Aero-engine turbine blade high-altitude state cooling test device and test method thereof

Abstract

The invention discloses an aero-engine turbine blade high-altitude state cooling test device which comprises a main gas pipeline used for being communicated with a gas source station, a main gas flow heating device is arranged on the main gas pipeline and communicated with a main test cavity, and the main test cavity is communicated with a main exhaust end. The main exhaust end is provided with a high-pressure air connecting pipe used for being connected with a high-pressure ejector, the main air pipeline is provided with a cold air taking opening, the cold air taking opening is connected with a cold air pipe through a flow adjusting valve, and the cold air pipe is communicated with the main test cavity. The test method comprises the following steps: S1, designing a test section, and installing a test piece; s2, air supply and drainage; s3, temperature control and pressure regulation; s4, detecting and measuring; s5, measuring the cold air temperature Tc; and S6, calculating the blade cooling effect eta. The turbine blade cooling effect test can be carried out under the high-altitude low Reynolds number, the real working condition of the turbine blade is simulated more truly, and a basis can be provided for design of a blade cooling structure.

Description

Technical Field

[0001] This invention relates to the field of ground testing technology for aero-engines, specifically to a high-altitude cooling test device and test method for aero-engine turbine blades. Background Technology

[0002] Turbine inlet temperature is a crucial parameter for aero-engine performance, and increasing this temperature is significant for improving engine performance, particularly thrust-to-weight ratio. This requires turbine blades to possess excellent high-temperature resistance and cooling efficiency. With the continuous improvement of aero-engine thrust-to-weight ratio, the rate of increase in turbine inlet gas temperature far exceeds the rate of increase in material temperature resistance. Therefore, effective cooling technology has become critical for the safe operation of turbine blades in high-temperature environments.

[0003] Developed countries have adopted advanced cooling structures and materials to ensure the reliability of turbine blades. Under the same materials and operating conditions, the reliability of air-cooled turbine blades depends on their cooling effect and temperature distribution, which in turn are closely related to the cooling structure and cooling method.

[0004] Due to the complexity of the hydrodynamics and heat exchange phenomena within the air-cooled turbine blades of aero-engines, theoretical analysis alone is insufficient to accurately describe these phenomena. Therefore, conducting comprehensive experiments according to design procedures and model specifications to obtain information on blade cooling effects and wall temperature distribution is particularly important. This can verify and correct theoretical calculation results, evaluate the rationality of the cooling structure design, and thus guide subsequent designs.

[0005] However, there is currently almost no technical data available in China on experimental research into the cooling effect of real turbine blades under high-altitude, low-Reynolds-number conditions, leaving this field largely unexplored. Furthermore, existing technologies lack specific testing methods for cooling effects at high altitudes and low Reynolds numbers, and related testing techniques and experience are insufficient. Given the significant differences between the conditions of aero-engine turbines at high altitudes and low Reynolds numbers and those on the ground, it is urgent to conduct cooling effect tests on turbine blades under these conditions to achieve more in-depth research and precise design development. Summary of the Invention

[0006] The purpose of this invention is to address the aforementioned problems by providing a high-altitude cooling test device and method for aero-engine turbine blades, which is used to conduct turbine blade cooling effect tests at high altitudes and low Reynolds numbers, more realistically simulating the actual working conditions of turbine blades, and providing a reference for subsequent research on the cooling effect of turbine blades at high altitudes and low Reynolds numbers.

[0007] The technical solution of the present invention: A high-altitude cooling test device for aero-engine turbine blades includes a main gas pipeline for connecting to a gas source station. A main gas flow heating device is installed on the main gas pipeline, and the main gas flow heating device is connected to a main test chamber. The main test chamber is connected to a main exhaust end, and the main exhaust end is equipped with a high-pressure air connection pipe for connecting a high-pressure ejector. A cold air intake is installed on the main gas pipeline, and the cold air intake is connected to a cold air pipe through a flow regulating valve. The cold air pipe is connected to the main test chamber. A flow meter A, a cold air heater, and a temperature sensor are sequentially installed on the cold air pipe. The cold air heater is equipped with a temperature control device, and the temperature sensor is located at one end of the cold air heater's outlet.

