An integrated temperature control and verification method for an aeroengine

Abstract

This application belongs to the field of aero-engine design, and particularly relates to an aero-engine integrated temperature control and verification method. By splitting the gas temperature field into three dimensions, including the average gas temperature, the peak gas temperature, and the gas temperature distribution; and separately carrying out the regulation of the average gas temperature, the regulation of the peak gas temperature, and the regulation of the gas temperature distribution; a high-quality gas temperature field is constructed to ensure the reliable operation of high-temperature components such as turbine blades, solve the problems of "over-temperature" of the average gas temperature, too high peak gas temperature, and unreasonable gas temperature distribution, reduce the required gas temperature while meeting the thrust requirement, greatly reduce the temperature load of hot-end components such as the combustor and turbine, solve the "over-temperature" problem of aero-engines, and achieve reliable operation at high gas temperatures.

Description

Technical Field

[0001] This application belongs to the field of aero-engine design, and particularly relates to an aero-engine integrated temperature control and verification method. Background Technique

[0002] High-temperature components such as aero-engine turbine blades work in high-temperature gas and bear high mechanical loads and aerodynamic loads. Building a high-quality gas temperature field to achieve the long-term reliable operation of high-temperature components such as turbine blades is the biggest technical problem faced in the development of aero-engines. During the engine development process, the problem of "overtemperature" usually occurs, which is manifested as the deviation of the gas temperature field from the design target. "Overtemperature" is the first obstacle that must be overcome in the development of advanced aero-engines at home and abroad.

[0003] For a long time, engines have improved the gas temperature field through a large number of tests of "trial and error for temperature adjustment", but the problem of "overtemperature" has not been completely solved, resulting in insufficient temperature margin of the engine and slow improvement in reliability. However, with the further increase of gas temperature, "trial and error for temperature adjustment" can no longer solve the problem of reliable engine operation. During the development of aero-engines, a large number of high-temperature components such as turbine blades and afterburners are ablated and fractured, making the engine unable to work normally and seriously hindering the development process of the model.

[0004] The existing technical solution is that after the engine is completed with trial production and processing, during the whole-engine test run, a large number of tests are carried out for debugging and repeated trial and error to find problems in the gas temperature field and optimize them. It consumes a lot of resources and the optimization results are uncertain. It is often difficult to achieve the design expectation, and the positive design and debugging of the gas temperature field have not been realized. The disadvantages include:

[0005] 1) The problem of "overtemperature" of the average gas temperature occurs. Since there is no method for obtaining the true characteristics of components and solving the core flow rate in the whole-engine environment, the performance design and optimization of engines at home and abroad are carried out based on the characteristics of individual component tests to control the average gas temperature. However, individual component tests cannot accurately simulate the actual conditions such as high temperature, high pressure and flow field of the whole engine, and there are differences between the obtained component characteristics and the true characteristics of components in the whole-engine environment, resulting in the deviation of the benchmark for engine performance design and optimization, making it difficult to achieve the efficient collaborative work of each component, resulting in low thermal efficiency of the whole engine, and the average gas temperature of the engine far exceeding the design value under the condition of reaching the thrust index;

[0006] 2) The problem of too high peak gas temperature occurs. The existing solutions mainly rely on the traditional double swirl flame tube head configuration and improve the gas temperature uniformity by increasing the core vortex size. There is a lack of an organized combustion theory and design method for effectively controlling the temperature at the edge of the core vortex, resulting in easy ablation of the wall surface when the vortex core is close to the flame tube wall, and it is impossible to effectively improve the gas temperature uniformity by increasing the vortex core size;

[0007] 3) There is a problem of unreasonable gas temperature distribution. The existing technical solutions simply pursue a as-uniform-as-possible radial gas temperature distribution, failing to achieve adaptation to the complex loads of the turbine. As a result, key areas such as the roots of the turbine rotor blades simultaneously bear large temperature and mechanical loads, significantly reducing reliability. At the same time, relying solely on the ability of the combustion chamber to mix air and gas to adjust the temperature distribution is limited, and it is difficult to accurately achieve the expected temperature distribution in design.

[0008] Therefore, it is necessary to further solve the "overtemperature" problem of aeroengines. Summary of the Invention

[0009] The purpose of this application is to provide a comprehensive temperature control and verification method for aeroengines to solve the problem that the "overtemperature" of existing aeroengines is difficult to effectively solve.

