High-pressure turbine guide vane cooling structure and aero-engine

Abstract

The invention 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, the interior of the high-pressure turbine guide vane is provided with a cavity, the basin side and the back side of the high-pressure turbine guide vane are provided with film holes communicating with the cavity, and the tail of the high-pressure turbine guide vane is provided with a tail crack; the impact guide pipe is arranged in the cavity, an airflow channel is formed between the body part of the impact guide pipe and the cavity, the tail part of the impact guide pipe is provided with a convex hull, the convex hull is in close contact with the cavity so as to separate the airflow channel, and the body part of the impact guide pipe is respectively provided with impact cooling holes communicated with the airflow channels on the basin side and the back side; and a tail impact hole communicated with the tail crack is formed in the convex hull. According to the high-pressure turbine guide vane cooling structure, the parallel flow path design of the basin side, the back side and the tail edge is achieved, through refined flow path control, the problem that the outflow amount of the tail edge is insufficient under the condition of the low expansion ratio is solved, and the cold air consumption and the cold air quality of each part per unit area are guaranteed.

Description

Technical Field

[0001] The present application relates to the field of aero-engines, and in particular to a high-pressure turbine guide vane cooling structure and an aero-engine. Background Art

[0002] Short-takeoff and vertical landing (STOVL) aircraft place special demands on engine performance. To meet these requirements, overall performance requires adaptive adjustments to the matching of various engine components. For the high-pressure turbine, the main change compared to existing models is a lower expansion ratio and higher reaction force. This change directly leads to increased pressure at the trailing edge of the high-pressure turbine guide vanes, which reduces the outflow of cold air, thus affecting the cooling of the trailing edge and rear cavity. Engineering practice has shown that the existing cooling structure design for the rear cavity and trailing edge of the high-pressure turbine guide vanes cannot meet the cooling requirements of low-expansion-ratio, high-reaction-force blade profiles, necessitating the development of a completely new cooling structure tailored to the characteristics of the blade profile. Summary of the Invention

[0003] The purpose of the present application is to provide a high-pressure turbine guide vane cooling structure and an aircraft engine to solve or alleviate at least one problem in the background technology.

[0004] The technical solution of the present application is: a high-pressure turbine guide blade cooling structure, comprising:

[0005] A high-pressure turbine guide blade having a cavity therein, and having air film holes on the basin side and the back side thereof communicating with the cavity, and a tail cleavage slot at the tail end; and

[0006] An impingement duct is arranged in a cavity, and an air flow channel is formed between the main body of the impingement duct and the cavity. The tail of the impingement duct has a bulge, which is in close contact with the cavity to isolate the air flow channel. The main body of the impingement duct is respectively provided with impingement cooling holes that are connected to the air flow channels on the basin side and the back side, and the bulge is provided with a tail impingement hole that is connected to the tail split seam.

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

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

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

[0010] In at least one embodiment of the present application, there is a smooth transition between the convex hull and the main body portion.

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

[0012] On the other hand, the technical solution provided by the present application is: an aircraft engine, comprising any of the high-pressure turbine guide blade cooling structures described above.

[0013] The high-pressure turbine guide blade cooling structure provided in the present application realizes a parallel flow path design of the basin side, back side, and trailing edge. Through refined flow path control, it solves the problem of insufficient outflow at the trailing edge under low expansion ratio conditions, ensures the cooling air consumption and quality per unit area of ​​each part, and solves the cooling design difficulties caused by low cooling air quality and insufficient outflow at the tail split. BRIEF DESCRIPTION OF THE DRAWINGS

[0014] In order to more clearly illustrate the technical solutions provided by this application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of this application.

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

[0016] Figure 2 This is a schematic diagram of the cooling structure of the high-pressure turbine guide blades of this application. DETAILED DESCRIPTION

[0017] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below in conjunction with the drawings in the embodiments of this application.

