Aerodynamic performance testing device for cooling turbine
By placing the bearing and lubricating oil chamber downstream of the turbine rotor, and utilizing the heat insulation chamber and cold air passage for cooling and sealing components, the problem of lubricating oil failure was solved, ensuring the safety and accuracy of the test data, and simplifying the structural design.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- AECC HUNAN AVIATION POWERPLANT RES INST
- Filing Date
- 2026-03-10
- Publication Date
- 2026-05-12
AI Technical Summary
In aero-engine cooling turbine aerodynamic performance tests, lubricating oil is prone to failure due to high-temperature combustion gas heat conduction and heat radiation, leading to bearing wear and jamming, which affects test safety and data accuracy.
Design a cooling turbine aerodynamic performance test device, place the bearing and lubricating oil chamber downstream of the turbine rotor, set up a heat insulation chamber and a cold air channel, reduce the lubricating oil temperature through reverse heat exchange of cold air, and set up sealing components in key parts to block heat and fluid leakage.
It effectively prevents lubricating oil and bearings from failing due to high temperatures, ensures test safety and data accuracy, simplifies structural design, and supports long-term high-temperature testing.
Smart Images
Figure CN122016221A_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of aero-engine testing technology, and more specifically to a cooling turbine aerodynamic performance testing device. Background Technology
[0002] The turbine is a core power component of an aero-engine, and its aerodynamic performance directly determines the engine's thrust, fuel efficiency, and reliability. To develop high-performance aero-engines, it is necessary to simulate real-world operating conditions through cooled turbine aerodynamic performance tests to analyze the impact of parameters such as cold gas injection methods and flow rates on turbine efficiency and aerodynamic stability. In these tests, the high-temperature combustion gases (reaching 800℃-1200℃ or even higher) in the turbine's mains can transfer heat to the bearing oil chamber through heat conduction and radiation. This causes the oil temperature to exceed its allowable operating temperature (typically ≤150℃), leading to oil carbonization, viscosity reduction, and even failure. This, in turn, causes bearing wear and seizure, seriously threatening test safety. Simultaneously, oil failure disrupts the stable operation of the turbine rotor, making it impossible to accurately simulate real engine operating conditions. This results in distorted research data on the impact of cold gas on turbine performance, making it difficult to accurately verify the turbine's true performance.
[0003] In related technologies, the test specimen structure is mostly arranged in a symmetrical layout. The two support points of the bearing are located at the front and rear ends of the turbine rotor blade. The airflow temperature at the front end of the turbine rotor blade is the highest. After the high-temperature airflow drives the turbine rotor blade to do work, the temperature will decrease. Therefore, the bearing and lubricating oil located at the front end of the turbine rotor blade are subject to a large amount of heat conduction and heat radiation from the high-temperature airflow, which makes them prone to failure. Summary of the Invention
[0004] In view of this, the present invention provides a cooling turbine aerodynamic performance testing device to solve the problem of lubricating oil failure in the aerodynamic performance testing of cooling turbines in aero engines.
[0005] This invention provides a cooling turbine aerodynamic performance testing device, comprising:
[0006] Mainstream high-temperature gas flow channel; A turbine rotor, wherein the blades of the turbine rotor are disposed in the mainstream high-temperature gas flow channel; A rotating shaft, one end of which is connected to the turbine rotor, and the other end of which is adapted to be connected to the dynamometer device of the test bench; A support structure surrounds the outside of the rotating shaft, and a bearing is provided between the support structure and the rotating shaft. The bearing is located downstream of the turbine rotor, and an oil lubrication cavity is formed between the support structure and the rotating shaft.
[0007] Beneficial Effects: The turbine rotor blades are positioned within the main high-temperature gas flow channel. As the high-temperature gas flows through this channel, it drives the turbine rotor to rotate. One end of the shaft is connected to the turbine rotor, and the other end is connected to the dynamometer on the test bench, allowing for the measurement of the turbine rotor's rotational speed. Since one end of the shaft is connected to the turbine rotor, a bearing is installed between the support structure and the shaft. The bearing is located downstream of the turbine rotor, and an oil chamber is formed between the support structure and the shaft. Therefore, the entire oil chamber is located downstream of the turbine rotor. Given the extremely high temperature at the inlet of the main high-temperature gas flow channel, the airflow temperature significantly decreases after the turbine rotor blades expand and perform work. Placing the bearing and oil chamber downstream of the turbine rotor keeps them away from the high-temperature core area, reducing direct heat radiation and structural heat conduction from the high-temperature gas flow, thus reducing heat input to the oil chamber at the source and preventing oil and bearing failure. Furthermore, placing the bearing and oil chamber downstream of the turbine rotor shortens the cantilever length of the shaft, improving system stability and providing structural assurance for experimental accuracy.
