Aero-engine Compressor Research Articles

Preliminary Optimization of Multi-Stage Axial-Flow Industrial Process Compressors Using Aero-Engine Compressor Design Strategy

Aero-engine core compressor preliminary design strategy has been successfully applied to the advanced design of gas turbines compressors. However, few researchers have addressed the application of the aero-engine core compressor preliminary design strategy in the preliminary optimal design of industrial process compressors. Here we embedded the aero-engine core compressor preliminary design strategy into a preliminary optimal design method, in which six types of design parameters widely used to define the aero-engine compressor configuration, i.e., aspect ratio, solidity, reaction, rotation speed, outlet axial Mach number, and inlet radius ratio, were used as the design variables. The 4-stage, 5-stage, 6-stage, and 7-stage compressor configuration with the same overall design requirements for a large-scale air separation main compressor were preliminarily optimized by the developed method, in which the 4-stage design has a stage pressure rise level of current aero-engine core compressors, whereas the 7-stage design has that of current industrial process compressors. The optimized compressor configurations were then refined with the throughflow-based detailed design method and finally verified with computational fluid dynamic simulations. It is found that the developed method can optimize design efficiency and accurately predict aerodynamic performance of compressors in a few minutes. Several design guidelines for the advanced industrial process compressors were also identified. This work is of significance in extending aero-engine core compressor design strategy to the design of advanced industrial process compressors.

Effect of Laser Surface Texturing on Coating Adherence and Tribological Properties of CuNiIn and MoS2 Coating

Copper nickel indium (CuNiIn) and Molybdenum disulphide (MoS2) duplex coating is used in aero engine compressor blades of Ti6Al4V to prevent fretting failure. Surface before coating is roughened by grit blasting process for coating adhesion. In present study, in place of grit blasting, laser surface texturing (LST) process is used for surface preparation and compared with grit blasted surface. Four different elliptical and square patterns are created by LST on Ti6Al4V. Difference of surface topography between LST and grit blasted surface is analyzed by SEM and white light interferometery. CuNiIn is deposited on Ti6Al4V by atmospheric plasma spray. MoS2 lubricating coating is deposited on top of CuNiIn by spray painting and curing. Effect of difference in surface preparation on coating microstructure as well as on tribological properties is studied. The results showed that geometry and dimensions of LST pattern influence the coating adherence and wear performance. LST process can be optimized for better performance and explored as an alternative surface preparation process in industry for thermal spray coatings.

Multiaxial fatigue damage and reliability assessment of aero-engine compressor blades made of TC4 titanium alloy

This paper discusses the multiaxial fatigue damage and reliability evaluation of aero-engine compressor blades using an improved strength degeneration scheme based on Chaboche damage model combined with stress-strength interference (SSI) approach. Firstly, the multiaxial proportion and non-proportion fatigue tests of TC4 (Ti-6Al-4V) alloy under a series of loading paths are carried out. The nonlinear fatigue damage calculation and life prediction of TC4 alloy is described by combining the Chaboche nonlinear damage criterion and critical plane method with proposed damage control parameter. Secondly, a nonlinear strength degradation model based on accumulative damage is further developed, and it is proved that the model can get greater precision under different loading conditions. Finally, a revised time-varying reliability model is established by viewing mechanical characteristic of compressor blade set as interconnected system with variant load component on each blade. According to the results of Monte Carlo simulation, the proposed revised model has a better prediction precision, with the error controlled less than 5%.

Impact of foreign object damage on the leading edge of TC11 titanium alloy aeroengine blade like specimen

ABSTRACT Foreign object damage (FOD) generally occurs when hard body particles are ingested into an aero-engine fan and compressor blade when the engine is taxing, flying, or moving on the runway with definite speed. Ingested foreign object debris induces severe structural damage on the leading edge of the aero-engine blade due to the time-dependent axial fatigue loading under service conditions. Induced residual stresses due to foreign object damage play a major role in mitigating the rate of fatigue crack initiation and propagation, resulting in failure of components. In this present investigation, an explicit module of commercial software package LS-DYNA was used to simulate the impact of the GCr15 stainless steel ball (52,100 steel equivalent) against leading-edge specimen (TC11 titanium alloy) by varying speeds like the in-service operating speeds. A comparative study was carried out among obtained deformed geometry (Crater) and its dimensions from the numerical simulation with experimentation results. Variation of residual stresses in three mutually perpendicular directions along with depth has been determined to predict the possible location of crack initiation at different projectile impact velocities. A stress concentration factor has also been calculated at the vicinity of the impact, which is very much an essential factor for the fatigue strength evaluation of TC11 titanium alloy.

