A turbine blade coating thermal shock test system

Abstract

The present disclosure relates to the technical field of material corrosion test, and particularly relates to a turbine blade coating thermal shock test system. The turbine blade coating thermal shock test system comprises an environment simulation subsystem and a test execution subsystem, the environment simulation subsystem is used for generating a gas flow, a gas outlet of the environment simulation subsystem is in communication with a gas inlet of the test execution subsystem, the test execution subsystem comprises a test part and a test piece bearing mechanism, the test piece bearing mechanism comprises a test piece fixing assembly and an angle adjusting assembly, the test piece fixing assembly comprises a test bench and a test piece clamp, a bottom end of the test piece clamp is fixedly connected with a driving end of the angle adjusting assembly, the test piece clamp is used for fixing a turbine blade test piece, and the angle adjusting assembly is used for driving the test piece clamp to deflect the turbine blade test piece to a preset angle. The turbine blade coating thermal shock test system provided by the present disclosure can realize deflection of the turbine blade test piece in a high-temperature gas flow, and truly simulates a working state of a rotatable guide vane.

Description

Technical Field

[0001] This disclosure relates to the field of material corrosion testing technology, and in particular to a thermal shock testing system for turbine blade coatings. Background Technology

[0002] Aero engines are the core power units of aircraft, and the turbine blades used in aero engines operate under particularly harsh conditions. With the development of aviation technology, the increase in the thrust-to-weight ratio of engines has led to a continuous rise in turbine inlet temperature, which places higher demands on the high-temperature resistance of turbine blades.

[0003] To ensure reliable turbine blade operation, thermal barrier coatings are typically applied to their surfaces. These coatings effectively isolate the turbine blades from high-temperature combustion gases, reducing the operating temperature of the blade substrate and thus extending their service life. However, under actual operating conditions, turbine blades not only endure high-temperature and high-pressure environments but also face the combined effects of various complex factors such as combustion gas erosion and particulate corrosion. Under these harsh conditions, the coating is prone to failure phenomena such as peeling and flaking, severely impacting the reliability and safety of the engine.

[0004] Therefore, it is crucial to study the failure mechanisms of turbine blade coatings. However, although testing thermal barrier coatings on real aero-engines can yield the most direct and reliable data, this method requires significant R&D resources and time, and has considerable limitations in practical applications. Summary of the Invention

[0005] This disclosure is made in view of the above-mentioned problems. This disclosure provides a turbine blade coating thermal shock testing system that can simulate the real working environment of turbine blades, conduct turbine blade thermal shock tests and hot corrosion tests under gas flow conditions, realize the deflection of turbine blade specimens in high-temperature gas flow, and realistically simulate the working state of rotatable guide vanes.

[0006] According to one aspect of this disclosure, a turbine blade coating thermal shock test system is provided, comprising an environmental simulation subsystem and a test execution subsystem, wherein the environmental simulation subsystem is used to generate a gas flow, and the gas outlet of the environmental simulation subsystem is connected to the gas inlet of the test execution subsystem.

[0007] The test execution subsystem includes a test section and a specimen carrying mechanism. The specimen carrying mechanism includes a specimen fixing component disposed within the test section and an angle adjustment component extending out of the test section. The specimen fixing component includes a test bench and a specimen clamp rotatably connected to the test bench. The bottom end of the specimen clamp is fixedly connected to the driving end of the angle adjustment component. The specimen clamp is used to fix the turbine blade specimen, and the angle adjustment component is used to drive the specimen clamp to deflect the turbine blade specimen to a preset angle.

[0008] In the technical solution provided in this embodiment, firstly, an environmental simulation subsystem is used to generate gas flow. The gas outlet of the environmental simulation subsystem is connected to the gas inlet of the test execution subsystem, which can realistically reproduce the high-temperature gas environment of the aero-engine turbine blade, providing a reliable environmental basis for subsequent coating performance testing. Then, the test execution subsystem includes a test section and a specimen carrying mechanism. The specimen carrying mechanism includes a specimen fixing component disposed within the test section and an angle adjustment component extending from the test section. This layout ensures both the sealing of the test environment and effective isolation of the angle adjustment mechanism from the high-temperature area. The specimen fixing component includes a test bench and a specimen clamp rotatably connected to the test bench. The specimen clamp is used to fix the turbine blade specimen. The test bench provides a stable mounting base, while the rotatable clamp allows the turbine blade specimen to flexibly adjust its angle. Simultaneously, the bottom end of the specimen clamp is fixedly connected to the drive end of the angle adjustment component, which drives the specimen clamp to deflect the turbine blade specimen to a preset angle. The angle adjustment component allows for real-time adjustment of the turbine blade specimen's installation angle during testing. It can be fixed at a specific angle to simulate the blade's stress state under static conditions, or the angle can be periodically adjusted to more realistically simulate the angle changes of the turbine blade during actual operation, thus obtaining more comprehensive test data. Attached Figure Description

[0009] The above and other objects, features, and advantages of this disclosure will become more apparent from the more detailed description of the embodiments thereof in conjunction with the accompanying drawings. The drawings are provided to further illustrate the embodiments of this disclosure and form part of the specification. They are used together with the embodiments of this disclosure to explain the disclosure and do not constitute a limitation thereof. In the drawings, the same reference numerals generally represent the same components or steps.

[0010] Figure 1 This is a structural block diagram of a turbine blade coating thermal shock test system according to an embodiment of the present disclosure;

[0011] Figure 2 This is a schematic diagram of the test execution subsystem of the turbine blade coating thermal shock test system according to an embodiment of the present disclosure;

[0012] Figure 3 This is a schematic diagram of the specimen support mechanism according to an embodiment of the present disclosure;

[0013] Figure 4 This is a rear view of the specimen support mechanism according to an embodiment of this disclosure;

[0014] Figure 5 This is a simplified structural diagram of the turbine blade coating thermal shock test system according to an embodiment of the present disclosure;

[0015] Figure 6This is a partial cross-sectional view of the connection between the burner and the test section in an embodiment of this disclosure.

