A gas turbine blade thermal barrier coating high temperature erosion test device and test method

Abstract

The present invention discloses a high-temperature erosion test device and a test method for thermal barrier coatings of gas turbine blades, wherein the high-temperature erosion test device includes a high-temperature spray gun system, a test system, a cooling system, a central control system, and a dust removal system. The test device and the test method provided by the present invention can simulate the high-temperature erosion environment of gas turbine blades under working conditions to the greatest extent, simulate the position, speed, and angle relative to the coating surface of solid particles impacting the surface of the thermal barrier coating, and simulate the service environment in which particles impacting the surface of the thermal barrier coating rebound and impact other positions of the thermal barrier coating or other blade surfaces in different motion states, causing multiple erosion and wear to the coating. The test results can provide a reliable experimental basis for studying the service life of thermal barrier coatings, the selection of thermal barrier coating materials, etc.

Description

Technical Field

[0001] The invention belongs to the technical field of protective coatings for gas turbine blades, and in particular relates to a high-temperature erosion testing device and a testing method for thermal barrier coatings for gas turbine blades. Background Art

[0002] Thermal barrier coating technology has been widely used in gas turbine blades. By coating the surface of the blade base alloy with a high-temperature thermal insulation protective coating, the temperature of the base alloy can be reduced, high-temperature corrosion of the blade can be prevented, the service life of the blade can be extended, and the efficiency of the gas turbine can be improved. The thermal barrier coating system has a complex structure and an extremely harsh service environment. During service, it is not only subjected to thermal stress, but also in some cases of incomplete combustion or the use of blast furnace gas as fuel, the solid particles mixed in the gas are accelerated by the combustion chamber and impact the surface of the thermal barrier coating of the turbine blade at high speed, causing erosion and wear on the coating surface, causing the coating to become thinner, denser, and fall off, and ultimately resulting in insufficient thermal barrier performance of the coating and loss of protection for the hot end components. Therefore, it is very necessary to study the damage mechanism of high-temperature erosion of thermal barrier coatings on gas turbine blades.

[0003] In order to study the erosion damage mechanism of thermal barrier coatings on gas turbine blades and evaluate the ability of thermal barrier coatings to resist high-temperature erosion, it is necessary to simulate the erosion environment of solid particles at high temperatures. However, due to the different sizes of solid particles in the gas, the movement trajectories and motion states of the particles in the gas channel are also different. The turbine blades have complex geometric shapes and maintain high-speed rotation, which causes the solid particles to impact the surface of the thermal barrier coating at different positions, speeds, and angles relative to the coating surface. In addition, the particles that impact the blade surface will rebound and impact other positions of the thermal barrier coating in different motion states, causing multiple erosion and wear on the coating. The above factors are all difficult problems that need to be solved in simulating the high-temperature erosion conditions of thermal barrier coatings on gas turbine blades. Summary of the invention

[0004] The purpose of the present invention is to provide a high-temperature erosion test device and a test method for thermal barrier coatings of gas turbine turbine blades. By simulating the high-temperature erosion environment of gas turbine turbine blades under working conditions, the high-temperature erosion performance of the thermal barrier coating is tested. The test results provide a reliable basis for studying the service life of the thermal barrier coating, the selection of thermal barrier coating materials, etc.

[0005] The present invention is achieved by adopting the following technical solutions:

[0006] A high-temperature erosion test device for thermal barrier coating of turbine blades of a gas turbine, comprising a high-temperature spray gun system, a test system, a cooling system, a central control system and a dust removal system; wherein the high-temperature spray gun system is installed on a base, and comprises a compressed air storage tank, a fuel storage tank, a mixing chamber, a combustion chamber and an extended nozzle, one end of the mixing chamber is provided with a compressed air inlet and two fuel inlets, and the top is provided with a solid particle inlet, two fuel storage tanks are respectively connected to the corresponding fuel inlets of the mixing chamber through fuel flow regulators, and a solid particle flow control valve is provided at the solid particle inlet of the mixing chamber; a multi-porous atomizing nozzle is installed at the fuel inlet; the combustion chamber is a cavity structure, one end of the combustion chamber is connected to the mixing chamber, and the other end is inserted into the extended nozzle, and an annular cooling water pipeline is provided outside the combustion chamber and the extended nozzle;

