Self-adaptive high-pressure turbine cooling system for aero-engine and control method thereof

Abstract

The invention provides a self-adaptive high-pressure turbine cooling system for an aero-engine, which is used for self-adaptively adjusting the guide vane cooling air flow of a high-pressure turbine behind a core engine, and comprises a high-pressure turbine cooling structure for constructing a guide vane cooling air flow path for the high-pressure turbine; the acquisition control system is used for self-adaptive control of flow; various working conditions of the aircraft are determined according to the accelerator, the aircraft height, the aircraft elevation angle and other parameters, the needle valve opening degree is determined through early-stage engine simulation and test results, the needle valve opening degree is corrected through a machine learning method, and the purpose of the maximum fuel saving rate of the aircraft is achieved.

Description

Technical Field

[0001] The present invention belongs to the field of aero-engines, and in particular relates to an adaptive high-pressure turbine cooling system for an aero-engine and a control method thereof. Background Art

[0002] Today's advanced aircraft engines, as the main power source for aircraft and other aircraft, are developing towards high thrust-to-weight ratio, high reliability, environmental protection and long life.

[0003] As thrust demands on aircraft engines increase, the temperature upstream of the high-pressure turbine rises significantly, reaching 1900K to 2050K, requiring more cooling air to cool the high-pressure turbine guide vanes. Turbine cooling airflow requirements vary during takeoff, descent, acceleration, deceleration, subsonic cruise, and supersonic cruise. For example, supersonic cruise, where temperatures are highest, requires the largest flow rate, while subsonic cruise, where temperatures are lowest, requires the smallest flow rate.

[0004] With the application of high-temperature resistant metal materials capable of withstanding temperatures of 1100°C and composite materials capable of withstanding temperatures of 1350°C in aircraft engines, when the aircraft is in subsonic cruise mode, the turbine blades can withstand inlet gas temperatures of approximately 1000°C due to their inherent material properties. At this time, the demand for cooling flow is significantly reduced, and cooling may even be completely unnecessary. Given this situation, the development of a high-pressure turbine cooling system that can flexibly adjust the cooling air flow rate over a wider range has become an urgent need for industry development. The existing engine cooling solution is to distribute the same proportion of cooling air regardless of operating conditions. Because cooling air consumes energy when passing through the blades, using a large proportion of cooling air during low operating conditions such as subsonic cruise reduces the efficiency of the aircraft engine and increases the aircraft's fuel consumption.

[0005] As high-pressure turbine guide vanes adopt new technologies such as high-temperature resistant coatings, single crystal alloys, and composite materials, the temperature resistance is increased by more than 100°C, even reaching more than 1200°C. The need for cooling flow under medium and low operating conditions will be further reduced, further reducing fuel consumption.

[0006] Patent document CN117231368A discloses a cooling structure for the turbine of an aircraft engine combustion chamber bleed air. This scheme returns the bleed air between the outer wall of the combustion chamber and the rear end of the outer wall of the flame tube to the turbine rotor through the diffuser, the channel between the inner wall of the combustion chamber and the inner wall of the flame tube, thereby cooling the turbine rotor blades and achieving the purpose of reducing the cooling air flow of the turbine rotor blades under low operating conditions. Due to the small pressure difference between the bleed air end and the high-pressure turbine, the bleed air volume is relatively small.

[0007] GE's adaptive cycle engine uses a third-channel solution to significantly reduce fuel consumption under low operating conditions by reducing core engine flow and increasing the bypass ratio. Since the cooling flow ratio in the core engine remains unchanged, excessive cooling air will still be allocated under low operating conditions.

[0008] Patent document CN113738453B discloses a turbine guide vane cooling air flow control device, comprising: a main body, a cylindrical structure with an open lower end and a closed upper end, a first air inlet hole provided at the upper end of the main body's cylindrical wall, and a second air inlet hole provided at the lower end of the main body's cylindrical wall, both of which are connected to the main body's inner cavity; an intake casing, open at both ends, with the upper opening of the intake casing connected to the lower opening of the main body; high-pressure turbine guide vanes, disposed at the lower end outlet of the intake casing; a piston, disposed in the inner cavity of the main body and capable of moving up and down relative to the main body, with a flow hole group provided at the lower end of the piston capable of communicating with the second air inlet hole, and the flow area of ​​the flow hole group being smaller than the flow area of ​​the second air inlet hole. By adjusting the cooling air supply window area by moving the piston up and down, cooling air throttling is achieved, thereby controlling the amount of cooling air entering the high-pressure turbine guide vanes.

