Multistage Axial Flow Compressor Research Articles

Optimization of Three-Dimensional Blade and Variable Stators for Efficiency and Stability Enhancement of Multistage Axial Flow Compressor at Variable Speeds

Abstract High efficiency and wide stability at variable speeds are equally important to the design and operation of multistage axial flow compressors. However, published research works on three-dimensional design optimization of compressors are so far mainly limited to a single blade row or stage at design speed due to the curse of dimensionality. Moreover, optimization of variable inlet guide vanes (IGV)/stators for off-design operations is carried out by using a rapid but low-fidelity prediction tool and is generally independent of design optimization of blade geometry. To tackle these issues, a three-dimensional holistic design and adjustment optimization method is developed in which both three-dimensional blade geometry and variable IGV/stators are optimized simultaneously for better efficiency and stability at design and off-design conditions. Metamodel-interpreted data mining method and adaptive infilling strategy are used respectively to enhance the capability of the metamodeling and optimization. The developed method is then applied to a modern highly loaded 3.5-stage transonic axial flow compressor at both design and part-design speeds. The results show that the stall margin is extended from 8.23% to 19.65% at 70% design speed while peak efficiency is slightly improved at design speed. The flow mechanisms responsible for the efficiency enhancement at design speed are mainly associated with the reduced total pressure loss in stators as well as inter- and intra-stage loading redistribution. The stability enhancement at 70% design speed is mainly achieved by loading the front blade rows while unloading the limiting rear blade row through variable IGV/stators adjustment. The developed holistic design and adjustment optimization method with the aid of metamodel-interpreted data mining is of great application value for the design and adjustment of advanced multistage axial flow compressors.

Investigation of Tip Leakage Vortex Breakdown in a High-Speed Multistage Axial Compressor

AbstractIn this article, an unsteady tip leakage flow instability is identified and investigated for an axial compressor at near-surge conditions. We describe the results of experimental verification of a new compressor developed by improving the blade geometry based on the criterion for the occurrence of this unsteady phenomenon. In a high-speed multistage axial flow compressor having a subsonic high stagger rotor blade, a surge test was carried out by changing the tip clearance. Under a condition of large tip clearance, a drastic decrease in the static pressure rise coefficient near the surge point was observed. At this operating condition, large, unsteady pressure fluctuation at the blade tip was confirmed, and the occurrence of tip leakage vortex breakdown was clarified by unsteady multipoint pressure measurement and detailed unsteady numerical simulations. Due to the blockage effect caused by vortex breakdown of the tip leakage, double leakage and axially reversed flow near the trailing edge were observed. It was found that the vortex breakdown region of the tip leakage vortex propagated in the circumferential direction and caused the rotating instability. In order to investigate the relationship among this unsteady flow phenomenon, tip clearance size, and flow pattern, unsteady calculation was conducted by changing the blade tip stagger and tip clearance size. A new concept of tip clearance of staggered pitch reference was proposed, which makes it possible to include the effect of blade loading on the clearance and clarifies that there exists a threshold at which vortex breakdown occurs/does not occur. On the basis of the aforementioned results, a high-speed multistage improved compressor was designed and manufactured to prevent tip leakage vortex breakdown. A clearance change test using active clearance control technology was conducted, and an increase in the static pressure rise coefficient near the surge point was confirmed for each clearance. The design concept of the improved blade, which suppressed the unsteady tip leakage flow instability, was tested and verified, and the effectiveness of the design guideline in actual gas turbines for power generation was confirmed.

Exploration of Flow Instability Characteristics in a Two-Stage Axial-Flow Compressor via Numerical Simulation Method

A four-blade-passage numerical model was developed for a two-stage axial-flow compressor with an inlet guide vane (IGV) for the purpose of studying the steady and dynamic characteristics of the compressor approaching its stability limit. The flow-field information indicated that the tip-leakage flow decreased more rapidly from the blade’s leading edge to the trailing edge, with a decreasing flow rate. The leakage flow was verified to be driven via the blade load over the whole operating range. Further research on the blade load was carried out. The magnitude of the highest blade load in the leading-edge portion of the first-stage rotor determined the lowest flow rate with steady-simulation analysis. The circumferential grooves on the rotor improved the rotor’s stable range via reducing the blade load. Unsteady-simulation results showed that the extreme blade load appeared first at the front of the first stage, with a decreasing flow rate. The second stage played a positive compensative role through releasing some of the load from the first stage. It can be generalized that the lowest flow rate at a specific speed is determined via not only any single stage but also other stages in a multistage axial-flow compressor.

