-based models for median and spread. However, modeling the particle-wall interaction and its data spread on a transparent physical basis allows separating the effect of target roughness from superimposed effects. The presented article pursues this objective by assessing the statistical spread of particle rebound data through multiple techniques and utilizing their interdependencies. It combines experimental, numerical, and analytical considerations. For the first time, coherent boundary conditions for the experimental, numerical and analytical setup allow the distinction of the effect of roughness from the integral effect of the superimposed phenomena. A sandblast test rig equipped with laser measurement equipment was used to measure particle rebound from flat titanium and stainless steel plates at different angles. The numerical setup is developed under OpenFOAM 6 using a RANS solver for transient simulations with compressible media in combination with one-way coupled particle flows. The numerical model includes the rebound spread model proposed by Altmeppen et al. combined with the quasi-analytical rebound model proposed by Bons et al. The statistical spread of particle rebound is investigated for roughness levels that are similar to the ones of deteriorated high-pressure compressor blades as discussed by Gilge et al. The measured surface roughness of the experimentally investigated targets is used as input parameters to the numerical framework. The rebound statistics obtained in the numerical simulation are compared to the rebound data measured in the experiment. Based on this study, conclusions are drawn about which part of the rebound spread is attributable to surface roughness and which is caused by superimposed effects. It was found that the effect of surface roughness as characterized by Altmeppen et al. contributes in the order of 46 % to the rebound spread for small impact angles. However, the share in spread due to roughness gradually decreases with increasing global impact angles to a level of 13 % for angles close to 90°. The remaining percentage of rebound spread is attributed to superimposing phenomena. In addition to the rolling and sliding of aspherical particles, further phenomena such as plastic deformation and erosion of the roughness peaks during contact and the associated dissipation of energy gain in importance for steeper impact angles. Therefore, the effect of surface roughness should not be neglected in numerical simulations of particle-laden flows. Modeling the superimposed phenomena which are observed to be dominating at high impact angles opens up a further field of research.

Preparation and Performance of an Abradable NiCrFeAlBN-YSZ-NiCrAl Layered Seal Coating for Aircraft Engines

Ti2AlNb alloys are often used to fabricate high-pressure compressor blades and casings in aircraft engines with high thrust/weight ratios. In this study, a three-layered seal coating, NiCrFeAlBN-YSZ-NiCrAl, was prepared by thermal spraying on a titanium alloy (Ti2AlNb) substrate. The microstructure, surface hardness, and bond strength of coatings as well as their thermal cycling performance and abradability under simulated working conditions were investigated. The coating surface hardness was 61 HR15Y and the bond strength was 7.13 MPa. The coating maintained a good bonding to the substrate with no cracks after 40 cycles of water-cooling from 750 °C to room temperature. The abradability of the as-sprayed coating and the coating heat-treated at 750 °C for 200 h was tested against a Ti2AlNb blade at 750 °C, with a tip line speed of 300 m/s, and feed rates of 5, 50 and 480 m/s. The coating wear mechanisms involved micro-ruptures. The Cr2O3 phase formation and a decrease in the number of micropores in the heat-treated coating decreased its abradability.

Residual stress in the surface layers of high-pressure compressor blades made of titanium alloys after surface treatment with high-current pulsed electron beams

The paper deals with the evaluation of residual stresses in the surface layer of titanium alloys as a result of electron beam processing and solving problems associated with increasing heat resistance, corrosion resistance, and fatigue strength.

Modeling crack growth of an aircraft engine high pressure compressor blade under combined HCF and LCF loading

This paper presents a 3D crack growth simulation of an aircraft engine high pressure compressor blade which fractured in service. Using FRANC3D, an initial crack is inserted and grown using the max tensile stress theory. A numerical method was developed to capture the stress field under low cycle fatigue (LCF) and high cycle fatigue (HCF) loading by superimposing test measured dynamic loading with steady stresses. In doing so, the predicted crack trajectory, aspect ratio, and shape more closely agreed with the fractured airfoils compared to performing the simulation under LCF loads alone. This supports findings from fractographic investigation that the crack exhibits HCF rather than LCF characteristics.

