Comprehensive temperature control and verification method for aero-engine

Abstract

The invention belongs to the field of aero-engine design, and particularly relates to a comprehensive temperature control and verification method for an aero-engine, and the method comprises the steps: dividing a gas temperature field into three dimensions, including gas average temperature, gas peak temperature and gas temperature distribution; fuel gas average temperature regulation and control, fuel gas peak temperature regulation and control and fuel gas temperature distribution regulation and control are carried out respectively; a high-quality gas temperature field is constructed, reliable work of high-temperature components such as turbine blades is guaranteed, the problems of over-temperature of average gas temperature, over-high gas peak temperature and unreasonable gas temperature distribution are solved, the required gas temperature is reduced while thrust is not reduced, temperature loads of hot end components such as a combustion chamber and a turbine are greatly reduced, and the combustion efficiency is improved. The problem of overtemperature of the aero-engine is solved, and reliable work at high gas temperature is achieved.

Description

Technical Field

[0001] The present application belongs to the field of aircraft engine design, and in particular relates to a comprehensive temperature control and verification method for an aircraft engine. Background Art

[0002] High-temperature components such as aircraft engine turbine blades work in high-temperature combustion gases and are subject to high mechanical and aerodynamic loads. Constructing a high-quality combustion gas temperature field to achieve long-term reliable operation of high-temperature components such as turbine blades is the biggest technical challenge facing aircraft engine development. "Overtemperature" problems often occur during engine development, manifested as the combustion gas temperature field deviating from the design target. "Overtemperature" is the first obstacle that must be overcome in the development of advanced aircraft engines at home and abroad.

[0003] For a long time, the engine has been improved through a large number of tests and "trial and error temperature adjustment" to improve the gas temperature field, but the "overheating" problem has not been completely solved, resulting in insufficient engine temperature margin and slow reliability improvement; however, as the gas temperature continues to increase, "trial and error temperature adjustment" can no longer solve the problem of reliable engine operation. During the development of aviation engines, a large number of high-temperature components such as turbine blades and afterburner combustion chambers have been burned and broken, making the engine unable to work normally, which has seriously hindered the model development process.

[0004] The existing technical solution is to discover and optimize the gas temperature field problems through a large number of tests and debugging, repeated trial and error during the whole machine test run after the engine is completed. This consumes a lot of resources and the optimization results are uncertain. It is often difficult to achieve the design expectations, and the forward design and debugging of the gas temperature field are not achieved. Disadvantages include: 1) The problem of "overheating" of the average gas temperature occurs. Since the method of obtaining the real characteristics of components and solving the intrinsic flow in the whole machine environment has not been established, the engine development at home and abroad is based on the test characteristics of individual components to carry out engine performance design and optimization to control the average gas temperature. However, the individual component test cannot accurately simulate the actual conditions of the whole machine such as high temperature, high pressure and flow field. The obtained component characteristics are different from the real characteristics of the components in the whole machine environment, resulting in the deviation of the benchmark of engine performance design and optimization, making it difficult to achieve efficient coordination of various components, resulting in low thermal efficiency of the whole machine. The average gas temperature of the engine far exceeds the design value under the condition of achieving thrust indicators; 2) The problem of excessively high peak gas temperature occurs. The existing solutions are mainly based on the traditional double swirl flame tube head configuration and improving the gas temperature uniformity by increasing the core vortex core size. There is a lack of organizational combustion theory and design methods to effectively control the core vortex core edge temperature. As a result, when the vortex core is close to the flame tube wall, it is easy to cause wall ablation, and it is impossible to effectively improve the gas temperature uniformity by increasing the vortex core size; 3) There is an unreasonable gas temperature distribution problem. The existing technical solution design simply pursues the most uniform radial gas temperature distribution, but fails to adapt to the complex load of the turbine. As a result, key areas such as the root of the turbine rotor blades are subjected to large temperature and mechanical loads at the same time, and the reliability is greatly reduced. At the same time, the ability to adjust the temperature distribution by simply relying on the flame tube mixed gas is limited, and it is difficult to accurately achieve the expected temperature distribution.

[0005] Therefore, it is necessary to further solve the "overheating" problem of aircraft engines. Summary of the invention

[0006] The purpose of this application is to provide a comprehensive temperature control and verification method for aircraft engines to solve the problem of "overheating" in existing aircraft engines that is difficult to effectively solve.