[0010] The technical solution of this application is: a comprehensive temperature control and verification method for aeroengines, including:

[0011] Regulation of the average gas temperature: Based on the true characteristics of components under the whole-engine conditions, establish a high-precision performance evaluation model for aeroengines. Calculate the average gas temperature and deviation through the whole-engine average gas temperature regulation test. Then, determine the adjustment amount of the aeroengine according to the deviation, and evaluate the deviation of the adjusted average gas temperature through the high-precision performance evaluation model of the aeroengine. Adjust the structural design parameters of the aeroengine according to the evaluation results until there is no overtemperature problem under the current deviation.

[0012] Regulation of the peak gas temperature: Establish a relationship model based on the non-uniformity δ of the oil-gas distribution in the core vortex of the main combustion zone, determine the swirl number S in the relationship model, perform multi-parameter coupling design according to the swirl number S to obtain multiple different design schemes. Then, conduct verification tests according to different design schemes until a design scheme that meets the design requirements is obtained, and obtain the peak gas temperature regulation parameters according to the design scheme that meets the design requirements; regulate the peak gas temperature according to the peak gas temperature regulation parameters.

[0013] Regulation of the gas temperature distribution: Obtain the gas temperature distribution design and blade surface temperature data of the current aeroengine for data processing to obtain the optimal gas temperature distribution; design the temperature distribution requirements of the aeroengine according to the optimal gas temperature distribution, and conduct verification tests on the gas temperature distribution of the main combustion chamber according to the design results of the temperature distribution requirements to determine whether the gas temperature distribution requirements are met. If not, re-design the temperature distribution requirements until the gas temperature distribution requirements are met.

[0014] Preferably, the specific design method of the regulation of the average gas temperature is:

[0015] Obtain the true characteristics of components such as the fan, compressor, combustor, and turbine under the overall environment of an aero-engine through special engine tests, and establish a high-precision performance evaluation model for the aero-engine based on the true characteristics of the components under the overall engine conditions;

[0016] Obtain the adjustable geometric parameters during the test run of the aero-engine, including the fan adjustable guide vane angle α1, the compressor adjustable guide vane angle α2, the nozzle throat area A8, and the nozzle exit area A9; conduct an overall engine gas average temperature T4 regulation test based on the adjustable geometric parameters of the aero-engine, calculate the gas average temperature T4 according to the test results, and compare the gas average temperature T4 with the design value T 4* to obtain the deviation ∆T4 = T 4* - T4;

[0017] Calculate and determine the geometric parameters of the aero-engine that cannot be adjusted in real time during the test run according to the deviation ∆T4, including: the adjustment amount of the high-pressure turbine guide vane throat area A th , the low-pressure turbine guide vane throat area A tl and the outer duct exit area A 16 ;

[0018] Input the geometric parameters that cannot be adjusted in real time during the test run into the high-precision performance evaluation model of the aero-engine to perform an over-temperature judgment, and judge whether |∆T4| < m, where m is the temperature standard value. If |∆T4| < m at this time, it indicates that there is no problem of over-temperature of the gas average temperature.

[0019] Preferably, if |∆T4| ≥ m, then re-conduct the overall engine gas average temperature T4 regulation test and adjust the fan adjustable guide vane angle α1, the compressor adjustable guide vane angle α2, the nozzle throat area A8, and the nozzle exit area A9 in real time online until |∆T4| < m.

[0020] Preferably, the specific method for establishing a relationship model based on the non-uniformity δ of the oil-gas distribution of the core vortex core in the main combustion zone is: conduct a numerical simulation of the main combustor to establish the relationship between the non-uniformity of the oil-gas distribution of the core vortex core in the main combustion zone δ and the air flow rate W a in its corresponding area, W f the fuel flow rate, S the swirl number, V the speed, A cv and the characteristic size of the core vortex core δ = f(W a ,W f , S, V, A cv )。

[0021] Preferably, the multi-parameter coupling design is carried out by selecting different air flow rates W a , fuel flow rates W f , speeds V and the characteristic size of the core vortex core A cv to obtain a variety of different design schemes.

[0022] Preferably, the specific design of the optimal gas temperature distribution is as follows: Obtain the gas temperature distribution design and blade surface temperature data of the current aero-engine, select the neural network method, establish the relationship between different gas temperature distributions and blade surface temperatures, calculate the corresponding blade strength reserve through the blade surface temperature, obtain the corresponding comprehensive strength coefficient through the blade strength reserve, and select the gas temperature distribution corresponding to the maximum value of the comprehensive strength coefficient, which is the optimal gas temperature distribution.