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

Laser shock peening (LSP) is an innovative surface treatment, which has been successfully applied in aero-engine compressor blades to improve high cycle fatigue performance. Deep compressive residual stress layer with a gradient distribution can be produced, which is the special advantage of LSP and plays a main role in high cycle fatigue performance improvement. Therefore, compressive residual stress, and especially the gradient distribution feature should be simultaneously considered in high cycle fatigue life prediction for laser-peened component. In this work, firstly, TC4 titanium alloy, a typical material used for aero-engine compressor blade, was treated by LSP with different power densities, overlapping rates and shocks respectively. High cycle fatigue life of TC4 titanium alloy is enhanced from 1473 × 103 cycles to 6148 × 103 cycles, by 317% increase with LSP treatment. Compressive residual stresses with a gradient distribution induced by LSP were disposed into an equivalent parameter according to critical plane method. At last, high cycle fatigue life of laser-peened TC4 titanium alloy specimens was successfully predicted within twice-fold error band, comprehensively considering equivalent compressive residual stress and FINDLEY model.

Analysis of Shroud and Disk Heat Transfer in Aero-Engine Compressor Rotors

Abstract Heat transfer within the rotating compressor cavity of an aero-engine is predominantly governed by buoyancy, which can be characterized by the Grashof number. Unsteady and unstable buoyancy-induced flow structures influence the temperatures and stresses in the compressor rotors, and these affect the radial growth of the disks. In addition, the heat transfer from the disks and shroud increases the temperature of the throughflow of cooling air. This paper contains two connected parts. First, a heat transfer correlation for the shroud of a rotating cavity was determined from steady-state heat flux measurements collected in the bath compressor-cavity rig at engine-simulated conditions. The Nusselt numbers were based on the cavity air temperature adjacent to the shroud, which was predicted using the Owen–Tang buoyancy model. Heat transfer from the shroud was consistent with free convection from a horizontal plate in a gravitational field. Maximum likelihood estimation was used with a Rayleigh–Bénard equation to correlate the shroud Nusselt number with the local Grashof number. Second, an energy balance was used to calculate the enthalpy rise of the axial throughflow from the measured disk and shroud heat fluxes. Disk fluxes were derived from radial distributions of measured steady-state disk temperatures using a Bayesian model and the equations for a circular fin. The calculated throughflow temperature rise was consistent with direct thermocouple measurements. The complex, three-dimensional flow near the cavity entrance can result in enthalpy exchange penetrating upstream in the throughflow, and rotationally induced flow can create upstream axial flow in the outer part of the annulus.

Effect of Radiation on Heat Transfer Inside Aeroengine Compressor Rotors

AbstractThe blade clearance in aero-engine compressors is mainly controlled by the radial growth of the compressor discs, to which the blades are attached. This growth depends on the radial distribution of the disc temperature, which in turn is determined by the heat transfer inside the internal rotating cavity between adjacent discs. The buoyancy-induced convection inside the cavity is significantly weaker than that associated with the forced convection in the external mainstream flow, and consequently radiation between the cavity surfaces cannot be ignored in the calculation of the disc temperatures. In this paper, both the Monte Carlo Ray-Trace (MCRT) method and the view factor (VF) method are used to calculate the radiative flux when the temperatures of the discs, shroud, and inner shaft of the compressor vary radially and axially. The Monte Carlo Ray-Trace method is computationally expensive, but it is able to incorporate the effect of complex geometries on radiation. The view factor method is quick to compute and, although the derivation becomes complicated when geometrical details are considered, it can be used as a first check of the effect of radiation in compressor cavities. Given distributions of surface temperatures, the blackbody and gray body heat fluxes were calculated for the discs, shroud, and inner shaft in two experimental compressor rigs and in a simulated compressor stage. For the experimental rigs, although the effect of radiation was relatively small for the case of large Grashof numbers, the relative effect of radiation increases as Gr (and consequently the convective heat transfer) decreases. For the simulated compressor, with a pressure ratio of 50:1 for state-of-the-art aircraft engines, radiation could have a significant effect on the disc temperature and consequently on the blade clearance; the effect is predicted to be more prominent for the next generation of aircraft engines with pressure ratios up to 70:1.

