Turbine and aircraft engines

Abstract

The present invention discloses a turbine and an aircraft engine, which relate to the field of aircraft engines and are used to improve the sealing effect of the turbine. The turbine provided by the embodiment of the present invention includes rotor blades and stator blades. The rotor blades have a first edge plate. The stator blades have a second edge plate, and the second edge plate and the first edge plate overlap in the axial direction of the turbine, and there is a gap between the two in the radial direction of the turbine. Among them, a concave portion is provided on the side of the second edge plate away from the first edge plate. The turbine provided by the above technical solution has rotor blades and stator blades, and the first edge plate of the rotor blade and the second edge plate of the stator blade form a channel for sealing. The second edge plate is provided with a concave portion to increase the cold air flow area at the channel and improve the gas intrusion phenomenon caused by the circumferential unevenness of the airflow at the rim seal outlet caused by factors such as the stagnation of the guide vane leading edge.

Description

Technical Field

[0001] The present invention relates to the field of aeroengines, and in particular to a turbine and an aeroengine. Background Art

[0002] The gas turbine engine includes a compressor, a combustion chamber and a turbine. The compressor compresses the gas, the combustion chamber burns to provide energy, and the turbine expands and does work to drive the compressor. Among them, the turbine has several stages, and each stage includes stator blades and rotor blades. The stator blades are fixed on the turbine casing, and the turbine rotor blades are fixed on the turbine disk. The turbine blades are in the mainstream gas environment of the engine, and the temperature resistance level is relatively high. However, the components other than the blades, especially the turbine disk, have a lower temperature resistance level and need to be cooled by cooling air. A rim sealing structure is designed at the rim of the root of the turbine blade to prevent the mainstream gas from entering the interior and causing overheating of components such as the turbine disk.

[0003] The inventors have found that there are at least the following problems in the prior art: In the related art, cooling is achieved by introducing cooling gas into the sealing channel, but this sealing structure is at risk of overheating due to the intrusion of mainstream gas during use, and the cooling effect is poor. Summary of the invention

[0004] The present invention provides a turbine and an aircraft engine, which are used to improve the sealing effect of the turbine.

[0005] An embodiment of the present invention provides a turbine, comprising:

[0006] A rotor blade having a first flange; and

[0007] The stator blade has a second edge plate, wherein the second edge plate overlaps with the first edge plate in the axial direction of the turbine, and there is a gap between the second edge plate and the first edge plate in the radial direction of the turbine;

[0008] Wherein, a concave portion is arranged on a side of the second edge plate away from the first edge plate.

[0009] In some embodiments, the rotor blades and the stator blades are arranged in one-to-one correspondence, and the second edge plate of each stator blade is provided with the inner recess.

[0010] In some embodiments, the second edge plate is provided with one or more recessed portions.

[0011] In some embodiments, each of the rotor blades is provided with a first edge plate, and in the radial direction of the turbine, the first edge plate is located on the side of the second edge plate close to the central axis of the turbine, and the inner recess is provided on the side of the second edge plate facing the first edge plate.

[0012] In some embodiments, each of the rotor blades is provided with two first edge plates; in the radial direction of the turbine, the second edge plate is located between the two first edge plates.

[0013] In some embodiments, the inner recess is provided on a side of the second edge plate facing the central axis of the turbine.

[0014] In some embodiments, the inner recess is disposed on a side of the second edge plate facing the central axis of the turbine, and the inner recess is also disposed on a side of the second edge plate facing away from the central axis of the turbine.

[0015] In some embodiments, the turbine further comprises:

[0016] The air collecting cavity is located on the side of the stator blade facing the central axis of the turbine, and the wall of the air collecting cavity is provided with air holes, which are communicated with the gap.

[0017] In some embodiments, the turbine further comprises:

[0018] The comb teeth sealing assembly is arranged on a side of the air collecting cavity away from the stator blades.

