High Pressure Compressor Rotor Research Articles

Non-uniform flow characteristics and rotating instability of a transonic high-pressure compressor rotor with cavity bleed

The presence of bleed in an aero engine’s compressor can significantly impact its flow characteristics and contribute to rotating instability. This study focuses on the impact of a typical cavity bleed structure on the internal flow characteristics of a compressor, specifically its circumferential non-uniformity and aerodynamic stability. A numerical simulation involving multiple flow passages was conducted on the transonic high-pressure rotor of the E3 compressor, considering its typical bleed structure. The study delves deep into the non-uniform flow characteristics and the mechanisms behind their generation in the compressor flow field. Furthermore, the influence of bleed on the rotating instability of the compressor is explored by comparing changes in compressor instability and adiabatic efficiency under uniform and non-uniform flow field conditions. The findings indicate that the axial position of the cavity bleed structure plays a crucial role in influencing key parameters such as rotor stall margin, peak efficiency, and total pressure ratio under near-stall conditions. The circumferential non-uniformity, resulting from the presence of the cavity bleed, intensifies with higher bleed air flow rates. For the upstream bleed configuration applied to the rotor, with a total bleed rate of 5%, the maximum variation in absolute flow angle at the inlet of different rotor channels can reach up to 1°. Additionally, the maximum difference in inlet flow coefficient can reach 0.0392. These findings demonstrate that the non-uniformity caused by the typical bleed structure leads to a loss in stall margin for the rotor when compared to a uniform flow field scheme.

A New Approach for Deviation Modeling in Compressors: Sensitivity-Correlated Principal Component Analysis

Studies on the geometry variation-related compressor uncertainty quantification (UQ) have often used dimension reduction methods, such as the principal component analysis (PCA), for the modeling of deviations. However, in the PCA method, the main eigenmodes were determined based only on the statistical behavior of geometry variations. While this process can cause some missing modes with a small eigenvalue, it is much more sensitive to blade aerodynamic performances, and thereby reducing the reliability of the UQ analysis. Hence, a novel geometry variation modeling method, named sensitivity-correlated principal component analysis (SCPCA), has been proposed. In addition, by means of the blade sensitivity analysis, the weighting factors for each eigenmode were determined and then used to modify the process of the PCA. As a result, by considering the covariance of geometry variations and the performance sensitivity, the main eigenmodes could be determined and used to reconstruct the blade samples in the UQ analysis. With 98 profile samples measured at the midspan of a high-pressure compressor rotor blade, both the PCA and SCPCA methods were employed for the UQ analysis. The results showed that, compared to the PCA method, the SCPCA method provided a more accurate reconstruction of sensitive deviations, leading to an 11.8% improvement in evaluating the scatter of the positive incidence range, while also maintaining the accuracy of the uncertainty assessment for other performances.

Geometric modeling analysis method for stacking assembly deviation of aero engine rotor

Purpose Assembly errors of aeroengine rotor must be controlled to improve the aeroengine efficiency. However, current method cannot truly reflect assembly errors of the rotor in working state owing to difficulties in error analysis. Therefore, the purpose of this study is to establish an optimization method for aeroengine rotor stacking assembly. Design/methodology/approach The assembly structure of aeroengine rotor is featured. Rotor eccentricity is optimized based on Jacobian–Torsor model. Then, an optimization method for assembly work is proposed. The assembly process of the high-pressure compressor rotor and the high-pressure turbine rotor as the rotor core assembly is mainly considered. Findings An aeroengine rotor is assembled to verify the method. The results show that the predicted eccentricity differed from the measured eccentricity by 6.1%, with a comprehensive error of 8.1%. Thus, the optimization method has certain significance for rotor assembly error analysis and assembly process optimization. Originality/value In view of the error analysis in the stacking assembly of aeroengine rotor, an innovative optimization method is proposed. The method provides a novel approach for the aeroengine rotor assembly optimization and is applicable for the assembly of high-pressure compressor rotor and high-pressure turbine rotor as the rotor core assembly.