Metamodel-Interpreted Data Mining for Stability and Efficiency Enhancement of Multistage Axial-Flow Compressors

Abstract The tradeoff between higher efficiency and wider stability of performance map is still one of the bottlenecks to hamper the further research and development of advanced multistage axial-flow compressor. The recent rapid growth of computational resources and artificial intelligence has enabled data mining as one of the most effective and potential ways to gain a deep insight into the complex correlations between aerodynamic performance and three-dimensional geometry parameters. In the open literatures, however, few research works have been found on using the data mining that is independent of design optimization to extract priori design guidelines for multistage axial-flow compressor mainly due to the lack of proper data mining method focused on the interpretation of metamodel with full use of limited time-consuming computational fluid dynamics dataset. To tackle this issue, a metamodel-interpreted data mining framework is developed in which extreme gradient boosting (XGBoost) metamodel combined with Shapley additive explanation (SHAP) model are employed to locally interpret the feature importance of each sample in the computational fluid dynamics dataset and then extract the design guidelines in terms of the most influential geometry parameters and their beneficial variation directions. The developed method is applied to data mining of design guidelines for efficiency and stability enhancement of a front 3.5-stage transonic axial-flow compressor in ship-board gas turbine usage. The results show that the aerodynamic performance of the investigated multistage compressor is most sensitive to three-dimensional geometry parameters related to blade lean, blade twist, and variable stators. Specially, the variable stators mainly affect the stall margin at part speed. The blade lean mainly influences the adiabatic efficiency at design speed as well as the stall margin at both speeds, while the blade twist mainly influences the aerodynamic performance at design speed. New designs followed by the design guidelines are obtained and critical performance indicators related to the goals of the data mining task are verified. The stall margin at part speed is widened to 5.87% with adjustment of blade lean and twist and further to 23.31% with additional adjustment of variable stators. The peak adiabatic efficiency at design speed is improved by 0.06% in spite of extremely limited potential for efficiency enhancement of the original design. The present work is of scientific significance as well as industrial application value in the three-dimensional design optimization of advanced multistage axial-flow compressor at the affordable computational cost.

Basic Nature of Behaviors of Surge Frequencies in Multi-Stage Axial Flow Compressors

The behaviors of surge frequencies in multi-stage axial-flow compressors are studied by use of several non-dimensional parameters on the basis of numerical-experimental results. One of the essential parameters proves to be a flowpath-average reduced surge frequency multiplied by the stalling pressure ratio and the tip-Mach number. Framework map in terms of such parameters against reduced resonance frequencies can provide a whole view of the surge frequency behaviors, including various characteristic zones of surge behaviors, i.e., convective surges, near-resonant surges, degenerated surges, and multi-stage effects. Particularly, for the convective surges, the surge frequency parameter comes to have a nearly constant value over a wide range of the reduced resonance frequencies, independently of number of stages, speeds, and pressure ratios, etc., of the compressors. The map is practically useful for the evaluation of deep-surge frequencies. At the same time, it could form a base for further investigations into the surge phenomena.