[0007] The technical solution of this application is: a comprehensive temperature control and verification method for an aircraft engine, comprising: Gas average temperature control: A high-precision performance evaluation model for aircraft engines is established based on the true characteristics of components under whole-machine conditions. The average gas temperature and deviation are calculated through the whole-machine gas average temperature control test. Then, the adjustment amount of the aircraft engine is determined based on the deviation. The deviation of the adjusted average gas temperature is evaluated through the high-precision performance evaluation model for aircraft engines. The structural design parameters of the aircraft engine are adjusted based on the evaluation results until there is no overtemperature problem at the current deviation. Regulation of gas peak temperature: establish a relationship model based on the oil and gas distribution unevenness δ of the core vortex core of the main combustion zone, determine the swirl number S in the relationship model, perform multi-parameter coupling design based on the swirl number S, obtain a variety of different design schemes, and then conduct verification tests according to different design schemes, until a design scheme that meets the design requirements is obtained, and obtain the gas peak temperature control parameters according to the design scheme that meets the design requirements; and control the gas peak temperature according to the gas peak temperature control parameters; Gas temperature distribution control: Obtain the gas temperature distribution design and blade surface temperature data of the current aircraft engine for data processing to obtain the optimal gas temperature distribution; design the temperature distribution requirements of the aircraft engine based on the optimal gas temperature distribution, and conduct verification tests on the main combustion chamber gas temperature distribution based on the temperature distribution requirement design results to determine whether the gas temperature distribution requirements are met. If not, redesign the temperature distribution requirements until the gas temperature distribution requirements are met.

[0008] Preferably, the specific design method for controlling the average gas temperature is: Through engine-specific tests, the true characteristics of the fan, compressor, combustion chamber, and turbine components under the complete environment of the aircraft engine are obtained, and based on the true characteristics of the components under complete conditions, a high-precision performance evaluation model for the aircraft engine is established; Obtain the adjustable geometric parameters during the test run of the aero-engine, including the angle α1 of the adjustable fan guide vane, the angle α2 of the adjustable compressor guide vane, the throat area A8 of the nozzle, and the exit area A9 of the nozzle; conduct a regulation test on the average gas temperature T4 of the whole engine according to the adjustable geometric parameters of the aero-engine, calculate the average gas temperature T4 based on the test results, and compare the average gas temperature T4 with the design value T 4* to obtain the deviation ΔT4 = T 4* - T4; 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 ΔT4, including: the adjustment amount of the throat area A th of the high-pressure turbine guide vane, the throat area A tl of the low-pressure turbine guide vane, and the exit area A 16 of the bypass duct; 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 |ΔT4| < m, where m is the temperature standard value. If |ΔT4| < m at this time, it indicates that there is no problem of over-temperature of the average gas temperature.

[0009] Preferably, if |ΔT4| ≥ m, then re-conduct the regulation test on the average gas temperature T4 of the whole engine, and adjust the angle α1 of the adjustable fan guide vane, the angle α2 of the adjustable compressor guide vane, the throat area A8 of the nozzle, and the exit area A9 of the nozzle in real time online until |ΔT4| < m.

[0010] Preferably, the specific method for establishing the 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 combustion chamber, and establish the relationship model between the non-uniformity δ of the oil-gas distribution of the core vortex core in the main combustion zone and the air flow rate W a in its corresponding area, the fuel flow rate W f the swirl number S the speed V and the characteristic size A cv of the core vortex core; δ = f(W a ,W f , S,V,A cv )。

[0011] Preferably, the multi-parameter coupling design is carried out by selecting different air flow rates W a the fuel flow rate W f the speed Vand the core vortex core characteristic size A cv Carry out design and obtain a variety of different design schemes.

[0012] Preferably, the optimal gas temperature distribution is specifically designed as follows: obtain the gas temperature distribution design and blade surface temperature data of the current aircraft engine, select a neural network method, establish the relationship between different gas temperature distributions and blade surface temperature, 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.

[0013] Preferably, the temperature distribution requirement design is performed through an integrated control method of volute control and root tip cooling and mixing.

[0014] Preferably, the integrated control method of 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.

[0015] The comprehensive temperature control and verification method for aircraft engines of the present application divides the gas temperature field into three dimensions, including average gas temperature, peak gas temperature, and gas temperature distribution; and respectively carries out average gas temperature control, peak gas temperature control, and gas temperature distribution control; constructs a high-quality gas temperature field to ensure the reliable operation of high-temperature components such as turbine blades, solves the problems of "overtemperature" of the average gas temperature, excessive peak gas temperature, and unreasonable gas temperature distribution, reduces the required gas temperature while ensuring that the thrust is not reduced, and greatly reduces the temperature load of hot end components such as the combustion chamber and turbine, solves the "overtemperature" problem of aircraft engines, and achieves reliable operation at high gas temperatures. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] In order to more clearly illustrate the technical solution provided by the present application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of the present application.