[0007] The test system includes turbine blades, a rotating shaft and a motor; both ends of the rotating shaft are fixed on a base, and a plurality of blade fixing holes are evenly arranged in the circumference of the rotating shaft, and the turbine blades are clamped on the rotating shaft through the blade fixing holes; the central control system is used to control the start of the motor, and the motor is used to drive the rotating shaft to rotate the turbine blades; the entire test system is located in a quartz cover of the test bench that can be opened and closed, and one side of the quartz cover of the test bench is connected to an extended nozzle, and the top is connected to a dust removal system;

[0008] The cooling system includes a cooling water source, a gas heat exchanger, a liquid heat exchanger, a compressed air pipeline and a cooling water pipeline; the outlet of the compressed air storage tank is divided into two routes, one route is connected to the compressed air inlet of the mixing chamber through a gas mass flow controller, and the other route is connected to the inlet of the gas heat exchanger through a gas pressure and flow controller. The outlet of the gas heat exchanger is divided into 6 routes, which are respectively connected to the inlets of 6 cooling gas channels on one side of the rotating shaft; the cooling water source is provided with a cooling water outlet and a cooling water inlet, and the cooling water outlet is connected to the inlet of the liquid heat exchanger through a cooling water pipeline. The outlet of the liquid heat exchanger is divided into two routes, one route is connected to an annular cooling water pipeline provided outside the combustion chamber and the extended nozzle, and the other route is connected to the cooling water inlet of the hollow cavity of the rotating shaft; all cooling water outlets of the high-temperature spray gun system and the test system are connected to the cooling water inlet of the cooling water source.

[0009] A further improvement of the present invention is that six blade fixing holes are arranged on average in the circumferential direction of the rotating shaft.

[0010] A further improvement of the present invention is that six cooling gas channels are provided on one side of the rotating shaft, and the six cooling gas channels are respectively communicated with six blade fixing holes to provide cooling gas for the cooling channels of the turbine blades.

[0011] A further improvement of the present invention is that the interior of the six blade fixing installation holes of the rotating shaft is designed with a flexible part clamping, which can adjust the inclination angle of the turbine blade relative to the axis of the rotating shaft, and the inclination angle ranges from 15° to 90°.

[0012] A further improvement of the present invention is that the speed range of the motor is 0-3600 r / min.

[0013] A further improvement of the present invention is that an automatic igniter is provided in the inner chamber of the combustion chamber.

[0014] A further improvement of the present invention is that a spring-type thermocouple is installed on the inner wall of the combustion chamber to measure the flame temperature inside the combustion chamber.

[0015] A further improvement of the present invention is that a platinum-rhodium thermocouple is installed at the outlet of the gas heat exchanger to measure the temperature of the cooling gas entering the cooling gas channel of the rotating shaft.

[0016] A method for testing high temperature erosion of thermal barrier coatings on gas turbine blades, comprising the following steps:

[0017] Step 1, open the cooling water source outlet, adjust the pressure regulator, and make the cooling water flow velocity of the cooling water pipes flowing into the combustion chamber and the outside of the extended nozzle and the hollow cavity of the rotating shaft be 0.2-0.8 m / s;

[0018] Step 2, open the compressed air storage tank and the fuel storage tank, adjust the gas mass flow controller and the fuel flow regulator to optimize the air-fuel ratio in the mixing chamber; adjust the gas pressure and flow controller to make the compressed air flow rate entering the rotating shaft cooling gas channel 0-150m / s and the pressure 0-5MPa; open the solid particle inlet at the top of the mixing chamber, and adjust the solid particle flow control valve to make the solid particle flow rate entering the combustion chamber 0-10g / min;

[0019] Step 3, install the turbine blades on the rotating shaft, adjust the blade inclination angle according to the test needs, plug in the power of the central control system and the motor, click the "Start" button on the central control system, the turbine blades begin to rotate, and the speed gradually increases from 0r / min to 3600r / min. At the same time, the combustion chamber is ignited and the test starts timing. The temperature of the nozzle outlet, the surface of the thermal barrier coating of the turbine blade and the surface of the rotating shaft is measured in real time by an infrared thermometer.

[0020] The present invention has at least the following beneficial technical effects:

[0021] The high-temperature erosion test device and test method for thermal barrier coating of gas turbine blades provided by the present invention can simulate the high-temperature erosion environment of gas turbine blades under working conditions to the greatest extent, simulate the position, speed and angle of solid particles impacting the surface of thermal barrier coating, and the service environment in which particles impacting the surface of thermal barrier coating rebound and impact other positions of thermal barrier coating or other blade surfaces in different motion states, causing multiple erosion and wear of the coating. The test results can provide a reliable experimental basis for studying the service life of thermal barrier coatings, the selection of thermal barrier coating materials, etc. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Figure 1 The present invention is a schematic diagram of the overall structure of a high-temperature erosion test device for thermal barrier coatings on gas turbine blades.