[0008] In one optional embodiment, a heat insulation cavity is provided between the mainstream high-temperature gas flow channel and the lubricating oil cavity.
[0009] Beneficial effects: By setting up a heat insulation cavity between the mainstream high-temperature gas flow channel and the lubricating oil cavity, the heat from the mainstream high-temperature gas flow channel can be further blocked from being conducted to the lubricating oil cavity, thereby preventing the failure of lubricating oil and bearings.
[0010] In one optional embodiment, the heat insulation cavity is provided with a cold air inlet and a cold air outlet, and the cold air inlet is connected to a directional airflow system.
[0011] Beneficial effects: The insulation cavity is equipped with a cold air inlet and a cold air outlet. The cold air inlet is connected to the directional flow system, so cold air can be introduced into the insulation cavity. The insulation cavity is located between the mainstream high-temperature gas flow channel and the lubricating oil cavity. Therefore, when the cold air flows through the insulation cavity, it can carry away the heat of the mainstream high-temperature gas flow channel and further prevent the heat of the mainstream high-temperature gas flow channel from being conducted to the lubricating oil cavity, thereby avoiding the failure of lubricating oil and bearings.
[0012] In one optional embodiment, the heat insulation cavity is provided with a heat insulation plate, which is parallel to the axis of the rotating shaft. The heat insulation plate divides the heat insulation cavity into a plurality of cooling channels distributed radially. Adjacent cooling channels are connected through a connecting port. The cold air inlet is connected to the innermost cooling channel, and the cold air outlet is connected to the outermost cooling channel.
[0013] Beneficial effects: The insulation chamber is equipped with an insulation plate, which is parallel to the axis of the rotating shaft. The insulation plate divides the insulation chamber into multiple cooling channels distributed radially. Adjacent cooling channels are connected through a connecting port. The cold air inlet is connected to the innermost cooling channel, and the cold air outlet is connected to the outermost cooling channel. Therefore, the cold air will first enter the innermost cooling channel through the cold air inlet, exchange heat with the lubricating oil chamber to remove the heat of the lubricating oil chamber, and then flow to the outside, finally flowing out through the outermost cooling channel. By setting multiple cooling channels and making the cold air flow through the set path, the temperature of the lubricating oil chamber can be cooled to the maximum extent, thereby maintaining the temperature of the lubricating oil chamber within the allowable range.
[0014] This design allows low-temperature cold air to flow from the inside to the outside, forming a reverse heat exchange with the temperature gradient of the insulation cavity (lower temperature inside, higher temperature outside), maximizing heat exchange efficiency and ensuring that the heat from the outer high-temperature insulation plate is carried away in time, preventing heat from being conducted to the lubricating oil cavity.
[0015] In one alternative implementation, the connecting openings on two adjacent insulation panels are offset along the axial direction.
[0016] Beneficial effects: By distributing the connecting ports on two adjacent heat insulation plates in an axially staggered manner, cold air can flow through each cooling channel in turn. By allowing the cold air to flow through the set path, the temperature of the lubricating oil cavity can be cooled to the maximum extent, thereby maintaining the temperature of the lubricating oil cavity within the allowable range.
[0017] In one alternative implementation, the cold air inlet is connected to the center of the innermost cooling channel.
[0018] Beneficial effects: The cold air inlet is connected to the center of the innermost cooling channel. After the cold air enters the innermost cooling channel, it flows to both sides, which can maximize the cooling of the lubricating oil cavity.
[0019] In one optional embodiment, a temperature sensor is provided in the lubricating oil chamber, and the directional flow guidance system includes a flow regulating valve and a controller, wherein the controller is communicatively connected to the temperature sensor and the flow regulating valve.
[0020] Beneficial effects: By installing a temperature sensor in the lubricating oil chamber, temperature changes can be monitored in real time. The controller is connected to the temperature sensor and the flow regulating valve, and can adjust the flow of cold air in real time according to the monitored temperature, thereby maintaining the temperature in the lubricating oil chamber within the permissible range.
[0021] In one optional embodiment, the mainstream high-temperature gas flow channel has an inner annular peripheral wall and an outer annular peripheral wall, and a first sealing assembly is provided between the turbine rotor and the inner annular peripheral wall. A second sealing assembly is provided between the connection between the turbine rotor and the shaft and the lubricating oil cavity; A third sealing assembly is provided between the end of the support structure away from the turbine rotor and the shaft.