[0009] However, patent document CN113738453B fails to accurately reflect all engine operating conditions. The comparative patent only features two operating conditions: throttled and unthrottled. In this comparative patent, the flow hole group and the second air inlet are connected to achieve throttling, while the piston is higher than the second air inlet to achieve unthrottled operation. This patent fails to provide stepless adjustment of the cooling air flow rate within a flow range of 0-100%, making it impossible to adjust the cooling air flow rate to the most fuel-efficient level for different flight conditions, making it difficult to achieve optimal fuel savings. Summary of the Invention

[0010] In view of the defects in the prior art, the purpose of the present invention is to provide an adaptive high-pressure turbine cooling system for an aircraft engine and a control method thereof.

[0011] According to the present invention, an adaptive high-pressure turbine cooling system for an aircraft engine includes:

[0012] High-pressure turbine 10;

[0013] A high-pressure turbine cooling structure is used to construct a cooling air flow path for the guide blades of the high-pressure turbine 10;

[0014] The acquisition control system is used for adaptive control of the guide blade cooling air flow rate.

[0015] Preferably, the high-pressure turbine cooling structure comprises: a needle valve 8 with adjustable opening;

[0016] The cooling air enters the guide vanes through the needle valve 8;

[0017] The acquisition control system includes: a communication interface 203, a cooling model 202, a control module 207 and a learning module 201;

[0018] The communication interface 203 is used to obtain the real-time parameters of the aircraft's throttle 204, flight altitude 205, and aircraft pitch angle 206;

[0019] The cooling model 202 receives the real-time parameters transmitted by the communication interface 203, combines the real-time parameters with the previous simulation model and the aircraft engine test results, thereby determining the opening of the needle valve 8 and issuing instructions to the control module 207;

[0020] The control module 207 adjusts the opening of the needle valve 8 according to the instruction issued by the cooling model 202, thereby adjusting the cooling air flow to the guide vanes;

[0021] The learning module 201 uses a neural network model to evaluate the operating results of the cooling model 202 and correct the parameters of the cooling model 202 based on the results of each flight and test.

[0022] Preferably, the high-pressure turbine cooling structure comprises: a compressor 1, a diffuser 2, a combustion chamber outer wall 3, a combustion chamber inner wall 12, a flame tube 11 and an air bleed pipe 5;

[0023] The cooling air is drawn out from compressor 1 and passes through diffuser 2 before being divided into path A and path B;

[0024] The cooling air of line A passes through the channel between the flame tube 11 and the inner wall 12 of the combustion chamber and enters the front cavity 301 of the guide vane. The flow rate of the cooling air of line A is a fixed value.

[0025] The cooling air in path B passes through the passage between the flame tube 11 and the outer wall 3 of the combustion chamber, enters the bleed air pipe 5 through the small hole in the outer wall 3 of the combustion chamber, passes through the bleed air pipe 5, enters the needle valve 8, and then enters the rear cavity 302 of the guide vane. The flow rate of the cooling air in path B is adjusted by the opening of the needle valve 8.

[0026] The pressures in the front cavity 301 and the rear cavity 302 of the guide vane are both greater than the mainstream pressure before the turbine, so that the cooling gases in the A path and the B path do not interfere with each other during the flow process.

[0027] Preferably, the rear end side wall of the combustion chamber outer wall 3 has a plurality of circumferentially distributed air inlet holes, and the number of the air inlet holes is the same as the number of the needle valves 8;

[0028] One end of the plurality of air bleed pipes 5 is fixedly connected to the outer wall 3 of the combustion chamber and corresponds to the air bleed holes on the outer wall 3 of the combustion chamber, and the other end of the air bleed pipe 5 is tightly connected to the needle valve 8.