[0016] Figure label:

[0017] 100-Environmental Simulation Subsystem, 110-Fuel Supply Unit, 120-Air Supply Unit, 121-Air Compressor, 122-Air Divider, 130-Burner, 200-Test Execution Subsystem, 210-Test Section, 210a-Inlet Section, 220-Specimen Fixing Assembly, 221-Test Bench, 222-Specimen Fixture, 2221-Fixture Base, 2222-Specimen Mounting Slot, 230-Angle Adjustment Assembly, 231-Rotation Shaft, 232-Connecting Rod, 233-Transmission Rod, 234-Drive Mechanism, 240- 300-Spray subsystem, 310-Spray device, 320-Corrosive solution tank, 330-Coolant tank, 400-Particulate matter addition device, 500-Tail gas recovery subsystem, 510-Tail gas treatment device, 520-Exhaust pipe, 610-Transition connection pipe, 620-First flange, 630-Second flange, 640-Seal, 650-First limiting element, 660-Second limiting element, 670-Third limiting element, 680-Guiding mechanism, 681-Guiding pin, 682-Guiding hole, 700-Waste liquid recovery device. Detailed Implementation

[0018] To make the technical problems, technical solutions, and beneficial effects to be solved by this disclosure clearer, the following detailed description is provided in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of this disclosure and are not intended to limit it.

[0019] It should be noted that when a component is referred to as being "fixed to" or "set on" another component, it can be directly on or indirectly on that other component. When a component is referred to as being "connected to" another component, it can be directly connected to or indirectly connected to that other component.

[0020] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this disclosure, "a plurality of" means two or more, unless otherwise expressly and specifically defined. "Several" means one or more, unless otherwise expressly and specifically defined.

[0021] In the description of this disclosure, it should be understood that the terms "upper", "lower", "front", "rear", "left", "right", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, and are only for the convenience of describing this disclosure and simplifying the description, and are not intended to indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this disclosure.

[0022] In the description of this disclosure, it should be noted that, unless otherwise expressly specified and limited, the terms "installation," "connection," and "joining" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components. Those skilled in the art can understand the specific meaning of the above terms in this disclosure according to the specific circumstances.

[0023] As the core power plant of an aircraft, the performance level of the aero-engine directly determines the overall capability of the aircraft. Turbine blades, as a key component of the engine's hot-end parts, operate in extremely harsh environments. With the development of aviation technology towards higher thrust-to-weight ratios and higher turbine inlet temperatures, the combustion gas temperatures of modern advanced engines have far exceeded the temperature limits of high-temperature alloy base materials. This poses unprecedented challenges to the high-temperature resistance and overall reliability of turbine blades.

[0024] To ensure the structural integrity and long-term stable operation of turbine blades in ultra-high temperature environments, the preparation of thermal barrier coating systems on their surfaces has become an indispensable key technology. Thermal barrier coatings can significantly isolate heat transfer from high-temperature combustion gases to the blade substrate, effectively reducing the substrate operating temperature by hundreds of degrees Celsius, thereby significantly improving blade service life and engine thermal efficiency. However, in harsh actual service conditions, turbine blades and their coating systems must withstand extreme environments with complex coupling effects from multiple physical and chemical fields, including high temperature, high pressure, high-cycle thermal cycling, high-speed combustion gas erosion, molten salt / deposit thermal corrosion, and erosion from inhaled particles. Under the long-term action of such coupled loads, thermal barrier coatings are prone to failure modes such as interfacial oxidation, sintering, phase transformation, cracking, and even large-area peeling, seriously threatening the operational safety and reliability of the engine.

[0025] Therefore, it is crucial to study the failure mechanisms of turbine blade coatings. However, although testing thermal barrier coatings on real aero-engines can yield the most direct and reliable data, this method requires significant R&D resources and time, and has considerable limitations in practical applications.

[0026] To address the aforementioned issues, this disclosure provides a turbine blade coating thermal shock testing system that can simulate the real working environment of turbine blades, conduct thermal shock and hot corrosion tests on turbine blades under gas flow conditions, and realize the dynamic deflection of turbine blade specimens in high-temperature gas flow, thus realistically simulating the working state of rotatable guide vanes.

[0027] Figure 1 A structural block diagram of a turbine blade coating thermal shock testing system according to an embodiment of this disclosure is shown. Figure 1 As shown, the turbine blade coating thermal shock test system of this disclosure includes: an environmental simulation subsystem 100 and a test execution subsystem 200. The environmental simulation subsystem 100 is used to generate a gas flow, and the gas outlet of the environmental simulation subsystem 100 is connected to the gas inlet of the test execution subsystem 200.

[0028] In practical applications, after the aforementioned environmental simulation subsystem 100 generates the gas flow, it can be transported to the interior of the test execution subsystem 200 through the gas outlet, thereby forming a high-temperature gas flow field inside the test execution subsystem 200. This allows the extreme gas environment that turbine blades actually experience during service to be simulated within the test execution subsystem 200, truly reproducing the high-temperature gas environment of aero-engine turbine blades and providing a reliable environmental basis for subsequent coating performance testing.

[0029] In one feasible way Figure 2 A schematic diagram of the test execution subsystem of the turbine blade coating thermal shock test system according to an embodiment of the present disclosure is shown. Figure 3 A schematic diagram of the specimen support mechanism according to an embodiment of the present disclosure is shown. Figure 4 A rear view of the specimen support mechanism according to an embodiment of this disclosure is shown. Figures 2-4 As shown, the above-mentioned test execution subsystem 200 includes a test section 210 and a specimen carrying mechanism. The specimen carrying mechanism includes a specimen fixing component 220 disposed in the test section 210 and an angle adjustment component 230 extending out of the test section 210. The specimen fixing component 220 includes a test bench 221 and a specimen clamp 222 rotatably connected to the test bench 221. The bottom end of the specimen clamp 222 is fixedly connected to the driving end of the angle adjustment component 230. The specimen clamp 222 is used to fix the turbine blade specimen, and the angle adjustment component 230 is used to drive the specimen clamp 222 to deflect the turbine blade specimen to a preset angle.

[0030] It should be noted that there can be multiple specimen clamps 222 for clamping multiple turbine blade specimens. Multiple thermocouples can be set on the surface of the turbine blade specimens. The thermocouple signal lines are led out along the reserved channel of the test bench 221 to the monitoring equipment outside the test section 210 for real-time acquisition of specimen surface temperature data.

[0031] In practice, the turbine blade specimen coated with a thermal barrier layer can be fixed onto the specimen clamp 222 to prevent displacement of the turbine blade specimen due to gas impact during the test. For static thermal shock tests, the angle adjustment component 230 remains stationary, and the specimen clamp 222 is fixed at a preset angle, such as facing the direction of gas flow. For dynamic tests to simulate the working state of rotatable guide vanes, the angle adjustment component 230 can be activated to drive the specimen clamp 222 to periodically deflect the turbine blade specimen within a preset angle range. The preset angle range refers to the deflection angle simulating the turbine blade under real operating conditions.