[0018] like Figure 1 The schematic diagram of a typical high-pressure turbine guide vane cooling structure is shown. This cooling structure 100 includes a high-pressure turbine guide vane 110 and an impingement duct 120. A cavity is defined within the high-pressure turbine guide vane 110. Impingement duct 120 is located within this cavity, having a substantially identical shape to the cavity, but smaller in size. A substantially uniform gap 130 is defined between the cavity and the impingement duct, forming an airflow channel. Impingement holes 121 are defined on the wall of impingement duct 120 for impingement cooling the blade body of high-pressure turbine guide vane 110. Film holes 111 are defined on the wall of high-pressure turbine guide vane 110, and a tail slit 112 is defined at the tail of high-pressure turbine guide vane 110. A portion of the cooling airflow flowing out of impingement duct 120 flows through film holes 111, forming film cooling on the surface of high-pressure turbine guide vane 110. The remaining portion flows along the airflow channel and out of tail slit 112.

[0019] However, in the existing cooling structure 100, at the trailing edge of the high-pressure turbine guide vane 110, the cold air outflow reaches a critical velocity, and the cold air flow rate reaches the maximum flow rate allowed by the structure. However, after adjusting the aerodynamic parameters, the expansion ratio is reduced, and the cold air outflow rate cannot reach the critical velocity, reducing the cold air outflow rate restriction. In the prior art, the use of end-throttling in the flow path results in insufficient cooling of the entire flow path. Furthermore, the cold air impinges on the inner wall surface of the high-pressure turbine guide vane 110 through the impingement holes 121 on the basin side and the back side of the impingement duct 120. The two streams of cooling air then flow toward the trailing edge and converge at the inlet of the trailing edge's tail split 112, cooling the tail split 112. However, for the tail split 112, the incoming cold air has already fully exchanged heat with the inner wall surface of the basin back, resulting in a large temperature rise in the cold air, significantly reducing the cooling quality, and adversely affecting the cooling of the trailing edge's tail split 112.

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

[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 impingement duct 220. A cavity is defined within the turbine guide vane 210, and the impingement duct 220 is disposed within this cavity. The main body of the impingement duct 220 conforms to the shape of the cavity and is slightly smaller than the cavity, thereby forming an airflow channel between the impingement duct 220 and the inner wall of the high-pressure turbine guide vane 210. Impingement cooling holes 221 are provided in the main body. A convex hump 222 is provided at the rear end of the main body of the impingement duct 220. The hump 222 fits tightly with the rear end of the cavity of the high-pressure turbine guide vane 210, forming a closed structure and separating the cavity in front of and behind the hump 222. The hump 222 is provided with a tail impingement hole 223. Film holes 212 are provided on the basin and back walls of the high-pressure turbine guide vane 210, and a tail slit 211 is provided at the trailing edge. A portion of the cooling gas in the impact duct 220 impacts the inner wall surface of the high-pressure turbine guide blade 210 from the impact cooling holes 221 in the main body, and flows out from the air film holes 212 on the high-pressure turbine guide blade 210 after cooling; another portion of the cooling gas in the impact duct 220 flows out from the tail impact holes 223 on the bulge 222, and flows out along the tail split slit 211 of the high-pressure turbine guide blade 210.

[0022] The high-pressure turbine guide vane cooling structure provided in this application reorganizes the cooling flow paths of the rear cavity and tail split slit 211 of the high-pressure turbine guide vane 210. The cold air outflow through the impact duct 220 can be divided into three parts: the basin-side outflow, the back-side outflow, and the trailing edge outflow. The basin-side outflow is responsible for cooling the inner wall of the basin side of the rear cavity. The cooled gas is discharged into the main channel through the air film holes 212 on the basin side. The back-side outflow is responsible for cooling the inner wall of the back cavity. The cooled gas is discharged into the main channel through the air film holes 212 on the back side. The trailing edge outflow is responsible for supplying air separately to the tail split slit 211. These three airflows are isolated from each other by the convex bulge 222 at the trailing edge of the impact duct 220, preventing the high-temperature gas cooled on the back side of the basin from merging into the trailing edge air supply, which would reduce the quality of the cold air supply at the trailing edge. Furthermore, the three separate air supply parts can effectively ensure the cold air distribution per unit area of ​​each part, facilitating more refined control of the cold air. The present application solves the problem of insufficient cooling of the flow path caused by the small outflow of cold air from the tail split joint, and the problem of difficulty in cooling design of the tail split joint caused by the low quality of cold air.