Abrupt Variations in Deep Surge Frequencies in High-Pressure-Ratio Multi-Stage Axial Flow Compressors

High-pressure-ratio multi-stage axial flow compressors tend to show complicated behaviors of frequencies in deep surges, including discontinuous changes. The phenomena for a nine-stage compressor were investigated by one-dimensional numerical simulations. The phenomena were found to occur somewhere in the range of roughly (90 – 100)% of the design speed, where incomplete surge recoveries, i.e., premature surge recoveries accompanied by immediate re-stallings, tend to occur, resulting in longer surge periods and lower surge frequencies. The incomplete surge recoveries are found to occur by premature unstalling owing to re-ingestions of the reversed hot air produced by the preceding surge, which is too hot for the compressor to work normally. The re-ingestion temperature, and the duration time of the hot re-ingestion relative to the normal suction appear to affect significantly the recovery process. It tends to occur essentially in the range of (90 -100)% of the design speed, where the compressor working conditions are changeable sensitively to small changes in the working conditions. The situations are found to be governed mainly by the following parameters; the representative tip Mach number, the reduced resonance frequency related with the suction flowpath, and the relative location of the compressor in the whole flowpath. Sufficiently below and above the sensitive speed range, the situations keep their respective stable and continuous surge behaviors. In the higher speed range, the deep surges having lowered frequencies at higher speeds become the normal state. The phenomena depend strongly on the design conditions, i.e., speed and pressure ratio and number of stages. The above conclusions suggest the natures of the discontinuous frequency behaviors in high-pressure-ratio multi-stage compressors, although they are qualitative in that they are strongly influenced by the present particular design conditions and the accuracies of the employed analysis.

Investigation on the circumferential grooves effects in a two-stage axial flow compressor

Abstract The performances of a two-stage industrial axial-flow compressor were tested to investigate the effects of circumferential grooves. The corresponding numerical models with four-blade cascade were set up. The numerical and test results showed that the operation range of the compressor near stall point can be extended obviously by the circumferential grooves. The internal flow field of the compressor was analysed numerically in detail, especially the tip leakage flow. It was found that stall occurred originally in the first stage and the local flow rate of the tip leakage was approximately proportional to the corresponding blade load. In addition, the local flow rate of tip leakage tended to decrease gradually along the flow direction near stall point. The tendency indicated that there were the most part of the blade load in the front of the rotor. It was proposed that the circumferential grooves could be set up at a specific stage in the multistage axial flow compressor where stall might occur originally to extend the operation range with lower manufacturing cost. Meanwhile, several circumferential grooves in the front segment of the rotor would be enough to play the effect on the stability enhancement for the axial compressor.

A Simplified Method for the Deposition Rate Assessment on the Vanes of a Multistage Axial-Flow Compressor

Abstract Gas turbine particle ingestion may lead to the deposition of contaminants in the compressor section, inducing the performance losses of the whole engine. The economic losses derived from this issue push great interest in the investigation of such a phenomenon from a numerical and experimental standpoint. This article describes a quantitative approach to predict particle deposition on the vanes of an axial compressor starting from the flow field obtained employing computational fluid dynamic (CFD) simulations. The results are then compared to the experiments performed on the Allison 250 C18 compressor unit subject to particle ingestion under controlled conditions. The results derived from the experimental and numerical investigations are presented, providing insight into the mass deposited on the vanes and the corresponding zones most affected by the particle deposition issue. The methodology showed good agreement in the estimation of the predicted values of the deposited mass and the corresponding patterns through the compressor stages. The low-complexity approach proposed here could help the designer to predict the contamination of the stationary rows starting from a simple set of single-phase numerical results. Furthermore, with the implementation of this approach into the design path, the designer could reduce the impact of fouling, looking at the effects of their solutions under the fouling-reduction light.

Stall Warning Strategy Based on Fast Wavelet Analysis in a Multistage Axial Flow Compressor

Abstract As a reliable stall warning strategy, the fast wavelet method was introduced to successfully predict the aerodynamic instability of a multistage axial flow compressor. One single sensor installed at each stage is proved to be sufficient to predict the stability status in a three-stage axial flow compressor. The whole prediction strategy includes the dynamic pressure signal capture, disturbance extraction using decomposition and reconstruction via fast wavelet transform, and stall warning index calculation based on statistical probability distribution. On this premise, the first occurrence of the stall in this three-stage axial flow compressor is predicted to be within the first stage, which is consistent with the stall route captured by the eight transducers around the casing wall. Thereafter, the stall warning index is used to monitor the stability status during the continuous throttling process. Furthermore, the validation using tip air injection and inlet radial distortion indicated that the stall warning index decreases as the compressor's stability improves. Conversely, the deterioration of stability causes the increase of the stall warning index. Thus, experimental results demonstrate that the stall warning method based on fast wavelet analysis can predict the aerodynamic instability in actual application.