[0023] Preferably, the temperature distribution requirement design is carried out by means of a vortex casing control and an integrated regulation method for root tip cooling and mixing.

[0024] Preferably, the vortex casing control and the integrated regulation method for root tip cooling and mixing are specifically as follows:

[0025] The gas temperature distribution is divided into five parts according to the radial height, namely the tip near-wall region, the upper-middle region, the middle region, the lower-middle region, and the root near-wall region;

[0026] The temperature in the tip near-wall region is regulated by controlling the size of the cooling air volume; the temperature in the upper-middle region is regulated by controlling the proportion of the mixing air volume; the temperature in the middle region is regulated by controlling the vortex core size; the temperature in the lower-middle region is regulated by controlling the proportion of the mixing air volume; the temperature in the root near-wall region is regulated by controlling the size of the cooling air volume.

[0027] The aero-engine comprehensive temperature control and verification method of the present application splits the gas temperature field into three dimensions, including the average gas temperature, the peak gas temperature, and the gas temperature distribution; and separately carries out the regulation of the average gas temperature, the regulation of the peak gas temperature, and the regulation of the gas temperature distribution; constructs a high-quality gas temperature field, ensures the reliable operation of high-temperature components such as turbine blades, solves the problems of "over-temperature" of the average gas temperature, too high peak gas temperature, and unreasonable gas temperature distribution, reduces the required gas temperature while meeting the thrust requirement, greatly reduces the temperature load of hot-end components such as the combustion chamber and the turbine, solves the "over-temperature" problem of aero-engines, and realizes reliable operation at high gas temperatures. Description of the Drawings

[0028] To more clearly illustrate the technical solution provided by this application, the accompanying drawings will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of this application.

[0029] Figure 1 This is the overall system flowchart of this application;

[0030] Figure 2 This is the flowchart for regulating the average gas temperature of this application;

[0031] Figure 3 This is the flowchart for regulating the peak gas temperature of this application;

[0032] Figure 4 This is the flowchart for regulating the gas temperature distribution of this application. Specific Embodiments

[0033] Next, the technical solutions in the embodiments of the present invention will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, rather than all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

[0034] An integrated temperature control and verification method for an aeroengine performs over-temperature control in the following four directions: First, the gas temperature field is split into three dimensions, namely the average gas temperature, the peak gas temperature, and the gas temperature distribution; second, the average gas temperature is regulated based on the true characteristics of components in the whole-engine environment; third, the peak gas temperature is regulated by the core vortex core size of the main combustion chamber; fourth, the gas temperature distribution is regulated by the core vortex core size of the main combustion chamber and the integration of root tip cooling and mixing; finally, a high-quality gas temperature field is constructed to ensure the reliable operation of high-temperature components such as turbine blades. The integrated temperature control logic diagram is shown in Figure 1 .

[0035] Specifically, it includes the following steps:

[0036] Step S100, regulation of the average gas temperature:

[0037] As Figure 2 , a high-precision performance evaluation model of the aeroengine is established based on the true characteristics of components under the whole-engine conditions. The average gas temperature and deviation are calculated through the whole-engine average gas temperature regulation test. Then, the adjustment amount of the aeroengine is determined according to the deviation, and the deviation of the adjusted average gas temperature is evaluated through the high-precision performance evaluation model of the aeroengine. The structural design parameters of the aeroengine are adjusted according to the evaluation results until there is no over-temperature problem under the current deviation.

[0038] Specifically:

[0039] Step S110: Obtain the true characteristics of components such as the fan, compressor, combustor, and turbine under the overall environment of the aero-engine through special engine tests, and establish a high-precision performance evaluation model of the aero-engine based on the true characteristics of the components under the overall engine conditions;

[0040] Step S120: Obtain the adjustable geometric parameters during the aero-engine test run, including the fan adjustable guide vane angle α1, the compressor adjustable guide vane angle α2, the nozzle throat area A8, and the nozzle exit area A9; conduct an overall engine gas average temperature T4 regulation test according to the adjustable geometric parameters of the aero-engine, calculate the gas average temperature T4 based on the test results, and compare the gas average temperature T4 with the design value T 4* to obtain the deviation amount ∆T4 = T 4* - T4;

[0041] Step S130: Calculate and determine the geometric parameters of the aero-engine that cannot be adjusted in real time during the test run according to the deviation amount ∆T4, including: the adjustment amount of the high-pressure turbine guide vane throat area A th , the low-pressure turbine guide vane throat area A tl and the outer bypass exit area A 16 ;

