A high pressure turbine guide vane cooling structure and an aeroengine

Abstract

The application provides a high-pressure turbine guide vane cooling structure and an aero-engine, and belongs to the technical field of aero-engines.The high-pressure turbine guide vane cooling structure comprises: a high-pressure turbine guide vane, which has a cavity inside, and has film holes on the basin side and the back side, which are connected with the cavity, and has a split gap at the tail; an impact guide pipe is arranged in the cavity, and a gas flow channel is formed between the body part of the impact guide pipe and the cavity, the tail of the impact guide pipe has a convex bump, the convex bump is in close contact with the cavity so as to separate the gas flow channel, the body part of the impact guide pipe is respectively provided with impact cooling holes which are connected with the gas flow channels on the basin side and the back side, and the convex bump is provided with a tail impact hole which is connected with the split gap.The high-pressure turbine guide vane cooling structure provided by the application realizes the parallel flow path design of the basin side, the back side and the tail edge, solves the problem of insufficient tail edge outflow under the condition of low expansion ratio through fine flow path control, and guarantees the cold gas consumption and the cold gas quality per unit area.

Description

Technical Field

[0001] This application belongs to the field of aero-engines, and specifically relates to a high-pressure turbine guide vane cooling structure and an aero-engine. Background Technology

[0002] Short takeoff and vertical landing (STOVL) aircraft place unique demands on engine performance. To meet these requirements, the overall performance needs to be adaptively adjusted in terms of the matching of various engine components. For high-pressure turbines, compared to existing models, the main change lies in the reduced expansion ratio and increased reaction force. This change directly leads to increased pressure at the trailing edge of the high-pressure turbine guide vanes, which reduces the flow rate of cool air, thus affecting the cooling of the trailing edge and the rear chamber. Engineering practice shows that the existing cooling structure design for the rear chamber and trailing edge of high-pressure turbine guide vanes cannot meet the cooling requirements of airfoils with low expansion ratios and high reaction forces, necessitating the development of a completely new cooling structure tailored to the characteristics of the airfoil. Summary of the Invention

[0003] The purpose of this application is to provide a high-pressure turbine guide vane cooling structure and an aero-engine to solve or mitigate at least one of the problems in the prior art.

[0004] The technical solution of this application is: a cooling structure for a high-pressure turbine guide vane, comprising:

[0005] A high-pressure turbine guide vane has an internal cavity, and the vane has film vents communicating with the cavity on its basin side and back side, and a tail slit at its tail end; and

[0006] An impact duct is disposed within a cavity, and an airflow channel is formed between the body portion of the impact duct and the cavity. The tail portion of the impact duct has a bulge that is in close contact with the cavity, thereby isolating the airflow channel. The body portion of the impact duct is provided with impact cooling holes that communicate with the airflow channels on the basin side and the back side, respectively. The bulge is provided with a tail impact hole that communicates with the tail slit.

[0007] In at least one embodiment of this application, a gap is provided between the convex bulge and the tail slot inlet.

[0008] In at least one embodiment of this application, the gap is the same as or similar to the distance between the body portion of the impact duct and the inner wall surface of the high-pressure turbine guide vane.

[0009] In at least one embodiment of this application, the axis of the tail impact hole on the convex bulge is collinear with the axis of the tail wedge.

[0010] In at least one embodiment of this application, the convex hull and the body portion have a smooth transition.

[0011] In at least one embodiment of this application, the impact conduit is made of a high-temperature alloy or composite material.

[0012] On the other hand, the technical solution provided in this application is: an aero-engine, the aero-engine including the high-pressure turbine guide vane cooling structure as described above.

[0013] The high-pressure turbine guide vane cooling structure provided in this application realizes the parallel flow path design of the basin side, back side, and trailing edge. Through refined flow path control, it solves the problem of insufficient trailing edge outflow under low expansion ratio conditions, ensures the amount and quality of cold air per unit area of ​​each part, and solves the cooling design difficulties caused by low cold air quality and insufficient trailing edge outflow. Attached Figure Description

[0014] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly described below. Obviously, the drawings described below are merely some embodiments of this application.

[0015] Figure 1 This is a schematic diagram of a typical high-pressure turbine guide vane cooling structure.

[0016] Figure 2 This is a schematic diagram of the cooling structure for the high-pressure turbine guide vanes of this application. Detailed Implementation

[0017] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings.