[0023] Figure 2 The present invention is a schematic diagram of the partial structure of a high-temperature erosion test device for thermal barrier coatings on gas turbine blades.

[0024] Description of reference numerals:

[0025] 1. Compressed air storage tank; 2. Fuel storage tank; 3. Mixing chamber; 4. Combustion chamber; 5. Extended nozzle; 6. Base; 7. Compressed air inlet; 8. Fuel inlet; 9. Solid particle inlet; 10. Multi-hole atomizing nozzle; 11. Annular cooling water pipeline; 12. Turbine blades; 13. Rotating shaft; 14. Motor; 15. Blade fixing mounting hole; 16. Test bench quartz cover; 17. Cooling water source; 18. Gas heat exchanger; 19. Liquid heat exchanger; 20. Cooling gas channel; 21. Cooling water inlet; 22. Automatic igniter; 23. Spring-type thermocouple; 24. Pressure regulator; 25. Gas mass flow controller; 26. Fuel flow regulator; 27. Gas pressure controller; 28. Solid particle flow control valve. DETAILED DESCRIPTION

[0026] The present invention is further described in detail below in conjunction with specific implementation modes, but the present invention is not limited to the following embodiments.

[0027] like Figure 1 and Figure 2As shown, a high-temperature erosion test device for thermal barrier coating of gas turbine turbine blades provided by the present invention includes a high-temperature spray gun system, a test system, a cooling system, a central control system and a dust removal system; wherein the high-temperature spray gun system is installed on a base 6, and includes a compressed air storage tank 1, a fuel storage tank 2, a mixing chamber 3, a combustion chamber 4 and an extended nozzle 5, one end of the mixing chamber 3 is provided with a compressed air inlet 7, two fuel inlets 8, and a solid particle inlet 9 is provided on the top, the two fuel storage tanks 2 are respectively connected to the corresponding fuel inlet 8 of the mixing chamber 3 through a fuel flow regulator 26, and the solid particles in the mixing chamber 3 are connected to the fuel inlet 8 of the mixing chamber 3. A solid particle flow control valve 28 is provided at the particle inlet 9; a porous atomizing nozzle 10 is installed at the fuel inlet; the combustion chamber 4 is a cavity structure, one end of the combustion chamber is connected to the mixing chamber 3, and the other end is inserted into the extended nozzle 5, and an annular cooling water pipeline 11 is provided outside the combustion chamber and the extended nozzle; the test system includes a turbine blade 12, a rotating shaft 13 and a motor 14; both ends of the rotating shaft 13 are fixed on the base 6, and the rotating shaft 13 is evenly provided with a plurality of blade fixing holes 15 in the circumferential direction, and the turbine blades are clamped on the rotating shaft 13 through the blade fixing holes 15; the central control system is used to control the motor 1 4 is turned on, the motor 14 is used to drive the rotating shaft 13 to rotate the turbine blades, and the speed range is 0-3600r / min; the entire test system is located in the test bench quartz cover 16 which can be opened and closed, one side of the test bench quartz cover 16 is connected to the expansion nozzle 5, and the top is connected to the dust removal system; the cooling system includes a cooling water source 17, a gas heat exchanger 18, a liquid heat exchanger 19, a compressed air pipeline and a cooling water pipeline; the outlet of the compressed air storage tank is divided into two routes, one is connected to the compressed air inlet 7 of the mixing chamber through the gas mass flow controller 25, and the other is connected to the compressed air inlet 7 of the mixing chamber through the gas pressure controller 27 It is connected to the inlet of the gas heat exchanger 18, and the outlet of the gas heat exchanger 18 is divided into 6 paths, which are respectively connected to the inlets of 6 cooling gas channels 20 on one side of the rotating shaft; the cooling water source is provided with a cooling water outlet and a cooling water inlet, and the cooling water outlet is connected to the inlet of the liquid heat exchanger 19 through a cooling water pipe. The outlet of the liquid heat exchanger 19 is divided into two paths, one of which is connected to the annular cooling water pipe 11 provided outside the combustion chamber and the extended nozzle, and the other is connected to the cooling water inlet 21 of the hollow cavity of the rotating shaft; all cooling water outlets of the high-temperature spray gun system and the test system are connected to the cooling water inlet of the cooling water source.

