Gas Turbine Engine Compressor Research Articles

Модельные технологические решения по обеспечению качества изделий при обработке точением

Model technological solutions for ensuring the quality of products during turning processing are proposed. A study of the workability of materials based on the ISO classification was carried out. To clarify the workability group, the MS code system was used, which allowed not only to clarify the workability group and the terms of delivery of the workpiece, but also to determine the specific cutting force. Based on the classifier-handbook, the parameters of the prefabricated cutting tool were selected and the technological modes for the prescribed processing conditions were determined. Based on the obtained specific cutting force, an algorithm for simplified determination of the cutting force is implemented. A study of the vibrations of the technological machining system was performed, which made it possible to formulate dynamic models for the system as a whole, subsystems of the workpiece and cutting tool. In relation to the dynamic model of the workpiece subsystem. The model has been simplified. A dynamic cutting characteristic was constructed, which made it possible to build an autonomous dynamic model of the technological system. In relation to this model, parameterization was performed, which made it possible to determine the stiffness coefficients of the fixed workpiece, as well as the inertia and dissipation coefficients. In relation to the selected processing conditions, the accuracy and roughness of the treated surface were checked. A study of the stability of the technological system of mechanical processing based on the algebraic Hurwitz criterion is carried out. In the general formulation, the Hurwitz determinant and stability conditions are obtained. The performed complex of calculations showed that the conditions determined by the Hurwitz stability criterion are met. This indicates that when processing in the selected modes, quiet cutting is carried out, which ensures the requirements for the quality of the product. The construction of the stability boundary can be performed in the space of variable technological parameters using the D-partitioning method or based on the R. Loeb algorithm. As a result of constructing the stability boundary in the space of variable parameters, the area of permissible technological modes is highlighted, which is the basis for solving the problem of processing control in accordance with the formed efficiency criteria. The complex of technological solutions presented in the article was deployed and implemented in relation to the turning of coating discs, as well as discs of turbines and compressors of gas turbine engines.

Equivalent diffusion ratio distribution in the UGT15000 low pressure compressor

With spatial and temporal periodicity approach the results of steady state CFD simulation of compressible air flow in each blade-to-blade row of low pressure compressor (LPC) of the UGT15000 gas turbine engine (Ukrainian SSR constr. and prod.) are presented. Near the nominal rotation speed compressor map and adiabatic efficiency of each stage and the flow angles at leading and trailing blade edges are predicted and calculated. At the average radius stage flow and stage load coefficients are determined. It has been established that the reason of the low efficiency of the last stage is associated with the low stage reaction (≈0.5) leading to high airflow swirling at outlet. For LPC optimization a calculation of the radial distributions of the equivalent diffuser coefficient was performed. It is shown that the greatest total pressure losses are in the stator rows of stages 9, 8, 5 and in the rotor rows of stage 9. High losses are also shown on the upper half of the 4th, 3rd, 2nd stator blades and the lower half of the 1st rotor blades. The identified potential for increasing the efficiency of the LPC is up to 2% without significant construction changes.

A flow field around a cylindrical probe in proximity to stator blades and its effect on the measurements

In multistage axial compressors of gas turbine engines, there is a need for a detailed understanding of the flow field in between the blade rows, which could be obtained by spanwise traversing. This requires a pneumatic probe to be immersed into the flow path between the blade rows, where the space is limited. Therefore, the probe will affect the flow field around it and in the blade channel, and the probe readings will be affected by that flow field. As a result, these probe readings cannot be translated to flow parameters based on just the freestream calibration characteristics, obtained in the idealized wind tunnel. In our paper, we provide a computational analysis of the flow field around the cylindrical probe in a constrained inter-blade-row environment at different circumferential locations and flow conditions both upstream and downstream of the stator blade row. It is shown that the flow angle measurement error can reach up to five degrees in the mid-pitch locations compared to undistorted flow, and dynamic head measurements can be up to 30% away from the actual mean values of the flow. These deviations are shown to be caused by flow field interaction inside the blade channel and, as a result, measured values, obtained during industrial compressor testing, could be corrected accordingly. The universal correction procedure is proposed for further use in the industry

