Turbine blade cooling effect test system and method

Abstract

The present invention discloses a turbine blade cooling effect test system and method. The turbine blade cooling effect test system includes a static turbine blade cooling test device, an aircraft engine, a cooling air source device, a test cooling water device, a control system and a data acquisition system. The static turbine blade cooling test device is used to test turbine static blades and static turbine moving blades. The gas generator includes a combustion chamber and a compressor. The cooling air source device is used to control the cooling air flow and pressure entering the turbine blade cooling test device. The test cooling water device is used to cool the exhaust duct shell of the inlet and outlet of the static turbine blade cooling test device. The control system is used to control the output of mainstream fuel gas, cooling air and cooling water. The data acquisition system is used to collect the operating parameters of the test system and send them to the control system. The turbine blade cooling effect test system provided by the present invention has the advantages of low test cost, no location restriction on the test device, and flexible activation.

Description

Technical Field

[0001] The present invention relates to the technical field of gas turbines, and in particular to a system and method for testing the cooling effect of turbine blades. Background Art

[0002] In order to improve market competitiveness, gas turbines need to continuously improve the output and efficiency of the units, resulting in a continuous increase in the inlet temperature of the turbine blades. In order to ensure that the turbine blades operate safely, stably and for a long life at higher temperatures, in addition to using more advanced materials, it is more important to introduce efficient turbine blade cooling design technology. Efficient turbine blade cooling design technology is becoming increasingly complex. Simply relying on design experience or numerical calculations cannot meet the accuracy requirements of blade cooling design and heat transfer analysis. Whether it is a newly developed turbine blade or a turbine blade after design optimization and modification, a turbine blade cooling effect test must be carried out and can only be finalized after passing the test verification. However, the gas turbine turbine blade cooling effect test device in the related technology has been unable to meet the needs of efficient turbine blade cooling testing as the inlet temperature of the turbine blades continues to increase. Summary of the invention

[0003] The present invention aims to solve one of the technical problems in the related art at least to a certain extent.

[0004] To this end, an embodiment of the present invention provides a turbine blade cooling effect test system, which has the advantages of low test cost, no location restriction on the test device, and flexible activation.

[0005] According to the turbine blade cooling effect test system of the embodiment of the present invention, the turbine blade cooling effect test system includes a static turbine blade cooling test device, an aircraft engine, a cooling air source device, a test cooling water device, a control system and a data acquisition system. The static turbine blade cooling test device is used to test turbine static blades and static turbine moving blades. The gas generator of the aircraft engine includes a combustion chamber for providing mainstream gas to the static turbine blade cooling test device and a compressor for providing cooling air to the static turbine blade cooling test device. The cooling air source device is connected to the static turbine blade cooling test device. The cooling air source device is used to control the cooling air flow and pressure entering the turbine blade cooling test device. The test cooling water device is connected to the static turbine blade cooling test device. The test cooling water device is used to cool the exhaust pipe shell of the inlet and outlet of the static turbine blade cooling test device. The control system is respectively connected to the static turbine blade cooling effect test device, the cooling air source device and the test cooling water device. The control system is used to control the output of mainstream gas, cooling air and cooling water. The data acquisition system is electrically connected to the control system. The data acquisition system is used to collect the operating parameters of the test system and send them to the control system.

[0006] The turbine blade cooling effect test system according to the embodiment of the present invention has the advantages of low test cost, no location restriction of the test device, and flexible activation.

[0007] In some embodiments, the static turbine blade cooling test device includes a mainstream gas intake valve, a mainstream gas regulating valve, a turbine blade test section, a spray section and a silencer tower, the mainstream gas intake valve is connected to the gas generator outlet exhaust pipe, the turbine test section, the spray section and the silencer tower are connected in sequence, the spray section is used to quickly cool down the mainstream gas, and the silencer tower is used to eliminate noise.

[0008] In some embodiments, the turbine blade test section includes 3-7 full-size turbine blades, and the multiple full-size turbine blades are distributed in a fan shape. The two outermost full-size turbine blades are supporting blades, and the multiple full-size turbine blades form 2-6 turbine blade flow channels.