[0019] An embodiment of the present invention further provides an aircraft engine, comprising a turbine provided by any technical solution of the present invention.

[0020] The turbine provided by the above technical solution has rotor blades and stator blades, and the first edge plate of the rotor blades and the second edge plate of the stator blades form a channel for sealing. The second edge plate is provided with an inner concave portion to increase the cold air flow area at the channel and improve the gas intrusion phenomenon caused by the circumferential non-uniformity of the airflow at the wheel rim seal outlet caused by factors such as the stagnation of the guide vane leading edge. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] The drawings described herein are used to provide a further understanding of the present invention and constitute a part of this application. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings:

[0022] Figure 1 A schematic diagram of a local structure of a turbine provided in an embodiment of the present invention;

[0023] Figure 2 A schematic diagram of pressure distribution at the sealing part of a turbine provided by an embodiment of the present invention;

[0024] Figure 3 A schematic diagram of the structure of a second turbine edge plate provided in an embodiment of the present invention;

[0025] Figure 4 for Figure 3 AA cross-sectional view of . DETAILED DESCRIPTION

[0026] Combine the following Figure 1 to Figure 4 The technical solution provided by the present invention is described in more detail.

[0027] The inventor has found that in the existing wheel rim sealing technology, due to the stagnation effect on the leading edge of the turbine stator blades, the circumferential pressure distribution of the mainstream at the wheel rim sealing outlet is uneven, and the high-pressure area has a periodic distribution, resulting in the risk of overheating caused by gas intrusion at the high-pressure location. In response to this phenomenon, the inventor proposes the following technical solution.

[0028] See also Figure 1 The embodiment of the present invention provides a turbine that can improve the wheel rim sealing effect. The wheel rim sealing is a sealing structure used by the wheel rim of the turbine blade of a gas turbine engine to prevent the gas from invading the internal structure of the turbine and causing overheating.

[0029] The turbine comprises a rotor blade 1 and a stator blade 2. The rotor blade 1 has a first edge plate 11. The stator blade 2 has a second edge plate 21, and the second edge plate 21 overlaps the first edge plate 11 in the axial direction of the turbine, and there is a gap between the second edge plate 21 and the first edge plate 11 in the radial direction of the turbine. The second edge plate 21 is provided with an inner recess 22 on a side away from the first edge plate 11.

[0030] See also Figure 1 , the rotor blade 1 is provided with edge plates on both sides of its axial direction, and the edge plates facing the stator blade 2 are called first edge plates 11. There may be one or two first edge plates 11. In some embodiments, each rotor blade 1 is provided with a first edge plate 11, and in the radial direction of the turbine, the first edge plate 11 is located on the side of the second edge plate 21 close to the central axis of the turbine, and the inner recess 22 is provided on the side of the second edge plate 21 facing the first edge plate 11.

[0031] See also Figure 1 In other embodiments, each rotor blade 1 is provided with two first edge plates 11; in the radial direction of the turbine, the second edge plate 21 is located between the two first edge plates 11. In this case, in some embodiments, the second edge plate 21 is provided with an inner recess 22 on the side facing the central axis of the turbine. Alternatively, in still other embodiments, the second edge plate 21 is provided with an inner recess 22 on the side facing the central axis of the turbine, and the second edge plate 21 is also provided with an inner recess 22 on the side facing away from the central axis of the turbine. The inner recess 22 of the second edge plate 21 is, for example, a groove of a regular shape, such as a rectangle, an arc, etc. Figure 3 and Figure 4 shown.

[0032] Similarly, the stator blade 2 is provided with edge plates on both sides of its axial direction, and the one facing the rotor blade 1 is called the second edge plate 21. The gap between the second edge plate 21 and the first edge plate 11 in the radial direction of the turbine is a channel that needs to be sealed. By providing an inner recess 22 on the second edge plate 21, the flow area of ​​the channel can be changed, so that more cold air can flow out through the channel, thereby improving the sealing effect.