Impact of surface roughness on gas turbine engine fan and compressor rotor

Impact of surface roughness on gas turbine engine fan and compressor rotor

Ідентифікатор передпомпажних станів компресора авіаційного ГТД

Surging is a dangerous operating mode of the compressor; in this mode, destabilization and flow stall occur, causing intense pressure fluctuations of the working fluid in the flow path or gas-dynamic shocks. Surging phenomena in gas turbine engines lead to a loss of thrust, are accompanied by an increase in the temperature of gases in front of the turbine and an increase in the level of vibration because of large amplitudes of pressure pulsations and mass flow rates along the engine path. The possibility of occurrence of these phenomena is a serious obstacle to increasing the reliability of gas turbine engines, including flight safety in general. The study of non-stationary processes in compressors shows that surge is preceded by such phenomena as pre-surge and rotating stall, which are called the pre-surge state of an aircraft GTE compressor. The study aims to develop informative criteria for the timely detection of the pre-surge state of the compressor of an aircraft gas turbine engine and to build an anti-surge algorithm for the functioning of the proposed identifier based on the developed criteria. To assess the stability of pressure pulsation periodicities, it is proposed to process and compare the signals received from fast-response sensors located in the compressor. In this article, informative criteria are formed for identifying the pre-surge state of an aircraft GTE compressor. The first criterion is the level of stability of the periodic component of pressure fluctuations. The second criterion is the limit of the relative range of pressure pulsations in the compressor. With the help of the obtained criteria, the pre-surge area is divided into pre-surge, weak stall, and intense stall. Such detailing makes it possible to prevent erroneous wear (in high-cost modes) and premature wear (in case of non-intensive disruption) of the anti-surge control system. Based on the criteria obtained, an algorithm for anti-surge adjustment was developed, according to which the anti-surge system is prepared for testing when a pre-surge or weak stall occurs in the compressor flow path, and is triggered when an intense stall occurs. Based on the anti-surge algorithm, a scheme for identifying the pre-surge states of an aircraft GTE compressor is proposed in the form of a three-level alarm and protection system. Recommendations for the practical implementation of the identifier are described.

Development of an anisokinetic particle sampling probe for use in a gas turbine engine compressor

Sand and dust particle ingestion is an inevitability for aircraft operating in arid environments. For conventional takeoff and landing aircraft, significant dust can be ingested into the gas turbine powerplant. Helicopters and vertical takeoff and landing (VTOL) aircraft are at especially high risk due to their tendency to blow significant debris into the air during takeoff and landing operations. The present study highlights the development of an anisokinetic particle sampling probe for use in aircraft engines to obtain real-time measurement of ingested particles often present in these harsh environments. Offtake of particles during engine operation in dusty conditions will provide researchers with an improved understanding of particle breakage tendency and component erosion susceptibility. Three foundational studies were conducted to establish a baseline understanding of probe performance: an aerodynamic study, a particle tracking study, and a particle sampling study. These studies were performed using the Free Jet rig at the Advanced Propulsion and Power Laboratory at Virginia Tech. Particles as large as 1.3 mm were sampled at Mach numbers where M = (0.25, 0.70) and yaw angles ranging from 0° to 45° relative to freestream, conditions not previously investigated. Results indicate that the probe operates sub-isokinetically throughout the full range of test conditions and that probe aerodynamic capture efficiency is inversely proportional to both Mach number and yaw angle. However, this efficiency limitation does not notably influence the sampling probe’s ability to capture the test dust of interest. While the presence of the probe in the flow does result in an airflow velocity reduction of up to 55%, due to their relatively large Stokes numbers the particles of interest only experience a decrease of roughly 5%. These results indicate that this probe is capable of providing researchers with valuable particle size and shape information through effective particle sampling at particle sizes (100 μm ≤ dp ≤ 1,300 μm) and speeds (M ≥ 0.25) not previously investigated.