[0022] Beneficial effects: The main high-temperature gas flow channel has an inner annular peripheral wall and an outer annular peripheral wall. A first sealing assembly is provided between the turbine rotor and the inner annular peripheral wall, which can prevent the high-temperature fluid in the main high-temperature gas flow channel from leaking out through the gap between the turbine rotor and the inner annular peripheral wall. A second sealing assembly is provided between the connection between the turbine rotor and the shaft and the lubricating oil cavity, which can prevent the high-temperature gas leaking from the gap between the turbine rotor and the inner annular peripheral wall from entering the lubricating oil cavity. A third sealing assembly is provided between the end of the support structure away from the turbine rotor and the shaft, which can prevent the lubricating oil in the lubricating oil cavity from leaking out.
[0023] In one optional embodiment, the inner annular peripheral wall includes a first peripheral wall segment and a second peripheral wall segment, the turbine rotor blades passing through the gap between the first peripheral wall segment and the second peripheral wall segment, and the first sealing assembly includes: A first connector is connected to the first peripheral wall section, and the first connector and the turbine form a first sealing cavity. The first connector is provided with a first groove and a first connecting part. A first protruding edge is provided on the side of the turbine rotor facing the first peripheral wall section, and the first protruding edge is inserted into the first groove; The second convex edge is provided on the side of the turbine rotor facing the second peripheral wall section, and the second convex edge extends to the outer peripheral side of the second peripheral wall section; The third convex edge is located on the side of the turbine rotor facing the first peripheral wall section, and a toothed sealing structure is formed between the third convex edge and the first connecting part.
[0024] Beneficial effects: The first convex edge is located on the side of the turbine rotor facing the first peripheral wall section. The first convex edge inserts into the first groove, preventing high-temperature gas in the main high-temperature gas flow channel from entering the first sealing cavity without affecting the normal rotation of the turbine rotor. The third convex edge forms a toothed sealing structure with the first connecting part, preventing the leakage of a small amount of high-temperature gas entering the first sealing cavity. The second convex edge is located on the side of the turbine rotor facing the second peripheral wall section, extending to the outer periphery of the second peripheral wall section. Without affecting the normal rotation of the turbine rotor, it can prevent the leakage of high-temperature gas in the main high-temperature gas flow channel.
[0025] In one alternative implementation, the second sealing component includes: The second connector is fixedly connected to the support structure. A second sealing cavity is formed between the second connector, the support structure, the second peripheral wall section and the turbine rotor. A third sealing cavity is formed between the second connector and the rotating shaft. The second connector is provided with a second connecting part and a third connecting part. A fourth protruding edge is provided on the side of the turbine rotor facing the support structure. A toothed sealing structure is formed between the fourth protruding edge and the second connecting part. One end of the rotating shaft is fixedly connected to the fourth protruding edge. A toothed sealing structure is formed between the third connecting part and the rotating shaft.
[0026] Beneficial effects: The second convex edge extends to the outer periphery of the second peripheral wall section, which can prevent the high-temperature gas in the mainstream high-temperature gas flow channel from leaking into the second sealing cavity without affecting the normal rotation of the turbine rotor. The fourth convex edge is located on the side of the turbine rotor facing the support structure. The fourth convex edge and the second connecting part form a grate sealing structure, which can prevent a small amount of gas entering the second sealing cavity from entering the third sealing cavity. The third connecting part and the rotating shaft form a grate sealing structure, which can prevent a small amount of gas entering the third sealing cavity from entering the lubricating oil cavity, and at the same time prevent the lubricating oil in the lubricating oil cavity from leaking. Attached Figure Description
[0027] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.
[0028] Figure 1 This is a schematic diagram of a cooling turbine aerodynamic performance testing device according to an embodiment of the present invention; Figure 2 for Figure 1 A magnified view of part A in the diagram.
[0029] Explanation of reference numerals in the attached figures: 1. Mainstream high-temperature gas flow channel; 101. Inner annular peripheral wall; 1011. First peripheral wall section; 1012. Second peripheral wall section; 102. Outer annular peripheral wall; 2. Turbine rotor; 201. First convex edge; 202. Second convex edge; 203. Third convex edge; 204. Fourth convex edge; 3. Rotating shaft; 4. Support structure; 5. Bearing; 6. Lubricating oil chamber; 7. Turbine guide; 8. First flange connection structure; 9. Second flange connection structure; 10. Insulation chamber; 1001. Cold air inlet; 1002. Cold air outlet; 11. Insulation plate; 1101. Connecting port; 12. First connecting piece; 1201. First groove; 1202. First connecting part; 13. First sealing chamber; 14. Second connecting piece; 1401. Second connecting part; 1402. Third connecting part; 15. Second sealing chamber; 16. Third sealing chamber; 17. Third connecting piece. Detailed Implementation
[0030] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.
[0031] In the description of this invention, it should be noted that the terms "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on the invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.
[0032] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" 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; and they can refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.
[0033] Furthermore, the technical features involved in the different embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.