Geometry scaling technique and application to aerodynamic redesign of multi-stage transonic axial-flow compressors

Geometry scaling technique, as one of the essential steps in aerodynamic redesign of multi-stage compressor, directly determines whether the new scaled geometry meets the required performance or not. It is also one of the effective measures in unsteady flow simulation of multi-stage compressor to reduce greatly the computational cost by scaling the computational domain from full-annulus passages to partial-annulus passages. The common critical issue arisen from both practices is the geometry dissimilarity in which some certain difference exists between the original and scaled geometries due to a diversity of redesign requirements or model simplifications. However, the extent of how the geometry dissimilarity influences the aerodynamic performance of multi-stage transonic compressor is unclear up to now. Particularly, how to keep the aerodynamic performance similarity in the case of geometry dissimilarity has not been explored yet. Here we propose a geometry scaling technique in which 3D parametric geometry modeling and 1D mean-line analysis are combined to define a series of key redesign criteria in order to keep as minor as possible variations in the aerodynamic performance of multi-stage compressor. The proposed technique is applied to redesign a 3.5 stage transonic axial-flow compressor with two geometry scaling strategies, i.e., adjustable blade/vane numbers and inlet guide vane (IGV)-rotor-stator axial spacing. The key aerodynamic parameters determined by the key redesign criteria are then quantified to verify the feasibility and accuracy of the proposed technique by computational fluid dynamics (CFD) simulations. The results show that by keeping the similarity of key redesign criteria, that is, blade solidity, stagger angle, inlet blade angle, outlet blade angle and thickness to chord distribution, the similarity of key aerodynamic parameters (incidence angle, deviation angle, loss coefficient and aerodynamic blockage) and thus of overall aerodynamic performance (mass flow rate, pressure ratio and efficiency) can be generally guaranteed. The variations of aerodynamic performance in the case of geometry dissimilarity with different blade/vane numbers and IGV-rotor-stator axial spacing are generally within 1%. Compared against the similarity of the other key aerodynamic parameters, a certain dissimilarity of aerodynamic blockage is observed, impacting slightly on the mass flow rate and thus stall margin. The present work is of effective assistance to improve the redesign methodology for multi-stage axial-flow compressors.

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.

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

Design Optimization of a Multistage Axial Flow Compressor Based on Full-Blade Surface Parameterization and Phased Strategy

A new design optimization method is proposed for the problem of high-precision aerodynamic design of multistage axial compressors. The method mainly contains three aspects: full-blade surface parametrization can significantly reduce the number of control variables per blade row and increase the degrees of freedom of the leading edge blade angle compared with the traditional semiblade parametric method; secondly, the artificial bee colony algorithm improved initialization and food source exploration and exploitation mechanism to enhance the global optimization ability and convergence speed, and a distributed optimization system is built on the supercomputing platform based on this method; finally, a phased optimization strategy based on the “synchronous change in multirow blades” is proposed, and expert experience is introduced to achieve a better balance between exploration and exploitation. The optimization method is applied to the AL-31F four-stage low-pressure compressor. As a result, the adiabatic efficiency is improved by 0.67% and the surge margin is improved by 3.1% under the premise that the total pressure ratio and mass flow rate satisfy the constraints, which verifies the effectiveness and engineering practicality of the proposed optimization method in the field of multistage axial flow compressor aerodynamic optimization.