Особливості варіювання просторової форми лопаток при аеродинамічному вдосконаленні компресорів авіаційних газотурбінних двигунів

This work is devoted to the development of procedures for the aerodynamic improvement of gas-turbine engine axial-flow compressor blade rows. The aim of the work is to compare the efficiency of two methods for impeller blade shape variation in the aerodynamic improvement of an aircraft gas-turbine engine two-stage fan. The first method consists only in varying the blade profile angle along the blade height, while the second consists in varying the blade profile angle and geometrical parameters. The features of the approach used in the solution of this problem are as follows: formulating quality criteria as the mean integral values of the power characteristics of each impeller of the fan over the operating range of the air flow rate through the impeller and searching for advisable values of the impeller blade parameters by scanning the independent variable range at points that form a uniformly distributed sequence of small length. The basic tool is a numerical method developed at the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine, which simulates 3D turbulent gas flows in the compressor stage blade channels using the complete averaged Navier¬–Stokes equations. The results of multiparameter calculations of 3D turbulent gas flows show that at the initial stage of the aerodynamic improvement of compressor blade rows varying the blade profile angle alone is more efficient; however, increasing the number of elements of the uniformly distributed sequence of points in the variable range increases the possibility of finding a point at which the blade profile aerodynamic characteristics significantly improve. The results obtained are expected to be used in the aerodynamic improvement of gas-turbine engine compressor blade rows.

АНАЛИТИЧЕСКОЕ ОПИСАНИЕ КОНСТРУКТОРСКО-ТЕХНОЛОГИЧЕСКИХ И ГЕОМЕТРИЧЕСКИХ ИСХОДНЫХ ДАННЫХ ДЛЯ ПРОГНОЗИРОВАНИЯ ПРОГИБА КОМПРЕССОРНЫХ ЛОПАТОК ГАЗОТУРБИННОГО ДВИГАТЕЛЯ ПРИ ПОПЕРЕЧНОМ СТРОЧНОМ ФРЕЗЕРОВАНИИ ПРОТОЧНОЙ ЧАСТИ НА СТАНКАХ С ЧИСЛОВЫМ ПРОГРАММНЫМ УПРАВЛЕНИЕМ

When milling the profile of the flow path of the vanes of a gas turbine engine on multi-axis machines, a technological scheme of transverse line shaping is used, according to which the vane is rotated around its own axis and processed with a tool with a spherical working surface that performs rotation and interpolated axial movement. The required indicators of the surface quality of the vane airfoil profile (profile accuracy and surface roughness) are provided by assigning a combination of controlled parameters of the milling mode. However, at present there are no recommendations on the calculation and assignment of a combi-nation of controlled parameters for the milling of complex-profile surfaces, which is the profile of the flow path of the gas turbine engine compressor vanes. The accuracy of the profile of the vane feather for each line and the angle of its rotation will be deter-mined by the value of the vane deformations, which should not exceed the tolerance for its manufacture. From the analysis of ge-ometric relationships in the contact zone of the mill with a spherical working surface and a curved profile of the flow path of the vane, dependencies were established for calculating the component of the milling force and its projection on the Y axis, as well as the effective diameter of the mill necessary to calculate the total deformation of the vane. The methodology is described and analyt-ical expressions are obtained for calculating and assigning a combination of controlled parameters of the milling mode that provide the required accuracy of the vane airfoil profile when developing a control program and milling a vane on a Computer Numerical Control machines.

Влияние формы заготовок на остаточные напряжения при линейной сварке трением

Linear friction welding is an advanced technology for manufacturing titanium blisks for gas-turbine engine compressors, which are subjected to stringent requirements for cyclic strength and dimensional accuracy. Substitution of conventional butt joints with more technological T-shape joints is a promising area, which provides reducing of the pre-welding machining costs. The introduction of T-form joints requires additional research of thermal distribution specifics and strain-stress state formation in the welding process and after its end. Therefore, the study of residual stresses in titanium alloy T-shape joints produced by linear friction welding is topical. The paper investigates the residual stresses in imitating welded blisk joints. The authors consider the results of welding where the blade imitator has a reamed relief of a smaller section. The finite element model covering forging, cooling, and disassembly of welded specimens is offered. The authors developed the model in ANSYS Workbench to describe the strain-stress state of welded specimens, which allows for estimating the residual stress levels and spreading. The main distinctive feature of the model is an accounting of asymmetric temperature distribution obtained by finite-difference solving of a T-shape joint thermal problem and weld shape simulation obtained as a result of welded joints metallographic research. The presented model allows the evaluation of the residual stresses in joints. The distribution of residual stresses in T-shaped welded joints is specific – compressive stresses existing in a weld are balanced by tensile stresses acting at a distance of 1 mm from the joint. The formation of compressive stresses in a weld is caused by plastic deformation due to the forging force action.