[0042] Step S140: Input the geometric parameters that cannot be adjusted in real time during the test run into the high-precision performance evaluation model of the aero-engine for over-temperature judgment. Specifically: judge whether |∆T4| < m, where m is the temperature standard value. If |∆T4| < m at this time, it indicates that the T4 during the overall engine test run is close to the design value and there is no problem of "over-temperature" of the gas average temperature; otherwise, if |∆T4| ≥ m, it is necessary to re-conduct the overall engine gas average temperature T4 regulation test and adjust the fan adjustable guide vane angle α1, the compressor adjustable guide vane angle α2, the nozzle throat area A8, and the nozzle exit area A9 in real time online until |∆T4| < m. The value of m is determined according to the margin of the engine gas average temperature and generally takes a value of 10 - 20 °C.

[0043] Step S200: Regulation of the gas peak temperature:

[0044] Such as Figure 3 , establish a relationship model of the non-uniformity of the oil-gas distribution in the core vortex of the main combustion zone δ , determine the swirl number S in the relationship model, conduct multi-parameter coupling design according to the swirl number S to obtain a variety of different design schemes, and then conduct verification tests according to different design schemes respectively until a design scheme that meets the design requirements is obtained. Obtain the gas peak temperature regulation parameters according to the design scheme that meets the design requirements; conduct gas peak temperature regulation according to the gas peak temperature regulation parameters.

[0045] Step S210: The non-uniformity of the oil-gas distribution in the core vortex of the main combustion zoneδ is a key factor affecting the peak temperature of the combustion chamber gas. In order to obtain the non-uniformity of the oil-gas distribution in the core vortex of the main combustion zone δ, First, a relational model based on the non-uniformity δ of the oil-gas distribution in the core vortex of the main combustion zone needs to be established ， The establishment method is as follows: Conduct numerical simulation of the main combustion chamber to establish the relationship model between the non-uniformity of the oil-gas distribution in the core vortex of the main combustion zone δ and the control parameters of the peak temperature of the combustion chamber gas in the corresponding area. The control parameters of the peak temperature of the combustion chamber gas include air flow rate W a , fuel flow rate W f , swirl number S , velocity V and the characteristic size of the core vortex A cv and other parameters δ = f(W a ,W f , S, V, A cv ) .

[0046] Step S220, determine the swirl number in the relational model S . Among them, the larger the swirl number (that is, the higher the swirl intensity) and the larger the core vortex size, the more uniform the oil-gas mixing, the smaller the non-uniformity, and the more conducive it is to reducing the peak temperature. However, considering the size of the main combustion chamber of existing high-performance aero-engines and the import flow rate limitation, the higher the swirl number, the more difficult it is to achieve

[0047] Step S230, after the swirl number S is determined, according to the swirl number S carry out multi-parameter coupling design of different air flow rates W a , fuel flow rates W f , velocities V and the characteristic size of the core vortex A cv . According to different parameter couplings, a variety of different design schemes are obtained. Considering the main combustion size of the high-performance main combustion chamber at this stage and the import flow rate limitation, generally, the swirl number S =3 is selected. For example, the "double-stage axial + annular cooling radial" triple swirl flame tube head configuration forms a third strong swirl with the cooling air in the axial swirler and the mounting seat ring cavity, greatly increasing the vortex core size, reducing the temperature at the edge of the vortex core through cooling, avoiding excessive heat load on the flame tube wall, and reducing the peak temperature; or, when the size of the main combustion chamber permits, it is designed as a "three-stage axial". The triple swirl flame tube head configuration can also well increase the vortex core size and reduce the peak temperature

[0048] Step S240: Conduct a verification test on the peak gas temperature in the main combustion chamber according to the current design scheme to verify whether the peak gas temperature in the main combustion chamber meets the design requirements. If not, repeat steps S220 - S240 until a design scheme that meets the design requirements is obtained. Determine the gas peak temperature control parameters according to this design scheme, including: swirl number S、 air flow rate W a , fuel flow rate W f , speed V and the characteristic size of the core vortex core A cv and other parameters.

[0049] Step S300: Gas temperature distribution control:

[0050] For example Figure 4 , obtain the gas temperature distribution design and blade surface temperature data of the current aero - engine for data processing to obtain the optimal gas temperature distribution; conduct a temperature distribution requirement design for the aero - engine according to the optimal gas temperature distribution, and conduct a verification test on the gas temperature distribution in the main combustion chamber according to the temperature distribution requirement design result to determine whether it meets the gas temperature distribution requirements. If not, re - conduct the temperature distribution requirement design until the gas temperature distribution requirements are met.

