Method and system for simulating influence of geometric deviation of compressor blade on aero-engine performance

Abstract

The invention provides a simulation method and system for the influence of the geometric deviation of a compressor blade on the performance of an aero-engine. The method comprises the steps that stage superposition is conducted on the characteristics of a multi-stage compressor based on the influence of the geometric deviation; superposing the characteristic change of the gas compressor caused by the geometric deviation of the blade to the general characteristic of the downstream gas compressor to obtain the change of the working point of the downstream gas compressor; after the characteristics of the gas compressor are obtained under the condition of blade geometric deviation, the characteristics are introduced into an engine performance calculation model, and changes of engine thrust, oil consumption rate and exhaust temperature are obtained. Computational analysis of gas compressor characteristics and engine performance is realized, influences of typical geometric deviations on gas compressor blade channel flow, engine surge margin, rotor power capability and working efficiency, engine thrust, oil consumption rate, exhaust temperature and the like are disclosed, and rapid prediction of engine performance changes under different blade geometric deviations is realized. And formulating the use specifications of the aero-engine under the condition that the compressor blades have geometric deviation in the support service process.

Description

Technical Field

[0001] The present invention belongs to the technical field of compressors, and particularly relates to a simulation method for the influence of geometric deviations of compressor blades on the performance of aero-engines. Background Art

[0002] In order to meet the requirements of the engine for high pressure ratio and high efficiency, compressor blades tend to be thinner and have an increased sweep, gradually having complex three-dimensional strong curvature geometric features. The complexity of the blade geometry makes it difficult to avoid geometric deviations during modeling, machining, and installation, resulting in typical deviations such as twist deviation, leading and trailing edge shape deviation, profile deviation, and installation angle error. Due to the high uncertainty of the machining deviations of the blades, the internal flow field of the blades becomes complex and diverse, and the influence on the loss coefficient and stagger angle of the blade profile is also complex and variable. Eventually, the aerodynamic performance of the compressor affected by the machining deviations also has a high degree of uncertainty. In the compressor stage environment, the influence of geometric deviations on the aerodynamic performance of components, such as total pressure ratio and adiabatic efficiency, needs to be studied more carefully.

[0003] By reviewing the research work of domestic and foreign research teams on geometric deviations and compressor stage performance, we found that the geometric deviations of the blades will affect the aerodynamic performance or stability margin of the compressor to varying degrees. Since China started relatively late in the field of aero-engines, domestic research scholars have only begun to pay attention to the influence of geometric deviations on the overall performance of components in recent years. Most of the research methods use numerical simulation methods, and then based on statistical thinking, a variety of uncertainty quantification analysis methods are proposed. For example, the team of Chu Wuli from Northwestern Polytechnical University [Guo Zhengtao, Chu Wuli, Yan Song, Shen Zhengjing, Wang Guang. Data mining on the influence of machining errors on the aerodynamic performance and stability of compressor cascades [J]. Journal of Propulsion Technology, 2022, 43(3): 133-145.] uses the artificial neural network prediction method to predict the influence of normal machining deviations on the aerodynamic performance and corner aerodynamic stability of the cascade. Random samples of normal machining deviations are generated by the pseudo Monte Carlo method, and then a geometric uncertainty reduction model is established through the standard deviation function. Chinese Patent with the publication number CN117034584A provides an analysis method for the influence of geometric deviations of rotor blades on the stability of the compressor. According to the type and distribution form of the geometric deviations of the target blade, a number of sample points are generated; the target blade is parameterized; a number of sample blade geometric models are generated; a linear relationship between the unsteady numerical simulation and steady numerical simulation results of the sample blades is fitted and established; the unsteady numerical simulation calculation results of all sample blades are obtained; a response model is constructed and initialized to obtain response values; the convergence of the response model is verified; the accuracy of the response model is verified; by statistically analyzing the response values, the probability distribution of the improvement amount of the compressor stability margin is obtained, and the influence effect of various geometric deviations on the compressor stability is analyzed. The present invention takes less time, has high model convergence and accuracy, and can meet certain engineering application requirements. Summary of the Invention

[0004] In order to provide reliable numerical simulation results, the present invention provides a simulation method for the influence of geometric deviations of compressor blades on the performance of aero-engines, to realize the calculation and analysis of compressor characteristics and engine performance, to reveal the influence of typical geometric deviations on the flow in the compressor blade passage, the engine surge margin, the rotor work capacity and working efficiency, the engine thrust, the specific fuel consumption and the exhaust gas temperature, etc., to realize the rapid prediction of the engine performance changes under different blade geometric deviations, and to support the formulation of the usage specifications of aero-engines in the case of geometric deviations of compressor blades during service.