Application of the vacuum casting technology in the production of small-size gas turbine engine blade machines

The application of casting technologies in the production of parts and assemblies of small-size gas turbine engines is justified in the paper. The technology of vacuum casting in gypsum molds was tested during the production of an experimental centrifugal compressor of a small-size gas turbine engine. On the basis of a 3D model of the designed centrifugal compressor, computational studies of vacuum casting were carried out and rational parameters of the technological process were determined. Prototypes of the developed centrifugal compressor of a small-size gas turbine engine were made. The results of calculations and the performed technological experiment confirmed the fill rate of the gating form and the absence of short pour. The distribution of shrinkage porosity and cavities corresponds to the design values and is concentrated in the central part of the casting that is subjected to subsequent machining. The area of the blades, disc and sleeve is formed without defects. The use of casting technologies in the production of parts and assemblies of small-size gas turbine engines assures the required quality with a comparatively low price of the finished product, making it possible to achieve the balance between the cost of the technology and the quality of the product made according to this technology.

Dependence of thermogasodynamic parameters of operation of a helicopter gas turbine engine on its operation and dust in the atmosphere

By scanning the blades of the impeller and the guide devices of the full-scale compressor of the helicopter gas turbine engine, a solid-state design model of the compressor with a tunable geometry in the height of the blades was developed. The empirical dependencies of the values of nonlinear wear of the blade of the 1st stage of the compressor on the operating time and the concentration of dust in the air are presented. A block diagram of the procedure for calculating the characteristics and parameters of a helicopter gas turbine compressor is presented. When calculating the flow in the gas–air path of the compressor, the numerical solution of the Navier–Stokes equations averaged by Reynolds, the finite element method in combination with the establishment of patterns of erosive wear of the blades depending on the operating conditions of the engine is used. Mathematical dependencies of the thermogasodynamic parameters of the helicopter gas turbine compressor operation on its operating time and the dustiness of the atmosphere are obtained using modeling. The obtained results can be used in the development of a methodology for automated monitoring of the wear condition of the compressor blades of a helicopter gas turbine engine according to its thermogasodynamic parameters, applied to various climatic operating conditions.

Метод оптимизации многоступенчатых компрессоров

The paper presents an algorithm for seeking an optimal blade configuration for multistage axial-flow compressors. The primary tool behind the algorithm is 3D CFD simulation, augmented by commercial optimisation software. The core of the algorithm involves feeding an initial data vector to the parametric simulation module so as to form a "new" blade geometry, which is then transferred to 3D computational software. The results obtained are further processed in a program that implements the algorithm for seeking the optimum and forms a new input data vector to achieve the set goal. We present a method of parametrically simulation the blade shape, implemented in a software package, making it possible to describe the shape of the compressor blade profiles using a minimum number of variables and to automatically change the shape in the optimisation cycle. The algorithm developed allows the main parameters of compressor operation (efficiency, pressure ratio, air flow rate, etc.) to be improved by correcting the profile shape and relative position of the blades. The algorithm takes into account various possible constraints. We used the method developed to solve practical problems of optimising multistage axial compressors of gas turbine engines for various purposes, with the number of compressor stages ranging from 3 to 15. As a result, the efficiency, pressure ratio and stability margin of gas turbine engines were increased

Дослідження напруженого стану міжпазових виступів диска робочого колеса компресора з урахуванням впливу відцентрових сил та розладу частот коливань лопаток

The authors present the results of the numerical experiments on the determination of the effect of compressor blade mistuning on the stress state of a blade-disk joint under the action of the centrifugal forces and their comparison with results of full-scale experiments. The analysis of the results of the previously performed experimental and numerical experiments on the determination of the possible sources of blade-disk joints failure shows their stress state strongly depends on the disk rim thickness and the blade vibration mode and their mistuning. It indicates the need for further studies to determine the causes of cracks and possible fracture of blade-disk joints in the compressor of a gas-turbine engine. The task is solved in the variation of the dovetail-rim region thickness of the disk of the 1st stage of the low-pressure compressor of the gas-turbine engine D-36, which has blades without shrouds. The finite element models of the blade assembly, are used to perform the investigation. Obtained results of the vibration stress analysis of blade-disk joints in the compressor both modifications under the action of the centrifugal forces have good agreement with the data of full-scale experiments and numerical data using rod models. The presented results of the numerical experiments indicate the possibility of reducing vibration stresses by analyzing and choosing parameters that affect the rigidity of the coupling of the blades, which is of practical importance for increasing the vibration reliability of the compressors of gas-turbine engines.