[0029] Preferably, the high-pressure turbine cooling structure further comprises: a stepping motor 6 and a bracket 7;

[0030] The outer side of the combustion chamber outer wall 3 is provided with a combustion chamber outer casing 4;

[0031] The stepper motor 6 is firmly connected to the bracket 7 , and the bracket 7 is firmly connected to the outside of the outer casing 4 of the combustion chamber.

[0032] Preferably, the needle valve 8 includes a needle 104, a valve core 101 and a housing 102;

[0033] The needle valve 8 is fixed on the outer casing 4 of the combustion chamber, and the needle valve 8 is arranged in the outer duct between the outer wall 3 of the combustion chamber and the outer casing 4 of the combustion chamber;

[0034] The shaft 103 of the stepper motor 6 is firmly connected to the needle 106 of the needle valve 8. The stepper motor 6 can drive the shaft 104 to move the needle 106 back and forth, thereby adjusting the opening of the needle valve 8.

[0035] The control module 207 controls the moving length of the shaft 103 by sending instructions to the stepping motor 6, thereby steplessly adjusting the opening of the control needle valve 8 between 0-100%;

[0036] The needle valve 8 is sealed by two polyimide sealing rings. The distance between the first sealing ring and the joint between the needle valve 8 and the outer casing 4 of the combustion chamber is 2 to 3 mm, and the distance between the two sealing rings is 2 to 3 mm.

[0037] Preferably, the high-pressure turbine cooling structure further comprises: a baffle 9;

[0038] The baffle 9 is between the high-pressure turbine 10 and the combustion chamber and is used to prevent the cooling gas from directly entering the high-pressure turbine 10 .

[0039] Preferably, the needle valve 8 is connected to the outer casing of the high-pressure turbine 10;

[0040] The front end of the outer casing is butted against the rear end of the outer wall 3 of the combustion chamber, the inner ring of the guide of the high-pressure turbine 10 is connected to the rear end of the inner wall 12 of the combustion chamber, and one end of a plurality of guide blades is mounted on the inner ring of the guide, and the other end is mounted on the outer casing;

[0041] The inner ring of the guide is provided with an air intake hole connected to the front cavity 301 , and the outer casing is provided with an air intake hole connected to the rear cavity 302 .

[0042] Preferably, the front cavity 301 of the guide vane and the rear cavity 302 of the guide vane are separated from each other; the front cavity 301 is close to the leading edge of the guide vane, and the rear cavity 302 is close to the trailing edge of the guide vane.

[0043] According to a control method for an adaptive high-pressure turbine cooling system for an aircraft engine provided by the present invention, when a throttle 204 is greater than a preset throttle threshold, a flight altitude 205 is less than or equal to a preset flight altitude threshold, and an aircraft pitch angle 206 is greater than a preset flight pitch angle threshold parameter, it is determined that the aircraft is in a takeoff condition, the opening of the needle valve 8 is adjusted to 50%-70%, and the parameters of the cooling model 202 are corrected using the learning module 201;

[0044] When the throttle 204 is at its maximum, the flight altitude 205 is higher than the flight altitude threshold, and the aircraft pitch angle 206 is flat, it is determined that the aircraft is in an accelerating state, and the opening of the needle valve 8 is adjusted to 100%;

[0045] When the throttle 204 is greater than the throttle threshold, the flight altitude 205 is higher than the flight altitude threshold, and the aircraft pitch angle 206 is flat, it is determined that the aircraft is in supersonic cruise mode, the opening of the needle valve 8 is adjusted to 70%-100%, and the parameters of the cooling model 202 are corrected through the learning module 201;

[0046] When the throttle 204 is less than or equal to the throttle threshold, the flight altitude 205 is higher than the flight altitude threshold, and the aircraft pitch angle 206 is flat, it is judged that the aircraft is in subsonic cruise condition. The material of the high-pressure turbine 10 is metal or composite material. For the guide vanes of the high-pressure turbine 10 made of metal, the opening of the needle valve 8 is adjusted to 10%-20%. For the guide vanes of the high-pressure turbine 10 made of composite material, the opening of the needle valve 8 is adjusted to 0%-20%. The parameters of the cooling model 202 are corrected through the learning module 201.