[0009] In some embodiments, a safety valve is provided at the inlet of the static turbine blade cooling test device, and the safety valve is used to test the overpressure protection of the turbine blade cooling test device.

[0010] In some embodiments, the static turbine blade cooling test device further includes a sliding section, which is disposed on the exhaust duct between the spray section and the silencer tower, and the sliding section is used to adjust the displacement of the exhaust duct.

[0011] In some embodiments, a back pressure valve is provided between the spray section and the sliding section, and the back pressure valve is used to adjust the back pressure of the turbine blade test section.

[0012] In some embodiments, the cooling air source device includes a cooling air intake valve and a cooling air regulating valve, the cooling air intake valve is connected to the exhaust cylinder of the compressor through an exhaust pipe, the cooling air intake valve and the cooling air regulating valve are connected to the turbine blade test section through a pipe, and the cooling air regulating valve is connected to the turbine blade test section, and the cooling air intake valve and the cooling air regulating valve are used to control the cooling air flow and pressure entering the turbine blade test section.

[0013] In some embodiments, a flow measuring point, a pressure measuring point and a temperature measuring point are provided on the pipeline between the cooling air regulating valve and the turbine blade test section.

[0014] In some embodiments, the test cooling water device includes a cooling tower, a water pump, a water inlet pipe, a water make-up pipe and a water outlet pipe. The cooling tower is connected to the double-layer pipe shell at the outlet of the gas generator and the inlet of the spray section through the water inlet pipe. The cooling tower is connected to the exhaust pipe shell at the inlet and outlet of the test turbine section through the water outlet pipe, and the water make-up pipe is connected to the cooling tower.

[0015] According to the turbine blade cooling effect test method of the embodiment of the present invention, the turbine blade cooling effect test method includes the following steps: extracting the outlet gas of the aircraft engine gas generator as the mainstream gas, and extracting the outlet air of the aircraft engine compressor as the cooling air; passing the mainstream gas and the cooling air into the mainstream channel of the turbine blade test section and the internal cooling channel of the turbine blade respectively, to conduct a turbine blade cooling effect test; adjusting the opening of the mainstream gas inlet valve and the mainstream gas regulating valve to control the mainstream gas flow, and adjusting the back pressure valve to adjust the back pressure of the turbine blade test section; collecting the outlet airflow of the turbine blade test section, quickly cooling the outlet airflow and discharging it into the silencer tower. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] Figure 1 1 is a schematic diagram of the structure of a turbine blade cooling effect test system according to an embodiment of the present invention.

[0017] Figure 2 Schematic diagram of a turbine blade flow passage of a turbine blade test section of a turbine blade cooling effect test system according to an embodiment of the present invention.

[0018] Figure numerals: 1. aircraft engine; 2. compressor; 3. combustion chamber; 4. gas turbine; 5. power turbine; 6. gas generator; 7. static turbine blade cooling test device; 8. cooling air source device; 9. test cooling water device; 10. control system; 11. data acquisition system; 12. mainstream gas inlet valve; 13. mainstream gas regulating valve; 14. bellows compensator; 15. turbine blade test section; 16. spray section; 17. back pressure valve; 18. sliding section; 19. silencer tower; 20. cooling air inlet valve; 21. cooling air regulating valve; 22. water pump; 23. water supply pipe; 24. water inlet pipe; 25. cooling tower; 26. water outlet pipe; 27. flow measuring point; 28. pressure measuring point; 29. ​​temperature measuring point; 30. hydraulic dynamometer; 31. full-size turbine blade. DETAILED DESCRIPTION

[0019] Embodiments of the present invention are described in detail below, and examples of the embodiments are shown in the accompanying drawings. The embodiments described below with reference to the accompanying drawings are exemplary and intended to be used to explain the present invention, but should not be understood as limiting the present invention.