[0018] like Figure 1 The diagram shows a typical cooling structure for a high-pressure turbine guide vane. This cooling structure 100 includes a high-pressure turbine guide vane 110 and an impact duct 120. The high-pressure turbine guide vane 110 has an internal cavity, and the impact duct 120 is disposed within this cavity. Its shape is basically the same as the cavity, only smaller in size. A roughly uniform gap 130 is provided between the two, forming an airflow channel. The wall of the impact duct 120 has impact holes 121 for impact cooling of the blade of the high-pressure turbine guide vane 110. The wall of the high-pressure turbine guide vane 110 has film cooling holes 111, and its tail end has a tail slit 112. Part of the cooling airflow from the impact duct 120 flows out through the film cooling holes 111, forming film cooling on the surface of the high-pressure turbine guide vane 110; the other part flows out along the airflow channel from the tail slit 112.

[0019] However, in the existing cooling structure 100, the cold air outflow reaches the critical velocity at the trailing edge of the high-pressure turbine guide vane 110, and the cold air flow rate reaches the maximum flow rate allowed by the structure. After the aerodynamic parameters are adjusted, the expansion ratio decreases, and the cold air outflow rate cannot reach the critical velocity, reducing the cold air outflow limit. In the existing technology, the end-throttling of the flow path causes insufficient cooling of the entire flow path. In addition, the cold air impacts and cools the inner wall surface of the high-pressure turbine guide vane 110 through the impact holes 121 on the basin side and back side of the impact duct 120, respectively. Then, the two streams of cooling air flow towards the trailing edge and converge at the inlet of the trailing edge wedge 112 to cool the trailing edge wedge 112. However, for the trailing edge wedge 112, the incoming cold air has already undergone sufficient heat exchange with the inner wall surface of the basin and back, resulting in a large temperature rise and a significant decrease in cold air quality, which is detrimental to the cooling of the trailing edge wedge 112.

[0020] Therefore, in order to overcome the above-mentioned defects, this application provides a new cooling structure for high-pressure turbine guide vanes.

[0021] like Figure 2 As shown, the high-pressure turbine guide vane cooling structure 200 provided in this application includes a high-pressure turbine guide vane 210 and an impact duct 220. The turbine guide vane 210 has an internal cavity, and the impact duct 220 is disposed within this cavity. The body portion of the impact duct 220 is adapted to the shape of the cavity, and its body portion is slightly smaller than the cavity, thereby forming an airflow channel between the impact duct 220 and the inner wall surface of the high-pressure turbine guide vane 210. The body portion has impact cooling holes 221. The rear end of the body portion of the impact duct 220 has a protrusion 222, which tightly fits against the rear end of the cavity of the high-pressure turbine guide vane 210 to form a closed structure, separating the cavities on the front and rear sides of the protrusion 222. The protrusion 222 has a tail impact hole 223. The basin-side and back-side walls of the high-pressure turbine guide vane 210 have film gas holes 212, and its tail edge has a tail slit 211. Part of the cooling gas in the impact duct 220 impacts and cools the inner wall of the high-pressure turbine guide vane 210 through the impact cooling hole 221 of the main body, and then flows out through the film cooling hole 212 on the high-pressure turbine guide vane 210; another part of the cooling gas in the impact duct 220 flows out through the tail impact hole 223 on the bulge 222 and flows out along the tail slit 211 of the high-pressure turbine guide vane 210.

[0022] The high-pressure turbine guide vane cooling structure provided in this application reorganizes the cooling flow path of the rear cavity and the tail slit 211 of the high-pressure turbine guide vane 210. The cold gas outflow through the impact duct 220 can be divided into three parts: basin-side outflow, back-side outflow, and trailing edge outflow. The basin-side outflow is responsible for cooling the inner wall of the basin side of the rear cavity, and the cooled gas is discharged into the main channel through the gas film hole 212 on the basin side. The back-side outflow is responsible for cooling the inner wall of the back side of the rear cavity, and the cooled gas is discharged into the main channel through the gas film hole 212 on the back side. The trailing edge outflow is responsible for supplying gas to the tail slit 211 separately. These three airflows are isolated from each other by the protrusion 222 at the trailing edge of the impact duct 220, which prevents the high-temperature gas cooled on the back side of the basin from merging into the trailing edge gas supply, thus avoiding a decrease in the quality of the trailing edge cold gas supply. At the same time, the separate gas supply of the three parts can effectively ensure the amount of cold gas distributed per unit area in each part, making it easier to control the cold gas more precisely. This application solves the problem of insufficient cooling of the flow path caused by the low cold air output of the tail split, and the problem of difficult tail split cooling design caused by low cold air quality.