[0023] In some embodiments of the present application, the impact tube 220 is made of a high-temperature alloy or a composite material. For example, the impact tube 220 can be made of a nickel-based high-temperature alloy material, or a ceramic-based composite material.

[0024] In some embodiments of the present application, the convex bulge 222 and the main body are transitioned with a smooth structure, thereby avoiding stress concentration in the shock tube 220 and improving the durability of the shock tube 220. For example, the smooth structure can be an arc structure or a chamfered structure.

[0025] In the present application, a certain gap is provided between the tail of the convex hump 222 and the inlet of the tail split slit 211. The cooling gas flowing out of the impingement hole 223 at the tail of the convex hump 222 can be retained in this gap for cooling. For example, this gap is approximately the same as the distance between the main body of the impingement duct 220 and the inner wall of the high-pressure turbine guide vane 210.

[0026] Furthermore, the axis of the tail impact hole 223 on the convex hump 222 is substantially colinear with the tail split slit 211 , so that the cooling gas in the impact duct 210 can smoothly enter the tail split slit 211 to reduce pressure loss.

[0027] Finally, the present application also provides an aircraft engine, which includes the above-mentioned high-pressure turbine guide blade cooling structure.

[0028] The high-pressure turbine guide vane cooling structure of the present application utilizes a convex hull at the tail of the impingement duct, which not only separates the flow path but also increases the outflow area of ​​the trailing-edge gas, ensuring the gas supply pressure at the trailing cleavage 211. Compared to a structure without the convex hull, the trailing edge of the high-pressure turbine guide vane 210 requires a shorter impingement duct and a longer trailing cleavage to maintain the same outflow area. However, a longer trailing cleavage degrades the cooling structure design.

[0029] In summary, simply increasing the throttling area at the air outlet still fails to increase the cooling air volume to the required value. It is necessary to readjust the cooling air flow path to solve the problem of insufficient cooling air outflow from the tail split. Furthermore, the series flow path causes a large temperature rise in the cooling air flowing into the tail split, which brings great difficulties to the cooling design of the tail split. In particular, the problem of not being able to improve the cooling air quality entering the tail split when the cooling air outflow is limited is caused.

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

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

Claims

1. A high-pressure turbine guide vane cooling structure (200), characterized in that: include: A high-pressure turbine guide blade (210) has a cavity inside, and the basin side and the back side of the high-pressure turbine guide blade (20) have air film holes (212) communicating with the cavity, and the tail portion has a tail splitting seam (211); and An impact duct (220) is provided in a cavity, and an air flow channel is formed between the main body of the impact duct (220) and the cavity. The tail of the impact duct (220) has a convex bulge (222), and the convex bulge (222) is in close contact with the cavity to isolate the air flow channel. The main body of the impact duct (220) is respectively provided with impact cooling holes (221) that are connected to the air flow channels on the basin side and the back side, and the convex bulge (222) is provided with a tail impact hole (223) that is connected to the tail split seam (211).

2. The high-pressure turbine guide vane cooling structure according to claim 1, wherein: A gap is provided between the convex bulge (222) and the inlet of the tail split seam (211).

3. The high-pressure turbine guide vane cooling structure according to claim 2, wherein: The gap is the same as or similar to the distance between the main body of the impact duct (220) and the inner wall surface of the high-pressure turbine guide blade (210).

4. The high-pressure turbine guide vane cooling structure according to claim 1, wherein: The axis of the tail impact hole (223) on the convex bump (220) is collinear with the axis of the tail split seam (112).

5. The high-pressure turbine guide vane cooling structure according to claim 1, wherein: There is a smooth transition between the convex hull (222) and the main body portion.

6. The high-pressure turbine guide vane cooling structure according to any one of claims 1 to 5, characterized in that: The impact conduit (220) is made of high-temperature alloy or composite material.

7. An aircraft engine, characterized in that: The aircraft engine comprises the high-pressure turbine guide vane cooling structure according to any one of claims 1 to 6.

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