[0020] According to the turbine blade cooling effect test system of the embodiment of the present invention, the turbine blade cooling effect test system includes a static turbine blade cooling test device 7, an aircraft engine 1, a cooling air source device 8, a test cooling water device 9, a control system 10 and a data acquisition system 11. The static turbine blade cooling test device 7 is used to test turbine static blades and static turbine moving blades. The gas generator 6 of the aircraft engine 1 includes a combustion chamber 3 for providing mainstream fuel gas to the static turbine blade cooling test device 7 and a compressor 2 for providing cooling air to the static turbine blade cooling test device 7. The cooling air source device 8 is connected to the static turbine blade cooling test device 7. The cooling air source device 8 is used to control the cooling air flow and pressure entering the turbine blade cooling test device. The test cooling water device 9 is connected to the static turbine blade cooling test device 7. The test cooling water device 9 is used to cool the exhaust pipe shell of the inlet and outlet of the static turbine blade cooling test device 7. The control system 10 is respectively connected to the static turbine blade cooling effect test device, the cooling air source device 8 and the test cooling water device 9. The control system 10 is used to control the output of the mainstream gas, cooling air and cooling water. The data acquisition system 11 is electrically connected to the control system 10. The data acquisition system 11 is used to collect the operating parameters of the test system and send them to the control system 10. The present invention uses an aircraft engine 1 to supply air to the static turbine blade cooling test device 7. The aircraft engine 1 uses a retired aircraft engine 1. The retired aircraft engine 1 includes a gas generator 6 and a power turbine 5. The gas generator 6 includes a compressor 2, a combustion chamber 3 and a gas turbine 4 connected in sequence. The output power of the retired aircraft engine 1 is consumed by a hydraulic dynamometer 30 located after the power turbine 5, and the exhaust gas is discharged to the atmosphere after being cooled and noise-reduced. Compared with a specially established working condition static turbine blade cooling effect test device, the static turbine blade cooling test device 7 of the present invention can reduce the air compressor and air electric heater for providing high-pressure mainstream air, the combustion chamber for providing high-temperature fuel gas, and the blower, filter and air heater for providing cooling air, thereby saving the equipment purchase cost of the test device and the power consumption cost of the corresponding equipment. The static turbine blade cooling test device 7 using the aircraft engine 1 for air supply is less limited by the spatial position, does not require power grid dispatching and unit transformation, and makes the test start-up flexible and time-saving.

[0021] The turbine blade cooling effect test system according to the embodiment of the present invention has the advantages of low test cost, no location restriction of the test device, and flexible activation.

[0022] In some embodiments, the static turbine blade cooling test device 7 includes a mainstream gas intake valve 12, a mainstream gas regulating valve 13, a turbine blade test section 15, a spray section 16, a sliding section 18 and a silencer tower 19. The mainstream gas intake valve 12 is connected to the outlet exhaust pipe of the gas generator 6, and the turbine test section, the spray section 16 and the silencer tower 19 are connected in sequence. The spray section 16 is used to quickly cool down the mainstream gas, and the silencer tower 19 is used to eliminate noise.

[0023] Specifically, the mainstream gas intake valve 12 extracts the gas in the rear exhaust cylinder of the gas generator 6 of the aircraft engine 1 as the mainstream gas, and the mainstream gas regulating valve 13 is arranged on the exhaust pipe. The mainstream gas intake flow and pressure are controlled by the opening of the mainstream gas intake valve 12 and the mainstream gas regulating valve 13. The exhaust gas of the spray section 16 enters the silencer tower 19 after passing through the exhaust butterfly valve to reduce noise and then is discharged into the atmosphere. A metal expansion joint can be arranged on the mainstream intake pipe.

[0024] In some embodiments, the turbine blade test section 15 includes 3-7 full-size turbine blades 31, and the multiple full-size turbine blades 31 are distributed in a fan shape. The two outermost full-size turbine blades 31 are supporting blades, and the multiple full-size turbine blades 31 form 2-6 turbine blade flow channels.