Інтелектуальний регулятор запасу газодинамічної стійкості компресора авіаційного ГТД

The gas-dynamic stability of compressors of aircraft gas turbine engines is one of the most important conditions that determine their reliability and level of flight safety. Unstable operation of the compressor in the engine system (surge) leads to loss of thrust accompanied by an increase in gas temperature in front of the turbine and increased vibration because of large amplitudes of pressure pulsations and mass flow through the engine path. To improve the parameters of ACS aviation gas turbine engines are increasingly using regulators built using fuzzy logic algorithms. The implementation of fuzzy control algorithms differs from classical algorithms, which are based on the concept of feedback and reproduce a given functional dependence or differential equation. These functions are related to the qualitative assessment of system behavior, analysis of the current changing situation, and the selection of the most appropriate for the situation supervision of the gas turbine engine. This concept is called advanced management. ACS gas turbine engines with fuzzy regulators are nonlinear systems in which stable self-oscillations are possible. Approximate methods are used to solve the problems of analysis of periodic oscillations in nonlinear systems. Among them, the most developed in theoretical and methodological aspects is the method of harmonic linearization. The scientific problem is solved in the work – methods of synthesis of intelligent control system with the fuzzy regulator as a separate subsystem based on the method of harmonic linearization and design on its basis of fuzzy ACS reserve of gas-dynamic stability of aviation gas turbine engine. Based on the analysis of the principles of construction of fuzzy control systems, it is shown that the use of fuzzy logic provides a new approach to the design of control systems for aviation gas turbine engines in contrast to traditional control methods. It is shown that the fuzzy controller, as the only essentially nonlinear element when using numerical integration methods, can be harmonically linearized. Harmonic linearization allows using the oscillation index to assess the quality in the separate channels of fuzzy ACS aviation gas turbine engines. A fuzzy expert system has been developed for optimal adjustment of the functions of belonging of typical fuzzy regulators according to quality criteria to transients.

Manufacture and Dimensional Verification of a Gas Turbine Engine Compressor Blade Produced via Additive Manufacturing Method using Ti6Al4V

Bu çalışmada, bir gaz türbin motoruna ait kompresör palesinin, Ti6Al4V alaşımından Eklemeli Üretim (EÜ) yöntemiyle üretimi ve boyutsal doğrulama ölçümleri gerçekleştirilmiştir. Havacılık endüstrisinde yaygın kullanılan Direkt Metal Lazer Sinterleme (DMLS) yöntemi seçilmiştir. İmalattan sonra her bir palede 87 adet olmak üzere toplam 174 yüzey ölçümü gerçekleştirilmiştir. Ölçümlerde, katı modele oranla, 1 Nu’lı palede ortalama +0,0944/-0,0809 mm arasında, 2 Nu’lı palede ortalama +0,1093/-0,0978 mm değişim olduğu belirlenmiştir. Her iki paledeki geometrik değişimin birbiri ile tutarlı sonuçlar verdiği gözlemlenmiştir. Sonuç olarak, bir uçar-parçanın uçağa takılmasına yönelik uçuşa elverişlilik çalışmaları konusuna girmeden, bir gaz türbin motoru kompresör palelerinin Ti6Al4V alaşımından EÜ yöntemi ile üretilebileceği ortaya konulmuştur.

Manufacture and Dimensional Verification of a Gas Turbine Engine Compressor Blade Produced via Additive Manufacturing Method using Ti6Al4V

Bu çalışmada, bir gaz türbin motoruna ait kompresör palesinin, Ti6Al4V alaşımından Eklemeli Üretim (EÜ) yöntemiyle üretimi ve boyutsal doğrulama ölçümleri gerçekleştirilmiştir. Havacılık endüstrisinde yaygın kullanılan Direkt Metal Lazer Sinterleme (DMLS) yöntemi seçilmiştir. İmalattan sonra her bir palede 87 adet olmak üzere toplam 174 yüzey ölçümü gerçekleştirilmiştir. Ölçümlerde, katı modele oranla, 1 Nu’lı palede ortalama +0,0944/-0,0809 mm arasında, 2 Nu’lı palede ortalama +0,1093/-0,0978 mm değişim olduğu belirlenmiştir. Her iki paledeki geometrik değişimin birbiri ile tutarlı sonuçlar verdiği gözlemlenmiştir. Sonuç olarak, bir uçar-parçanın uçağa takılmasına yönelik uçuşa elverişlilik çalışmaları konusuna girmeden, bir gaz türbin motoru kompresör palelerinin Ti6Al4V alaşımından EÜ yöntemi ile üretilebileceği ortaya konulmuştur.