Probabilistic Risk of Titanium Disks Considering Residual Stresses of Hard Alpha Inclusions

In aeroengine titanium alloy components, hard alpha (HA) inclusions strongly affect material performance, especially in terms of fatigue properties. Residual stresses caused by thermal expansion mismatch during the cooling process affect the fatigue life of metal components. This study proposes a method of probabilistic risk assessment for aeroengine disks considering the residual stresses of HA inclusions. Specifically, numerical procedures are adopted to investigate the residual stresses of an inclusion under surface effects, and a functional relationship model of residual stresses under surface effects is proposed. To deal with the randomness of inclusion size and position, a calculation criterion for the residual stresses of an inclusion is established. Analysis of an aeroengine compressor disk shows that the expected decline in the failure risk of the disk is approximately 27% under the consideration of residual stresses. The proposed method predicts the failure risk more accurately, and is thus valuable for the safety assessment of an aeroengine titanium disk.

Design and Testing of a Rig to Investigate Buoyancy-Induced Heat Transfer in Aero-Engine Compressor Rotors

Abstract The change in compressor blade-tip clearance across the flight cycle depends on the expansion of the rotor, which in turn depends on the temperature and stress in the disks. The radial distribution of temperature is directly coupled to the buoyancy-driven flow and heat transfer in the rotating disk cavities. This paper describes a new test rig specifically designed to investigate this conjugate phenomenon. The rig test section includes four rotating disks enclosing three cavities. Two disks in the central cavity are instrumented with thermocouples to provide the radial distribution of temperature; the two outer cavities are thermally insulated to create appropriate boundary conditions for the heat transfer analysis. An axial throughflow of air is supplied between a stationary shaft and the bore of the disks. The temperature of the throughflow air is measured by thermocouples in rakes upstream and downstream of the central cavity. For a cold throughflow, the outer shroud of the central cavity is heated. Two independently controlled radiant heaters allow differential shroud temperatures for the upstream and downstream disks, as found in aero-engine compressors. Alternatively, the throughflow can be heated above the shroud temperature to simulate the transient conditions during engine operation where stratified flow can occur inside the cavity. The rig is designed to operate in conditions where both convective and radiative heat transfer dominate; all internal surfaces of the cavity are painted matt black to allow the accurate calculation of the radiant heat transfer. Separate attachments can be fitted to the cobs of both central disks; the attachments reduce the axial gap between the cobs—reducing the gap to zero creates a closed cavity, which can occur in some compressor designs. Other instrumentation includes heat-flux gages on the shroud and high-frequency pressure transducers embedded into the disk diaphragm to capture unsteady flow structures. Attention has been given to experimental uncertainty, including the computation of the thermal-disturbance errors, caused by thermocouples embedded in the rotating disks; a Bayesian statistical model is used to reduce the effect of uncertainties in temperature measurements on the calculation of the Nusselt number. The effect of relevant nondimensional parameters on the radial distribution of the disk and throughflow temperatures has been shown for some typical cases.

Measurement and Analysis of Buoyancy-Induced Heat Transfer in Aero-Engine Compressor Rotors