[0025] Specifically, the turbine blade test section 15 is connected to the cooling air source device 8, and the cooling air enters the internal cooling channel of the turbine blade from the top of the stationary blade or the root of the moving blade. The shell of the turbine blade test section 15 is a double-layer shell, and heat insulation materials are installed between the double-layer shells. The corrugated compensator 14 is installed on the mainstream air intake duct to absorb the expansion of the air intake duct. The air intake duct and exhaust duct before and after the turbine blade test section 15 are designed as double-layer pipes and connected to the test cooling water device 9.

[0026] In some embodiments, a safety valve is provided at the inlet of the static turbine blade cooling test device 7, and the safety valve is used to test the overpressure protection of the turbine blade cooling test device.

[0027] Therefore, the safety valve improves the safety of the static turbine blade cooling test device and prevents the pressure from exceeding the limit range and endangering the safety of the equipment.

[0028] In some embodiments, the static turbine blade cooling test device 7 further includes a sliding section 18, which is disposed on the exhaust duct between the spray section 16 and the silencer tower 19, and is used to adjust the displacement of the exhaust duct.

[0029] Specifically, during the disassembly and assembly process of the turbine blade test section 15, the flange compression makes it difficult to disassemble and assemble the turbine blade test section 15. The sliding section 18 on the exhaust pipe is connected to the pipe and the like by a flange as a whole, and a sliding platform for supporting the sliding section 18 is provided below the sliding section 18. The sliding section 18 can move axially in the sliding bushing to facilitate the axial movement of the test section when it is disassembled and assembled. When the exhaust pipe is thermally deformed due to heating, the sliding section 18 can move axially in the sliding bushing to compensate for the displacement of the exhaust pipe.

[0030] In some embodiments, a back pressure valve 17 is provided between the spray section 16 and the sliding section 18, and the back pressure valve 17 is used to adjust the back pressure of the fan-shaped blade grid formed by multiple turbine blades in the turbine blade test section 15. The rated pressure of the back pressure valve 17, i.e., the set pressure, is adjustable. When the system pressure is lower than the set pressure, the diaphragm inside the back pressure valve 17 blocks the pipeline under the action of the spring force, thereby increasing the system pressure; when the system pressure is higher than the set pressure, the diaphragm inside the back pressure valve 17 compresses the spring, and the pipeline is connected, thereby reducing the system pressure.

[0031] Therefore, adjusting the back pressure of the turbine blade test section by adjusting the opening of the back pressure valve to achieve pressure control is helpful to achieve pressure regulation and stabilization of the turbine blade cooling effect test system.

[0032] In some embodiments, the cooling air source device 8 includes a cooling air intake valve 20 and a cooling air regulating valve 21. The cooling air intake valve 20 is connected to the exhaust cylinder of the compressor 2 through an exhaust pipe. The cooling air intake valve 20 and the cooling air regulating valve 21 are connected through a pipe. The cooling air regulating valve 21 is connected to the turbine blade test section 15. The cooling air intake valve 20 and the cooling air regulating valve 21 are used to control the cooling air flow and pressure entering the turbine blade test section 15.

[0033] Specifically, the cooling air inlet valve 20 is connected to the exhaust cylinder of the compressor 2 of the aircraft engine 1, and the cooling air regulating valve 21 is connected to the turbine blade test section 15 through a pipeline. The cooling air enters the internal cooling channel of the turbine blade from the top of the stationary blade or the root of the moving blade. The air flow rate at the outlet of the compressor 2 of the aircraft engine 1 extracted by the cooling air source device 8 accounts for 0.52% to 3.54% of the inlet flow rate of the compressor 2.

[0034] In some embodiments, a flow measuring point 27 , a pressure measuring point 28 and a temperature measuring point 29 are provided on the pipeline between the cooling air regulating valve 21 and the turbine blade test section 15 .

[0035] Specifically, flow collection devices, pressure collection devices and temperature collection devices are set at the flow measuring point 27, the pressure measuring point 28 and the temperature measuring point 29. The temperature collection device collects the temperature of the cooling air, the flow collection device collects the flow of the cooling air, and the pressure collection device collects the pressure of the cooling air.