Optimization Method of the Multistage Axial Compressors Using CFD Simulation

The current level of numerical methods of gas dynamics makes it possible to optimize compressors using 3D CFD models. However, the methods and means are not sufficiently developed for their wide application. This paper describes a new method for the optimization of multistage axial compressors based on 3D CFD modeling and summarizes the experience of its application. The developed method is a complex system of interconnected components (an effective mathematical model, a parameterizer, and an optimum search algorithm). The use of the method makes it possible to improve or provide the necessary values of the main gas-dynamic parameters of the compressor by changing the shape of the blades and their relative position. The method was tested in solving optimization problems for multistage axial compressors of gas turbine engines (the number of stages from 3 to 15). As a result, an increase in efficiency, pressure ratio, and stability margins was achieved. The presented work is a summary of a long-years investigation of the research team and aims at creating a complete picture of the obtained results for the reader. A brief description of the results of industrial compresses optimization contained in the paper is given as an illustration of the effectiveness of the developed methods.

Mixed-mode crack growth simulation in aviation engine compressor disk

This study focuses on a mixed mode crack growth analysis of an aviation gas turbine engine (GTE) compressor disk based on simulation principles. The operation damages occur via inclined surface cracks in a “dovetail type” disk-and-blade attachment. Based on the sizes and configuration of this attachment, three geometries for simulation models of the GTE compressor disk are proposed. The biaxial loading conditions of the simulation models are determined and numerically verified to reproduce the stress-strain state in critical zones of the compressor disk during operation. To determine the elastic-plastic fracture mechanical parameters, a 3D full-size finite element method (FEM) analysis is conducted for the GTE compressor disk and simulation models, based on operating loading conditions. As a result, the distributions of the new mixed-mode fracture resistance parameters in the form of equivalent elastic and plastic stress intensity factors along the crack front in the disk-and-blade dovetail-type attachment are applied to the experimental uniaxial and biaxial data for crack growth rate interpretation. The possibility of using the proposed simulation principles for an assessment of the structural integrity of GTE rotating disks is discussed.

Influence of chrome electroplating on fatigue strength of parts

The article is devoted to the results of studies, which have been conducted on parts with electrolytic chromium in order to determine the effect of the coating on fatigue strength of their. The work was performed in observance of standards, which are fixed in GOST RV 2840-001-2008. Samples for the tests were made from standard gas turbine engine compressor blades. We used a VEDS- 1500 electrodynamic vibration stand with an UMK-12K power amplifier to excite vibrations. It has been shown that the minimum endurance limit of 46 kgf/mm2 based on 2 ∙ 107 cycles, established on uncoated parts, didn't decrease during fatigue tests of compressor blades with an electroplated chrome layer. It should be stressed that the influence of the geometry of the chrome-plated part on the reduction of the endurance limit has been established.

Effect of the Rotor Blades Part Span Shrouds on the Transonic Axial Compressor Operation Map

The paper presents the results of the investigation of the part span shroud effect on the transonic axial compressor operational map. The objective of the paper is a 4.5-stage transonic low-pressure compressor of the aero-derivative gas turbine engine. The study was performed with using 3D RANS calculations via Ansys CFX commercial solver. The numerical model was validated with experimental data for 4 speedlines. The standard k-ɛ turbulence model demonstrated good agreement with the experimental data while SST model unexpectedly overpredicted mass flow rate. The strong effect of the mesh aspect ratio in the radial direction was observed. A significant decrease of the mass flow rate and the compressor efficiency was shown for the rotation speed range 85% – 95%. The maximum mass flow rate decreasing was estimated as 3.5%, maximum efficiency deterioration – 1.5%.