[0032] In one possible implementation, the turbine blade coating thermal shock test system of this disclosure embodiment further includes a control subsystem, which is communicatively connected to the aforementioned environmental simulation subsystem 100 and test execution subsystem 200, respectively.

[0033] In one alternative embodiment, the specimen fixture 222 includes a fixture base 2221 located inside the test section 210 and a specimen mounting groove 2222 fixed on the fixture base 2221. The specimen mounting groove 2222 is used to fix the turbine blade specimen. The fixture base 2221 is provided on the test bench 221 and is rotatably connected to the test bench 221.

[0034] In practice, the turbine blade specimen to be tested can be placed into the specimen mounting slot 2222 of the specimen fixture 222, ensuring that the specimen fits snugly against the inner wall of the mounting slot. High-temperature alloy bolts can be used to pass through the reserved holes in the mounting slot to fasten the turbine blade specimen to the fixture base 2221, preventing the turbine blade specimen from shifting due to gas impact during the test.

[0035] It should be understood that the profile of the above-mentioned specimen mounting groove 2222 matches the blade root profile of the turbine blade specimen to be tested, in order to ensure installation accuracy and contact stiffness.

[0036] In some examples, such as Figures 2-4 As shown, the angle adjustment assembly 230 includes a rotating shaft 231, a connecting rod 232, a transmission rod 233, and a drive mechanism 234. The bottom of the fixture base 2221 is fixedly connected to one end of the rotating shaft 231, and the other end of the rotating shaft 231 extends to the outside of the test section 210 and is fixedly connected to the connecting rod 232. The connecting rod 232 is rotatably connected to the transmission rod 233, and the transmission rod 233 is drively connected to the drive mechanism 234. The drive mechanism 234 is used to drive the transmission rod 233 to drive the connecting rod 232 and the rotating shaft 231 to rotate, so that the fixture base 2221 and the fixed turbine blade specimen deflect synchronously.

[0037] It should be noted that there can be multiple rotating shafts 231 and connecting rods 232. The number of rotating shafts 231 and connecting rods 232 corresponds to the number of specimen fixtures 222. Each rotating shaft 231 is fixedly connected to the corresponding specimen fixture 222 and the corresponding connecting rod 232. Multiple connecting rods 232 are rotatably connected to the transmission rod 233 through a rotating pair.

[0038] Specifically, the hinge points between each connecting rod 232 and the transmission rod 233 are arranged at equal intervals along the axial direction of the transmission rod 233, and the spacing between adjacent hinge points matches the installation spacing of the specimen fixture 222 in the test bench 221. This symmetrical structure ensures that the thrust or tension on each connecting rod 232 is uniform and consistent when the transmission rod 233 moves linearly, avoiding differences in deflection angles caused by force deviations. When the servo motor drives the transmission rod 233 to move linearly, the transmission rod 233 synchronously pushes or pulls each connecting rod 232 through the revolute joint, and each connecting rod 232 then drives each rotating shaft 231 to rotate synchronously, thereby driving the specimen fixture 222 and the turbine blade specimen to achieve completely consistent deflection movements.

[0039] In practical implementation, when a turbine blade deflection experiment is required, the control subsystem can control the drive mechanism 234 to drive the transmission rod 233 to perform linear reciprocating motion along the axial direction. The transmission rod 233 synchronously pushes or pulls each connecting rod 232 to rotate, thereby causing the rotating shaft 231 to rotate around its own axis. When the transmission rod 233 moves forward along the axial direction, the connecting rod 232 is pushed forward, and the connecting rod 232 will drive the rotating shaft 231 to rotate clockwise around its own axis. When the transmission rod 233 moves backward, the connecting rod 232 is pulled backward, causing the rotating shaft 231 to rotate counterclockwise around its own axis. Through the reciprocating linear motion of the transmission rod 233, the rotating shaft 231 achieves periodic rotation of alternating forward and reverse directions, thereby driving the fixture base 2221 and the turbine blade specimen to deflect synchronously. Finally, all specimens are synchronously adjusted to a preset angle in a high-temperature gas environment, realizing the testing of the thermal shock performance of the turbine blade under different angles of attack.

[0040] For example, three specimen clamp bases 222 can be installed on the test bench 221. Specimen clamps 222 are installed on the specimen clamp bases 222. The specimen clamp bases 222 are fixed together with the rotating shaft 231 and can rotate together with the rotating shaft 231. The three transmission shafts are respectively connected to the transmission rod 233 through three connecting rods 232. The rotating shaft 231 and the connecting rods 232 are fixed by welding. The connecting rods 232 and the transmission rods 233 are connected by a revolute joint. The connecting rods 232 can rotate as the transmission rods 233 move, thereby driving the rotating shaft 231 to rotate, thus enabling the turbine blade test specimen to rotate.

[0041] For example, the drive mechanism 234 may include a servo motor and a transmission gearbox. The input end of the transmission gearbox is connected to the output end of the servo motor, and the output end of the transmission gearbox is fixedly connected to the transmission rod 233. The transmission rod 233 may be hinged to multiple connecting rods 232 simultaneously, and the other end of each connecting rod 232 is fixed to the corresponding rotation shaft 231.

[0042] In practical implementation, when the control subsystem controls the operation of the aforementioned servo motor, its output rotational motion is first transmitted to the transmission gearbox. After the transmission gearbox adjusts the transmission ratio, it drives the transmission rod 233 to generate corresponding motion, which in turn drives all rotating shafts 231 to achieve synchronous deflection through each connecting rod 232. This transmission structure not only provides sufficient output torque to overcome the resistance encountered by multiple specimen clamps 222 during rotation, but also achieves precise control of the deflection angle by adjusting the transmission ratio, thereby ensuring that all turbine blade specimens maintain strictly synchronized deflection motion during dynamic testing. The adjustable transmission ratio of the gearbox also allows for flexible adaptation to different deflection speed requirements; a smaller transmission ratio is used when rapid angle scanning is needed, while a larger transmission ratio is used when precise angle positioning is required. Simultaneously, the multi-link 232 synchronous drive mechanism ensures a high degree of consistency among all turbine blade specimens during deflection, providing reliable angle control for thermal shock testing.