Physics-informed distributed modeling for CCF reliability evaluation of aeroengine rotor systems

Reliability evaluation of aeroengine rotor systems is often characterized by multiple correlated frail sites and multiple coupled failure modes, leading to the traditional integral modeling methods or separate modeling methods prone to unacceptable computing efficiency or accuracy. In this case, by absorbing the physics-informed thought into the distributed modeling, a novel physics-informed distributed modeling (PIDM) method is presented. In PIDM, based on the physical information of frail sites and failure modes, the complex reliability problem is first decomposed into several distributed sub-problems (i.e., multi-variate, multi-disciplinary and multi-objective); moreover, based on the extreme gradient boosting (XGB) algorithm, the mutually nested distributed sub-models of multi-variate mapping, static/dynamic multi-disciplinary and multi-objective reliability are established; finally, the multi-variate synchronous mapping, combined cycle fatigue (CCF) prediction, and integrated reliability analysis are achieved by the built distributed sub-models. To verify the effectiveness of the proposed method, the CCF reliability analysis of a typical high-pressure compressor rotor is performed. The comparisons of the direct Monte Carlo method, support vector regression (SVR), multilayer perceptron (MLP), extreme gradient boosting (XGB), MLP-based PIDM (MLP-PIDM), and XGB-based PIDM (XGB-PIDM) reveal that the proposed PIDM method holds the computing benefits on accuracy and efficiency. The current effort shed the light on the theoretical development of physics-informed modeling approaches for complex reliability evaluations.

Application of Adjoint-Enhanced First Order Second Moment Method for Robust Design Optimization of a High Pressure Compressor Rotor

Abstract This paper applies the computationally advantageous combination of a differential analysis technique and the adjoint method in the context of uncertainty quantification to the gradient-based robust design optimization of aerofoils. First, the accuracy and feasibility of the method is evaluated using an analytical test function. The method is subsequently applied to the robust design optimization of a rotor blade under geometric uncertainty, that is based on a parametrized population of optically 3D scanned blades. A significant improvement of the robustness w.r.t. the isentropic efficiency of the rotor is reported.

IMPROVEMENT IN DURABILITY OF WELDED DRUMS COMPRESSOR ROTORS BY TREATMENT IN FLUDIZED BED OF ABRASIVE MATERIAL

Purpose. Improvement in durability of welded drums of compressor rotors of the gas turbine engines by treatment in fluidized bed of abrasive material.
 Methods of research. The investigations were carried out on the welded drums of rotors of high pressure compressor (HPC) of the D-36 aero engine. The drums were treated in the fluidized bed of abrasive material of the АПС-600 installation without the nozzles and with the use of special air-blast nozzles.
 Test specimens for checking durability were cut of HPC stages 1 and 4 discs after carrying out various machining processes.
 Durability testing of test specimens cut of HPC stage 4 disc was carried out at the УМЭ-1 ТМ machine at temperature Т = 400 °С.
 Durability testing of test specimens cut of HPC stage 1 disc was carried out at the ЭДЦ-20 machine at temperature Т = 20 °С.
 Results. It was found that averaged cyclic life of the test specimens made of titanium alloy ВТ-9, treated in the fluidized bed of abrasive material being a part of the drum with the use of air-blast nozzles, is 2.2 times higher than the averaged cyclic life of the test specimens treated in the fluidized bed of abrasive material without the use of air-blast nozzles at temperature Т = 400 °С.
 When treating the drum of the HPC rotor in the fluidized bed of abrasive material without the use of air-blast nozzle, better treatment of the bottom of the slot and corners of its end faces is ensured.
 Low-temperature annealing of the drum at temperature Т = 550 °С is preferable than at Т = 750 °С, since cyclic life of the test specimens annealed at Т = 550 °С is somewhat higher.
 Annealing at Т = 750 °С removes completely strengthening effect obtained by treatment in the fluidized bed of abrasive material.
 Successive treatment of the disc as a part of the drum in the fluidized bed restores its durability to the initial state.
 Scientific novelty. It was demonstrated that treatment of the welded drums of the compressor rotors in the fluidized bed of abrasive material with the use of special air-blast nozzles improves the quality of treatment and cyclic life if compared to the treatment without air-blast nozzles.
 Annealing temperature for the compressor rotor drum, which ensures high average cyclic life, was established.
 Practical value. There were offered technological procedure and modes of treatment of the welded drums of the HPC of the D-36 aero engine in the fluidized bed of abrasive material with the use of special air-blast nozzles, ensuring durability 2.2 times higher than without the use of the air-blast nozzles.

A New assembly precision prediction method of aeroengine high-pressure rotor system considering manufacturing error and deformation of parts

The accurate prediction of high-pressure rotor system assembly precision before assembly is the premise of improving aeroengine assembly quality and performance. The existing assembly precision prediction models (APPM) only consider the manufacturing error factors of parts, but rarely involve the deformation of parts under load, so there is a certain gap between the prediction results and the actual situation. This article studies the construction method of APPM considering the manufacturing error and deformation factors of parts. Firstly, the fitting algorithm is used to obtain the fitting deformation surface(FDS) of each mating surface under load, which provides the basis for constructing assembly error model considering manufacturing error and deformation of parts; secondly, according to the relative position relationship between the FDS and the datum plane, the error model of each mating surface of the assembly is effectively constructed by the small displacement torsor theory; thirdly, according to the different errors of each fitting end face, a prediction model of assembly precision for two rough surfaces is constructed by homogeneous coordinate transformation method; finally, a high-pressure compressor rotor system is used as an example to verify the effectiveness of the precision prediction model. The results show that the prediction results based on the proposed model are closer to the actual conditions. This paper provides an effective prediction model for high-pressure rotor system assembly precision, and has important application value for improving the assembly quality and performance of aeroengine.