[0047] Compared with the prior art, the present invention has the following beneficial effects:

[0048] 1. It can adaptively and steplessly adjust the cooling air flow of the guide blades of the high-pressure turbine behind the core engine to achieve more precise adjustment.

[0049] 2. It can adjust the most appropriate cooling air flow according to various flight conditions to achieve the maximum fuel saving rate for the aircraft. For example, in high operating conditions such as supersonic cruise, a larger proportion of cooling air can be allocated. In low operating conditions such as subsonic cruise, a smaller proportion of cooling air can be allocated or even the cooling air can be turned off. This takes into account the thrust and efficiency requirements of the aircraft engine under different operating conditions.

[0050] 3. Strong versatility and interchangeability: Although the flow, pressure, and temperature of different engines are different, the flow adjustment range of the cooling system of the present invention is very large. It only needs to match the needle valve opening under different flight conditions and can be used on different aircraft engines without changing the design.

[0051] 4. Use the solution of needle valve opening throttling, especially the opening can be zero, to ensure the working condition of zero cooling flow demand and save fuel to the greatest extent. BRIEF DESCRIPTION OF THE DRAWINGS

[0052] Other features, objects and advantages of the present invention will become more apparent upon reading the detailed description of non-limiting embodiments with reference to the following drawings:

[0053] Figure 1 Schematic diagram of a cooling structure for a high-pressure turbine of an aircraft engine according to an embodiment of the present invention;

[0054] Figure 2 Schematic diagram of a needle valve and a stepping motor connected together in an embodiment of the present invention;

[0055] Figure 3 Schematic diagram of a control system according to an embodiment of the present invention;

[0056] Figure 4 Schematic diagram of the internal cavity of the high-pressure turbine guide vane in an embodiment of the present invention.

[0057] The figure shows:

[0058] DETAILED DESCRIPTION

[0059] The present invention will be described in detail below with reference to specific embodiments. The following examples will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those skilled in the art, several changes and improvements can be made without departing from the scope of the present invention. These all fall within the scope of protection of the present invention.

[0060] The present invention aims to provide an adaptive high-pressure turbine cooling system for an aircraft engine, adapted to adaptively adjust the flow rate of cooling air for the guide vanes of the high-pressure turbine 10 behind the core engine. The adaptive high-pressure turbine cooling system includes a high-pressure turbine cooling structure that establishes a cooling air flow path for the guide vanes of the high-pressure turbine 10; and a data acquisition and control system for adaptive flow rate control. The system determines various aircraft operating conditions based on parameters such as the throttle 204, flight altitude 205, and aircraft pitch angle 206, determines the opening of the needle valve 8 based on previous engine simulation and test results, and modifies the needle valve 8 opening using machine learning methods to maximize fuel savings.

[0061] like Figure 1 As shown, the high-pressure turbine cooling structure includes a compressor 1, a diffuser 2, a combustion chamber outer wall 3, a combustion chamber inner wall 12, a flame tube 11, an air bleed pipe 5, a needle valve 8, a stepper motor 6, a baffle 9, a high-pressure turbine 10, and a bracket 7.

[0062] The cooling air is drawn out from the compressor 1 and passes through the diffuser 2 and is divided into two paths: A and B. The cooling air in path A passes through the channel between the flame tube 11 and the outer wall 12 of the combustion chamber and enters the front cavity 301 of the guide blade of the high-pressure turbine 10, and the flow rate of the cooling air in path A is a fixed value; the cooling air in path B passes through the annular channel between the flame tube 11 and the outer wall 3 of the combustion chamber, enters the air bleed pipe 5 through the small hole on the outer wall 3 of the combustion chamber, and then enters the needle valve 8 and enters the rear cavity 302 of the guide blade of the high-pressure turbine 10. The flow rate of the cooling air in path B is adjusted by the opening of the needle valve 8.