[0005] In order to achieve the above object, the technical solution adopted by the present invention is as follows: On the one hand, a simulation method for the influence of geometric deviations of compressor blades on the performance of aero-engines is provided, including the following steps: Performing stage superposition based on the characteristics of a multi-stage compressor under the influence of geometric deviations; Calculating the engine performance under the influence of geometric deviations according to the stage superposition result.

[0006] Further, when performing stage superposition based on the characteristics of a multi-stage compressor under the influence of geometric deviations, the general characteristics of the compressor are based on the theory of similarity of the gas flow state in the compressor. When two gas flows satisfy the similarity conditions, it is considered that the flow states of these two flows are similar, and the corresponding physical quantities of both are proportional, or their ratios are equal.

[0007] Further, convert the original working condition parameters into parameters describing the similar state, where is the rotational speed, is the flow rate, and the general characteristics of the compressor are summarized as the functional relationships F1 and F2. F1 and F2 are similarity parameters. The rotational speed similarity parameter of the compressor is: and the flow rate similarity parameter of the compressor is: Then the general characteristics of the compressor are: and .

[0008] Further, when performing stage superposition based on the characteristics of a multi-stage compressor under the influence of geometric deviations, assume that the change in the characteristics of a single-stage compressor only affects the downstream compressor, and superimpose the change in the compressor characteristics caused by the geometric deviation of the blade on the general characteristics of the downstream compressor, so as to obtain the change in its operating point.

[0009] Further, in the case of geometric deviations of the blade, through calculation and analysis, the change in the characteristics of the first-stage compressor can be obtained, and then the change in the outlet total pressure and total temperature can be obtained. Assume that the engine intake flow rate and physical rotational speed do not change, and the above changes in the total pressure and total temperature will cause the similar rotational speed of the second-stage compressor and similar flow rates It changes, causing its operating point to deviate from the operating condition point A to reach the operating condition point A1. The pressure ratio and efficiency of the second-stage compressor change. After the changes in the pressure ratio and efficiency of the second-stage compressor are determined, the calculation formulas for the compression work and efficiency of the compressor are used to determine the changes in its outlet total pressure and total temperature. Repeat the above calculation steps to obtain the new operating point B1 of the third-stage compressor, and so on to obtain the influence result of the blade geometric deviation on the characteristics of the multi-stage compressor.

[0010] Furthermore, when calculating the engine performance under the influence of geometric deviation according to the stage superposition result, the compressor characteristics under the condition of blade geometric deviation are introduced into the engine performance calculation, and the common working equations of the engine components are solved to obtain the changes in the engine thrust, specific fuel consumption, and exhaust gas temperature.

[0011] Furthermore, calculating the engine performance under the influence of geometric deviation according to the stage superposition result includes solving the rotational speed balance equation, flow rate balance equation, power balance equation, and the relationship of aerodynamic and thermal parameters.

[0012] On the other hand, the present invention provides a simulation system for the influence of compressor blade geometric deviation on the performance of an aeroengine, including a stage characteristic superposition module and a performance solution module; the stage characteristic superposition module performs stage superposition based on the characteristics of the multi-stage compressor under the influence of geometric deviation; the performance solution module is used to calculate the engine performance under the influence of geometric deviation according to the stage superposition result.

[0013] The present invention can also provide a computer device, including a processor and a memory. The memory is used to store computer-executable programs. The processor reads the computer-executable programs from the memory and executes them. When the processor executes the computer-executable programs, it can implement the simulation method for the influence of compressor blade geometric deviation on the performance of an aeroengine described in the present invention.

[0014] At the same time, a computer-readable storage medium is provided. A computer program is stored in the computer-readable storage medium. When the computer program is executed by a processor, it can implement the simulation method for the influence of compressor blade geometric deviation on the performance of an aeroengine described in the present invention.

[0015] Compared with the prior art, the present invention has at least the following beneficial effects: Since the machining deviation of the blade has a high degree of uncertainty and the influence on the blade profile is also complex and variable, ultimately resulting in a high degree of uncertainty in the aerodynamic performance of the compressor under the influence of the machining deviation. The present invention completes the evaluation of the influence of compressor blade geometric deviation on the performance of an aeroengine through calculation and analysis of the compressor characteristics and engine performance, and reveals the law that typical blade geometric deviations exacerbate the flow blockage in the compressor blade passage, reduce the engine surge margin, change the rotor work capacity and working efficiency, and affect the engine thrust, specific fuel consumption, and exhaust gas temperature. Description of the Drawings

[0016] Figure 1 It is a general characteristic diagram of a typical fan.