Analisis Pengaruh High Pressure Compressor Rotor Clearance Terhadap Exhaus Gas Temperature Margin pada CFM56-7

Exhaust Gas Temperatue is an parameter where the hot gases’s temperature leave the gas turbine. Exhaust gas temperature margin is the difference between highest temperature at take off phase with redline on indicator (???????????? ???????????????????????? °????=???????????? ????????????????????????????−???????????? ???????????????? ????????????). EGTM is one of any factor to determine engine performance. A good perfomance of an engine when it has a big margin (EGTM), during operation of an engine the EGTM could decrease untill 0 (zero). So many factors could affect EGTM deteroration there are: distress hardware such as airfoil erosion, leak of an airseals, and increase of clearance between tip balde and shroud. Increase of clearance happens in high pressure compressor rotor clearance. In CFM56-7 have 9 stage(s) of high pressure compressor and each stage give the EGT Loses. The calculation of EGT Effect/Losses is actual celarance – minimum clearance x 1000 x EGT Effect °C, where actual clearance define by the substraction of outside diameter’s rotor with inside diameter’s shroud, minimum clearance define in the manual, 1000 is adjustment from mils/microinch to inch, and EGT Effect is temperature that define in the manual. The analysist had done with 6 (six) engine serial number and proceed by corelation that shown linkage between clearance and EGT Effect, the corelation is strong shown the result of corelation (r) is 0.994275999 or nearest 1.

Model for evaluating the end runouts of a rotor with parallel connections of parts

The existing methods for calculating the assembly dimensional chains of aircraft engine rotors are analyzed. The factors that have a significant impact on the reliability of the calculation of the controlled assembly parameters of the product are identified. One of these factors is the existence of parallel connections of parts in the rotor. In the drum disk rotors, parallel rotor connections are formed by mating their parts along several end surfaces in the axial direction. A mathematical model is proposed that allows taking into account the parallel connections of the rotor parts. The form of relationship between rotor end run-outs and amplitudes of deviations of the shape of the mating surfaces of the parts and their angular positions in the unit is determined. The determined dependence includes many coefficients that allow taking into account the amplitudes of deviations of the shape of the mating surfaces, parallel connections of parts in the rotor, and their angular position. Determination of dependence coefficients values is solved as a problem of regression analysis. The initial data for obtaining the dependence are formed using the developed parameterized finite element model (FEM) of a part of the rotor of an aircraft engine high-pressure compressor (HPC). The results of research of end run-outs of control surfaces of disks of the considered HPC rotor assembly part are presented. The values of the dependence coefficients for assessing the end run-outs of the rotor are determined.

Surge Fault Detection of Aeroengines Based on Fusion Neural Network

Aeroengine surge fault is one of the main causes of flight accidents. When a surge occurs, it is hard to detect it in time and take anti-surge measures correctly. Recently, people have been applying detection methods based on mathematical models and expert knowledge. Due to difficult modeling and limited failure-mode coverage of these methods, early surge detection cannot be achieved. To address these problems, firstly, this paper introduced the data of six main sensors related to the aeroengine surge fault, such as, total pressure at compressor (high pressure rotor) outlet (Pt3), high pressure compressor rotor speed (N2), power level angle (PLA), etc. Secondly, aiming at preprocessing of data sets, this paper proposed a data standardization preprocessing algorithm based on batch sliding window (DSPABSW) to build a training set, validation set and test set. Thirdly, aeroengine surge fault detection fusion neural network (ASFDFNN) was provided and optimized to improve the detection accuracy of aeroengine surge faults. Finally, the experimental results showed that the model achieved 95.7%, 93.6% and 94.7% in precision rate, recall rate and F1_Score respectively and consequently it can detect the aeroengine surge fault 260 ms in advance.