[0028] The rotating shaft 13 is evenly provided with 6 blade fixing holes 15 in the circumferential direction. One side of the rotating shaft 13 is provided with 6 cooling gas channels 20. The 6 cooling gas channels are respectively connected with the 6 blade fixing holes 15 to provide cooling gas for the cooling channels of the turbine blades. The interior of the 6 blade fixing holes of the rotating shaft 13 is designed with flexible parts clamping, which can adjust the inclination angle of the turbine blade relative to the axis of the rotating shaft, and the inclination angle ranges from 15° to 90°.

[0029] An automatic igniter 22 is provided in the inner chamber of the combustion chamber, and a spring-type thermocouple 23 is installed on the inner wall of the combustion chamber to measure the flame temperature inside the combustion chamber.

[0030] A platinum-rhodium thermocouple is installed at the outlet of the gas heat exchanger 18 to measure the temperature of the cooling gas entering the cooling gas channel of the rotating shaft.

[0031] The present invention provides a high temperature erosion test method for thermal barrier coating of a gas turbine blade, comprising:

[0032] First, open the cooling water source outlet and adjust the pressure regulator 24 so that the cooling water flow velocity in the annular cooling water pipe 11 outside the combustion chamber and the extended nozzle and the hollow cavity of the rotating shaft is 0.2-0.8 m / s. The optimal choice is that the cooling water flow velocity in the annular cooling water pipe is 0.42 m / s and the cooling water flow velocity in the hollow cavity of the rotating shaft is 0.65 m / s.

[0033] Secondly, open the compressed air storage tank 1 and the fuel storage tank 2, adjust the gas mass flow controller 25 and the fuel flow regulator 26 to make the air-fuel ratio in the mixing chamber reach 14.7:1; adjust the gas pressure controller 27 to make the compressed air flow rate entering the rotating shaft cooling gas channel 0-150m / s and the pressure 0-5MPa. According to the aerodynamic heat transfer calculation and analysis of the first-stage turbine moving blades, the optimal choice is a compressed air flow rate of 98m / s and a pressure of 1.4MPa; open the solid particle inlet at the top of the mixing chamber, and adjust the solid particle flow control valve to make the solid particle flow entering the combustion chamber 0-10g / min, the optimal choice is 2g / min, and the solid particle diameter range is 10-600μm.

[0034] Finally, install 6 turbine blades on the rotating shaft 13 at the same interval and around the rotating shaft 13 for one circle. According to the test needs, adjust the blade inclination angles to 15°, 30°, 45°, 60°, 75° and 90° respectively; plug in the power of the central control system and the motor, click the "Start" button on the central control system, the turbine blades begin to rotate, and the combustion chamber is ignited and the test starts timing. The motor 14 controls the speed of the rotating shaft 13 to increase from 0r / min to 3600r / min within 30min. The infrared thermometer is used to measure the temperature of the nozzle outlet, the thermal barrier coating surface of the turbine blade and the rotating shaft surface in real time, and the central control system is fed back in real time.

Claims

1. A high temperature erosion test device for thermal barrier coating of gas turbine blades, characterized in that: It includes high temperature spray gun system, test system, cooling system, central control system and dust removal system; among them, The high-temperature spray gun system is installed on a base (6), and comprises a compressed air storage tank (1), a fuel storage tank (2), a mixing chamber (3), a combustion chamber (4) and an extended nozzle (5). One end of the mixing chamber (3) is provided with a compressed air inlet (7), two fuel inlets (8), and a solid particle inlet (9) is provided on the top. The two fuel storage tanks (2) are respectively connected to the corresponding fuel inlets (8) of the mixing chamber (3) through a fuel flow regulator (26). A solid particle flow control valve (28) is provided at the solid particle inlet (9) of the mixing chamber (3); a multi-hole atomizing nozzle (10) is installed at the fuel inlet; the combustion chamber (4) is a cavity structure, one end of the combustion chamber is connected to the mixing chamber (3), and the other end is inserted into the extended nozzle (5), and an annular cooling water pipeline (11) is provided outside the combustion chamber and the extended nozzle; The test system comprises a turbine blade (12), a rotating shaft (13) and a motor (14); both ends of the rotating shaft (13) are fixed on a base (6); a plurality of blade fixing holes (15) are evenly arranged on the rotating shaft (13) in a circumferential direction; the turbine blade is clamped on the rotating shaft (13) through the blade fixing holes (15); the central control system is used to control the motor (14) to start, and the motor (14) is used to drive the rotating shaft (13) to rotate the turbine blade; the entire test system is located in a quartz cover (16) of a test bench that can be opened and closed; one side of the test bench quartz cover (16) is connected to an extended nozzle (5), and the top is connected to a dust removal system; The cooling system comprises a cooling water source (17), a gas heat exchanger (18), a liquid heat exchanger (19), a compressed air pipeline and a cooling water pipeline; the outlet of the compressed air storage tank is divided into two paths, one path is communicated with the compressed air inlet (7) of the mixing chamber through a gas mass flow controller (25), and the other path is communicated with the inlet of the gas heat exchanger (18) through a gas pressure and flow controller (27); the outlet of the gas heat exchanger (18) is divided into six paths, which are respectively communicated with the inlets of six cooling gas channels (20) on one side of the rotating shaft; the cooling water source is provided with a cooling water outlet and a cooling water inlet, the cooling water outlet is communicated with the inlet of the liquid heat exchanger (19) through a cooling water pipeline, the outlet of the liquid heat exchanger (19) is divided into two paths, one path is communicated with an annular cooling water pipeline (11) provided outside the combustion chamber and the expansion nozzle, and the other path is communicated with the cooling water inlet (21) of the hollow cavity of the rotating shaft; all cooling water outlets of the high-temperature spray gun system and the test system are communicated with the cooling water inlet of the cooling water source.