[0063] This invention is feasible for application in next-generation engines. The throttling elements (such as the needle valve 8 and stepper motor 6) are located away from high-temperature sections, allowing for a wider range of materials and improving feasibility. In response to the characteristics of next-generation engines, the present invention positions moving parts away from the combustion chamber and high-pressure turbine 10, with the needle valve 8 positioned in the duct between the outer combustion chamber wall 3 and the outer casing 4.

[0064] Bracket 7 is a metal structure with multiple holes on its side. It is screwed onto the outside of the combustor casing 4, isolating it from the high temperatures of the combustor casing 4 and preventing damage to the stepper motor 6. A needle valve 8 is connected to the combustor casing 4 and connected to the high-pressure turbine guide vanes via a pipeline. The combustor casing temperature is approximately 200°C, while the ambient temperature of the stepper motor 6 is below 100°C. Therefore, a stepper motor with a temperature resistance of 250°C can be used. The temperature of the secondary cooling air in the pipeline is approximately 650°C, and the needle 104, which comes into contact with the secondary cooling air, can be made of stainless steel.

[0065] Since the needle valve 8 can conduct heat to the outer casing 4 of the combustion chamber, the temperature of the sealing position in the needle valve 8 is expected to be only 300°C. The present invention will use a new type of polyimide sealing ring, which has a self-lubricating function and an operating temperature of 500°C. In addition, since the adjustment speed of the stepper motor 6 is relatively slow, the friction between its output shaft and the sealing ring is relatively small, and the new type of polyimide sealing ring can meet the use requirements. The rear end side wall of the combustion chamber outer wall 3 has a plurality of bleed holes distributed along the circumference, and the number of holes is the same as the number of needle valves 8. One end of the plurality of bleed pipes 5 is fixed to the combustion chamber outer wall 3 by a flange, corresponding to the bleed holes on the combustion chamber outer wall 3, and the other end of the bleed pipe 5 is threadedly connected to the needle valve 8.

[0066] The front end of the outer casing of the high-pressure turbine 10 is butted against the rear end of the outer wall 3 of the combustion chamber. The inner ring of the guide vane is connected to the rear end of the combustion chamber wall 12. The front end of the guide vane, where the front cavity 301 is located, is mounted on the inner ring of the guide vane, while the rear end, where the rear cavity 302 is located, is mounted on the outer casing of the high-pressure turbine 10. The number of guide vanes can be dozens. A baffle 9 is located between the high-pressure turbine 10 and the combustion chamber to prevent cooling gas from directly entering the high-pressure turbine 10.

[0067] Because the guide vane has a front cavity 301 and a rear cavity 302, as long as the pressure in both cavities is greater than the prevailing pressure in front of the turbine, the coolant entering from the inner ring of the guide vane and the coolant entering from the outer casing do not affect each other. If the guide vane has only one cavity, when needle valve 8 is not fully open, the coolant will experience a pressure drop when passing through needle valve 8. Therefore, the outlet pressure of needle valve 8 will be lower than the pressure of the coolant from the lower part of the guide vane, which may prevent the coolant from passing through needle valve 8.

[0068] The inner ring of the guide vane is provided with an air inlet hole communicating with the front cavity 301 , and the outer casing of the high-pressure turbine 10 is provided with an air inlet hole communicating with the rear cavity 302 .

[0069] like Figure 2 As shown, the stepper motor 6 is threadedly connected to the bracket 7, and the shaft 103 of the stepper motor 6 is threadedly connected to the needle 105 of the needle valve 8. When the stepper motor 6 receives a command from the control system, the drive shaft 103 drives the needle 105 to move back and forth to adjust the flow area of ​​the needle valve 8.

[0070] The needle valve 8 can be sealed with two polyimide sealing rings. The distance between the first sealing ring and the joint between the needle valve 8 and the outer casing 4 of the combustion chamber is 2 to 3 mm, and the spacing between the two sealing rings is 2 to 3 mm.