A Coupled Effect Model of Two-Position Local Geometric Deviations on Subsonic Blade Aerodynamic Performance

The coupled aerodynamic performance influence of multiple deviations in the blade profile is a complex phenomena. In this paper, a high-pressure compressor rotor blade mid-span profile is studied with the design of experiments (DOE), numerical simulation and surrogate model to analyze the influence of the deviations on the blade aerodynamic performance. The purpose of this work is to provide a new rapid evaluation approach for the blade aerodynamic performance under multiple geometric deviations influence. First, the Hicks-Henne function was used to model the local geometric deviations of the blade profile, and the blade aerodynamic performance was calculated by using the computational fluid dynamics tools. By analyzing the calculation results, the momentum thickness of the boundary layer, the deviations height and the distance between the deviations are combined into a coupled effect model. Then, the coupled effect model was used to rapidly evaluate the blade aerodynamic performance when the two-position local geometric deviations exist on the blade surface. Finally, the evaluated performance were compared with the results predicted by a high-precision surrogate model, which verifies the high accuracy of the coupled effect model in evaluating the positive incidence range of the blade profile.

Study on Propagation Depth of Ultrasonic Longitudinal Critically Refracted (LCR) Wave.

The accurate measurement of stress at different depths in the end face of a high-pressure compressor rotor is particularly important, as it is directly related to the assembly quality and overall performance of aero-engines. The ultrasonic longitudinal critically refracted (LCR) wave is sensitive to stress and can measure stress at different depths, which has a prominent advantage in stress non-destructive measurements. In order to accurately characterize the propagation depth of LCR waves and improve the spatial resolution of stress measurement, a finite element model suitable for the study of LCR wave propagation depths was established based on a wave equation and Snell law, and the generation and propagation process of LCR waves are analyzed. By analyzing the blocking effect of grooves with different depths on the wave, the propagation depth of the LCR wave at seven specific frequencies was determined in turn. On this basis, the LCR wave propagation depth model is established, and the effects of wedge materials, piezoelectric element diameters, and excitation voltages on the propagation depth of LCR waves are discussed. This study is of great significance to improve the spatial resolution of stress measurements at different depths in the end face of the aero-engine rotor.

Measurement of fastening force using dry-coupled ultrasonic waves

The accurate measurement of assembly fastening force of the high-pressure compressor rotor is of great significance to improve the structural connection quality and comprehensive performance of aero-engine. To solve the problem that liquid couplant reduces the measurement accuracy and causes the surface of the bolt to rust and corrode, a method of measuring the fastening force with dry-coupled ultrasonic wave was proposed in this paper. The measurement model of the fastening force including the thickness of the protective film of the ultrasonic transducer and the couplant was established, and the influence of the thickness variation in the couplant on the measurement accuracy of the fastening force was analyzed. Based on the propagation model of the ultrasonic wave on the heterogeneous interface and the principle of ultrasonic dry coupling, a coupling device was designed. At last, the experiment of fastening force measurement based on dry coupling and liquid coupling ultrasonic wave was carried out and compared. The experimental results show that the average relative errors of the fastening force measurement based on dry coupling and liquid coupling ultrasonic wave are 2.13% ± 0.42% and 3.15% ± 0.80%, respectively. Therefore, the dry coupling method can be as good or better in measuring the accuracy of the fastening force. Furthermore, it also overcomes the limitations of the liquid coupling method, which should make it more suitable to the measurement of fastening force in the aerospace field.

Optimization of a Welded Portion of the High Pressure Compressor Rotor

An integrated method for optimization of a welded portion of a high pressure compressor rotor as a single object is developed. The minimum mass criterion is used in the optimization method; limitations are imposed in accordance with strength and stiffness requirements. An example is given of optimization of a typical welded portion of a high pressure compressor rotor resulting in mass reduction and increase of cyclic durability.

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.

A brief overview and metallography for commonly used materials in aero jet engine construction

Abstract Aluminium, titanium, and nickel base alloys are mostly and widely used for aircraft jet engine construction. A proper evaluation of its microstructure is important from working safety point of view. To receive a well prepared sample of microstructure, some important steps have to be undertaken. Except for proper grinding and polishing of a sample, structure developing is a significant step, too. In order to develop microstructure various chemical reagents were used to achieve the best results for microstructure evaluation. The chemical reagents were used according to the previous knowledge and some new ones were also tested. Aluminium AK4-1č, titanium VT – 8, and nickel VŽL – 14 and ŽS6 – U alloys were used as an experimental materials. Alloy AK4-1č is used for fan blade production with working temperatures up to 300°C. It is a forged piece of metal machined down into final shape by five-axe milling machine. Alloy VT – 8 is used for high pressure compressor rotor blade production with working temperatures up to 500°C. Blades are forged as well and finally grinded. Finally nickel base alloys VŽL – 14 and ŽS6 – U are used for turbine blade production with working temperatures up 950°C. Blades for turbine are casted into mould with reducible models.