[0036] In some embodiments, the test cooling water device 9 includes a cooling tower 25, a water pump 22, a water inlet pipe 24, a water supply pipe 23 and a water outlet pipe 26. The cooling tower 25 is connected to the double-layer pipe shell of the outlet of the gas generator 6 and the inlet of the spray section 16 through the water inlet pipe 24, the cooling tower 25 is connected to the exhaust pipe shell of the inlet and outlet of the test turbine section through the water outlet pipe 26, and the water supply pipe 23 is connected to the cooling tower 25. The water pump 22 is arranged on the water inlet pipe 24.

[0037] Specifically, the test cold water device is used to cool the inlet and outlet of the turbine blade test section 15. The cooling water of the cooling tower 25 is sprayed into the spray section 16 to reduce the exhaust temperature after the turbine test section to below 100°C, ensuring that the downstream back pressure valve 17 and the silencer tower 19 do not exceed the use temperature; the pipe section before the spray section 16 adopts a high-temperature alloy double-layer water jacket structure, and the pipe material after the water spray is 304 stainless steel.

[0038] In some embodiments, the data acquisition system 11 includes a pressure acquisition device, a temperature acquisition device and a flow acquisition device. The pressure acquisition device acquires the mainstream gas inlet pipeline pressure, the outlet pressure of the turbine component, and the pressure of the cooling air. The temperature acquisition device acquires the temperature of the mainstream gas and the cooling air. The flow acquisition device acquires the flow of the mainstream gas and the cooling air.

[0039] Specifically, the pressure acquisition device includes a pressure scanning valve and a pressure transmitter for acquisition, and the temperature acquisition device includes a thermocouple and a temperature indicating paint for acquisition; the flow acquisition device uses a flow nozzle and a mass flow meter for acquisition; the data acquisition system 11 acquires data such as the mainstream gas inlet temperature, the mainstream gas inlet total pressure, the test section outlet static pressure, the mainstream air flow, the fuel temperature and pressure of the combustion chamber 3, and the cooling air flow, temperature and pressure.

[0040] According to the turbine blade cooling effect test method of the embodiment of the present invention, the turbine blade cooling effect test method includes the following steps: extracting the outlet gas of the gas generator 6 of the aircraft engine 1 as the mainstream gas, and extracting the outlet air of the compressor 2 of the aircraft engine 1 as the cooling air; respectively passing the mainstream gas and the cooling air into the mainstream channel of the turbine blade test section 15 and the internal cooling channel of the turbine blade to conduct a turbine blade cooling effect test; adjusting the opening of the mainstream gas inlet valve 12 and the mainstream gas regulating valve 13 to control the mainstream gas flow, and adjusting the back pressure valve 17 to adjust the back pressure of the turbine blade test section 15; collecting the outlet airflow of the turbine blade test section 15, quickly cooling the outlet airflow and discharging it into the silencer tower 19.

[0041] An exhaust port is installed at the outlet of the gas generator 6 of the retired aircraft engine 1 to extract high-temperature and high-pressure gas as the mainstream gas for the turbine cooling effect test. An exhaust port is installed at the outlet of the compressor 2 of the retired aircraft engine 1 to extract air as the cooling air for the turbine cooling effect test. The mainstream intake flow rate is controlled by adjusting the opening of the mainstream gas intake valve 12 and the mainstream gas regulating valve 13, and the back pressure behind the blade grid of the turbine blade test piece is adjusted by adjusting the opening of the back pressure valve 17 to achieve pressure control. The flow and pressure of the cooling air are controlled by adjusting the opening of the cooling air intake valve 20 and the cooling air regulating valve 21.