[0034] The turbine is a core power component of an aero-engine, and its aerodynamic performance directly determines the engine's thrust, fuel efficiency, and reliability. To develop high-performance aero-engines, it is necessary to simulate real-world operating conditions through cooled turbine aerodynamic performance tests to analyze the impact of parameters such as cold gas injection method and flow rate on turbine efficiency and aerodynamic stability. In these tests, the high-temperature combustion gases (reaching 800℃-1200℃ or even higher) in the turbine's mains flow channel transfer heat to the lubricating oil chamber 6 of bearing 5 through heat conduction and radiation. This causes the lubricating oil temperature to exceed its allowable operating temperature (typically ≤150℃), leading to oil carbonization, viscosity reduction, and even failure. This, in turn, causes wear and jamming of bearing 5, seriously threatening test safety. Simultaneously, lubricating oil failure disrupts the stable operation of the turbine rotor 2, making it impossible to accurately simulate real engine operating conditions. This results in distorted research data on the impact of cold gas on turbine performance, making it difficult to accurately verify the turbine's true performance.
[0035] In related technologies, the test specimen structure is mostly arranged in a symmetrical layout. The two support points of the bearing 5 are located at the front and rear ends of the turbine rotor 2 blade. The airflow temperature at the front end of the turbine rotor 2 blade is the highest. After the high-temperature airflow pushes the turbine rotor 2 blade to do work, the temperature will decrease. Therefore, the bearing 5 and lubricating oil located at the front end of the turbine rotor 2 blade are subject to a large amount of heat conduction and heat radiation from the high-temperature airflow, which makes them prone to failure.
[0036] In addition, the relevant technology does not take into account heat insulation. The heat insulation between the high-temperature gas flow channel and the bearing 5 support structure 4 is mainly through the metal wall of the inner flow channel casing, which cannot support the long-term operation of the test. As time goes by, the heat of the mainstream high-temperature gas is gradually transferred to the bearing 5 and the heat radiation effect causes the temperature of the lubricating oil chamber 6 to gradually rise, eventually exceeding the lubricating oil operating temperature, resulting in lubricating oil failure and causing test risks.
[0037] Therefore, how to prevent the high-temperature combustion gas in the main channel from transferring heat to the lubricating oil chamber 6 has become a core technical challenge in the design of the cooling turbine test piece, directly affecting the feasibility of the test and the validity of the data. With the increase in the thrust-to-weight ratio of aero engines, the turbine inlet temperature continues to rise. If the thermal insulation fails, it will not only lead to damage to the test piece and a longer test cycle, but may also cause data distortion that misleads the direction of turbine design optimization, resulting in wasted R&D costs.
[0038] The following is combined Figures 1 to 2 The following describes embodiments of the present invention.
[0039] According to an embodiment of the present invention, a cooling turbine aerodynamic performance testing device is provided, comprising a main high-temperature gas flow channel 1, a turbine rotor 2, a rotating shaft 3, and a support structure 4.
[0040] Among them, the blades of the turbine rotor 2 are located in the main high-temperature gas flow channel 1; one end of the rotating shaft 3 is connected to the turbine rotor 2, and the other end is suitable for connecting to the dynamometer device of the test bench; the support structure 4 surrounds the outside of the rotating shaft 3, and a bearing 5 is provided between the support structure 4 and the rotating shaft 3. The bearing 5 is located downstream of the turbine rotor 2, and a lubricating oil cavity 6 is formed between the support structure 4 and the rotating shaft 3.
[0041] In this embodiment, the blades of the turbine rotor 2 are located in the main high-temperature gas flow channel 1. When the high-temperature gas passes through the main high-temperature gas flow channel 1, it drives the turbine rotor 2 to rotate. One end of the shaft 3 is connected to the turbine rotor 2, and the other end is connected to the dynamometer of the test bench, so the rotational speed of the turbine rotor 2 can be measured. Since one end of the shaft 3 is connected to the turbine rotor 2, a bearing 5 is provided between the support structure 4 and the shaft 3. The bearing 5 is located downstream of the turbine rotor 2, and a lubricating oil cavity 6 is formed between the support structure 4 and the shaft 3. Therefore, the entire lubricating oil cavity 6 is located downstream of the turbine rotor 2. The inlet temperature of the main high-temperature gas flow channel 1 is extremely high. After the blades of the turbine rotor 2 expand and do work, the airflow temperature will be significantly reduced. Placing the bearing 5 and the lubricating oil cavity 6 downstream of the turbine rotor 2 can keep the bearing 5 and the lubricating oil cavity 6 away from the high-temperature core area. From a spatial layout perspective, this reduces the direct heat radiation and structural heat conduction of the high-temperature gas, reduces the heat input of the lubricating oil cavity 6 from the source, and thus avoids the failure of the lubricating oil and the bearing 5. In addition, placing the bearing 5 and the lubricating oil chamber 6 downstream of the turbine rotor 2 can shorten the cantilever length of the shaft 3, improve the stability of the system, and provide structural assurance for the accuracy of the test.