[0008] The main airflow heating device is a single-tube combustion chamber, and the single-tube combustion chamber is equipped with a high-altitude igniter.

[0009] The main gas pipeline is connected to the main gas flow heating device via a flow control valve and a flow meter B. The main gas flow heating device is connected to the test inlet of the main test chamber. The main test chamber is connected to the high-pressure ejector via the main exhaust end. The high-pressure ejector is matched with a silencer tower.

[0010] The cold air heater is an electric furnace heater, and the temperature control device is a heating power controller.

[0011] A pressure stabilizing tank is connected to the air conditioning pipe, and the pressure stabilizing tank is located between the flow regulating valve and the flow meter A.

[0012] A method for high-altitude cooling test of aero-engine turbine blades includes the following steps: S1: Design the test section and install the test specimen; first, connect the gas source station, main gas pipeline, main airflow heating device, main test chamber, and main exhaust end in sequence; then connect the high-pressure air connection pipe outlet to the main exhaust end, and then connect the cold air pipe as a branch pipe to the main gas pipeline and the main test chamber, and connect the flow meter A, cold air heater, and temperature sensor from the air inlet to the air outlet to the cold air pipe in sequence; finally, connect the temperature control device to control the temperature of the cold air heater; wherein, the gas source station is used to supply gas, the main gas pipeline is used to connect pipelines, the main airflow heating device is a single-tube combustion chamber, and the single-tube combustion chamber is equipped with a high-altitude igniter; S2: Gas supply and drainage; the main gas pipeline supplies gas, and high-pressure air is dredged to create a negative pressure state required for the test. S3: Temperature and pressure control; ensure that the inlet gas pressure, flow rate, and inlet gas temperature Tg of the test section meet the test requirements; S4: Detection and measurement; Under the condition that step S3 is satisfied, measure the surface temperature Ts of the turbine blade; S5: Measures the air conditioning temperature Tc; S6: Calculate the blade cooling effect η according to the formula η=(Tg-Ts) / (Tg-Tc); Wherein, Tg is the inlet gas temperature of the test section, Ts is the surface temperature of the turbine blade, Tc is the temperature of the cooling gas, and η is the blade cooling effect.

[0013] The temperature and voltage control mentioned in step S3 includes the following steps: a. Ignite the single-tube combustion chamber and heat the air in the main gas pipeline to the required test section inlet gas temperature Tg by adjusting the fuel flow rate; b. Draw a line of cold air from the main gas duct and heat the cold air through a cold air heater; c. By adjusting the main gas pipeline supply pressure and high-pressure air pressure, ensure that the inlet gas pressure of the test section meets the test requirements; d. By adjusting the flow rate of the cold air and the heating power of the cold air heater, the flow rate of the cold air and the surface temperature Ts of the turbine blades are made to meet the test requirements.

[0014] In step a, the air flow rate is adjusted by the flow control valve and the air flow rate is measured by the flow meter B to meet the requirements of the main gas flow rate at the inlet of the test piece in each test state. The air-fuel mixing ratio is adjusted by the flow control valve to ensure that the gas temperature meets the requirements of the total inlet temperature of the test piece in each test state. The gas inlet pressure is made to meet the requirements of the total inlet pressure of the test piece by ejecting the gas through the high-pressure ejector. The flow control valve is used to control the pipeline flow rate, and the flow meter B is used to display the flow rate in real time.

[0015] In step b, the air diverted from the main gas pipeline is pressurized by the pressure stabilizing tank and then enters the cold air heater. The air flow rate is adjusted by the flow regulating valve. Cooling air with different temperatures, pressures and flow rates is provided to the test blades according to the test conditions to meet the requirements of the cooling air flow rate ratio for each test condition. The pressure stabilizing tank is used to stabilize the pressure of the rapidly flowing gas.