Abstract The flow inside cavities between corotating compressor disks of aero-engines is driven by buoyancy, with Grashof numbers exceeding 1013. This phenomenon creates a conjugate problem: the Nusselt numbers depend on the radial temperature distribution of the disks, and the disk temperatures depend on the Nusselt numbers. Furthermore, Coriolis forces in the rotating fluid generate cyclonic and anticyclonic circulations inside the cavity. Such flows are three-dimensional, unsteady, and unstable, and it is a challenge to compute and measure the heat transfer from the disks to the axial throughflow in the compressor. In this paper, Nusselt numbers are experimentally determined from measurements of steady-state temperatures on the surfaces of both disks in a rotating cavity of the Bath compressor-cavity rig. The data are collected over a range of engine-representative parameters and are the first results from a new experimental facility specifically designed to investigate buoyancy-induced flow. The radial distributions of disk temperature were collected under carefully controlled thermal boundary conditions appropriate for analysis using a Bayesian model combined with the equations for a circular fin. The Owen-Tang buoyancy model has been used to compare predicted radial distributions of disk temperatures and Nusselt numbers with some of the experimentally determined values, taking account of radiation between the interior surfaces of the cavity. The experiments show that the average Nusselt numbers on the disk increase as the buoyancy forces increase. At high rotational speeds, the temperature rise in the core, created by compressibility effects in the air, attenuates the heat transfer, and there is a critical rotational Reynolds number for which the Nusselt number is a maximum. In the cavity, there is an inner region dominated by forced convection and an outer region dominated by buoyancy-induced flow. The inner region is a mixing region, in which entrained cold throughflow encounters hot flow from the Ekman layers on the disks. Consequently, the Nusselt numbers on the downstream disk in the inner region tend to be higher than those on the upstream disk.

KGAssembly: Knowledge graph-driven assembly process generation and evaluation for complex components

ABSTRACT The semantic information of process documents for the assembly of complex components plays an important role in the guidance of assembly operations and the feasibility evaluation of process plans. There are many types of semantic elements contained in assembly process documents. The semantic relationship between assembly elements is complex. Additionally, there is a lack of effective modeling method to deal with the implicit associative semantic knowledge that exists among existing assembly process cases. In this case, it is a great challenge to use the professional knowledge in assembly process documents to guide the intelligent decision-making of assembly process planning and the intelligent analysis of assembly process enforceability evaluation. In this paper, a knowledge graph-driven assembly process generation and evaluation method for complex components is proposed. An APKG (assembly process knowledge graph) model is built. A distributed graph embedding-based model SKGCN (sequence knowledge graph convolutional network) is designed to generate assembly process planning. Furthermore, model and knowledge dual-driven evaluation method for assembly sequences is presented. It provides assembly expert knowledge support for the evaluation method of interference detection of assembly sequence based on point cloud assembly feature recognition. Finally, the approach is evaluated by assembling an aero-engine compressor rotor.

Effect of mistuning parameters on dynamic characteristics of mistuned bladed disk

The aim of this paper is to study the nonlinear dynamics of the tenon and mortise of aero-engine compressor blades, based on lumped parameter model of tenon and mortise of blade disk system, considering the influence of gap and friction, The nonlinear vibration characteristics of mistuned bladed disk with different mistuning parameters are studied. The effects of tenon and mortise gap, dry friction, blade damping mistuning and tenon damping mistuning on the vibration characteristics of the mistuned bladed disk under strong coupling and weak coupling are obtained respectively. The results show that the gap and friction has little effect on the vibration response of the weakly coupled bladed disk, but the gap and friction has a great influence on the strongly coupled bladed disk, the larger the gap, the more complex the nonlinear dynamic characteristics of tenon. Under the same gap and coupling strength, the mistuning parameters have a certain influence on the vibration amplitude of the bladed disk.

High velocity measurements of particle rebound characteristics under erosive conditions of high pressure compressors

Ingested sand particles may lead to severe erosion in aero engine compressors resulting in performance deterioration. Predicting particle trajectories through a turbo machine requires accurate estimation of the particle rebound. The experimental investigation of particle impact is a challenging task with high demands on measurement equipment due to the combination of small particle sizes with high impact velocities. Therefore high velocity impacts with a realistic combination of both, impact velocity and particle size, have not been investigated experimentally yet. In this work high velocity impacts have been measured for 60 μm sized quartz particles impacting stainless steel (1.4301) and Ti6Al4V targets. The experiments were conducted in a sand-blast type erosion test rig designed to meet compressor specific boundary conditions representative for a third stage axial compressor of civil jet engines. Particle Shadow Velocimetry is used to measure the particle velocities in the test rig. A dimensional analysis of normal impact is presented to discuss the effect of particle size and impact velocity on the rebound characteristics. It was found that the widely used approach to implement the energy loss via a simple curve-fit that solely depends on the impact angle is not valid throughout the broad range of possible impact events in aero engine compressors. In addition, for a realistic combination of particle size and impact velocity, particle fracture was observed. The fracture alters the Stokes Number of the particles and with that the particle trajectory after first impact. Hence the accurate prediction of the particle-wall interaction includes not only the change of particle direction and velocity upon impact but the treatment of the particle fracture. This is especially relevant for high impact velocities and high normal impact velocities.