[0042] In this embodiment, the bearing 5 and the lubricating oil chamber 6 are placed downstream of the turbine rotor 2, which eliminates the need to modify the core components of the turbine, thus simplifying the structure and eliminating the need for separate heat insulation and cooling of the front end support.
[0043] Specifically, the inlet temperature of the mainstream high-temperature gas flow channel 1 is extremely high, reaching 800℃-1200℃ or even higher. After the blades of the turbine rotor 2 expand and do work, the airflow temperature will be significantly reduced, with a reduction of 30%-40%.
[0044] In one embodiment, such as Figure 1 As shown, a turbine guide 7 is provided upstream of the turbine rotor 2. The turbine guide 7 is stationary and can pressurize the airflow. The pressurized airflow passes through the blades of the turbine rotor 2, thereby driving the turbine rotor 2 to rotate.
[0045] In one embodiment, the cooling turbine aerodynamic performance testing device includes a first flange connection structure 8 and a second flange connection structure 9. The first flange connection structure 8 and the second flange connection structure 9 are used to connect with the test bench, and the space between the first flange connection structure 8 and the second flange connection structure 9 is connected to the mainstream high-temperature gas flow channel 1.
[0046] In one embodiment, a heat insulation cavity 10 is provided between the main high-temperature gas flow channel 1 and the lubricating oil cavity 6.
[0047] In this embodiment, by providing a heat insulation cavity 10 between the mainstream high-temperature gas flow channel 1 and the lubricating oil cavity 6, the heat of the mainstream high-temperature gas flow channel 1 can be further blocked from being conducted to the lubricating oil cavity 6, thereby preventing the lubricating oil and bearing 5 from failing.
[0048] In this embodiment, by placing the bearing 5 and the lubricating oil chamber 6 downstream of the turbine rotor 2, and providing a heat insulation chamber 10 between the mainstream high-temperature gas flow channel 1 and the lubricating oil chamber 6, the temperature of the lubricating oil chamber 6 can be significantly reduced, thereby preventing the failure of the lubricating oil and the bearing 5.
[0049] In one embodiment, the heat insulation cavity 10 is provided with a cold air inlet 1001 and a cold air outlet 1002, and the cold air inlet 1001 is connected to a directional airflow system.
[0050] In this embodiment, the heat insulation cavity 10 is provided with a cold air inlet 1001 and a cold air outlet 1002. The cold air inlet 1001 is connected to the directional flow system, so cold air can be introduced into the heat insulation cavity 10. The heat insulation cavity 10 is located between the mainstream high-temperature gas flow channel 1 and the lubricating oil cavity 6. Therefore, when the cold air flows through the heat insulation cavity 10, it can carry away the heat of the mainstream high-temperature gas flow channel 1, further blocking the heat of the mainstream high-temperature gas flow channel 1 from being conducted to the lubricating oil cavity 6, thereby preventing the lubricating oil and bearing 5 from failing.
[0051] Specifically, the directional airflow system includes an external cold air source, a flow regulating valve, a temperature sensor, cold air inlet / outlet pipes, and an exhaust device. The external cold air source provides clean compressed cold air (which can be ambient temperature compressed air or -20℃ low temperature nitrogen). After the flow rate is precisely controlled by the flow regulating valve, the air is introduced into the heat insulation chamber 10 through the cold air inlet 1001.
[0052] In one embodiment, the heat insulation cavity 10 is provided with a heat insulation plate 11, which is parallel to the axis of the rotating shaft 3. The heat insulation plate 11 divides the heat insulation cavity 10 into a plurality of cooling channels distributed radially. Two adjacent cooling channels are connected through a connecting port 1101. The cold air inlet 1001 is connected to the innermost cooling channel, and the cold air outlet 1002 is connected to the outermost cooling channel.
[0053] In this embodiment, a heat insulation plate 11 is provided inside the heat insulation cavity 10. The heat insulation plate 11 is parallel to the axis of the rotating shaft 3. The heat insulation plate 11 divides the heat insulation cavity 10 into multiple cooling channels distributed radially. Two adjacent cooling channels are connected through a connecting port 1101. The cold air inlet 1001 is connected to the innermost cooling channel, and the cold air outlet 1002 is connected to the outermost cooling channel. Therefore, the cold air will first enter the innermost cooling channel through the cold air inlet 1001, exchange heat with the lubricating oil cavity 6 to remove the heat of the lubricating oil cavity 6, and then flow to the outside. Finally, it flows out through the outermost cooling channel. By setting multiple cooling channels and making the cold air flow through the set path, the lubricating oil cavity 6 can be cooled to the maximum extent, thereby keeping the temperature of the lubricating oil cavity 6 within the allowable range.