Compressor Surge Frequencies Modeled by the Concept of Emptying and Filling Actions

Surge frequencies in compressors can vary much in dependence of many factors, such as compressor conditions, such as sizes, speeds, number of stages, and flowpath geometries, relative compressor locations, etc. The present investigation shows the basic behaviors of surge frequencies in multi-stage axial flow compressors, governed by a small number of essential parameters, which are obtained by physical insight to the phenomena and examinations in a trial-and-error manner on a n umber of numerical-experimental data. One of the essential parameters is a flowpath-average reduced resonance frequency, implying the number of excitations by the resonance frequency on the fluid particle in passing through the flowpath. Another one is also a reduced frequency formulated by modelling the emptying and filling actions in surge. A generalized basic framework of surge frequencies against resonance frequencies has been constructed by applying the numerical experimental data on a wide range of conditions of the compressors and the flowpaths in terms of both parameters. The framework thus prepared can be useful for practical purposes, such as for estimation of surge frequencies and surge degeneration (stall stagnation) boundaries, and also for understanding the effects of various related factors. It can provide a rule of thumb for estimation of surge frequency in the condition of sufficiently large values of the reduced resonance frequency parameter. It could serve also as a clue to further study on the details.

Research on Aerodynamic Optimization Method of Multistage Axial Compressor under Multiple Working Conditions Based on Phased Parameterization Strategy

Multistage axial compressor is the key component of aeroengine and gas turbine to realize energy conversion. In order to avoid the “curse of dimensionality” problem in the global optimization process of AL-31F four-stage low-pressure compressor under multiple working conditions, an optimization method based on phased parameterization strategy is proposed. The method uses the idea of “exploration before exploitation” for reference and divides the optimization process into two phases. In the first phase, the traditional parametric modification method based on stacking line is adopted; in the second phase, the full-blade surface parametric modification method with significant low-dimensional characteristics is adopted. Based on the improved artificial bee colony algorithm, a multitask concurrent optimization system is built on the supercomputing platform, and the engineering optimization solution is obtained within 91 hours. The optimization results are as follows: under the condition of meeting the constraints, the adiabatic efficiency is increased by 0.3% and the surge margin is 4.0% at the design speed; the adiabatic efficiency is increased by 0.8% and the surge margin is 2.3% at the off-design speed. These results verify the usefulness and reliability of the optimization method in the field of aerodynamic optimization of a multistage axial flow compressor.

Implementation of stability-enhancement with tip air injection in a multi-stage axial flow compressor

Tip air injection for stall margin extension is experimentally investigated in a multi-stage axial flow compressor. By comparing the stall margin improvement (SMI) under tip air injection in different stages, injected air only upstream from the first stage can causes a 20% SMI. This is because stall inception appears in the first stage. From the trend curve of SMI versus injected momentum ratios applied in the first stage, two demarcation points are found that separate the curve into micro, macro, and overlarge injections. Micro injection with a small SMI only weakens the tip leakage flow. Macro and overlarge injections influence both the tip leakage flow and mainstream flow, which contribute to a large SMI. However, overlarge injection causes the second stage to stall first, and causes the SMI to remain constant with further increases in the injected momentum ratio. The outlet axial velocity and flow angle measured using a five-hole probe indicate that with the overlarge injection in the first stage, the separation of the rotor blade passage in the second stage is more adverse than those in the other stages. The adverse passage separation originating from the hub region results in the second stage losing its working capability and stalling first. Moreover, at this point, the blade tip injection in the second stage has no additional effect on the extension of the stall margin.

Dust Ingestion in a Rotorcraft Engine Compressor: Experimental and Numerical Study of the Fouling Rate

Helicopters often operate in dusty sites, ingesting huge amounts of contaminants during landing, take-off, hover-taxi, and ground operations. In specific locations, the downwash of the rotor may spread soil particles from the ground into the environment and, once ingested by the engine, may stick to the compressor airfoils. In the present work, the Allison 250 C18 engine’s multistage axial-flow compressor is employed to study the fouling rate on rotor blades and stator vanes from both numerical and experimental standpoints. The compressor is operated in a typical ground-idle operation, in terms of the rotational regime and contaminant concentration, in laboratory-controlled conditions. The mass of deposits is collected from the airfoil surfaces at the end of the test and compared to that estimated through the numerical model. The experimental test shows that the airfoils collect almost 1.6% of the engine’s total mass ingested during a ground-idle operation. The capability of numerical methods to predict the fouling rate on the rotating and stationary airfoils of a multistage compressor is tested through the implementation of literature based deposition models. Sticking models show a good agreement in terms of the relative results; nevertheless, an overestimation of the deposited mass predicted is observed.