[0043] In one example, the aforementioned transmission gearbox is composed of multiple stages of cylindrical or bevel gears meshing together. By matching gears with different gear ratios, speed reduction and torque increase or speed increase and torque reduction can be achieved, ensuring that the output torque is sufficient to drive the specimen fixture 222 and the turbine blade specimen to overcome the gas flow resistance and complete the deflection. The output end of the aforementioned transmission gearbox and the transmission rod 233 can be connected via a screw-nut mechanism or a gear and rack mechanism. For example, if a screw-nut mechanism is used, the output shaft of the transmission gearbox can be connected to a ball screw, and the transmission rod 233 can be fixedly connected to the nut fitted on the screw. When the screw rotates, the nut fitted on the screw moves linearly along the screw axis, thereby driving the transmission rod 233 to move linearly synchronously with the nut. Both methods can achieve smooth reciprocating motion of the transmission rod 233, and the direction of movement of the transmission rod 233 can be switched by controlling the forward and reverse rotation of the servo motor.

[0044] In some instances, such as Figure 4As shown, a collar 240 can be machined at the through-hole of the rotating shaft 231 extending outside the test section 210 to provide external support and fixation for the rotating shaft 231. The contact surface between the collar 240 and the outer shell of the test section 210 is sealed with a ceramic fiber gasket, thereby maintaining continuous airtight isolation between the internal and external environments of the test section 210 during the dynamic deflection of the rotating shaft 231. Correspondingly, a collar can be machined between the rotating shaft 231 and the test bench 221. This collar is used for radial positioning and axial fixation of the rotating shaft 231, and also serves as the mounting base for the specimen clamp 222. A ceramic fiber gasket is also provided between the mating surfaces of the collar and the test bench 221, utilizing its high-temperature elasticity and sealing properties to ensure that the high-temperature combustion gas inside the test section 210 does not leak circumferentially along the rotating shaft 231. The sealing structure of the double collar and ceramic fiber gasket ensures both the stability and accuracy of the rotating support and effectively solves the dynamic sealing problem of moving parts under high-temperature conditions, ensuring the reliability and safety of the test system during long-term operation.

[0045] For example, Figure 5 A simplified structural diagram of a turbine blade coating thermal shock testing system according to an embodiment of this disclosure is shown, as follows: Figure 5 As shown, the above-mentioned environmental simulation subsystem 100 includes a fuel supply unit 110, an air supply unit 120, and a burner 130. The fuel outlet of the fuel supply unit 110 is connected to the inlet of the burner 130 for supplying fuel to the burner 130. The air outlet of the air supply unit 120 is connected to the inlet of the burner 130 for supplying combustion air into the burner 130. The burner 130 is used to generate a gas flow.

[0046] For example, the fuel supply unit 110 may include a fuel tank, a delivery pump, and a flow control valve for controlling fuel supply. The air supply unit 120 is used to provide clean and pressure-stable combustion air. The burner 130 preferably adopts an aircraft engine combustion chamber structure, capable of generating stable and temperature-controlled high-temperature combustion gas to simulate the actual service environment of turbine blades.

[0047] In practice, the control subsystem controls the delivery pump in the fuel supply unit 110 to pump aviation kerosene from the fuel tank. After being regulated by a flow control valve, the kerosene is delivered to the burner 130 via pipeline. Simultaneously, the air supply unit 120 provides the required combustion air flow, which is also delivered to the burner 130. After the fuel and combustion air mix at the head of the burner 130, they are ignited inside the combustion chamber and maintained in a stable combustion state, generating a high-temperature gas flow. The burner 130 is based on the structural design of an aero-engine combustion chamber and typically contains components such as swirlers and flame tubes, enabling it to form a stable recirculation zone and organize efficient combustion, thereby producing high-temperature gas with a uniform and controllable temperature distribution. This high-temperature gas flow is then guided into the test execution subsystem 200 to simulate the high-temperature service environment faced by real aero-engine turbine blades.

[0048] In some examples, such as Figure 5 As shown, the air supply unit 120 includes an air compressor 121 and an air splitter 122. The compressed gas outlet of the air compressor 121 is connected to the inlet of the air splitter 122. The air splitter 122 is configured to split the compressed air into two independent airflow channels, a first airflow channel and a second airflow channel. The outlet of the first airflow channel is connected to the inlet of the burner 130 and is used to provide combustion air to the burner 130. The outlet of the second airflow channel is directly connected to the inlet of the test section 210 and is used to provide mixed cooling air to the test section 210.

[0049] In practical applications, by adjusting the flow distribution ratio of the first and second airflow channels, it is possible to ensure stable combustion by maintaining the optimal air-fuel ratio within the burner 130, and to control the temperature of the gas entering the test chamber 210 by changing the flow rate of the mixed cooling air. This split-flow design allows the system to maintain stable combustion while enabling rapid and continuous adjustment of the gas temperature within the test chamber 210 within the range of 300℃ to 1200℃, meeting the requirements of thermal shock tests under different operating conditions.

[0050] For example, in this embodiment of the present disclosure, the fuel supply unit 110 further includes an oil circuit valve to regulate the supply of aviation kerosene. The air diversion device is an air circuit three-way valve. The air circuit three-way valve and the oil circuit valve are independently controlled by servo motors. The air circuit three-way valve adjusts its valve core opening via the servo motor, thereby precisely controlling the flow ratio between the main air path to the burner 130 and the branch air path to the test section. The servo motors controlling the air circuit three-way valve and the oil circuit valve are centrally coordinated by the aforementioned control subsystem to achieve coordinated regulation of air and fuel flow, thereby completing a wide-range and rapid adjustment of the gas temperature within the test section 210.

[0051] The inventors discovered that the gas temperature depends on the air-fuel ratio of aviation kerosene to air within the burner 130, as well as the mixing ratio of the high-temperature gas produced by combustion with the cooling air entering the test section via the air branch. To achieve rapid and continuous adjustment of the gas temperature over a wide range, it is necessary to first ensure stable combustion of fuel and air within the burner 130, which is achieved by coordinating and controlling the air-fuel ratio, the main air flow rate, and the air branch flow rate. For example, when aviation kerosene burns under an ideal air-fuel ratio of approximately 15:1, the gas temperature can reach 1250℃~1300℃. When the air-fuel ratio decreases, incomplete combustion occurs, and the gas temperature range is 1000℃~1200℃. When the air-fuel ratio increases, excess air is formed, and the temperature drops to 800℃~1000℃.

[0052] Neglecting heat loss, the temperature of the mixed gas can be calculated using the following formula based on the law of conservation of heat:

[0053] .