Sensitivity Analysis of BLISK Airfoil Wear

The decreasing performance of jet engines during operation is a major concern for airlines and maintenance companies. Among other effects, the erosion of high-pressure compressor (HPC) blades is a critical one and leads to a changed aerodynamic behavior, and therefore to a change in performance. The maintenance of BLISKs (blade-integrated-disks) is especially challenging because the blade arrangement cannot be changed and individual blades cannot be replaced. Thus, coupled deteriorated blades have a complex aerodynamic behavior which can have a stronger influence on compressor performance than a conventional HPC. To ensure effective maintenance for BLISKs, the impact of coupled misshaped blades are the key factor. The present study addresses these effects on the aerodynamic performance of a first-stage BLISK of a high-pressure compressor. Therefore, a design of experiments (DoE) is done to identify the geometric properties which lead to a reduction in performance. It is shown that the effect of coupled variances is dependent on the operating point. Based on the DoE analysis, the thickness-related parameters, the stagger angle, and the max. profile camber as coupled parameters are identified as the most important parameters for all operating points.

Failure Analysis of High-Pressure Compressor Blade in an Aero Gas Turbine Engine

Failure of high-pressure compressor rotor blade in an aero gas turbine engine is analyzed to determine its root cause. Forensic and metallurgical investigations are carried out on the blade and failed parts. The failure of the platform ladder is found to the first in the chain of events that led to the compressor blade failure. The mode of failure in the blade is found to be fatigue and has originated from the damaged region on the leading edge caused by dislodgement of platform ladder. The failure has caused extensive damages in high-pressure compressor module and also in downstream turbine blades as a secondary effect.

Determination of Johnson–Cook Plasticity Model Parameters for Inconel718

In order to simulate foreign object damage (FOD) phenomenon in aircraft high-pressure compressor blades made of a nickel-based super-alloy, Johnson–Cook (J–C) plasticity model was used. For prediction of material’s plastic behavior at temperature of 400 °C (working temperature of the blades) in the range of strain rates associated with the FOD phenomenon (in order of 106 s−1), material parameters of A, B, C, n and m for the J–C plasticity model had to be determined experimentally. Parameters of A, B and n with values of 1108, 699 MPa and 0.5189, respectively, were obtained from quasi-static tensile tests. Moreover, m was determined to be 1.2861, also through quasi-static tensile tests with a strain rate of 1 s−1 at three temperatures of 475, 550 and 625 °C. However, in order to determine C, firstly a steel ball was impacted on the surface of a flat specimen made of a precipitation-hardening alloy, and then, the impact site was 3D scanned to obtain the induced crater profile. Finally, the impact test (ballistic) was simulated using Abaqus, and a C value of 0.0085 was determined by comparing the actual crater profile with the one obtained from the simulation through a trial-and-error approach.

Mechanical characterization of craters induced by impacting foreign objects on Inconel718

In this study, foreign object damage on aircraft high pressure compressor blades made of Inconel718 nickel-based superalloy was experimentally simulated at the working temperatures of the blades. Steel balls (as a foreign objects) with velocities of 292, 208, 133 m/s were impacted on the surface of Inconel718 flat specimens at various temperatures (400, 475, 550, and 625℃). In order to characterize induced craters, impact sites were studied by scanning electron microscopy to identify mechanical features such as microcracks and notches. Then, the specimens were 3D scanned to measure the depth of each crater. It was found that variation of stress concentration factors and induced damages (including microcracks and notches, loss of material, and shear bands) could occur as a result of steel ball impact with velocities of 292 and 133 m/s at temperatures of 625 and 400℃, respectively.

Application of structure-energy failure theory for determining the probability of failure and durability of turbine and compressor blades of aircraft engines

The aim of the present work was to develop a methodology for assessing reliability on the basis of physical characteristics of an element. To this goal, we used the structure-energy power failure theory. The use of the theory makes it possible to determine the probability of failure using the information on the internal damage of the material. The gas turbine engine blades of the first, second and third turbine stages, new and reconstructed high-pressure compressor blades of stage 5 were the object of the analysis. We used an industrial computer tomograph XTH 450 LC to carry out non-destructive testing. We carried out destructive tests on stands of the electrodynamic type. The probabilities of failure calculated by using the structure-energy failure theory correspond to the values of reliability indices obtained by the destructive experiments. We developed a computer program based on a statistical method for estimating the reliability of an aircraft engine. The program takes into account the reliability index of rotor blades calculated with the use of the structure-energy failure theory.