[0017] Figure 1 This is the overall system flow chart of this application; Figure 2 This is the average gas temperature control flow chart for this application; Figure 3A flow chart for regulating the peak gas temperature for this application; Figure 4 This is the flow chart of gas temperature distribution control for this application. DETAILED DESCRIPTION

[0018] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0019] A comprehensive temperature control and verification method for aircraft engines is used to control over-temperature in the following four directions: first, the gas temperature field is divided 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 real characteristics of the components under the whole machine environment; third, the peak gas temperature is regulated by the size of the core vortex core of the main combustion chamber; fourth, the gas temperature distribution is regulated by the integration of the size of the core vortex core of the main combustion chamber and the 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 comprehensive temperature control logic diagram is shown in Figure 1 .

[0020] The specific steps include: Step S100, average gas temperature control: like Figure 2 A high-precision performance evaluation model for aircraft engines is established based on the true characteristics of components under whole-machine conditions. The average gas temperature and deviation are calculated through the whole-machine gas average temperature control test. Then, the adjustment amount of the aircraft engine is determined according to the deviation. The deviation of the adjusted average gas temperature is evaluated through the high-precision performance evaluation model of the aircraft engine. According to the evaluation results, the structural design parameters of the aircraft engine are adjusted until there is no over-temperature problem under the current deviation.

[0021] Specifically: Step S110, obtaining the real characteristics of the fan, compressor, combustion chamber, turbine and other components under the whole machine environment of the aircraft engine through engine special tests, and establishing a high-precision performance evaluation model of the aircraft engine based on the real characteristics of the components under the whole machine condition; Step S120, obtaining the adjustable geometric parameters of the aircraft engine during the 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 outlet area A9; conducting a whole-machine gas average temperature T4 control test according to the adjustable geometric parameters of the aircraft engine, calculating the gas average temperature T4 according to the test results, and comparing the gas average temperature T4 with the design value T 4*Compare to obtain the deviation amount ∆T4 = T 4* - T4; Step S130, calculate and determine the geometric parameters of the aero - engine that cannot be adjusted in real - time during the test run, including: the adjustment amount of the throat area A th of the high - pressure turbine guide vane, the throat area A tl of the low - pressure turbine guide vane, and the exit area A 16 of the bypass duct; 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 T4 during the whole - machine test run is close to the design value, and there is no problem of "over - temperature" of the average gas temperature; conversely, if |∆T4| ≥ m, it is necessary to re - conduct the regulation test of the average gas temperature T4 of the whole machine and adjust the angles α1 of the fan adjustable guide vanes, α2 of the compressor adjustable guide vanes, the throat area A8 of the nozzle, and the exit area A9 of the nozzle in real - time online until |∆T4| < m. m is determined according to the margin of the average gas temperature of the engine, and generally takes a value of 10 - 20 °C.

[0022] Step S200, regulation of the peak gas temperature: For example 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 until a design scheme that meets the design requirements is obtained. 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.

[0023] Step S210, the non - uniformity of the oil - gas distribution in the core vortex of the main combustion zone δ is a key factor affecting the peak gas temperature in the combustion chamber. In order to obtain the non - uniformity of the oil - gas distribution in the core vortex of the main combustion zone δ, , first, a relationship 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, and establish a relationship model between the non - uniformity of the oil - gas distribution in the core vortex of the main combustion zone δ and the peak gas temperature regulation parameters in its corresponding area. The peak gas temperature regulation parameters include the air flow rate W a , the fuel flow rate W f , the swirl number S , the velocity V and the characteristic size of the core vortexA cv And other parameters, δ = f(W a ,W f ,S,V,A cv ) .

[0024] Step S220, determining the swirl number in the relational model S The larger the swirl number (that is, the higher the swirl intensity) and the larger the core vortex size, the more uniform the oil and gas mixing and the smaller the unevenness, which is more conducive to reducing the peak temperature. However, considering the size of the main combustion chamber of existing high-performance aircraft engines and the inlet flow restrictions, the higher the swirl number, the more difficult it is to achieve.

[0025] Step S230, swirl number S After determination, according to the swirl number S Carry out different air flow W a , fuel flow W f ,speed V and the core vortex core characteristic size A cv According to the multi-parameter coupling design, a variety of different design schemes are obtained according to different parameter coupling. Considering the main combustion size of the high-performance main combustion chamber at this stage, the inlet flow restriction, etc., generally speaking, the swirl number is selected. S =3, such as the "two-stage axial + annular cooling radial" triple swirl flame tube head configuration, which forms a third strong swirl through the cooling air in the axial swirler and the mounting ring cavity, greatly increasing the size of the vortex core, reducing the edge temperature of the vortex core through cooling, avoiding excessive heat load on the flame tube wall, and reducing the peak temperature; or if the size of the main combustion chamber allows, it can be designed as a "three-stage axial". The triple swirl flame tube head configuration can also effectively increase the size of the vortex core and reduce the peak temperature; Step S240, conduct a verification test of the peak temperature of the main combustion chamber gas according to the current design scheme to verify whether the peak temperature of the main combustion chamber gas meets the design requirements. If not, repeat steps S220-S240 until a design scheme that meets the design requirements is obtained. According to the design scheme, determine the gas peak temperature control parameters, including: swirl number S、 Air flow W a , fuel flow W f ,speed V and the core vortex core characteristic size A cv And other parameters.