[0054] Where T is the temperature of the mixed gas in the test section 210, 100 is the percentage of total air, T1 is the air temperature, T2 is the gas temperature at the outlet of the burner 130, x is the percentage of air entering the burner 130, and 100-x is the percentage of air entering the test section.

[0055] Based on this, the system of this embodiment can achieve gas temperature regulation within the range of 300℃ to 1200℃. By integrating the control of the servo motors of the air circuit three-way valve and the oil circuit valve based on the control subsystem, the air-fuel ratio and the flow rates of the main air circuit and branch circuits are regulated, ultimately achieving the goal of rapid, continuous, and controllable gas temperature regulation over a wide range.

[0056] In some examples, such as Figure 5 As shown, the turbine blade coating thermal shock test system of this disclosure embodiment also includes a spray subsystem 300. The spray subsystem 300 includes a spray device 310 and a medium supply unit. The spray device 310 is disposed at the gas inlet of the test section 210. The nozzle of the spray device 310 faces the turbine blade specimen and is used to spray the medium onto the turbine blade specimen. The medium outlet of the medium supply unit is connected to the spray device 310. The medium supply unit includes a corrosive solution tank 320 and a coolant tank 330 and is used to provide spray media with different functions to the spray device 310. The spray subsystem 300 is used to perform a thermal corrosion test on the turbine blade specimen or to perform system cooling after the test.

[0057] It should be noted that the corrosive solution tank 320 is used to store solutions of corrosive media such as sulfates in a simulated engine environment, and the coolant tank 330 is used to store the coolant required for system cooling after the test.

[0058] For example, the spray subsystem 300 also includes control valves respectively installed on the corrosive solution pipeline and the coolant pipeline. These valves control the timing, duration, and flow rate parameters of the spraying of the corresponding medium according to the test requirements. The spray subsystem 300 is communicatively connected to the control subsystem. The control subsystem controls the spray subsystem 300 to perform a thermal corrosion environment simulation test on the turbine blade specimen according to a preset program or instruction, or to start a cooling program to safely cool the system after the test process is completed.

[0059] In some examples, such as Figure 5 As shown, the turbine blade coating thermal shock test system of this embodiment further includes a particulate matter adding device 400. The particulate matter adding device 400 includes a particulate material bin and a conveying pipeline. The outlet of the particulate material bin is connected to the inlet of the conveying pipeline, and the outlet of the conveying pipeline is connected to the inlet of the burner 130.

[0060] It should be understood that the particulate matter adding device 400 is used to add hard particulate matter of a set particle size and concentration into the burner 130 to simulate the erosion effect on the turbine blades caused by foreign objects such as dust and sand particles inhaled by the aircraft engine during actual operation, as well as internal impurities such as carbon particles generated by combustion.

[0061] In practice, the particulate material silo can be equipped with a vibrating feeder or a screw feeder to ensure that the particulate matter is output at a constant rate. The conveying pipeline can adopt a pneumatic conveying method, using compressed air to uniformly mix the particulate matter into the combustion air, and then enter the burner 130 together. After being accelerated by the high-temperature combustion gas in the burner 130, the particulate matter impacts the surface of the turbine blade specimen, thus realistically reproducing the particulate erosion conditions in the actual service environment. By controlling the material, particle size, concentration, and addition rate of the particulate matter, the damage mechanism and failure behavior of the thermal barrier coating under different erosion conditions can be systematically studied.

[0062] In some examples, such as Figure 5 As shown, the turbine blade coating thermal shock test system of this embodiment further includes an exhaust gas recovery subsystem 500. The exhaust gas recovery subsystem 500 includes an exhaust gas treatment device 510 and an exhaust pipe 520. The outlet of the test section 210 is connected to the inlet of the exhaust gas treatment device 510 for purifying the exhaust gas after the test. The outlet of the exhaust gas treatment device 510 is connected to the inlet of the exhaust pipe 520 for discharging the purified gas.

[0063] Specifically, the exhaust gas treatment device 510 can adopt a multi-stage purification structure, such as: a quench tower for rapidly reducing the temperature of the exhaust gas, a bag filter or electrostatic precipitator for removing particulate matter, an alkaline spray tower for neutralizing acidic gas components, a selective catalytic reduction device for removing nitrogen oxides, and a device for desulfurization and denitrification.

[0064] The aforementioned exhaust gas recovery subsystem 500 can effectively remove pollutants such as solid particles, sulfur oxides, and nitrogen oxides from exhaust gas, ensuring that the emitted gas meets environmental protection requirements, while also guaranteeing the safe and stable operation of the test system.

[0065] In some instances, such as Figure 5 As shown in the figure, the turbine blade coating thermal shock test system of this disclosure also includes a waste liquid recovery device 700. The waste liquid recovery device includes a waste liquid collection tank, a waste liquid pipeline, and a waste liquid valve. The waste liquid collection tank is connected to the drain port of the test section 210 through the waste liquid pipeline. The waste liquid valve is installed on the waste liquid pipeline to control the flow of waste liquid generated during the test to the waste liquid collection tank. The waste liquid recovery device 700 is used to collect various types of waste liquid generated during the test, including corrosive solutions remaining after the hot corrosion test, coolant generated after the system cools down, and cleaning fluid used to rinse the pipeline. By centrally collecting and treating these waste liquids, environmental protection requirements are met, and the continuous operation of the test system is guaranteed.

[0066] In some examples, Figure 6 A partial cross-sectional view of the connection between the burner and the test section according to an embodiment of this disclosure is shown, as follows: Figure 6 As shown, the turbine blade coating thermal shock test system of this embodiment further includes a floating sealing assembly. The burner 130 and the test section 210 are connected by the floating sealing assembly. The floating sealing assembly includes a transition connecting pipe 610, a first flange 620, a second flange 630, and a seal 640. One end of the transition connecting pipe 610 is fixedly connected to the outlet of the burner 130. The test section 210 has an inlet section 210a. The other end of the transition connecting pipe 610 extends into the inlet section 210a of the test section. The first flange 620 is loosely fitted on the outer periphery of the transition connecting pipe 610, the second flange 630 is loosely fitted on the outer periphery of the inlet section 210a, and the seal 640 is loosely fitted on the outer periphery of the transition connecting pipe 610. The seal 640 is located between the first flange 620 and the second flange 630. The first flange 620 and the second flange 630 are connected by fasteners, so that the seal 640 is subjected to uniform compression force in the axial direction, thereby realizing a dynamic sealing connection between the burner 130 and the test section 210.