[0054] This design allows the low-temperature cold air to flow from the inside to the outside, forming a reverse heat exchange with the temperature gradient of the insulation cavity 10 (lower temperature inside, higher temperature outside), maximizing the heat exchange efficiency and ensuring that the heat of the outer high-temperature insulation plate 11 is carried away in time, preventing heat from being conducted to the lubricating oil cavity 6.
[0055] In one embodiment, the connecting openings 1101 on two adjacent heat insulation plates 11 are staggered along the axial direction.
[0056] In this embodiment, by distributing the connecting ports 1101 on two adjacent heat insulation plates 11 in an axially staggered manner, the cold air can flow through each cooling channel in turn, and the cold air can flow through the set path to maximize the cooling of the lubricating oil cavity 6, thereby keeping the temperature of the lubricating oil cavity 6 within the allowable range.
[0057] In one specific embodiment, each of the connecting ports 1101 is a ring.
[0058] Specifically in one embodiment, such as Figure 1 As shown, the heat insulation cavity 10 is provided with two heat insulation plates 11, which divide the heat insulation cavity 10 into three cooling channels. The connecting port 1101 on the inner heat insulation plate 11 is located at the right end, and the connecting port 1101 on the outer heat insulation plate 11 is located at the left end. The cold air inlet 1001 is connected to the innermost cooling channel. The cold air enters the innermost cooling channel through the cold air inlet 1001, exchanges heat with the lubricating oil cavity 6 to remove the heat of the lubricating oil cavity 6, and then flows radially outward through the connecting port 1101 to the middle cooling channel. In the middle cooling channel, it flows to the left, and then flows radially outward through the connecting port 1101 to the outermost cooling channel, and then flows out from the cold air outlet 1002.
[0059] In one embodiment, the cold air inlet 1001 is connected to the center of the innermost cooling channel.
[0060] In this embodiment, the cold air inlet 1001 is connected to the center of the innermost cooling channel. After the cold air enters the innermost cooling channel, it flows to both sides, which can cool the lubricating oil chamber 6 to the maximum extent.
[0061] In one embodiment not shown in the figure, the cold air inlet 1001 may also be connected to one end of the innermost cooling channel.
[0062] In one embodiment, the surface of the heat insulation panel 11 is polished.
[0063] In one embodiment, a temperature sensor is provided in the lubricating oil chamber 6, and the directional flow guidance system includes a flow regulating valve and a controller, with the controller communicating with the temperature sensor and the flow regulating valve.
[0064] In this embodiment, by installing a temperature sensor in the lubricating oil chamber 6, temperature changes can be monitored in real time. The controller is connected to the temperature sensor and the flow regulating valve, and can adjust the cold air flow in real time according to the monitored temperature, so that the temperature in the lubricating oil chamber 6 is maintained within the permissible range.
[0065] Specifically, when the temperature in the lubricating oil chamber 6 approaches the alarm value (e.g., 120°C), the controller controls the flow regulating valve to increase the cold air flow. When the temperature is below the safe value (e.g., 80°C), the controller controls the flow regulating valve to decrease the cold air flow, thereby achieving on-demand supply of cold air and balancing heat insulation and energy-saving requirements.
[0066] In this embodiment, a heat insulation cavity 10 is provided and a heat insulation plate 11 is provided inside the heat insulation cavity 10. Cold air is introduced into the heat insulation cavity 10. The high-efficiency heat insulation capability of the heat insulation cavity 10 supports the gas temperature in the mainstream high-temperature gas flow channel 1 to reach the actual level of the engine. It can accurately simulate the impact of cold air on turbine performance, provide a reliable basis for turbine design optimization, and solve the defect of existing technology that cannot simulate real high-temperature working conditions.
[0067] In one embodiment, the main high-temperature gas flow channel 1 has an inner annular peripheral wall 101 and an outer annular peripheral wall 102. A first sealing assembly is provided between the turbine rotor 2 and the inner annular peripheral wall 101. A second sealing assembly is provided between the connection between the turbine rotor 2 and the shaft 3 and the lubricating oil cavity 6. A third sealing assembly is provided between the end of the support structure 4 away from the turbine rotor 2 and the shaft 3.