[0017] Figure 2 It is a flow chart of the method implementation of the present invention. Specific implementation manners

[0018] The present invention will be further described below in conjunction with the accompanying drawings.

[0019] Step 1: The stage superposition method for the characteristic analysis of a multi-stage compressor under the influence of geometric deviation; The general characteristics of a compressor are based on the theory of similar airflow flow states in the compressor. Similar flow states refer to flow states that satisfy the conditions of geometric similarity, flow field similarity (kinematic similarity), and dynamic similarity. When two airflow flows satisfy the similarity conditions, it is considered that the flow states of these two flows are similar, and their corresponding physical quantities are proportional, or their ratios are equal. Taking a compressor as an example, although the compressor operates under different conditions, when its flow state is similar, there are , , where the subscripts 1 and 2 represent two flow states respectively, and the subscripts 2 and 3 in the parentheses represent two operating points respectively. The superscript * represents that the parameter is a stagnation parameter, p is the inlet pressure, and T is the inlet temperature. Therefore, it can be known that , , is the pressure ratio, is the efficiency, indicating that the performance parameters of the compressor can remain unchanged after the flow state is similar. In this way, the performance parameters of the compressor are only related to the similar flow state, and the original working condition parameters can be converted into parameters describing the similar state, is the rotational speed, is the flow rate, so that these parameters have universality, and the characteristics formed in this way become general characteristics. Therefore, the general characteristics of the compressor can be simply summarized as a new functional relationship, that is F1(similarity parameters), F2(similarity parameters).

[0020] According to the conclusion of the similarity theory, for the same compressor, when operating under different working conditions, the condition for its flow state to be similar is that the airflow Mach numbers and Reynolds numbers at the corresponding points are equal. And within most of the working range of the compressor, the Reynolds number is in the self-modeling range, and its influence can be ignored. Therefore, the condition for state similarity is only the Mach number. After conversion, the similarity parameters expressed by the working condition parameters are finally obtained.

[0021] The rotational speed similarity parameter of the compressor is: 。 Its meaning is that for two different operating conditions, when the operating states are similar, there is 。

[0022] The flow similarity parameter of the compressor is: 。 Its meaning is that for two different operating conditions, when the operating states are similar, there is 。

[0023] In this way, the universal characteristics of the compressor can be expressed as: and 。

[0024] This application assumes that when the characteristics of a single-stage compressor change, it only affects the downstream compressor. The changes in the compressor characteristics caused by blade geometric deviations are superimposed on the universal characteristics of the downstream compressor, and then the changes in its operating point are obtained. Figure 1 shows the typical universal characteristics of a fan. Assume that the second-stage and third-stage compressors operate at operating conditions A and B respectively under a certain operating condition.

[0025] In the case of blade geometric deviations, through calculation and analysis, the changes in the characteristics of the first-stage compressor can be obtained, and then the changes in the outlet total pressure and total temperature can be obtained. Assuming that the engine inlet air flow and physical speed do not change, the above changes in total pressure and total temperature will cause the similarity speed and similarity flow rate of the second-stage compressor to change, and then its operating point deviates from the operating point A to reach the operating point A1, causing changes in the pressure ratio and efficiency of the second-stage compressor. After the changes in the pressure ratio and efficiency of the second-stage compressor are determined, using the calculation formulas for the compression work and efficiency of the compressor, the changes in its outlet total pressure and total temperature are determined. Repeating the above calculation steps can obtain the new operating point B1 of the third-stage compressor, and so on to obtain the influence of blade geometric deviations on the characteristics of multi-stage compressors.

[0026] Step 2: Engine performance calculation method under the influence of geometric deviations; After obtaining the compressor characteristics under the condition of blade geometric deviations, introduce them into the engine performance calculation model, and according to the process as Figure 2 shown, obtain the changes in engine thrust, specific fuel consumption, and exhaust gas temperature. The engine performance calculation process is mainly to solve the simultaneous working equations of engine components. Taking a dual-duct and dual-rotor turbofan engine as an example, the form of the simultaneous working equations is as follows.