[0067] It is understood that the aforementioned transition connection pipe 610 adopts a fan-shaped cross-section pipe section adapted to the outlet of the burner 130, and one end of it is fixedly connected to the outlet of the burner 130 by welding. The inlet section 210a of the aforementioned test section 210 is designed with a socket that matches the shape of the transition connection pipe 610, allowing the other end of the transition connection pipe 610 to extend into it while maintaining a certain radial clearance. The aforementioned sealing element 640 is preferably a high-temperature resistant ceramic fiber reinforced composite gasket, the cross-sectional shape of which corresponds to the fan-shaped cross-section of the transition connection pipe 610. When the burner 130 and the test section 210 experience axial and radial relative displacement due to temperature changes, the sealing element 640 can compensate for the clearance change in real time through its own compression or rebound, always maintaining a tight fit with the flange and the transition connection pipe 610, and preventing high-temperature gas from leaking from the gap.

[0068] In practical applications, fasteners bring the sealing surfaces of the first flange 620 and the second flange 630 closer together, applying a uniform clamping force to the intermediate sealing element 640. The elastic deformation of the sealing element 640 fills the gap between the flange and the transition connecting pipe 610, forming an initial seal. When the system expands due to heat, the transition connecting pipe 610 can freely expand and contract relative to the inlet section 210a of the test section 210. Simultaneously, the first flange 620 and the second flange 630 can axially float along the outer circumference of the pipe section, and the elastic deformation of the sealing element 640 continuously maintains the sealing state. This effectively solves the problem of sealing failure caused by thermal deformation under high-temperature conditions, ensuring the long-term sealing reliability of the test system in a high-temperature and high-pressure gas environment.

[0069] In some examples, the above-described floating sealing assembly further includes a first limiting member 650 and a second limiting member 660. The first limiting member 650 is fixed to the outer periphery of the transition connecting pipe 610, and the second limiting member 660 is fixed to the outer periphery of the inlet section 210a. The first limiting member 650 is located between the burner 130 and the first flange 620, and the second limiting member 660 is located between the second flange 630 and the test section 210. The end of the inlet section 210a has a third limiting member 670, which is located between the sealing member 640 and the second flange 630.

[0070] It should be understood that the aforementioned first limiting member 650, second limiting member 660, and third limiting member 670 are all annular retaining ring structures, which are fixed to the corresponding positions of the transition connecting pipe 610 by welding or threaded connection. The first limiting member 650 is used to limit the axial displacement of the first flange 620 towards the burner 130, the second limiting member 660 is used to limit the axial displacement of the second flange 630 towards the test section 210, and the third limiting member 670 is used to limit the axial displacement of the second flange 630 towards the burner 130, preventing the second flange 630 from falling off the inlet section 210a.

[0071] In practical applications, during testing, the temperature difference between the burner 130 and the test section 210 leads to different amounts of thermal expansion and contraction (e.g., the burner 130 side has a higher temperature, resulting in greater elongation of the transition connecting pipe section). The floating design allows the first flange 620 to slide along the transition connecting pipe 610 and the second flange 630 to slide along the inlet section 210a. Combined with the radial clearance between the transition connecting pipe 610 and the socket, this simultaneously compensates for axial and radial displacement due to pipe section expansion and contraction, avoiding tensile or compressive stresses generated during rigid connections and preventing flange cracking or weld breakage. Therefore, the first limiting member 650, the second limiting member 660, and the third limiting member 670 together constitute the axial constraint system of the floating sealing assembly, ensuring the stability and safety of the connection structure while allowing thermal deformation compensation. By controlling the floating range of each component, sealing performance is maintained while preventing component separation or seal failure due to excessive displacement, ensuring long-term reliable operation of the test system under high temperature and high pressure conditions.

[0072] In some examples, the floating sealing assembly described above also includes a guide mechanism 680 connection, which includes a plurality of guide pins 681 disposed on the first flange 620 and a plurality of guide holes 682 disposed on the second flange 630. Each guide hole 682 cooperates with a corresponding guide pin 681 to limit radial and circumferential displacement between the first flange 620 and the second flange 630.

[0073] In practical applications, the guide pins 681 are evenly distributed circumferentially along the sealing surface of the first flange 620. The guide pins 681 and the first flange 620 can be connected by threads or manufactured as an integral structure; this is not limited here. The second flange 630 has guide holes 682 machined at corresponding positions to match the guide pins 681. The diameter of the guide holes 682 is slightly larger than the diameter of the guide pins 681, forming a precise clearance fit. This fit allows the first flange 620 and the second flange 630 to float relatively freely in the axial direction to compensate for thermal deformation, while effectively limiting the relative movement of the first flange 620 and the second flange 630 in the radial and circumferential directions, avoiding misalignment leading to uneven wear or localized compression failure of the seal 640. Therefore, the guide mechanism 680 ensures that the two flanges maintain good alignment under thermal and vibration loads, guaranteeing the reliability of the seal, extending the service life of the seal 640, and improving the overall stability of the floating sealing assembly under high-temperature conditions.

[0074] In some examples, the aforementioned test section 210 also includes multiple temperature sensors located in the inlet section 210a and pressure sensors evenly distributed inside the test section 210. The temperature sensors are preferably thermocouples, arranged radially at intervals along the inlet section 210a, for real-time monitoring of the temperature distribution of the high-temperature gas entering the test section 210 at different locations. The pressure sensors are evenly distributed at the inlet, middle, and tail of the test section to monitor gas pressure changes in different areas inside the test section 210.

[0075] The aforementioned pressure sensor is connected to the data acquisition system via a signal line, transmitting the monitored temperature and pressure data to the control subsystem in real time.

[0076] For example, the outlet gas passage of the burner 130 can be a fan-shaped cross-section to simulate the radial non-uniformity of the temperature distribution at the combustion chamber outlet in an engine. Specifically, the geometry of the fan-shaped cross-section passage corresponds to the annular passage at the outlet of a real aero-engine combustion chamber, and its radially varying cross-sectional dimensions can form a temperature gradient distribution that matches reality. This design allows the high-temperature gas to reproduce the radially non-uniform thermal load experienced by the turbine blades in the actual working environment when it enters the test section, thereby more realistically simulating the thermal shock conditions experienced by the thermal barrier coating at different radial locations.