Behaviors of Surge Frequencies in Multi-stage Axial Flow Compressors over a Wide Range of Speeds

The fundamental behaviors of deep surge frequencies in axial-flow compressors have come to be described as a rather simple framework for low pressure-ratio compressors, in terms of two essential reduced frequencies concerning resonance frequencies and surge frequencies, on the basis of numerical-experimental results. For high-pressure-ratio multi-stage compressors, however, the frequency behaviors over a wide range of compressor speeds tend to give drastic deviations from the simple framework, including abrupt changes and apparent irregularities. The cause was investigated here by numerical simulations on a nine-stage axial flow compressor. Generally, deep surge frequencies of high-pressure ratio compressors tend to lower gradually, from low speeds toward higher ones. At somewhat below the design speed, the frequencies tend to decrease precipitously, settling to some very low levels of values. The discontinuous drops in the frequencies are shown to result from occurrences of short-term incomplete surge recoveries in the stalling phase of surge. The phenomenon tends to elongate the duration of the stalled phase and the whole surge duration, resulting in the lowered surge frequency. The phenomenon is caused by the flow reversal of high-temperature air in the stalling phase of surge and the re-suction of the hot air in the unstalling phase. The degree of the peak temperature and the intensity of the interferences between the hot-air re-suction phenomenon and the intrinsic surge appear to affect significantly the additional short-term premature stalling of the compressor. It could be strongly related with the high average temperature due to the hot air re-suction, and the duration time of the suction. The complicated behaviors thus become qualitatively explainable, though only by examinations on numerical-simulation results.

Performance simulation of multi-stage axial-flow compressor of turbo-shaft engine with account for erosive wear nonlinearity of its blades

The construction and useful practice of gas-turbine engine diagnosis systems depend largely on the availability of the engine mathematical models and its certain components in their structure. Utilization of multi-stage axial flow compressor performance with account for erosive wear of its parts during the operation fundamentally raises possibilities of such systems as erosive wear of flow channel, blade rings of impellers and vane rings of multi-stage compressor is a common cause of preschedule gas-turbine engine detaching from an aircraft. As evidenced by various contributions presented in the article, special emphasis on abrasive wear impact assessment on axial flow compressor performance is placed upon rotor-wing turbo-shaft engine due to their particular operating conditions. One of the main tasks in the process of mathematic simulation of an axial flow compressor blade ring is consideration of its wear type that again has a nonlinear distribution along the level of the blade. In addition, wear rate at entry and exit blade edges often have different principles. Detecting of these principles and their consideration when constructing the compressor mathematical model is a crucial task in diagnostic assessment and integrity monitoring of rotor-wing turbo-shaft engine in operation. The article represents a concept to an estimate nonlinear erosive wear effect of axial flow compressor blades on its performance based on the three-dimensional flow approach in the gas-air flow duct of compressor with a formulation of the blade rings. This approach renders possible to take into account the nonlinearity of the compressor blades wear during their operation. Through the example of the inlet compressor stage of a rotor-wing aircraft gas-turbine engine, the engine pump properties predictions with different kind of rotor blade wear have been presented.

SYNTHESIS OF ROBUST OPTIMAL CONTROL PROGRAM FOR AXIAL FLOW COMPRESSOR TURNING GUIDE VANES

The method of solving the selecting turning guide vanes law of control problem for the multistage axial flow compressor is proposed in order to ensure maximum efficiency along the operating line while maintaining specified stability margins under the uncertainty of input data. The problem under consideration belongs to a class of multi-objective stochastic optimization problems with mixed conditions. Evolutionary computational method of solution synthesis of the problems belonging to this class is developed based on the genetic algorithm. The implementation example of proposed method for selecting law of control for turning inlet at first four stages guide vanes of multistage axial flow compressor (MSAFC) of modern helicopter jet engine is considered. The problem of technological tolerance selecting the stagger angles of compressor blade rows in order to provide the specified confidence intervals based on integral performance parameters is solved. The examples of solving the problem using determined and stochastic formulations are given.