[0016] The beneficial effects of this invention are: By adopting this experimental method, turbine blade cooling effect tests can be carried out at high altitudes and low Reynolds numbers, which more realistically simulates the actual working conditions of turbine blades. This provides a basis for the design of blade cooling structures and a reference for subsequent experimental research on the cooling effect of turbine blades at high altitudes and low Reynolds numbers. Attached Figure Description

[0017] Figure 1 This is a schematic diagram of the experimental apparatus of the present invention; Figure 2 This is a schematic diagram of the intake heating principle of the present invention; Figure 3 This is a schematic diagram of the cold air heating principle of the present invention; Figure 4 This is a flowchart of the test results for the high-altitude cooling effect of the turbine blades in this invention.

[0018] Attached reference numerals: 1-Main gas pipeline, 2-Main gas flow heating device, 3-Main test chamber, 4-Main exhaust end, 5-High pressure air connection pipe, 6-Cold air intake, 7-Flow regulating valve, 8-Cold air pipe, 9-Flow meter A, 10-Cold air heater, 11-Temperature sensor, 12-Temperature control device, 13-Pressure stabilizing tank, 14-Flow control valve, 15-Flow meter B. Detailed Implementation

[0019] refer to Figures 1 to 4 A high-altitude cooling test device for aero-engine turbine blades includes a main air pipeline 1 for connecting to an air source station, a main airflow heating device 2 installed on the main air pipeline 1, a main test chamber 3 connected to the main test chamber 3 and a main exhaust end 4, a high-pressure air connection pipe 5 for connecting to a high-pressure ejector installed on the main exhaust end 4, a cold air intake port 6 installed on the main air pipeline 1, a cold air pipe 8 connected to the cold air intake port 6 via a flow regulating valve 7, the cold air pipe 8 connected to the main test chamber 3, and a flow meter A9, a cold air heater 10 and a temperature sensor 11 installed sequentially on the cold air pipe 8, a temperature control device 12 installed on the cold air heater 10, and a temperature sensor 11 installed at one end of the outlet of the cold air heater 10.

[0020] The main airflow heating device 2 is a single-tube combustion chamber, and the single-tube combustion chamber is equipped with a high-altitude igniter.

[0021] The main gas pipeline 1 is connected to the main gas flow heating device 2 via the flow control valve 14 and the flow meter B15. The main gas flow heating device 2 is connected to the test inlet of the main test chamber 3. The main test chamber 3 is connected to the high-pressure ejector via the main exhaust end 4. The high-pressure ejector is matched with a silencer tower.

[0022] The air heater 10 is an electric furnace heater, and the temperature control device 12 is a heating power controller.

[0023] A pressure stabilizing tank 13 is connected to the air conditioning pipe 8. The pressure stabilizing tank 13 is located between the flow regulating valve 7 and the flow meter A9.

[0024] This application utilizes experimental equipment to conduct tests on the high-altitude cooling effect of aero-engine turbine blades. Clean air supplied by a gas source station enters the main gas flow heating device 2 via the main gas pipeline 1 for heating. A single-tube combustion chamber is used as the main gas flow heating device 2, equipped with a high-altitude igniter. The air supplied by the gas source station is mixed with fuel and directly combusted to produce the high-temperature gas required for the test. At the main exhaust end 4, high-pressure air is used to divert the low-pressure gas in the main gas pipeline 1, reducing the gas pressure to meet the requirements of the inlet main gas flow rate, temperature, and pressure for each test condition. The air diverted from the main gas pipeline 1 enters the cold air heater 10, and the air flow rate is regulated by a flow control valve 7, providing cooling air of different temperatures, pressures, and flow rates to the test blades according to the test conditions.