[0077] For example, the aforementioned control subsystem is also used to coordinate and control the fuel supply unit 110, air supply unit 120, burner 130, and spray subsystem 300 in the environmental simulation subsystem 100, adjust the gas temperature, flow rate, and medium injection parameters, perform motion control on the angle adjustment component 230 in the test execution subsystem 200, realize the static positioning or dynamic deflection of the turbine blade specimen, receive and process detection signals from various sensors, including temperature, pressure, angle, and flow parameters, realize real-time monitoring and data acquisition of the test process, automatically execute the test process according to the preset test program, including system startup, parameter adjustment, test execution, safety interlock, and system shutdown, provide a human-machine interface for setting test parameters, process monitoring and data visualization, and store test data and generate test reports, providing data support for the performance analysis of turbine blade coatings.

[0078] Example 1

[0079] A method for conducting static thermal shock tests on turbine blade coatings using the turbine blade coating thermal shock testing system according to embodiments of the present disclosure includes the following steps:

[0080] 1. System inspection and preparation: Inspect the material status of the particulate material silo, fuel storage tank, corrosive solution tank 320, coolant tank 330 and waste liquid collection tank, and confirm the sealing performance of the entire test device.

[0081] 2. Specimen installation and positioning: The turbine blade specimen is installed on the specimen fixture 222, and the fixed angle of the turbine blade specimen relative to the direction of gas flow is set by the control subsystem.

[0082] 3. Test environment establishment and monitoring: Start the air compressor 121, set the opening parameters of the air circuit three-way valve and the oil circuit valve, the gas temperature change curve and the test duration in the control subsystem, supply air and aviation kerosene to the burner 130 and ignite it, and at the same time add particulate matter from the particulate material bin into the burner 130, and monitor the temperature and pressure distribution in the test section and the surface temperature of the turbine blade specimen in real time.

[0083] 4. Hot corrosion test execution: After the combustion in the burner 130 is stable, open the corrosive solution circuit valve, and supply corrosive salt solution from the corrosive solution tank 320 to the spray device 310. The corrosive salt solution is sprayed onto the surface of the turbine blade specimen through the spray device 310. The exhaust gas generated during the test is introduced into the exhaust gas recovery device for treatment before being discharged.

[0084] 5. Test completion and system cooling: After the test is completed, turn off the air compressor 121. After the system stabilizes, open the coolant circuit valve and supply coolant from the coolant tank 330 to the spray device 310 to spray and cool the test section.

[0085] 6. Waste liquid recovery: After the cooling process is completed, open the waste liquid valve and discharge the waste liquid in the test section into the waste liquid collection tank.

[0086] 7. Specimen Removal and Analysis: The turbine blade specimens were removed for subsequent testing and analysis.

[0087] Example 2

[0088] A method for conducting dynamic thermal shock tests on turbine blade coatings using the turbine blade coating thermal shock testing system according to embodiments of this disclosure includes the following steps:

[0089] 1. System inspection and preparation: Inspect the material status of the particulate material silo, fuel storage tank, corrosive solution tank 320, coolant tank 330 and waste liquid collection tank, and confirm the sealing performance of the entire test device.

[0090] 2. Specimen installation and dynamic parameter setting: The turbine blade specimen is installed on the specimen fixture 222. The rotation angle range and rotation period of the turbine blade specimen are set through the control subsystem to simulate the actual working process of the rotatable guide vane and start the periodic rotation.

[0091] 3. Test environment establishment and monitoring: Start the air compressor 121, set the opening parameters of the air circuit three-way valve and the oil circuit valve, the gas temperature change curve and the test duration in the control subsystem, supply air and aviation kerosene to the burner 130 and ignite it, and at the same time add particulate matter from the particulate material bin into the burner 130, and monitor the temperature and pressure distribution in the test section and the surface temperature of the turbine blade specimen in real time.

[0092] 4. Hot corrosion test execution: After the combustion in the burner 130 is stable, the corrosive salt solution circuit valve is opened, and the corrosive salt solution is supplied from the corrosive solution tank 320 to the spraying device 310. The corrosive salt solution is sprayed onto the surface of the turbine blade specimen that is in a periodic rotation state through the spraying device 310. The exhaust gas generated during the test is introduced into the exhaust gas recovery device for treatment before being discharged.

[0093] 5. Test completion and system cooling: After the test is completed, turn off the air compressor 121. After the system stabilizes, open the coolant circuit valve and supply coolant from the coolant tank 330 to the spray device 310 to spray and cool the test section.

[0094] 6. Waste liquid recovery: After the cooling process is completed, open the waste liquid valve and discharge the waste liquid in the test section into the waste liquid collection tank.

[0095] 7. Specimen Removal and Analysis: The turbine blade specimens were removed for subsequent testing and analysis.

[0096] The basic principles of this disclosure have been described above with reference to specific embodiments. However, it should be noted that the advantages, benefits, and effects mentioned in this disclosure are merely examples and not limitations, and should not be considered as essential features of each embodiment of this disclosure. Furthermore, the specific details disclosed above are for illustrative and facilitative purposes only, and are not limitations. These details do not limit the scope of this disclosure to the necessity of employing the aforementioned specific details for implementation.

[0097] The block diagrams of devices, apparatuses, devices, and systems disclosed herein are merely illustrative examples and are not intended to require or imply that they must be connected, arranged, or configured in the manner shown in the block diagrams. As those skilled in the art will recognize, these devices, apparatuses, devices, and systems can be connected, arranged, and configured in any manner. Words such as “comprising,” “including,” “having,” etc., are open-ended terms meaning “including but not limited to,” and are used interchangeably with them. The terms “or” and “and” as used herein refer to the terms “and / or,” and are used interchangeably with them unless the context clearly indicates otherwise. The term “such as” as used herein refers to the phrase “such as but not limited to,” and is used interchangeably with it.

[0098] Additionally, as used herein, the "or" used in a list of items beginning with "at least one" indicates a separate list, such that a list of, for example, "at least one of A, B, or C" means A or B or C, or AB or AC or BC, or ABC (i.e., A and B and C). Furthermore, the word "exemplary" does not imply that the described example is preferred or better than other examples.

[0099] It should also be noted that in the systems and methods of this disclosure, the components or steps can be decomposed and / or recombined. These decompositions and / or recombinations should be considered as equivalent solutions to this disclosure.

[0100] Various changes, substitutions, and modifications can be made to the technology herein without departing from the teachings defined by the appended claims. Furthermore, the scope of the claims of this disclosure is not limited to the specific aspects of the processes, machines, manufactures, events, means, methods, and actions described above. Currently existing or later-developed processes, machines, manufactures, events, means, methods, or actions that perform substantially the same function or achieve substantially the same result as the corresponding aspects herein can be utilized. Therefore, the appended claims include such processes, machines, manufactures, events, means, methods, or actions within their scope.