Study on Blade Fatigue Life of Rotating Power Machinery

The linear damage model (LDM) is widely applied in engineering calculation, but it does not consider the relationship between damage variable and load parameters. Therefore, the life prediction based on LDM is not satisfied for the aero-engine blades. Besides, it easily brings about error in predicting fatigue life by common nonlinear damage model which neglect the influence of torsional stress. Hence, a modified nonlinear continuum damage model (CDM) is put forward based on Chaboche nonlinear damage model in this research. And to determine the damage and fatigue life of TC4 material used in aero-engine blades, axial tension and compression fatigue test is conducted. Compared with LDM results, the fatigue life prediction results of the modified CDM in this work show a good agreement with the tests data. So the correctness of the modified model is verified. Finally, the fatigue life of a certain aero-engine high pressure compressor blade is predicted by the modified nonlinear continuum damage model.

Fatigue Fracture of a Compressor Blade of an Aeroengine: What Caused this Failure?

The success of failure analysis depends on the use of appropriate analytical techniques in proper sequence. The approach and the methodology to be adopted for a particular failure investigation largely depend on the nature of failure. Collection of background information leading to development of a complete case history of the failure is, therefore, essential. This is followed by examination of the components, to establish the physical cause of failure, mostly through metallurgical investigation and evaluation of the material of construction. Metallurgical investigation can pinpoint the physical cause of failure but often falls short of identifying the true cause of failure. It has been seen that many a time, the true cause of failure is not established due to overlooking of minor details and not because of lack of advanced knowledge. This aspect of failure analysis has been highlighted in this article through a practical example. The failure involved fatigue fracture of a high pressure compressor blade of an aeroengine. The failure occurred within 12.2 h of engine run on the test bed.

The study of the structural and acoustic anisotropy of the material of high-pressure turbine compressor blades

Non-destructive testing of carbide phase anisotropy in the material (steel EI961) of the rotor blades of a high-pressure turbine compressor (HPTC) by optical metallography and ultrasonic pulse-echo techniques directly after operation and after tempering has been carried out. A correlation between the parameter of intrinsic acoustic anisotropy and the coefficient of structural anisotropy of the carbide phase was established

Concept description and assessment of the main features of a geared intercooled reversed flow core engine

Intercooled turbofan cycles allow higher overall pressure ratios to be reached which gives rise to improved thermal efficiency. Intercooling also allows core mass flow rate to be reduced which facilitates higher bypass ratios. A new intercooled core concept is proposed in this paper which promises to alleviate limitations identified with previous intercooled turbofan designs. Specifically, these limitations are related to core losses at high overall pressure ratios as well as difficulties with the installation of the intercooler. The main features of the geared intercooled reversed flow core engine are described. These include an intercooled core, a rear-mounted high-pressure spool fitted rearwards of the low-pressure spool as opposed to concentrically as well as a mixed exhaust. In these studies, the geared intercooled reversed flow core engine has been compared with a geared intercooled straight flow core engine with a more conventional core layout. This paper compares the mechanical design of the high-pressure spools and shows how different high-pressure compressor and high-pressure turbine blade heights can affect over-tip leakage losses. In the reversed configuration, the reduction in high-pressure spool mean diameter allows for taller high-pressure compressor and turbine blades to be adopted which reduces over-tip leakage losses. The implication of intercooler sizing and configuration, including the impact of different matrix dimensions, is assessed for the reversed configuration. It was found that a 1-pass intercooler would be more compact although a 2-pass would be less challenging to manufacture. The mixer performance of the reversed configuration was evaluated at different levels of mixing effectiveness. This paper shows that the optimum ratio of total pressure in the mixing plane for the reversed flow core configuration is about 1.02 for a mixing effectiveness of 80%. Lower mixing effectiveness would result in a higher optimum ratio of total pressure in the mixing plane and fan pressure ratio.