[0042] like Figure 2 As shown, when there are 7 full-size turbine blades 31 in the turbine blade test section 15, for a certain type of 300MW F-class gas turbine, the inlet air flow rate of compressor 2 is 730kg / s, and the inlet air flow rate of compressor 2 of a certain type of retired aircraft engine 1 is 138kg / s. The calculation results of the air flow rate of the cooling effect test of the first-stage stationary blades and the first-stage moving blades with 2 turbine blade flow passages formed by 3 full-size turbine blades 31 and 6 turbine blade flow passages formed by 7 full-size turbine blades 31 are listed in Table 1. The air flow rate of compressor 2 extracted in the turbine blade cooling effect test accounts for approximately 0.77% to 3.06% of the inlet flow rate of compressor 2 of the retired aircraft engine 1, which will not affect the safe operation of the retired aircraft engine 1.

[0043] [Table 1]

[0044]

[0045] The present invention provides a turbine blade cooling effect test system and method for retired aircraft engine air supply. Compared with an independent full-pressure, full-temperature, full-size turbine blade cooling effect test bench, the system can reduce the number of air compressors and air electric heaters for providing high-pressure mainstream air, combustion chambers for providing high-temperature fuel gas, and blowers, filters, air heaters and other equipment for providing cooling air.

[0046] The present invention saves 55-95 million yuan in equipment purchase costs. The power consumption of these equipments is 26,000-32,000 kW. A turbine blade cooling effect test is conducted under 6 working conditions, each working condition is calculated as 1 hour, and the test power consumption is 156,000 kWh-192,000 kWh. If industrial electricity consumption is calculated at 0.9 yuan / kWh, the method provided by the present invention can save 140,400-172,800 yuan.

[0047] In the description of the present invention, it is to be understood that the terms “center”, “longitudinal”, “lateral”, “length”, “width”, “thickness”, “up”, “down”, “front”, “back”, “left”, “right”, “vertical”, “horizontal”, “top”, “bottom”, “inside”, “outside”, “clockwise”, “counterclockwise”, “axial”, “radial”, “circumferential”, etc., indicating orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as limiting the present invention.

[0048] In addition, the terms "first" and "second" are used for descriptive purposes only and should not be understood as indicating or implying relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first" and "second" may explicitly or implicitly include at least one of the features. In the description of the present invention, the meaning of "plurality" is at least two, such as two, three, etc., unless otherwise clearly and specifically defined.

[0049] In the present invention, unless otherwise clearly specified and limited, the terms "installed", "connected", "connected", "fixed" and the like should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral one; it can be a mechanical connection, an electrical connection, or communication with each other; it can be a direct connection, or an indirect connection through an intermediate medium, it can be the internal connection of two elements or the interaction relationship between two elements, unless otherwise clearly defined. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0050] In the present invention, unless otherwise clearly specified and limited, a first feature being "above" or "below" a second feature may mean that the first and second features are in direct contact, or the first and second features are in indirect contact through an intermediate medium. Moreover, a first feature being "above", "above" or "above" a second feature may mean that the first feature is directly above or obliquely above the second feature, or simply means that the first feature is higher in level than the second feature. A first feature being "below", "below" or "below" a second feature may mean that the first feature is directly below or obliquely below the second feature, or simply means that the first feature is lower in level than the second feature.

[0051] In the present invention, the terms "one embodiment", "some embodiments", "examples", "specific examples", or "some examples" etc. mean that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present invention. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the described specific features, structures, materials or characteristics may be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art may combine and combine the different embodiments or examples described in this specification and the features of the different embodiments or examples, without contradiction.

[0052] Although the embodiments of the present invention have been shown and described above, it is to be understood that the above embodiments are illustrative and are not to be construed as limitations on the present invention. Changes, modifications, substitutions and variations of the above embodiments by those of ordinary skill in the art are all within the scope of protection of the present invention.