[0051] Step S310: Determine the optimal gas temperature distribution requirements as follows: Obtain the gas temperature distribution design and blade surface temperature data of the current aero - engine, select an appropriate data - processing method, establish the relationship between different gas temperature distributions and blade surface temperatures, calculate the corresponding blade strength reserve through the blade surface temperature, obtain the corresponding comprehensive strength coefficient through the blade strength reserve, and select the gas temperature distribution corresponding to the maximum comprehensive strength coefficient as the optimal gas temperature distribution. The preferred data - processing method is preferably a neural network or other methods.

[0052] Step S320: According to the optimal gas temperature distribution requirements, conduct a temperature distribution requirement design through the integrated control method of volute control and root tip cooling mixing to obtain the temperature distribution requirement design result. The specific method is as follows: Divide the gas temperature distribution into five parts according to the radial height, namely the tip near - wall region, the upper - middle region, the middle region, the lower - middle region, and the root near - wall region. The control methods are as follows:

[0053] Tip near - wall region: Close to the upper wall surface, the cooling aerodynamic quantity of the wall cooling holes is sufficient, and the temperature is controlled by adjusting the size of the cooling air volume;

[0054] Upper - middle region: The gas in this region is mainly the mixing gas, and the temperature is controlled by adjusting the proportion of the mixing gas volume;

[0055] Middle region: This region is the core vortex core area, and the temperature is regulated by controlling the vortex core size;

[0056] Lower-middle region: This region is mainly composed of the premixed gas, and the temperature is regulated by controlling the proportion of the premixed gas volume;

[0057] Root near-wall region: Close to the upper wall surface, the cooling aerodynamic momentum of the wall cooling holes is sufficient, and the temperature is regulated by controlling the size of the cooling gas volume.

[0058] Step S330: Conduct a verification test on the gas temperature distribution in the main combustion chamber according to the design result of the temperature distribution requirement. Obtain the comprehensive strength coefficient according to the verification test result, and then verify whether the gas temperature distribution in the main combustion chamber meets the best gas temperature distribution requirement according to the comprehensive strength coefficient. If not, repeat steps S320 - S330 until the best gas temperature distribution requirement is achieved and the gas temperature distribution regulation is completed.

[0059] The designs in the above steps S100, S200, and S300 are not in sequence, and can be carried out simultaneously or separately.

[0060] In summary, in this application, the gas temperature field is split into three dimensions, including the average gas temperature, the peak gas temperature, and the gas temperature distribution; and the regulation of the average gas temperature, the regulation of the peak gas temperature, and the regulation of the gas temperature distribution are carried out respectively; a high-quality gas temperature field is constructed to ensure the reliable operation of high-temperature components such as turbine blades, solve the problems of "over-temperature" of the average gas temperature, too high peak gas temperature, and unreasonable gas temperature distribution, reduce the required gas temperature while meeting the thrust requirement, greatly reduce the temperature load of hot-end components such as the combustion chamber and turbine, solve the "over-temperature" problem of aero-engines, and achieve reliable operation at high gas temperatures.

[0061] Finally, it should be noted that: in the attached drawings of the disclosed embodiments of the present invention, only the structures related to the disclosed embodiments are involved. Other structures can refer to the general design. Without conflict, the same embodiment and different embodiments of the present invention can be combined with each other;