Redesign of a High-Pressure Compressor Blade Accounting for Nonlinear Structural Interactions

Recent numerical developments dedicated to the simulation of rotor/stator interaction involving direct structural contacts have been integrated within the Snecma industrial environment. This paper presents the first attempt to benefit from these developments and account for structural blade/casing contacts at the design stage of a high-pressure compressor blade. The blade of interest underwent structural divergence after blade/abradable coating contact occurrences on a rig test. The design improvements were carried out in several steps with significant modifications of the blade stacking law while maintaining aerodynamic performance of the original blade design. After a brief presentation of the proposed design strategy, basic concepts associated with the design variations are recalled. The iterated profiles are then numerically investigated and compared with respect to key structural criteria such as: (1) their mass, (2) the residual stresses stemming from centrifugal stiffening, (3) the vibratory level under aerodynamic forced response, and (4) the vibratory levels when unilateral contact occurs. Significant improvements of the final blade design are found: the need for an early integration of nonlinear structural interactions criteria in the design stage of modern aircraft engines components is highlighted.

Investigation and analysis of deterioration in high pressure compressors due to operation

Due to environmental and operational effects, the deterioration of all gas path-related jet engine components is a highly influencing parameter leading to an increase in exhaust gas temperature and specific fuel consumption over time. As a particularly strained engine module, the high-pressure compressor (HPC) is prone to a variety of abrasive and deforming effects that are responsible for a considerable part of overall engine performance losses. During HPC overhaul, new airfoils are typically mixed with reused parts whose refurbishment workscopes typically range from “passed inspection” up to “full leading and/or trailing edge restoration” as well as “tip weld repair”. Hence, a wide spread of airfoil geometries with a distinct statistical distribution can be found within an engine after operation. To allow a statistical analysis, a geometric survey on two full ex-service HPC blade sets and an assortment of equivalent new production parts was conducted. The blades were digitalized by a structured-light 3D scanner in conjunction with a photogrammetry system. Based on the measured three-dimensional data, a CAD model was generated for each blade. Afterwards, airfoil sections on different blade heights were analyzed to generate aerodynamic design parameters such as leading and trailing edge geometries (radius and thickness), their maximum profile thickness, their chord length and their stagger angle. By assessing the statistical results of both used and new parts, the effects of manufacture variations and in-service wear on HPC blade geometry can be compared in detail.

Surface Quality of Amilled Gamma Titaniumaluminide for Aeronautical Applications

Abstract Gamma titanium aluminides are an interesting alternative for nickel, iron or cobalt matrix superalloys. Due to the advantageous strength properties at high temperatures they can successfully replace superalloys in applications such as high pressure compressor blades, low pressure turbine blades, high pressure compressor case, low pressure turbine case. Milling is one of the processes that can be applied in the forming elements made from this type of alloys for the aviation industry. Research included the selection of tool, the process kinematics and the range of milling gamma titanium aluminide (Ti-45Al-5Nb-0.2B-0.2C) process parameters were carried out. Milling can be an effective method of forming of elements made of gamma TiAL in the range of processing parameters: vc = 20-70 m/min, ap = 0.3-0.7 mm, fz = 0.1- 0.45 mm/tooth. In the tests carried out the best results were obtained using a R300-016A20L- 08L milling cutter, S30T tool coating and in-cut milling.

Curvature driven two-dimensional multi-objective optimization of compressor blade sections

This paper introduces a new curvature based design parameterization of two-dimensional high pressure compressor blade sections to be used in a multi-criteria aerodynamic design optimization process. The suction side of the airfoil section is represented by its curvature distribution which is described by a B-spline curve. The coordinates are then derived by numerical integration. The camber line as well as pressure side are obtained by adding a thickness distribution perpendicularly to the camber line. The thickness distribution itself is modeled as a B-spline curve as well and varied during optimization. In order to achieve smoothness between leading edge, trailing edge and pressure side, respectively, specific conditions for the control points of the thickness distribution are derived. The resulting compressor blade section is optimized w.r.t. pressure loss and working range using process integration and a genetic algorithm. The results show major improvements over a manually “optimized” datum design regarding both criteria.