[0027] Step1: Speed balance; The high-pressure compressor of the core engine and the high-pressure turbine are coaxial to form a high-pressure rotor, so their speeds are equal, collectively referred to as the high-pressure rotor speed, simply referred to as the high-pressure speed, usually represented by 。

[0028]

[0029] The fan (low-pressure compressor) and the low-pressure turbine are coaxially assembled to form the fan (low-pressure) rotor. Therefore, their rotational speeds are equal, collectively referred to as the rotational speed of the low-pressure rotor, simply called the low-pressure speed, usually represented by :

[0030] Step2: Flow balance; In the core engine, the flow balance relationship between components is:

[0031] Among them, is the gas flow rate of the high-pressure turbine, is the air flow rate at the inlet of the high-pressure compressor, is the fuel flow rate added in the combustion chamber, is the bleed air flow rate (meeting aircraft requirements) or anti-surge bleed air flow rate extracted from the high-pressure compressor, is the turbine cooling air flow rate extracted from the high-pressure compressor.

[0032] For a turbofan engine, "Ⅰ" is used to represent the flow rate of the core flow path, that is , There is:

[0033] The balance relationship between the flow rate of the fan (low-pressure compressor) and the flow rate of the low-pressure turbine is: and

[0034] Step3: Power balance; The power required by the high-pressure compressor and the output power of the high-pressure turbine should satisfy:

[0035] Among them, is the mechanical efficiency of the high-pressure rotor, is the power extracted from the engine through the high-pressure rotor to drive load devices such as generators, fuel pumps, and hydraulic pumps.

[0036] The power required by the fan (low-pressure compressor) and the output power of the low-pressure turbine should satisfy:

[0037] Among them, is the mechanical efficiency of the low-pressure rotor.

[0038] In a mixed-flow turbofan engine, the flow matching relationship between the low-pressure turbine, the mixer, and the nozzle is as follows:

[0039] Wherein, is the gas flow rate at the outlet of the mixer.

[0040] Step4: Relationship of aerodynamic and thermodynamic parameters; The compressor, the combustor, and the turbine are upstream and downstream components of each other. The outlet parameters of the upstream component become the inlet parameters of the downstream component, and the outlet parameters of each component are jointly determined by the performance parameters and the inlet parameters of the component under this operating condition.

[0041] During the engine performance calculation process, it is also necessary to calculate the characteristics of each engine component. The specific calculation process is as Figure 2 shown. Taking the outlet aerodynamic parameters of the previous component as the inlet aerodynamic parameters of the next component, the gas flows in from the inlet duct, following the flow balance relationship, and calculating the aerodynamic performance of the low-pressure compressor affected by the blade geometric deviation as the inlet parameters of the high-pressure compressor. Taking the outlet parameters of the high-pressure compressor as the inlet parameters of the combustor, and so on, obtaining the aerodynamic parameters of the high-pressure turbine, the low-pressure turbine, the mixer, the afterburner, and the nozzle respectively. According to the above process, the engine exhaust temperature is calculated. Based on the obtained engine aerodynamic parameters, using the thrust calculation formula where in, out represent the inlet and outlet cross-sections respectively, A is the area, W is the gas flow rate, c is the gas velocity, and the specific fuel consumption calculation formula where, F is the thrust, W is the gas flow rate, the thrust and specific fuel consumption of the engine can be obtained.

[0042] Based on the same concept, the present invention also provides a simulation system for the influence of compressor blade geometric deviation on the performance of an aeroengine, including a stage characteristic superposition module and a performance solution module; the stage characteristic superposition module performs stage superposition based on the characteristics of a multi-stage compressor affected by geometric deviation; the performance solution module is used to calculate the engine performance affected by geometric deviation according to the stage superposition result; When the stage characteristic superposition module performs stage superposition based on the characteristics of a multi-stage compressor affected by geometric deviation, the general characteristics of the compressor are based on the theory of similarity of the gas flow state in the compressor. When two gas flows satisfy the similarity conditions, it is considered that the flow states of these two flows are similar, and the corresponding physical quantities of the two are proportional, or their ratios are equal; converting the original working condition parameters into parameters describing the similar state, is the rotational speed, is the flow rate. The general characteristics of the compressor are summarized as a functional relationship F1 and F2. F1 and F2 are similarity parameters. The rotational speed similarity parameter of the compressor is: , and the flow rate similarity parameter of the compressor is: . The general characteristics of the compressor are then: and .