[0068] In this embodiment, the main high-temperature gas flow channel 1 has an inner annular peripheral wall 101 and an outer annular peripheral wall 102. A first sealing assembly is provided between the turbine rotor 2 and the inner annular peripheral wall 101, which can prevent the high-temperature fluid in the main high-temperature gas flow channel 1 from leaking out through the gap between the turbine rotor 2 and the inner annular peripheral wall 101. A second sealing assembly is provided between the connection between the turbine rotor 2 and the shaft 3 and the lubricating oil cavity 6, which can prevent the high-temperature gas leaking from the gap between the turbine rotor 2 and the inner annular peripheral wall 101 from entering the lubricating oil cavity 6. A third sealing assembly is provided between the end of the support structure 4 away from the turbine rotor 2 and the shaft 3, which can prevent the lubricating oil in the lubricating oil cavity 6 from leaking out.
[0069] In one embodiment, the inner annular peripheral wall 101 includes a first peripheral wall segment 1011 and a second peripheral wall segment 1012, and the blades of the turbine rotor 2 pass through the gap between the first peripheral wall segment 1011 and the second peripheral wall segment 1012, such as Figure 2 As shown, the first sealing assembly includes a first connector 12, a first protruding edge 201, a second protruding edge 202, and a third protruding edge 203. The first connector 12 is connected to the first peripheral wall section 1011, and forms a first sealing cavity 13 between the first connector 12 and the turbine. The first connector 12 has a first groove 1201 and a first connecting portion 1202. The first protruding edge 201 is located on the side of the turbine rotor 2 facing the first peripheral wall section 1011 and is inserted into the first groove 1201. The second protruding edge 202 is located on the side of the turbine rotor 2 facing the second peripheral wall section 1012 and extends to the outer periphery of the second peripheral wall section 1012. The third protruding edge 203 is located on the side of the turbine rotor 2 facing the first peripheral wall section 1011, and forms a toothed sealing structure between the third protruding edge 203 and the first connecting portion 1202.
[0070] In this embodiment, the first protruding edge 201 is located on the side of the turbine rotor 2 facing the first peripheral wall section 1011. The first protruding edge 201 is inserted into the first groove 1201. Without affecting the normal rotation of the turbine rotor 2, it can prevent the high-temperature gas in the mainstream high-temperature gas flow channel 1 from entering the first sealing cavity 13. The third protruding edge 203 and the first connecting part 1202 form a toothed sealing structure, which can prevent a small amount of high-temperature gas entering the first sealing cavity 13 from leaking out. The second protruding edge 202 is located on the side of the turbine rotor 2 facing the second peripheral wall section 1012. The second protruding edge 202 extends to the outer peripheral side of the second peripheral wall section 1012. Without affecting the normal rotation of the turbine rotor 2, it can prevent the high-temperature gas in the mainstream high-temperature gas flow channel 1 from leaking out.
[0071] In one embodiment, the second sealing assembly includes a second connector 14 and a fourth protruding edge 204. The second connector 14 is fixedly connected to the support structure 4, and a second sealing cavity 15 is formed between the second connector 14, the support structure 4, the second peripheral wall section 1012, and the turbine rotor 2. A third sealing cavity 16 is formed between the second connector 14 and the rotating shaft 3. The second connector 14 is provided with a second connecting portion 1401 and a third connecting portion 1402. The fourth protruding edge 204 is provided on the side of the turbine rotor 2 facing the support structure 4, and a toothed sealing structure is formed between the fourth protruding edge 204 and the second connecting portion 1401. One end of the rotating shaft 3 is fixedly connected to the fourth protruding edge 204. A toothed sealing structure is formed between the third connecting portion 1402 and the rotating shaft 3.
[0072] In this embodiment, the second protruding edge 202 extends to the outer periphery of the second peripheral wall section 1012. Without affecting the normal rotation of the turbine rotor 2, it can prevent the high-temperature gas in the mainstream high-temperature gas flow channel 1 from leaking into the second sealing cavity 15. The fourth protruding edge 204 is located on the side of the turbine rotor 2 facing the support structure 4. A toothed sealing structure is formed between the fourth protruding edge 204 and the second connecting part 1401, which can prevent a small amount of gas entering the second sealing cavity 15 from entering the third sealing cavity 16. A toothed sealing structure is formed between the third connecting part 1402 and the rotating shaft 3, which can prevent a small amount of gas entering the third sealing cavity 16 from entering the lubricating oil cavity 6, and at the same time prevent the lubricating oil in the lubricating oil cavity 6 from leaking.
[0073] In one embodiment, the third sealing assembly includes a third connector 17, which is fixedly connected to the end of the support structure 4 away from the turbine rotor 2. The third connector 17 and the rotating shaft 3 form a toothed sealing structure to prevent the lubricating oil in the lubricating oil cavity 6 from leaking out.
[0074] Although embodiments of the invention have been described in conjunction with the accompanying drawings, those skilled in the art can make various modifications and variations without departing from the spirit and scope of the invention, and such modifications and variations all fall within the scope defined by this application.