[0071] Needle valve 8 consists of a needle 106, a valve core 101, and a housing 102. The needle valve is sealed with two polyimide sealing rings 104. The height of the bypass determines the volume of needle valve 8 and the diameter of valve core 101, thus determining its overall dimensions. The throat diameter of valve core 101 is determined based on numerical flow simulation results. This ensures that when needle valve 8 is fully open, the needle valve flow resistance is less than 0.05 MPa. Under these conditions, the flow rate of a single needle valve 8 can be determined through numerical flow simulation.

[0072] The relationship between the flow rate of a single needle valve 8 and the total cooling flow rate determines the number of needle valves 8, which is 4-8. If the number of needle valves 8 is more than 8, it is necessary to coordinate with the overall engine and adjust the height of the outer duct to ensure that the number of needle valves 8 is not more than 8.

[0073] like Figure 3As shown, the control system includes a communication interface 203, a cooling model 202, a control module 207, and a learning module 201. The communication interface 203 is used to obtain parameters such as the aircraft's throttle 204, flight altitude 205, and pitch angle 206. Based on previous simulation models, aircraft engine test results, and real-time parameters such as the throttle 204, flight altitude 205, and pitch angle 206, the cooling model 202 issues instructions to the control module 207 regarding the number of movement steps of the stepper motor 6. Based on the instructions from the cooling model 202, the control module 207 controls the number of movement steps of the stepper motor 6 and the travel length of the control shaft 103, thereby steplessly adjusting the opening of the needle valve 8 between 0-100%. The learning module 201 uses a neural network model to evaluate the operating results of the cooling model 202 and adjusts the parameters of the cooling model 202 based on each flight and test results.

[0074] like Figure 4 As shown, the guide vane of the high-pressure turbine 10 has two inner cavities: a front cavity 301 and a rear cavity 302, and the two inner cavities are not connected to each other; the front cavity 301 is close to the leading edge of the guide vane and has a relatively high temperature, and the rear cavity 302 is close to the trailing edge of the guide vane and has a relatively low temperature.

[0075] The present invention also provides a control method applicable to the adaptive high-pressure turbine cooling system for the aero-engine of the present invention.

[0076] In some optional examples, when parameters such as the throttle 204 being greater than a preset throttle threshold, the flight altitude 205 being lower than a preset flight altitude threshold, and the aircraft pitch angle 206 being higher than a preset flight pitch angle threshold occur, it is determined that the aircraft is in a takeoff condition. According to simulation results, the combustion chamber temperature is medium at this time and will increase in the future. The total temperature of the mainstream gas at the blade inlet is medium, and a large proportion of cooling flow is required on the B path. The recommended opening of the needle valve 8 is 50%-70%. Through the learning module 201, the opening of the needle valve 8 is gradually optimized to the most reasonable ratio to minimize fuel consumption.

[0077] In some optional examples, when the throttle 204 is at maximum, the flight altitude 205 is higher than the flight altitude threshold, and the aircraft pitch angle 206 is equal to the parameters, it is determined that the aircraft is in an acceleration condition. According to simulation results, the combustion chamber temperature is high at this time, the total temperature of the mainstream gas at the blade inlet is the highest, and a large proportion of cooling flow is required in the B path. It is recommended that the opening of the needle valve 8 be 100%;

[0078] In some optional examples, when the throttle 204 is greater than the throttle threshold, the flight altitude 205 is higher than the flight altitude threshold, and the aircraft pitch angle 206 is equal to the parameters, it is determined that the aircraft is in supersonic cruise operation. According to simulation results, the combustion chamber temperature is high at this time, the B path requires a large proportion of cooling flow, and the opening of the needle valve 8 is 70%-100%. Through the learning module, the opening of the needle valve 8 is gradually optimized to the most reasonable ratio;

[0079] In some optional instances, when the throttle 204 is less than or equal to the throttle threshold, the flight altitude 205 is higher than the flight altitude threshold, and the aircraft pitch angle 206 is equal to the parameters, it is judged to be the aircraft subsonic cruise condition. According to the simulation results, the combustion chamber temperature is the lowest at this time, and the total temperature of the mainstream gas at the blade inlet is the lowest.