[0043] When the stage characteristic superposition module performs stage superposition based on the characteristics of a multi-stage compressor under the influence of geometric deviation, it is assumed that the change in the characteristics of a single-stage compressor only affects the downstream compressor. The change in the characteristics of the compressor caused by the blade geometric deviation is superimposed on the general characteristics of the downstream compressor, thereby obtaining the change in its operating point. In the case of geometric deviation of the blade, through calculation and analysis, the change in the characteristics of the first-stage compressor can be obtained, and then the change in the outlet total pressure and total temperature can be obtained. Assuming that the engine intake flow rate and physical rotational speed do not change, the above changes in the total pressure and total temperature will cause the similarity rotational speed and similarity flow rate of the second-stage compressor to change, causing its operating point to deviate from the operating condition A point to reach the operating condition A1 point. The pressure ratio and efficiency of the second-stage compressor change. After the pressure ratio and efficiency changes of the second-stage compressor are determined, using the compressor compression work and efficiency calculation formulas, the changes in its outlet total pressure and total temperature are determined. Repeating the above calculation steps, the new operating point B1 of the third-stage compressor is obtained, and so on to obtain the influence result of the blade geometric deviation on the characteristics of the multi-stage compressor.

[0044] When the performance solution module calculates the engine performance under the influence of geometric deviation according to the stage superposition result, it introduces the compressor characteristics in the case of blade geometric deviation into the engine performance calculation, solves the engine component co-working equation, and obtains the changes in the engine thrust, specific fuel consumption, and exhaust gas temperature. Calculating the engine performance under the influence of geometric deviation according to the stage superposition result includes solving the rotational speed balance equation, flow rate balance equation, power balance equation, and the relationship of aerodynamic and thermodynamic parameters.

[0045] The compressor blade geometric deviation simulation method described in the present invention accurately simulates various geometric deviations that the blade may encounter during the manufacturing and usage processes, helps to analyze how the deviations specifically affect the performance of the aeroengine, and provides guidance for the practice of blade design and manufacturing; during the design stage, this method enables designers to foresee and evaluate the potential impacts of different geometric deviations on the engine performance. Based on the simulation results, designers can implement more refined optimization design strategies to minimize the performance losses caused by the deviations, thereby significantly improving the overall efficiency and reliability of the engine; in the manufacturing field, the simulation method helps to promote process innovation. Through the detailed analysis of the distribution laws of blade geometric deviations under different manufacturing processes, it not only helps to identify the key factors leading to the deviations, but also can targetedly adjust and optimize the manufacturing process parameters. It can not only reduce the geometric deviations during the manufacturing process, but also achieve a leapfrog improvement in the blade processing accuracy, laying a solid foundation for the excellent performance of the aeroengine.

[0046] The simulation method of the present invention can simulate and predict in advance the specific impacts of these deviations on the engine performance. We can plan ahead and provide a scientific and reliable basis for subsequent fault diagnosis and formulation of preventive measures. This not only helps to extend the service life of the engine, but also greatly reduces the operation risks and costs caused by sudden failures.

[0047] In summary, the compressor blade geometric deviation simulation method described in the present invention demonstrates its indispensable importance in every link of aeroengine design and manufacturing. It not only provides strong support for technological innovation and process upgrade in the industry, but also contributes wisdom and strength to ensuring the safe and efficient operation of aeroengines. The present invention can also provide a computer device, including a processor and a memory. The memory is used to store computer-executable programs. The processor reads the computer-executable programs from the memory and executes them. When the processor executes the computer-executable programs, it can implement the simulation method of the impact of the compressor blade geometric deviation on the aeroengine performance described in the present invention.

[0048] On the other hand, the present invention provides a computer-readable storage medium. A computer program is stored in the computer-readable storage medium. When the computer program is executed by a processor, it can implement the simulation method of the impact of the compressor blade geometric deviation on the aeroengine performance described in the present invention.

[0049] The computer device can be a laptop computer, a desktop computer or a workstation.

[0050] The processor can be a central processing unit (CPU), a digital signal processor (DSP), an application-specific integrated circuit (ASIC) or a field-programmable gate array (FPGA) off-the-shelf.

[0051] For the memory described in the present invention, it can be an internal storage unit of a laptop, a desktop computer or a workstation, such as a memory or a hard disk; or an external storage unit can be adopted, such as a mobile hard disk or a flash card.

[0052] A computer-readable storage medium can include a computer storage medium and a communication medium. The computer storage medium includes volatile and non-volatile, removable and non-removable media implemented by any method or technology for storing information such as computer-readable instructions, data structures, program modules or other data. The computer-readable storage medium can include: read-only memory (ROM), random access memory (RAM), solid state drives (SSD) or optical discs, etc. Among them, the random access memory can include resistive random access memory (ReRAM) and dynamic random access memory (DRAM).