Claims

1. A turbine blade cooling effect test system, characterized in that: include: A static turbine blade cooling test device, wherein the static turbine blade cooling test device is used to test turbine static blades and static turbine moving blades; An aircraft engine, wherein a gas generator of the aircraft engine comprises a combustion chamber for providing mainstream fuel gas to the static turbine blade cooling test device and a compressor for providing cooling air to the static turbine blade cooling test device; A cooling air source device, the cooling air source device is connected to the static turbine blade cooling test device, and the cooling air source device is used to control the flow rate and pressure of cooling air entering the turbine blade cooling test device; A test cooling water device, the test cooling water device is connected to the static turbine blade cooling test device, and the test cooling water device is used to cool the exhaust duct shell side of the inlet and outlet of the static turbine blade cooling test device; A control system, wherein the control system is respectively connected to the static turbine blade cooling effect test device, the cooling gas source device and the test cooling water device, and the control system is used to control the output of the mainstream fuel gas, the cooling air and the cooling water; and A data acquisition system, the data acquisition system is electrically connected to the control system, and the data acquisition system is used to collect operating parameters of the test system and send them to the control system; The static turbine blade cooling test device includes a mainstream gas intake valve, a mainstream gas regulating valve, a turbine blade test section, a spray section, a sliding section and a silencer tower. The mainstream gas intake valve is connected to the gas generator outlet exhaust pipe, the turbine blade test section, the spray section and the silencer tower are connected in sequence, the spray section is used to quickly cool down the mainstream gas, the silencer tower is used to eliminate noise, the sliding section is arranged on the exhaust pipe between the spray section and the silencer tower, and the sliding section is used to adjust the displacement of the exhaust pipe.

2. The turbine blade cooling effect test system according to claim 1, characterized in that: The turbine blade test section includes 3-7 full-size turbine blades, and the multiple full-size turbine blades are distributed in a fan shape. The two outermost full-size turbine blades are supporting blades, and the multiple full-size turbine blades form 2-6 turbine blade flow channels.

3. The turbine blade cooling effect test system according to claim 1, characterized in that: A safety valve is arranged at the inlet of the static turbine blade cooling test device, and the safety valve is used to test the overpressure protection of the turbine blade cooling test device.

4. The turbine blade cooling effect test system according to claim 1, characterized in that: A back pressure valve is arranged between the spray section and the sliding section, and the back pressure valve is used to adjust the back pressure of the turbine blade test section.

5. The turbine blade cooling effect test system according to claim 1, characterized in that: The cooling air source device includes a cooling air intake valve and a cooling air regulating valve. The cooling air intake valve is connected to the exhaust cylinder of the compressor through an exhaust pipe. The cooling air intake valve and the cooling air regulating valve are connected to the turbine blade test section through a pipe. The cooling air intake valve and the cooling air regulating valve are used to control the cooling air flow and pressure entering the turbine blade test section.

6. The turbine blade cooling effect test system according to claim 5, characterized in that: A flow measuring point, a pressure measuring point and a temperature measuring point are arranged on the pipeline between the cooling air regulating valve and the turbine blade test section.

7. The turbine blade cooling effect test system according to claim 1, characterized in that: The test cooling water device includes a cooling tower, a water pump, a water inlet pipe, a water supply pipe and a water outlet pipe. The cooling tower is connected to the double-layer pipe shell at the outlet of the gas generator and the inlet of the spray section through the water inlet pipe. The cooling tower is connected to the exhaust pipe shell at the inlet and outlet of the turbine blade test section through the water outlet pipe, and the water supply pipe is connected to the cooling tower.

8. A method for testing the cooling effect of turbine blades, characterized in that: Using the turbine blade cooling effect test system according to any one of claims 1 to 7, the test method comprises the following steps: Extract the gas from the outlet of the aircraft engine gas generator as the mainstream gas, and extract the air from the outlet of the aircraft engine compressor as the cooling air; The mainstream fuel gas and cooling air are respectively introduced into the mainstream passage of the turbine blade test section and the internal cooling passage of the turbine blade to conduct a turbine blade cooling effect test; adjusting the opening of the mainstream gas inlet valve and the mainstream gas regulating valve to control the mainstream gas flow, and adjusting the back pressure valve to adjust the back pressure of the turbine blade test section; The outlet airflow of the turbine blade test section is collected, the outlet airflow is rapidly cooled and discharged into a silencer tower.

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