[0025] A test method for a high-altitude cooling test device for aero-engine turbine blades includes the following steps: S1: Design the test section and install the test specimen; first, connect the gas source station, main gas pipeline, main airflow heating device, main test chamber, and main exhaust end in sequence; then connect the high-pressure air connection pipe outlet to the main exhaust end, and then connect the cold air pipe as a branch pipe to the main gas pipeline and the main test chamber, and connect the flow meter A, cold air heater, and temperature sensor from the air inlet to the air outlet to the cold air pipe in sequence; finally, connect the temperature control device to control the temperature of the cold air heater; wherein, the gas source station is used to supply gas, the main gas pipeline is used to connect pipelines, the main airflow heating device is a single-tube combustion chamber, and the single-tube combustion chamber is equipped with a high-altitude igniter; Furthermore, in this application, the test section is designed according to the blade structure; the design concept of the high-altitude test section and the ground-level test section is the same. The inlet and outlet angles of the test section are designed according to the inlet and outlet airflow angles of the test blade. The main function of the test section is to connect the front transition measurement section, fix the test blade and the supporting blade, ensure that the combustion gas can smoothly enter the blade, and provide cooling air to the test blade and the supporting blade respectively.

[0026] S2: Gas supply and diversion; the main gas pipeline supplies gas, and high-pressure air is diverted to create a negative pressure state required for the test; clean gas supplied by the gas source station enters the main gas pipeline, and high-pressure air is diverted behind the test section. By adjusting the flow rate and pressure of the high-pressure air, the inlet pressure of the test section is ensured to be the negative pressure required for the test.

[0027] S3: Temperature and pressure control; ensure that the inlet gas pressure, flow rate, and inlet gas temperature Tg of the test section meet the test requirements; S4: Detection and measurement; Under the condition that step S3 is satisfied, measure the surface temperature Ts of the turbine blade; that is, after ensuring that the inlet gas flow rate and temperature of the test section and the flow rate and temperature of the cold air meet the requirements, measure the surface temperature Ts of the turbine blade, thereby obtaining the inlet gas temperature Tg of the test section and the surface temperature Ts of the turbine blade.

[0028] S5: Measures the air conditioning temperature Tc; S6: Calculate the blade cooling effect η according to the formula η=(Tg-Ts) / (Tg-Tc); Wherein, Tg is the inlet gas temperature of the test section, Ts is the surface temperature of the turbine blade, Tc is the temperature of the cooling gas, and η is the blade cooling effect.

[0029] Regarding the aforementioned blade cooling effect η, this application provides the following embodiments for illustration: The test conditions were as follows: the inlet gas temperature of the test section was Tg = 1600 ℃, which was controlled and kept constant by the main airflow heating device 2; the cold air temperature was Tc = 350 ℃ (provided by the external cooling air system, and the flow meter ensured a constant flow rate); thermocouples were arranged at typical positions such as the leading edge, trailing edge, blade base, and blade back of the blade to collect the surface temperature Ts in real time.

[0030] Measurement and Calculation: After the experiment has stabilized for 5 minutes, record the results simultaneously. Taking Tg = 1600 ℃; Tc = 350 ℃; and Ts = 1050 ℃ at the front edge measuring point as an example... Substitute into the formula: η = (Tg − Ts) / (Tg − Tc)= (1 600 − 1050) / (1 600 − 350)= 550 / 1250= 0.44 η = 0.44 indicates that, under the existing cooling structure, the blade surface temperature is potentially "lowered" by 44% relative to the combustion gas temperature. If the same blade is replaced with a control component without film cooling holes and only with internal impact, and Ts is measured at 1120 ℃ under the same operating conditions, then η = (1600−1120) / (1600-350) = 0.384. The increase in η from 0.384 to 0.44 demonstrates that adding film cooling holes significantly improves the cooling effect.

[0031] In other words, η directly quantifies the performance of the cooling design: The closer η is to 1, the closer the blade surface temperature is to the air temperature, and the more efficient the cooling. The closer η is to 0, the less effective the cooling system is, and the cooling structure needs to be optimized or the amount of cooling air needs to be increased.

[0032] This example demonstrates that the formula η=(Tg-Ts) / (Tg-Tc) normalizes the "temperature difference" into a dimensionless index between 0 and 1, which is both objective and convenient for quantitative comparison between different schemes.