[0026] Step S300, gas temperature distribution control: like Figure 4 , obtain the current aircraft engine's gas temperature distribution design and blade surface temperature data for data processing to obtain the optimal gas temperature distribution; design the aircraft engine's temperature distribution requirements based on the optimal gas temperature distribution, and conduct a verification test of the main combustion chamber gas temperature distribution based on the temperature distribution requirement design results to determine whether the gas temperature distribution requirements are met. If not, redesign the temperature distribution requirements until the gas temperature distribution requirements are met.

[0027] Step S310, determining the optimal gas temperature distribution requirement, the method is as follows: obtaining the gas temperature distribution design and blade surface temperature data of the current aircraft engine, selecting an adaptive data processing method, establishing the relationship between different gas temperature distributions and blade surface temperature, calculating the corresponding blade strength reserve through the blade surface temperature, obtaining the corresponding comprehensive strength coefficient through the blade strength reserve, and selecting the gas temperature distribution corresponding to the maximum value of the comprehensive strength coefficient, which is the optimal gas temperature distribution. The adaptive data processing method preferably uses a neural network or other method.

[0028] Step S320, according to the optimal gas temperature distribution requirement, the temperature distribution requirement is designed through the integrated control method of vortex casing control and root tip cooling and mixing, and the temperature distribution requirement design result is obtained. The specific method is 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, and the control method is as follows: Near-wall area at the tip: close to the upper wall, the cooling air flow of the wall cooling holes is sufficient, and the temperature can be regulated by controlling the cooling air flow; Middle and upper area: This area is mainly mixed gas, and the temperature is controlled by controlling the proportion of mixed gas; Middle area: This area is the core vortex core area, and the temperature is regulated by controlling the vortex core size; Middle and lower area: This area is mainly mixed gas, and the temperature is controlled by controlling the proportion of mixed gas; Root near-wall area: close to the upper wall, the cooling air momentum of the wall cooling holes is sufficient, and the temperature can be regulated by controlling the amount of cooling air.

[0029] Step S330, carry out verification test of gas temperature distribution in the main combustion chamber according to the design result of temperature distribution requirement, obtain comprehensive strength coefficient according to the verification test result, and then verify whether gas temperature distribution in the main combustion chamber meets the optimal gas temperature distribution requirement according to the comprehensive strength coefficient. If not, repeat steps S320-S330 until the optimal gas temperature distribution requirement is achieved and the gas temperature distribution control is completed.

[0030] The designs in the above steps S100, S200 and S300 are not sequentially ordered and can be performed simultaneously or separately.

[0031] In summary, the present application divides the gas temperature field into three dimensions, including the average gas temperature, the peak gas temperature, and the gas temperature distribution; and respectively carries out the control of the average gas temperature, the control of the peak gas temperature, and the control of the gas temperature distribution; constructs a high-quality gas temperature field to ensure the reliable operation of high-temperature components such as turbine blades, solves the problems of "overtemperature" of the average gas temperature, excessive peak gas temperature, and unreasonable gas temperature distribution, reduces the required gas temperature while ensuring that the thrust is not reduced, and greatly reduces the temperature load of hot end components such as the combustion chamber and turbine, solves the "overtemperature" problem of aircraft engines, and achieves reliable operation at high gas temperatures.

[0032] Finally, it should be noted that: the drawings of the embodiments disclosed in the present invention only involve structures related to the embodiments disclosed in the present invention, and other structures can refer to common designs. In the absence of conflicts, the same embodiment and different embodiments of the present invention can be combined with each other; Finally: The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should 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. Calculate the average gas temperature and deviation through the whole-engine average gas temperature regulation test. 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 perform 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. Regulate the peak gas temperature 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 carry out 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 not, re-conduct the temperature distribution requirement design until the gas temperature distribution requirements are met.

2. The aircraft engine comprehensive temperature control and verification method according to claim 1, characterized in that: 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 for 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.

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

4. The integrated temperature control and verification method for an aero-engine 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 )。 5. The aircraft engine comprehensive temperature control and verification method according to claim 1, characterized in that: Multi-parameter coupling design is achieved by selecting 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.

6. 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.

7. 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.

8. The aircraft engine comprehensive temperature control and verification method according to claim 7, 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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