[0053] The above content is only to illustrate the technical idea of the present invention and cannot be used to limit the protection scope of the present invention. Any modification made on the basis of the technical solution according to the technical idea proposed by the present invention falls within the protection scope of the claims of the present invention.

Claims

1. A simulation method for the effect of compressor blade geometry deviation on aero-engine performance, characterized in that: The following steps are involved: Perform stage superposition based on the characteristics of multi-stage compressors under the influence of geometric deviations; The engine performance under the influence of geometric deviation is calculated based on the stage superposition results.

2. The simulation method for the effect of compressor blade geometry deviation on aeroengine performance according to claim 1, characterized in that: When the multi-stage compressor characteristics under the influence of geometric deviation are superimposed, the general characteristics of the compressor are based on the similarity of the airflow flow state in the compressor. When two airflows meet the similarity conditions, the flow states of the two flows are considered to be similar, and the same-name physical quantities of the two are proportional, or their ratios are equal.

3. The simulation method for the effect of compressor blade geometry deviation on aero-engine performance according to claim 1, characterized in that: The original working condition parameters Converted into parameters describing similar states, is the rotation speed, The general characteristics of the compressor can be summarized as a function of flow rate. F1 and F2, F1 and F2 are similar parameters, and the similar parameters of the compressor speed are: , the similarity parameters of the compressor flow are: , the general characteristics of the compressor are: and .

4. The simulation method for the effect of compressor blade geometry deviation on aeroengine performance according to claim 1, characterized in that: When the multi-stage compressor characteristics under the influence of geometric deviation are superimposed, it is assumed that the change of the single-stage compressor characteristics will only affect the downstream compressor. The change of the compressor characteristics caused by the blade geometric deviation is superimposed on the general characteristics of the downstream compressor, and then the change of its working point is obtained.

5. The simulation method for the effect of compressor blade geometry deviation on aero-engine performance according to claim 4, characterized in that: When there is geometric deviation in the blades, the change in the characteristics of the first-stage compressor can be obtained through calculation and analysis, and then the change in the outlet total pressure and total temperature can be obtained. Assuming that the engine intake flow and physical speed do not change, the above changes in total pressure and total temperature will lead to a similar speed of the second-stage compressor. and similar traffic The operating point of the second-stage compressor changes, causing it to deviate from the operating point A to the operating point A1. The pressure ratio and efficiency of the second-stage compressor change. After the changes in the pressure ratio and efficiency of the second-stage compressor are determined, the changes in the total pressure and total temperature at its outlet are determined using the compressor compression work and efficiency calculation formula. The above calculation steps are repeated to obtain the new operating point B1 of the third-stage compressor. This can be deduced by analogy to obtain the results of the influence of blade geometry deviation on the characteristics of a multi-stage compressor.

6. The simulation method for the effect of compressor blade geometry deviation on aero-engine performance according to claim 1, characterized in that: When calculating the engine performance under the influence of geometric deviation according to the stage superposition results, the compressor characteristics under the condition of blade geometric deviation are introduced into the engine performance calculation, and the common working equation of the engine components is solved to obtain the changes in engine thrust, fuel consumption rate and exhaust temperature.

7. The simulation method for the effect of compressor blade geometry deviation on aero-engine performance according to claim 1, characterized in that: Calculating the engine performance under the influence of geometric deviation according to the stage superposition result includes solving the speed balance equation, the flow balance equation, the power balance equation and the relationship between aerodynamic and thermal parameters.

8. A simulation system for the effect of compressor blade geometry deviation on aero-engine performance, characterized in that: It includes a stage characteristic superposition module and a performance solution module; the stage characteristic superposition module performs stage superposition based on the characteristics of multi-stage compressors under the influence of geometric deviation; The performance solving module is used to calculate the engine performance under the influence of geometric deviation according to the stage superposition result.

9. A computer device, characterized in that: The invention comprises a processor and a memory, wherein the memory is used to store a computer executable program, and the processor reads the computer executable program from the memory and executes it. When the processor executes the computer executable program, the simulation method of the influence of the geometric deviation of the compressor blade on the performance of the aircraft engine as described in any one of claims 1 to 7 can be realized.

10. A computer-readable storage medium, characterized in that: A computer program is stored in a computer-readable storage medium. When the computer program is executed by a processor, the simulation method of the influence of compressor blade geometry deviation on aircraft engine performance as described in any one of claims 1 to 7 can be implemented.

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