Claims
1. A cooling turbine aerodynamic performance testing device, characterized in that, include: Mainstream high-temperature gas flow channel (1); Turbine rotor (2), the blades of the turbine rotor (2) are located in the main high-temperature gas flow channel (1). A rotating shaft (3), one end of which is connected to the turbine rotor (2), and the other end is adapted to be connected to the dynamometer of the test bench; A support structure (4) surrounds the outside of the rotating shaft (3), and a bearing (5) is provided between the support structure (4) and the rotating shaft (3). The bearing (5) is located downstream of the turbine rotor (2), and an oil lubrication chamber (6) is formed between the support structure (4) and the rotating shaft (3).
2. The cooling turbine aerodynamic performance testing device according to claim 1, characterized in that, A heat insulation cavity (10) is provided between the main high-temperature gas flow channel (1) and the lubricating oil cavity (6).
3. The cooling turbine aerodynamic performance testing device according to claim 2, characterized in that, The heat insulation cavity (10) is provided with a cold air inlet (1001) and a cold air outlet (1002), and the cold air inlet (1001) is connected to the directional airflow system.
4. The aerodynamic performance testing device for a cooling turbine according to claim 3, characterized in that, The heat insulation cavity (10) is provided with a heat insulation plate (11). The heat insulation plate (11) is parallel to the axis of the rotating shaft (3). The heat insulation plate (11) divides the heat insulation cavity (10) into multiple cooling channels distributed radially. Two adjacent cooling channels are connected through a connecting port (1101). The cold air inlet (1001) is connected to the innermost cooling channel, and the cold air outlet (1002) is connected to the outermost cooling channel.
5. The cooling turbine aerodynamic performance testing device according to claim 4, characterized in that, The connecting openings (1101) on two adjacent insulation boards (11) are staggered along the axial direction.
6. The cooling turbine aerodynamic performance testing device according to claim 4, characterized in that, The cold air inlet (1001) is connected to the center of the innermost cooling channel.
7. The cooling turbine aerodynamic performance testing apparatus according to any one of claims 3 to 6, characterized in that, A temperature sensor is provided in the lubricating oil chamber (6). The directional flow system includes a flow regulating valve and a controller. The controller is communicatively connected to the temperature sensor and the flow regulating valve.
8. The cooling turbine aerodynamic performance testing apparatus according to any one of claims 1 to 6, characterized in that, The main high-temperature gas flow channel (1) has an inner annular peripheral wall (101) and an outer annular peripheral wall (102), and a first sealing assembly is provided between the turbine rotor (2) and the inner annular peripheral wall (101); A second sealing assembly is provided between the connection between the turbine rotor (2) and the shaft (3) and the lubricating oil chamber (6); A third sealing assembly is provided between the end of the support structure (4) away from the turbine rotor (2) and the shaft (3).
9. The aerodynamic performance testing apparatus for a cooling turbine according to claim 8, characterized in that, The inner annular peripheral wall (101) includes a first peripheral wall segment (1011) and a second peripheral wall segment (1012). The blades of the turbine rotor (2) pass through the gap between the first peripheral wall segment (1011) and the second peripheral wall segment (1012). The first sealing assembly includes: The first connector (12) is connected to the first peripheral wall section (1011), and the first connector (12) and the turbine form a first sealing cavity (13). The first connector (12) is provided with a first groove (1201) and a first connecting part (1202). The first protruding edge (201) is provided on the side of the turbine rotor (2) facing the first peripheral wall section (1011), and the first protruding edge (201) is inserted into the first groove (1201). The second protruding edge (202) is provided on the side of the turbine rotor (2) facing the second peripheral wall section (1012), and the second protruding edge (202) extends to the outer peripheral side of the second peripheral wall section (1012); The third protruding edge (203) is provided on the side of the turbine rotor (2) facing the first peripheral wall section (1011), and the third protruding edge (203) and the first connecting part (1202) form a toothed sealing structure.
10. The aerodynamic performance testing apparatus for a cooling turbine according to claim 9, characterized in that, The second sealing component includes: The second connector (14) is fixedly connected to the support structure (4). A second sealing cavity (15) is formed between the second connector (14), the support structure (4), the second peripheral wall section (1012) and the turbine rotor (2). A third sealing cavity (16) is formed between the second connector (14) and the rotating shaft (3). The second connector (14) is provided with a second connecting part (1401) and a third connecting part (1402). The fourth protruding edge (204) is provided on the side of the turbine rotor (2) facing the support structure (4). The fourth protruding edge (204) and the second connecting part (1401) form a toothed sealing structure. One end of the rotating shaft (3) is fixedly connected to the fourth protruding edge (204). A toothed sealing structure is formed between the third connecting part (1402) and the rotating shaft (3).