2. A gas turbine blade thermal barrier coating high temperature erosion test device according to claim 1, characterized in that: Six blade fixing and mounting holes (15) are evenly arranged in the circumferential direction of the rotating shaft (13).

3. A gas turbine blade thermal barrier coating high temperature erosion test device according to claim 2, characterized in that: Six cooling gas channels (20) are arranged on one side of the rotating shaft (13), and the six cooling gas channels are respectively communicated with six blade fixing holes (15) to provide cooling gas for the cooling channels of the turbine blades.

4. A gas turbine blade thermal barrier coating high temperature erosion test device according to claim 2, characterized in that: The six blade fixing and mounting holes of the rotating shaft (13) are designed with flexible parts for clamping, and the inclination angle of the turbine blade relative to the axis of the rotating shaft can be adjusted, and the inclination angle ranges from 15° to 90°.

5. A gas turbine blade thermal barrier coating high temperature erosion test device according to claim 1, characterized in that: The speed range of the motor (14) is 0 to 3600 r / min.

6. A gas turbine blade thermal barrier coating high temperature erosion test device according to claim 1, characterized in that: An automatic igniter (22) is provided in the inner chamber of the combustion chamber.

7. A gas turbine blade thermal barrier coating high temperature erosion test device according to claim 1, characterized in that: A spring-type thermocouple (23) is installed on the inner wall of the combustion chamber to measure the flame temperature inside the combustion chamber.

8. A gas turbine blade thermal barrier coating high temperature erosion test device according to claim 1, characterized in that: A platinum-rhodium thermocouple is installed at the outlet of the gas heat exchanger (18) to measure the temperature of the cooling gas entering the cooling gas channel of the rotating shaft.

9. A method for testing high temperature erosion of thermal barrier coatings on gas turbine blades, characterized in that: The method is based on a high temperature erosion test device for thermal barrier coating of a gas turbine blade according to claim 1, and comprises the following steps: Step 1, opening the cooling water source outlet, adjusting the pressure regulator (24) so ​​that the cooling water flow velocity flowing into the combustion chamber and the cooling water pipeline outside the extended nozzle and the hollow cavity of the rotating shaft is 0.2-0.8 m / s; Step 2, open the compressed air storage tank and the fuel storage tank, adjust the gas mass flow controller (25) and the fuel flow regulator (26) to optimize the air-fuel ratio in the mixing chamber; adjust the gas pressure and flow controller (27) to make the compressed air flow rate entering the rotating shaft cooling gas channel 0-150m / s and the pressure 0-5MPa; open the solid particle inlet at the top of the mixing chamber, and adjust the solid particle flow control valve (28) to make the solid particle flow rate entering the combustion chamber 0-10g / min; Step 3, install the turbine blades on the rotating shaft, adjust the blade inclination angle according to the test needs, plug in the power of the central control system and the motor, click the "Start" button on the central control system, the turbine blades begin to rotate, and the speed gradually increases from 0r / min to 3600r / min. At the same time, the combustion chamber is ignited and the test starts timing. The temperature of the nozzle outlet, the surface of the thermal barrier coating of the turbine blade, and the surface of the rotating shaft is measured in real time by an infrared thermometer.

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