Effect of tip injection on the flow characteristics of an aeroengine compressor rotor

Existing experiments, simulations, and analysis have shown that the flow field of the blade tip has a significant effect on rotating stall and acoustic resonance in aeroengine compressors. Based on the numerical simulation of NASA rotor 37, the steady influence of air injection at blade tip was investigated in this investigation. The results show that the injected flow could obviously extend the stability margin of the axial compressor to varying degrees. The tip injection reduces the blockage caused by the tip leakage, and reduces the backflow in the trailing edge area of the blade tip, thereby delaying the occurrence of stall. Furthermore, in this paper, different types of rotors are constructed to verify the effect of injection parameters on the stability extension. The results show that the improvement of rotor stall margin is positively correlated with the injector inlet total pressure, and when the injector inlet flow conditions are of the same magnitude, the improvement of the rotor stall margin is negatively related to the distance between the injector and the leading edge of the blade.

Surface roughness prediction model of GH4169 superalloy abrasive belt grinding based on multilayer perceptron(MLP)

GH4169 nickel-based alloy materials are widely used in the design of aero-engine compressor terminal stators and rotor blades because of their high strength, oxidation resistance and gas corrosion characteristics in high-temperature environments. In addition, the advantages of abrasive belt grinding in the processing of nickel-based alloy blades, such as low temperature and strong vibration absorption, can effectively improve the surface integrity and be popularized and applied by aviation manufacturers. However, due to the random distribution of abrasive grains, grinding heat and machining chatter during the abrasive belt grinding process, there is a coupling and strong nonlinear model relationship between the grinding process parameters and the surface roughness after grinding. At present, the strong nonlinear model relationship of multi-dimensional parameters and their coupling relationship has not been resolved, so it is difficult to achieve accurate prediction of surface roughness.We propose a study on the multi-dimensional parameter model of abrasive belt grinding for the decision-making of the surface integrity characteristics of nickel-based alloy blades. Based on the MLP multi-layer perceptron, a prediction model of abrasive belt grinding surface roughness is proposed. According to the multi-layer perceptron model and the input parameter characteristics of the belt grinding system, the initial connection weight and paranoia, the backward propagation algorithm, and the neural network layer are optimized. Through the above algorithm, a multi-dimensional parameter surface integrity nonlinear model is obtained and combined with boundary conditions to achieve accurate prediction of the surface roughness of nickel-based alloy abrasive belt grinding.

Effect of heat treatment on the microstructure, microhardness and impact toughness of TC11 and TC17 linear friction welded joint

The service life of whole blisks for the aero-engine compressor is sharply decreased due to the dissimilar linear friction welded (LFWed) joint of Ti–6.5Al–3.5Mo–1.5Zr–0.3Si (TC11) and Ti–4Mo–4Cr–5Al–2Sn–2Zr (TC17) alloys, which have the different microstructure and mechanical properties. To optimize the microstructure and performance of the joint, and improve service life of the whole blisks, the post-weld heat treatment (PWHT) with annealing, solution plus aging were adopted for the welded joint. The results show that the nucleation rate of primary αP phase increases obviously, which are spheres, rods and strips, and orthorhombic martensite αs phases transform into hexagonal martensite αs phases interlaced evenly on the β phase matrix on the WZ of TC17 side; meanwhile, acicular hexagonal martensite αs phases transform into lamellar αs phases on the WZ of TC11 side after solid solution and dual aging. The difference of the microhardness on both sides of the joint is narrowed significantly. The impact toughness of the joint is improved sharply from 9.3 J/cm2 after welding to 27.6 J/cm2 with an increase of 200%. In addition, the dissolution and growth model of primary αp phases were established, which can reasonably predict the nucleation and grow process of α phase. Furthermore, the narrowed discrepancy of the microhardness and increased impact toughness can be attributed to the decreased slice thickness of martensite αs phase, based on the dislocation slip mechanism and distorted discordant model.