[0101] The above description of the disclosed aspects is provided to enable any person skilled in the art to make or use this disclosure. Various modifications to these aspects will be readily apparent to those skilled in the art, and the general principles defined herein may be applied to other aspects without departing from the scope of this disclosure. Therefore, this disclosure is not intended to be limited to the aspects shown herein, but rather to be carried out within the widest scope consistent with the principles and novel features disclosed herein.

[0102] The above description has been given for purposes of illustration and description. Furthermore, this description is not intended to limit the embodiments of this disclosure to the forms disclosed herein. Although numerous exemplary aspects and embodiments have been discussed above, those skilled in the art will recognize certain variations, modifications, alterations, additions, and sub-combinations thereof.

Claims

1. A thermal shock testing system for turbine blade coatings, characterized in that, It includes an environmental simulation subsystem and a test execution subsystem. The environmental simulation subsystem is used to generate a gas flow, and the gas outlet of the environmental simulation subsystem is connected to the gas inlet of the test execution subsystem. The test execution subsystem includes a test section and a specimen carrying mechanism. The specimen carrying mechanism includes a specimen fixing component disposed in the test section and an angle adjustment component extending out of the test section. The specimen fixing component includes a test bench and a specimen clamp rotatably connected to the test bench. The bottom end of the specimen clamp is fixedly connected to the driving end of the angle adjustment component. The specimen clamp is used to fix the turbine blade specimen, and the angle adjustment component is used to drive the specimen clamp to deflect the turbine blade specimen to a preset angle. The specimen fixture includes a fixture base located inside the test section. The fixture base is disposed on the test bench and rotatably connected to the test bench. The angle adjustment assembly includes a rotating shaft, a connecting rod, a transmission rod, and a drive mechanism. The bottom of the fixture base is fixedly connected to one end of the rotating shaft. The other end of the rotating shaft extends to the outside of the test section and is fixedly connected to the connecting rod. The connecting rod is rotatably connected to the transmission rod. The transmission rod is drively connected to the drive mechanism. The drive mechanism is used to drive the transmission rod to drive the connecting rod and the rotating shaft to rotate, so that the fixture base and the fixed turbine blade specimen deflect synchronously.

2. The turbine blade coating thermal shock testing system according to claim 1, characterized in that, The specimen fixture also includes a specimen mounting slot fixed on the fixture base, the specimen mounting slot being used to fix the turbine blade specimen.

3. The turbine blade coating thermal shock testing system according to claim 1, characterized in that, The environmental simulation subsystem includes a fuel supply unit, an air supply unit, and a burner. The fuel outlet of the fuel supply unit is connected to the inlet of the burner to supply fuel to the burner. The air outlet of the air supply unit is connected to the inlet of the burner to supply combustion air to the burner. The burner is used to generate a gas flow.

4. The turbine blade coating thermal shock testing system according to claim 3, characterized in that, The air supply unit includes an air compressor and an air distribution device. The compressed gas outlet of the air compressor is connected to the inlet of the air distribution device. The air distribution device includes a first airflow channel and a second airflow channel. The outlet of the first airflow channel is connected to the inlet of the burner, and the outlet of the second airflow channel is connected to the inlet of the test section.

5. The turbine blade coating thermal shock testing system according to claim 1, characterized in that, It also includes a spray subsystem, which comprises a spray device and a media supply unit. The spray device is located at the gas inlet of the test section and is used to spray media onto the turbine blade specimen. The media outlet of the media supply unit is connected to the spray device. The media supply unit includes a corrosive solution tank and a coolant tank, which are used to provide spray media with different functions to the spray device. The spray subsystem is used to perform hot corrosion tests on the turbine blade specimen or to cool the system after the test.

6. The turbine blade coating thermal shock testing system according to claim 3, characterized in that, It also includes a particulate matter adding device, which includes a particulate material bin and a conveying pipeline. The outlet of the particulate material bin is connected to the inlet of the conveying pipeline, and the outlet of the conveying pipeline is connected to the inlet of the burner. It is used to add hard particulate matter of a set particle size and concentration into the burner.

7. The turbine blade coating thermal shock testing system according to claim 1, characterized in that, It also includes an exhaust gas recovery subsystem, which includes an exhaust gas treatment device and an exhaust pipe. The outlet of the test section is connected to the inlet of the exhaust gas treatment device for purifying the exhaust gas after the test. The outlet of the exhaust gas treatment device is connected to the inlet of the exhaust pipe for discharging the purified gas.

8. The turbine blade coating thermal shock testing system according to claim 3, characterized in that, It also includes a floating sealing assembly, through which the burner and the test section are connected; The floating sealing assembly includes a transition connecting pipe, a first flange, a second flange, and a seal. One end of the transition connecting pipe is fixedly connected to the outlet of the burner. The test section has an inlet section, and the other end of the transition connecting pipe extends into the inlet section. The first flange is loosely fitted around the outer periphery of the transition connecting pipe, the second flange is loosely fitted around the outer periphery of the inlet section, and the seal is loosely fitted around the outer periphery of the transition connecting pipe. The seal is located between the first flange and the second flange. The first flange and the second flange are connected by fasteners to make the burner and the test section sealed together.

9. The turbine blade coating thermal shock testing system according to claim 8, characterized in that, The end face of the first flange near the seal is a convex face, and the end face of the second flange near the seal is a concave face that matches the convex face, forming a concave-convex sealing structure.

10. The turbine blade coating thermal shock testing system according to claim 8, characterized in that, The floating sealing assembly further includes a first limiting member and a second limiting member. The first limiting member is fixed to the outer periphery of the transition connecting pipe, and the second limiting member is fixed to the outer periphery of the inlet section. The first limiting member is located between the burner and the first flange, and the second limiting member is located between the second flange and the test section. The end of the inlet section has a third limiting member, which is located between the seal and the second flange.

11. The turbine blade coating thermal shock testing system according to claim 8, characterized in that, The floating sealing assembly also includes a guide mechanism connection, which includes a plurality of guide pins disposed on the first flange and a plurality of guide holes disposed on the second flange. Each guide hole cooperates with a corresponding guide pin to limit the radial and circumferential displacements between the first flange and the second flange.

12. The turbine blade coating thermal shock testing system according to any one of claims 1 to 11, characterized in that, The test section also includes multiple temperature detection elements and pressure sensors evenly distributed inside the test section.

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