[0062] Finally: The above are only the preferred embodiments of the present invention and are not used to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A method for comprehensive temperature control and verification of an aircraft engine, characterized in that: Including: Average gas temperature regulation: Based on the actual characteristics of components under the whole-engine conditions, establish a high-precision performance evaluation model for aero-engines. Through the whole-engine average gas temperature regulation test, calculate the average gas temperature and the deviation, then determine the adjustment amount of the aero-engine according to the deviation, and evaluate the deviation of the adjusted average gas temperature through the high-precision performance evaluation model of the aero-engine. Adjust the structural design parameters of the aero-engine according to the evaluation results until there is no over-temperature problem under the current deviation. Peak gas temperature regulation: Establish a relationship model based on the non-uniformity δ of the oil-gas distribution in the core vortex core of the main combustion zone, determine the swirl number S in the relationship model, and conduct multi-parameter coupling design according to the swirl number S to obtain a variety of different design schemes. Then, conduct verification tests according to different design schemes until a design scheme that meets the design requirements is obtained, and obtain the peak gas temperature regulation parameters according to the design scheme that meets the design requirements. Conduct peak gas temperature regulation according to the peak gas temperature regulation parameters. Gas temperature distribution regulation: Obtain the gas temperature distribution design and blade surface temperature data of the current aero-engine for data processing to obtain the optimal gas temperature distribution; conduct the temperature distribution requirement design of the aero-engine according to the optimal gas temperature distribution, and conduct the verification test of the gas temperature distribution in the main combustion chamber according to the temperature distribution requirement design result to judge whether it meets the gas temperature distribution requirements. If it does not meet, re-conduct the temperature distribution requirement design until it meets the gas temperature distribution requirements. The specific design method of the average gas temperature regulation is as follows: Obtain the actual characteristics of components such as the fan, compressor, combustion chamber, and turbine under the whole-engine environment of the aero-engine through special engine tests, and establish a high-precision performance evaluation model for the aero-engine based on the actual characteristics of components under the whole-engine conditions. The adjustable geometric parameters of the aircraft engine during the test run are obtained, including the fan adjustable guide vane angle α1, the compressor adjustable guide vane angle α2, the nozzle throat area A8 and the nozzle outlet area A9; the whole machine gas average temperature T4 control test is carried out according to the adjustable geometric parameters of the aircraft engine, the gas average temperature T4 is calculated according to the test results, and the gas average temperature T4 is compared with the design value T 4* By comparison, we can get the deviation ∆T4=T 4* -T4; The geometric parameters of the aircraft engine that cannot be adjusted in real time during the test run are calculated based on the deviation ∆T4, including: the throat area of ​​the high-pressure turbine guide vane A th , Low-pressure turbine guide vane throat area A tl and the outer culvert outlet area A 16 The amount of adjustment; Input the geometric parameters that cannot be adjusted in real time during the test run into the high-precision performance evaluation model of the aero-engine to conduct over-temperature judgment, and judge whether |∆T4| < m, where m is the temperature standard value. If |∆T4| < m at this time, it indicates that there is no over-temperature problem with the average gas temperature.

2. The aircraft engine comprehensive temperature control and verification method according to claim 1, characterized in that: If |∆T4| ≥ m, then re-conduct the whole-engine average gas temperature T4 regulation test, and adjust the fan variable guide vane angle α1, compressor variable guide vane angle α2, nozzle throat area A8, and nozzle exit area A9 in real time online until |∆T4| < m.

3. The aero-engine comprehensive temperature control and verification method according to claim 1, characterized in that The specific method of establishing the relationship model based on the oil and gas distribution unevenness δ of the core vortex core of the main combustion zone is: numerical simulation of the main combustion chamber is carried out to establish the oil and gas distribution unevenness δ of the core vortex core of the main combustion zone. δ The air flow in the corresponding area W a , fuel flow W f , swirl number S ,speed V and the core vortex core characteristic size A cv The relationship model between δ = f(W a ,W f , S, V, A cv )。 4. The aircraft engine comprehensive temperature control and verification method according to claim 1, characterized in that: Multi-parameter coupling design uses different air flow rates W a , fuel flow W f ,speed V and the core vortex core characteristic size A cv Carry out design and obtain a variety of different design schemes.

5. The aircraft engine comprehensive temperature control and verification method according to claim 1, characterized in that: The specific design of the optimal gas temperature distribution is as follows: Obtain the gas temperature distribution design and blade surface temperature data of the current aero-engine, select the neural network method, establish the relationship between different gas temperature distributions and blade surface temperatures, calculate the corresponding blade strength reserve through the blade surface temperature, obtain the corresponding comprehensive strength coefficient through the blade strength reserve, and select the gas temperature distribution corresponding to the maximum comprehensive strength coefficient as the optimal gas temperature distribution.

6. The aircraft engine comprehensive temperature control and verification method according to claim 1, characterized in that: Conduct the temperature distribution requirement design through the integrated regulation method of volute control and root tip cooling mixing.

7. The aircraft engine comprehensive temperature control and verification method according to claim 6, characterized in that: The integrated control method for volute control and root tip cooling and mixing is specifically as follows: The gas temperature distribution is divided into five parts according to the radial height, namely, the tip near wall area, the middle upper area, the middle area, the middle lower area, and the root near wall area; The temperature of the tip near-wall area is regulated by controlling the amount of cooling air; the temperature of the middle and upper area is regulated by controlling the proportion of the mixed air; the temperature of the middle area is regulated by controlling the size of the vortex core; the temperature of the middle and lower area is regulated by controlling the proportion of the mixed air; the temperature of the root near-wall area is regulated by controlling the amount of cooling air.

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