Adaptive fuzzy control of the parameters of an aircraft engine high pressure compressor

Multimode control of a turbojet bypass engine (TBE) is carried out by means of an electronic controller (EC). The EC program controls the flight dynamic if controlled disturbances take place in a predetermined range. Uncontrolled disturbances (exorbitant variation of outboard air parameters, fuel flow rate caused by fuel leakage in a dosing device, etc.) can spoil the normal conditions of an aircraft flight. A way to correct EC programs by means of adaptive fuzzy control of all loops is presented. We show how to replace a normal air pressure controller with an adaptive fuzzy controller (AFC) using the example of a loop for controlling the air pressure behind a high pressure compressor (HPC). The adaptive fuzzy controller consists of a fuzzifier and an activation function unit. A means to select term sets for a fuzzifier with respect to the dosing device’s properties, which excludes static error in the loop of the HPC rotor frequency, is presented. A neuron with step-by-step learning is used in adapting the fuzzifier. An algorithm for correcting the activated grades of the fuzzifier’s membership at a certain time moment varies the synapse’s numerical value in each iteration. The presented oscillograms show that it is possible to use adaptive fuzzy control. It is also possible to replace the normal controller with an adaptive fuzzy controller in other control loops (for controlling the temperature in a combustion chamber, the rotor frequency of a high pressure compressor, etc.).

Mixing layer resonance under high-speed stream forcing

In the majority of fluid–structure interaction problems, the biggest challenge lies in the fundamental understanding of the flow physics. Forced mixing layers is an important phenomenon found in many cases of flow-induced vibrations and acoustics. The response of a mixing layer to high-speed stream acoustic forcing is investigated with a theoretical and experimental approach. Two different experiments demonstrating the fluid mechanic phenomenon are presented. The first experiment consists of a circular jet impinging on a vibrating plate. The second experiment demonstrates the mixing layer resonance in the context of a fluidelastic instability causing high-amplitude vibrations in gas turbine high-pressure compressor rotor blades. Both the plate and the adjacent blade vibration induce an acoustic feedback that propagates within the jet and blade tip clearance flow, respectively. The resonance was found to occur when the feedback wavelength matched either the jet-to-plate or the inter-blade distance. In both experimental cases, the resonance condition has been simply modeled by the coincidence of a 1D feedback wave, which propagates upstream at reduced velocity by the high-speed flow. The coupling between the jet induced mixing layer and the feedback wave is assumed to naturally occur when one of the wave crests reaches the separation edge. The objective of this study is to improve the understanding of the coupling mechanism between an emanating shear layer and the acoustic forcing originating within a fast flow stream. The study is based on a simplified analytical model in order to enlarge the current understanding of the mixing layer receptivity to the more specific case of its response to high-speed stream forcing. To identify the mixing layer resonant modes, an analytical resonance condition is proposed. It is found that the mixing layer response becomes spatially resonant for specific source locations downstream in the high-speed flow. The study also provides an analytical mean to capture the critical source location periodicity that has been experimentally observed. The resulting theoretical prediction of the resonant source locations is in good agreement with the experimental data. Therefore, it supports the stream forced mixing layer analytical model and the proposed spatial resonance condition. The simple 1D reduced speed feedback wave model, which has been used to identify the experimental resonance conditions, is also in good agreement, and thus validated, with the results of this study.

Effect of combined cycle fatigue on Ti6242 fatigue strength

Gas turbine blades are subjected to very high levels of stress and temperature during each engine operating cycle. Superimposed upon this quasy-steady state condition, blades also endure vibrational stresses induced by local perturbations. When characterizing advanced manufacturing candidate materials, conventional fatigue tests (low cycle and high cycle fatigue standard tests independently) fail to represent the complex set of stresses generated in gas turbine blades. Combined cycle fatigue (CCF) laboratory tests comprising LCF and HCF at service high temperatures have been completed on a titanium alloy used in intermediate pressure and high pressure compressor rotor and stator components. The aim of the study was to determine the influence of the combined damage mechanisms on the expected fatigue strength of gas turbine components. Standard test rigs have been modified in order to achieve this new and severe testing scenario. The effect CCF stresses was evaluated at a testing temperature of 350°C. Results showed that the effect of superimposing LCF to HCF should be considered in blade design, especially at fillet radius.

THE FORGING PROCESS OF AIRCRAFT ENGINES TURBINE BLADES

The aim of this paper was to obtain information about optimum conditions of the forging process of a blade made from high-temperature creep resisting alloy called Inconel 718. A three-dimensional rigid-plastic finite element method (FEM) of forging process of high-pressure compressor rotor's blade analysis has been performed. Various forging process variants were analysed, including isothermal conditions.