[0080] The high-pressure turbine 10 can be made of metal or composite materials. If the turbine guide vanes are made of metal, a low proportion of cooling flow is required in path B, and the opening of needle valve 8 is 10%-20%. If the turbine guide vanes are made of composite materials, the opening of needle valve 8 is 0%-20%. Through the learning module 201, the opening of needle valve 8 is gradually optimized to the most reasonable ratio.

[0081] In the description of this application, it should be understood that the terms "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", etc., indicating the orientation or position relationship, are based on the orientation or position relationship shown in the accompanying drawings, and are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on this application.

[0082] The above describes specific embodiments of the present invention. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art may make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. The embodiments of this application and the features in the embodiments may be combined with each other in any manner unless there is a conflict.

Claims

1. An adaptive high-pressure turbine cooling system for an aircraft engine, characterized in that: include: High-pressure turbine (10); A high-pressure turbine cooling structure for constructing a cooling airflow path for guide blades of the high-pressure turbine (10); The acquisition control system is used for adaptive control of the guide blade cooling air flow rate.

2. The adaptive high-pressure turbine cooling system for an aircraft engine according to claim 1, characterized in that: The high-pressure turbine cooling structure comprises: a needle valve (8) with adjustable opening; The cooling air enters the guide vanes through the needle valve (8); An acquisition and control system comprises: a communication interface (203), a cooling model (202), a control module (207) and a learning module (201); The communication interface (203) is used to obtain real-time parameters of the aircraft's throttle (204), flight altitude (205) and aircraft pitch angle (206); The cooling model (202) receives the real-time parameters transmitted by the communication interface (203), combines the real-time parameters with the previous simulation model and the aircraft engine test results, thereby determining the opening of the needle valve (8), and issues a command to the control module (207); The control module (207) adjusts the opening of the needle valve (8) according to the instruction issued by the cooling model (202), thereby adjusting the cooling air flow to the guide vanes; The learning module (201) uses a neural network model to evaluate the operating results of the cooling model (202) and corrects the parameters of the cooling model (202) based on each flight and test result.

3. The adaptive high-pressure turbine cooling system for an aircraft engine according to claim 2, characterized in that: A high-pressure turbine cooling structure comprises: a compressor (1), a diffuser (2), an outer wall of a combustion chamber (3), an inner wall of a combustion chamber (12), a flame tube (11) and an air bleed pipe (5); The cooling air is drawn out from the compressor (1), passes through the diffuser (2), and is then divided into a path A and a path B; The cooling air of the A-path passes through the channel between the flame tube (11) and the inner wall (12) of the combustion chamber and enters the front cavity (301) of the guide vane. The flow rate of the cooling air of the A-path is a fixed value. The B-path cooling air passes through the passage between the flame tube (11) and the outer wall (3) of the combustion chamber, enters the bleed pipe (5) through the hole on the outer wall (3) of the combustion chamber, enters the needle valve (8) through the bleed pipe (5), and then enters the rear cavity (302) of the guide vane. The flow rate of the B-path cooling air is adjusted by the opening of the needle valve (8); The pressures in the front cavity (301) and the rear cavity (302) of the guide vane are both greater than the main flow pressure before the turbine, so that the cooling gases in the A path and the B path do not interfere with each other during the flow process.

4. The adaptive high-pressure turbine cooling system for an aircraft engine according to claim 3, characterized in that: The rear end side wall of the combustion chamber outer wall (3) has a plurality of air inlet holes distributed along the circumferential direction, and the number of the air inlet holes is the same as the number of the needle valves (8); One end of the plurality of air bleed pipes (5) is fixedly connected to the outer wall (3) of the combustion chamber and corresponds to the air bleed holes on the outer wall (3) of the combustion chamber, and the other end of the air bleed pipe (5) is tightly connected to the needle valve (8).

5. The adaptive high-pressure turbine cooling system for an aircraft engine according to claim 1, characterized in that: The high-pressure turbine cooling structure further comprises: a stepping motor (6) and a bracket (7); A combustion chamber outer casing (4) is provided on the outer side of the combustion chamber outer wall (3); The stepping motor (6) and the bracket (7) are tightly connected, and the bracket (7) is tightly connected to the outside of the outer casing (4) of the combustion chamber.