Foreign object damage simulation of aero-engine blade

To study the anti-foreign object damage (FOD) ability of aero-engine compressor blades, finite element model of compressor blades was established. Blades damage was evaluated with different steel ball diameters and shock angles. Results show that the damage width and depth becomes bigger with the steel ball’s diameter increases. The damage width doesn’t change apparently with the shock angles, and the damage width nearly equals to the steel ball diameter. The damage depth mostly becomes smaller with the increase of shock angle.

The Objective Space and the Formulation of Design Requirement in Natural Laminar Flow Optimization

Design requirement is as important in aerodynamic design as in other industries because it sets up the objective for the samples in design space to approach. Natural Laminar Flow (NLF) optimization belongs to the type of aerodynamic design problems featured by the combination of distinct aerodynamic performance, where the design requirement is often formulated in form of summation of laminar-related performance and pressure drag performance with different weight assignment according to different perspectives. However, the formulations are rather experience-oriented and are decided non-quantitatively. Inspired by data manipulation approaches in design space (spanned by design variables of geometrical representation parameters) in many aerodynamic designs, this paper proposes new formulations of design requirement in NLF optimization via consideration of objective space (projection of design space through aerodynamics) and shows the impact of the corresponding formulation of design requirement to the result of NLF optimization in cases of transonic airfoil and aero engine compressor blade design from two perspectives: Pareto front convergence and improving effect of accessory performance. The paper uses Principal Component Analysis (PCA) to obtain the eigenvectors of objective space to extract the intrinsic information about specific problem. The method is realized in two cases with satisfactory result.

Numerical Analysis of High-Altitude Inlet Air on Boundary Layer Flow Loss in an Aero-Engine Compressor

A numerical analysis is performed to explore the high altitude and high Mach flight on the effect of wall boundary layer loss in the compressor. The accuracy for solution results by the application of the similarity criterion and parameter definition of the air inlet is compared with the existing experimental test result. The results indicate that the radial adverse pressure gradient in the rotor domain gradually increases along the span direction and decreases as flight Mach number increases; meanwhile, the circumferential adverse pressure gradient on the pressure side of the rotor blade is correspondingly larger and less than that on the suction side. In particular, the entropy increase along the streamwise shows a decreasing trend and an increasing trend inside the hub and shroud wall boundary layers, respectively. At 2.1 Ma, the entropy increase in the rotor domains enhances by 24.36–27.80% inside the shroud boundary layer, relative to the hub boundary layer; however, it decreases by 0.97–8.54% in the stator domain. With the increase in flight Mach number from 2.1 to 3.4, the average entropy increase reductions in the rotor domain decrease by 18.99–24.97% within the hub boundary layer and 5.71–8.1% within the shroud boundary layer. In the stator domain, it drops by 18.45–9.03% inside the hub boundary layer and 6.88–8.67% inside the shroud boundary layer. It was therefore found that, as Mach number increases from 2.1 to 3.4, the entropy increase reduction is larger inside the hub boundary layer than inside the shroud boundary layer.

Development and experimental study on linear friction welding titanium alloy blisk

Taking the advanced aeroengine compressor blisk as the application object, the development and experimental study on the TC11/TC17 different materials linear friction welding integral blisk was carried out. The microstructure of welded joint was analyzed, the tensile and stress rupture properties of base metal and welded joint were compared, the vibration fatigue test on welded blade test pieces, and the overrunning test and low-cycle fatigue test on welding blisk were developed. The results show that the TC11/TC17 linear friction welding joint consists of five typical zones, they are TC11 base metal zone, TC11 thermal influence zone, weld zone, TC17 thermal influence zone and TC17 base metal zone. The tensile strength of the welded joint is basically equivalent to that of the TC11 base metal, elongation is equivalent to TC17 base metal; endurance strength lower than TC17 base metal, but less higher than TC11 base metal. The median fatigue life of welded blade test pieces with three weld positions is respectively 1.501 × 106, 0.344 × 106, and 0.132 × 106 cycles, the fatigue crack isn’t at welding seam and thermal influence zone. The welding blisk pass the 115% overrunning test and 1000 cycles low-cycle fatigue test, the experiment results meet the requirement.