[0033] Step S3, temperature and voltage control, includes the following steps: a. Ignition of the single-tube combustion chamber. By adjusting the fuel flow rate, the air in the main gas pipeline is heated to the required inlet gas temperature Tg of the test section. In this application, clean gas supplied by the gas source station enters the main gas flow heating device via the main gas pipeline for heating. The main gas flow heating device mixes the air supplied by the gas source station with fuel and then directly combusts it into the high-temperature gas required for the test. At the main exhaust end, a high-pressure ejector is used to eject the low-pressure gas flow in the main gas pipeline, reducing the inlet gas pressure to negative pressure. Gas heating adopts a single-stage heating method. The main gas flow heating device uses a single-tube combustion chamber. The air flow rate is adjusted by a flow control valve, and the air flow rate is measured by a flow meter B to meet the main gas flow rate requirements at the inlet of the test specimen in each test state. A regulating valve is used to adjust the air-fuel mixing ratio to ensure the gas temperature meets the total inlet temperature requirements of the test specimen in each test state. A high-pressure ejector is used to ensure the gas inlet pressure meets the total inlet pressure requirements of the test specimen.

[0034] b. Draw a line of cold air from the main gas duct and heat the cold air through a cold air heater; c. By adjusting the main gas pipeline supply pressure and high-pressure air pressure, ensure that the inlet gas pressure of the test section meets the test requirements; d. By adjusting the flow rate of the cold air and the heating power of the cold air heater, the flow rate of the cold air and the surface temperature Ts of the turbine blades are made to meet the test requirements.

[0035] In step a, the air flow rate is adjusted by the flow control valve and the air flow rate is measured by the flow meter B to meet the requirements of the main gas flow rate at the inlet of the test piece in each test state. The air-fuel mixing ratio is adjusted by the flow control valve to ensure that the gas temperature meets the requirements of the total inlet temperature of the test piece in each test state. The gas inlet pressure is made to meet the requirements of the total inlet pressure of the test piece by ejecting the gas through the high-pressure ejector. The flow control valve is used to control the pipeline flow rate, and the flow meter B is used to display the flow rate in real time.

[0036] In step b, the air diverted from the main gas pipeline is pressurized by the pressure stabilizing tank and then enters the cold air heater. The air flow rate is adjusted by the flow regulating valve. Cooling air with different temperatures, pressures and flow rates is provided to the test blades according to the test conditions to meet the requirements of the cooling air flow rate ratio for each test condition. The pressure stabilizing tank is used to stabilize the pressure of the rapidly flowing gas.

[0037] By adopting this experimental method, turbine blade cooling effect tests can be carried out at high altitudes and low Reynolds numbers, which more realistically simulates the actual working conditions of turbine blades. This provides a basis for the design of blade cooling structures and a reference for subsequent experimental research on the cooling effect of turbine blades at high altitudes and low Reynolds numbers.

Claims

1. A high-altitude cooling test device for aero-engine turbine blades, characterized in that: The system includes a main gas pipeline (1) for connecting to a gas source station. A main gas flow heating device (2) is installed on the main gas pipeline (1). The main gas flow heating device (2) is connected to a main test chamber (3). The main test chamber (3) is connected to a main exhaust end (4). The main exhaust end (4) is equipped with a high-pressure air connection pipe (5) for connecting a high-pressure ejector. A cold air intake port (6) is installed on the main gas pipeline (1). The cold air intake port (6) is connected to a cold air pipe (8) through a flow regulating valve (7). The cold air pipe (8) is connected to the main test chamber (3). A flow meter A (9), a cold air heater (10), and a temperature sensor (11) are also installed on the cold air pipe (8) in sequence. The cold air heater (10) is equipped with a temperature control device (12). The temperature sensor (11) is located at one end of the outlet of the cold air heater (10).

2. The high-altitude cooling test device for aero-engine turbine blades according to claim 1, characterized in that: The main airflow heating device (2) is a single-tube combustion chamber, and the single-tube combustion chamber is equipped with a high-altitude igniter.

3. The high-altitude cooling test device for aero-engine turbine blades according to claim 2, characterized in that: The main gas pipeline (1) is connected to the main gas flow heating device (2) in sequence through the flow control valve (14) and the flow meter B (15). The main gas flow heating device (2) is connected to the test inlet of the main test chamber (3). The main test chamber (3) is connected to the high-pressure ejector through the main exhaust end (4). The high-pressure ejector is matched with a silencer tower.