6. The adaptive high-pressure turbine cooling system for an aircraft engine according to claim 2, characterized in that: The needle valve (8) comprises a needle (104), a valve core (101) and a housing (102); The needle valve (8) is fixed on the outer casing (4) of the combustion chamber, and the needle valve (8) is arranged in the outer duct between the outer wall (3) of the combustion chamber and the outer casing (4) of the combustion chamber; The shaft (103) of the stepping motor (6) and the needle (106) of the needle valve (8) are tightly connected, and the stepping motor (6) can drive the shaft (104) to drive the needle (106) to move forward and backward, thereby adjusting the opening of the needle valve (8); The control module (207) controls the moving length of the shaft (103) by sending instructions to the stepping motor (6), thereby steplessly adjusting the opening of the needle valve (8) between 0-100%. The needle valve (8) is sealed by two polyimide sealing rings. The distance between the first sealing ring and the joint of the needle valve (8) and the outer casing (4) of the combustion chamber is 2 to 3 mm, and the spacing between the two sealing rings is 2 to 3 mm.

7. The adaptive high-pressure turbine cooling system for an aircraft engine according to claim 1, characterized in that: The high-pressure turbine cooling structure further includes: a baffle (9); The baffle (9) is between the high-pressure turbine (10) and the combustion chamber and is used to prevent the cooling gas from directly entering the high-pressure turbine (10).

8. The adaptive high-pressure turbine cooling system for an aircraft engine according to claim 1, characterized in that: The needle valve (8) is connected to the outer casing of the high-pressure turbine (10); The front end of the outer casing is butted against the rear end of the outer wall (3) of the combustion chamber, the inner ring of the guide of the high-pressure turbine (10) is connected to the rear end of the inner wall (12) of the combustion chamber, and one end of a plurality of guide blades is mounted on the inner ring of the guide, and the other end is mounted on the outer casing; The inner ring of the guide is provided with an air inlet hole communicating with the front cavity (301), and the outer casing is provided with an air inlet hole communicating with the rear cavity (302).

9. The adaptive high-pressure turbine cooling system for an aircraft engine according to claim 3, characterized in that: The front cavity (301) of the guide blade and the rear cavity (302) of the guide blade are separated from each other; the front cavity (301) is close to the leading edge of the guide blade, and the rear cavity (302) is close to the trailing edge of the guide blade.

10. A control method for the adaptive high-pressure turbine cooling system for an aircraft engine according to claim 1, characterized in that: When the throttle (204) is greater than a preset throttle threshold, the flight altitude (205) is less than or equal to a preset flight altitude threshold, and the aircraft pitch angle (206) is higher than a preset flight pitch angle threshold parameter, it is determined that the aircraft is in a take-off condition, the opening of the needle valve (8) is adjusted to 50%-70%, and the parameters of the cooling model (202) are corrected using the learning module (201); When the throttle (204) is at its maximum, the flight altitude (205) is higher than the flight altitude threshold, and the aircraft pitch angle (206) is flat, it is determined that the aircraft is in an accelerating state, and the opening of the needle valve (8) is adjusted to 100%; When the throttle (204) is greater than the throttle threshold, the flight altitude (205) is higher than the flight altitude threshold, and the aircraft pitch angle (206) is flat, it is determined that the aircraft is in supersonic cruise mode, the opening of the needle valve (8) is adjusted to 70%-100%, and the parameters of the cooling model (202) are corrected through the learning module (201); When the throttle (204) is less than or equal to the throttle threshold, the flight altitude (205) is higher than the flight altitude threshold, and the aircraft pitch angle (206) is flat, it is determined that the aircraft is in subsonic cruise condition, the material of the high-pressure turbine (10) is metal or composite material, for the high-pressure turbine (10) guide blades made of metal material, the opening of the needle valve (8) is adjusted to 10%-20%, and for the high-pressure turbine (10) guide blades made of composite material, the opening of the needle valve (8) is adjusted to 0%-20%, and the parameters of the cooling model (202) are corrected through the learning module (201).

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