[0023] In some embodiments of this application, the impact conduit 220 is made of high-temperature alloy or composite material. For example, the impact conduit 220 may be made of nickel-based high-temperature alloy material or ceramic matrix composite material.

[0024] In some embodiments of this application, the convex hull 222 and the body portion are transitioned by a smooth structure, thereby avoiding stress concentration in the impact conduit 220 and improving the durability of the impact conduit 220. For example, the smooth structure can be an arc-shaped structure or a chamfered structure, etc.

[0025] In this application, there is a certain gap between the tail of the convex bulge 222 and the inlet of the tail slit 211, and the cooling gas flowing out from the tail impact hole 223 of the convex bulge 222 can stay in this gap for cooling. For example, this gap is approximately the same as the distance between the body of the impact duct 220 and the inner wall surface of the high-pressure turbine guide vane 210.

[0026] Furthermore, the axis of the tail impact hole 223 on the convex bulge 222 is approximately collinear with the tail wedge 211, thereby allowing the cooling gas in the impact guide 210 to smoothly enter the tail wedge 211, thus reducing pressure loss.

[0027] Finally, this application also provides an aircraft engine that includes the aforementioned high-pressure turbine guide vane cooling structure.

[0028] In the high-pressure turbine guide vane cooling structure of this application, a bulge is provided at the tail of the impact duct, which not only serves to separate the flow path but also increases the outflow area of ​​the trailing edge gas, ensuring the supply pressure of the tail slit 211. Compared to a structure without a bulge, the trailing edge of the high-pressure turbine guide vane 210 would require a shorter impact duct and a longer tail slit for the same outflow area. A longer tail slit is less conducive to the design of the cooling structure.

[0029] In summary, simply increasing the throttling area of ​​the cold air outlet still cannot increase the cold air volume to the required value. It is necessary to readjust the cold air flow path to solve the problem of insufficient cold air output from the tail slit. Furthermore, the series flow path will cause a large temperature rise in the cold air flowing into the tail slit, which brings great difficulty to the cooling design of the tail slit—especially the problem of not being able to improve the quality of the cold air entering the tail slit when the cold air output is limited.

[0030] The high-pressure turbine guide vane cooling structure provided in this application realizes the parallel flow path design of the basin side, back side, and trailing edge. Through refined flow path control, it solves the problem of insufficient trailing edge outflow under low expansion ratio conditions, ensures the amount and quality of cold air per unit area of ​​each part, and solves the cooling design difficulties caused by low cold air quality and insufficient trailing edge outflow.

[0031] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A cooling structure (200) for a high-pressure turbine guide vane, characterized in that, include: The high-pressure turbine guide vane (210) has a cavity inside, and the high-pressure turbine guide vane (210) has a film vent (212) on the basin side and the back side that communicates with the cavity, and has a tail slit (211) at its tail end. and An impact duct (220) is disposed within a cavity, and an airflow channel is formed between the body portion of the impact duct (220) and the cavity. The tail portion of the impact duct (220) has a bulge (222), which is in close contact with the cavity to isolate the airflow channel. The body portion of the impact duct (220) is provided with impact cooling holes (221) communicating with the airflow channels on the basin side and the back side, respectively. The bulge (222) is provided with... There is a tail impact hole (223) communicating with the tail wedge (211). The axis of the tail impact hole (223) on the bulge (222) is collinear with the axis of the tail wedge (211). There is a gap between the bulge (222) and the inlet of the tail wedge (211). The gap is the same as or similar to the distance between the body part of the impact guide tube (220) and the inner wall surface of the high-pressure turbine guide vane (210). The bulge (222) and the body part have a smooth transition.

2. The high-pressure turbine guide vane cooling structure as described in claim 1, characterized in that, The impact conduit (220) is made of high-temperature alloy or composite material.

3. An aircraft engine, characterized in that, The aero-engine includes a high-pressure turbine guide vane cooling structure as described in any one of claims 1 to 2.

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