4. The high-altitude cooling test device for aero-engine turbine blades according to claim 1, characterized in that: The cold air heater (10) is an electric furnace heater, and the temperature control device (12) is a heating power controller.

5. The high-altitude cooling test apparatus for aero-engine turbine blades according to claim 1, characterized in that: A pressure stabilizing tank (13) is connected to the air cooling pipe (8), and the pressure stabilizing tank (13) is located between the flow regulating valve (7) and the flow meter A (9).

6. A method for high-altitude cooling test of aero-engine turbine blades, characterized in that: Includes the following steps: S1: Design the test section and install the test specimen; first, connect the gas source station, main gas pipeline, main airflow heating device, main test chamber, and main exhaust end in sequence; then connect the high-pressure air connection pipe outlet to the main exhaust end, and then connect the cold air pipe as a branch pipe to the main gas pipeline and the main test chamber, and connect the flow meter A, cold air heater, and temperature sensor from the air inlet to the air outlet to the cold air pipe in sequence; finally, connect the temperature control device to control the temperature of the cold air heater; wherein, the gas source station is used to supply gas, the main gas pipeline is used to connect pipelines, the main airflow heating device is a single-tube combustion chamber, and the single-tube combustion chamber is equipped with a high-altitude igniter; S2: Air supply and drainage; the main air pipeline supplies air, and high-pressure air is dredged to create a negative pressure state required for the test. S3: Temperature and pressure control; ensure that the inlet gas pressure, flow rate, and inlet gas temperature Tg of the test section meet the test requirements; S4: Detection and measurement; Under the condition that step S3 is satisfied, measure the surface temperature Ts of the turbine blade; S5: Measures the air conditioning temperature Tc; S6: Calculate the blade cooling effect η according to the formula η=(Tg-Ts) / (Tg-Tc); Wherein, Tg is the inlet gas temperature of the test section, Ts is the surface temperature of the turbine blade, Tc is the temperature of the cooling gas, and η is the blade cooling effect.

7. The test method of the high-altitude cooling test device for aero-engine turbine blades according to claim 6, characterized in that: The temperature and voltage control mentioned in step S3 includes the following steps: a. Ignite the single-tube combustion chamber and heat the air in the main gas pipeline to the required test section inlet gas temperature Tg by adjusting the fuel flow rate; b. Draw a line of cold air from the main gas duct and heat the cold air through a cold air heater; c. By adjusting the main gas pipeline supply pressure and high-pressure air pressure, ensure that the inlet gas pressure of the test section meets the test requirements; d. By adjusting the flow rate of the cold air and the heating power of the cold air heater, the flow rate of the cold air and the surface temperature Ts of the turbine blades are made to meet the test requirements.

8. The test method of the high-altitude cooling test device for aero-engine turbine blades according to claim 7, characterized in that: In step a, the air flow rate is adjusted by the flow control valve and the air flow rate is measured by the flow meter B to meet the requirements of the main gas flow rate at the inlet of the test piece in each test state. The air-fuel mixing ratio is adjusted by the flow control valve to ensure that the gas temperature meets the requirements of the total inlet temperature of the test piece in each test state. The gas inlet pressure is made to meet the requirements of the total inlet pressure of the test piece by ejecting the gas through the high-pressure ejector. The flow control valve is used to control the pipeline flow rate, and the flow meter B is used to display the flow rate in real time.

9. The test method of the high-altitude cooling test device for aero-engine turbine blades according to claim 7, characterized in that: In step b, the air diverted from the main gas pipeline is pressurized by the pressure stabilizing tank and then enters the cold air heater. The air flow rate is adjusted by the flow regulating valve. Cooling air with different temperatures, pressures and flow rates is provided to the test blades according to the test conditions to meet the requirements of the cooling air flow rate ratio for each test condition. The pressure stabilizing tank is used to stabilize the pressure of the rapidly flowing gas.

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