[0033] See also Figure 1 and Figure 2 Furthermore, due to the stagnation effect at the leading edge of the turbine stator blade 2 and other reasons, the circumferential pressure distribution of the mainstream at the rim seal outlet is uneven, the high-pressure area A is periodically distributed, and there is a risk of overheating caused by the invasion of mainstream gas S1 at the high-pressure position. In view of this phenomenon, an inner recess 22 is provided at the location of the second rim plate 21 corresponding to the high-pressure area, which can specifically increase the cold air flow area in the high-pressure area, increase the inflow of cold air, improve the sealing effect, reduce the risk of overheating caused by the invasion of mainstream gas, and improve the sealing effect. Mainstream gas S1 refers to the gas in the main flow channel of the gas turbine engine.

[0034] See also Figure 1 In some embodiments, the rotor blades 1 and the stator blades 2 are arranged one by one, and the second edge plate 21 of each stator blade 2 is provided with an inner recess 22. The sealing effect of the rotor blades 1 and the stator blades 2 can be improved, thereby improving the sealing effect of the entire turbine.

[0035] See also Figure 3 and Figure 4 In some embodiments, the second edge plate 21 is provided with one or more inner recesses 22. The more the number of the inner recesses 22 is, the larger the airflow area of ​​the inner recesses 22 is, and the better the sealing effect is.

[0036] Back to Figure 1 In some embodiments, the turbine further includes an air collecting chamber 3, which is located on the side of the stator blade 2 facing the central axis of the turbine. The wall of the air collecting chamber 3 is provided with air holes, which are connected to the gap. The air supply effect of the slotted airflow is further improved through the air collecting chamber 3 and the air holes, and the ability to resist the intrusion of fuel gas is improved. The setting position of the air holes and the position of the inner recess 22 of the second edge plate 21 correspond to each other in the circumferential direction of the turbine; the number of the air holes corresponds to the number of the inner recess 22 of the second edge plate 21, which can improve the flow path of the cooling airflow and improve the sealing effect.

[0037] The flow path of the cooling gas (also called sealing gas) S2 is as follows: the cooling gas S2 in the gas collecting cavity 3 flows out through the air holes into the disc cavity formed by the rotor blades 1 and the stator blades 2. The cooling gas in the disc cavity is divided into two paths. The first path of cooling gas S2 enters the channel between the first edge plate 11 and the second edge plate 21 as a sealing gas; the second path of cooling gas flows to the side of the stator blade 2 away from the rotor blade 1.

[0038] Continue to see Figure 1 In some embodiments, the turbine further comprises a grate seal assembly 4, which is arranged on the side of the gas collecting chamber 3 away from the stator blades 2. The second cooling gas flows through the grate seal assembly 4 and is supplied to the wheel rim seal downstream of the stator blades 2. The rotor member 41 of the grate seal assembly 4 is fixed to the same assembly as the rotor blades 1, and the honeycomb 42 of the grate is fixed to the wall of the gas collecting chamber 3.

[0039] An embodiment of the present invention further provides an aircraft engine, comprising a turbine provided by any technical solution of the present invention.

[0040] In the description of the present invention, it is necessary to understand that the terms "center", "longitudinal", "lateral", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside" and the like indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on the protection content of the present invention.

[0041] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein, but these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A turbine, characterized in that: include: A rotor blade (1) having a first edge plate (11); as well as The stator blade (2) has a second edge plate (21), the second edge plate (21) and the first edge plate (11) overlap in the axial direction of the turbine, and there is a gap between the two in the radial direction of the turbine; Wherein, a concave portion (22) is provided on a side of the second edge plate (21) away from the first edge plate (11) to increase the flow area, increase the amount of cold air flowing in, and improve the sealing effect; The stator blade (2) has a high-pressure area (A), the high-pressure area (A) is distributed periodically, and an inner concave portion (22) is provided at a location of the second edge plate (21) corresponding to the high-pressure area, so as to increase the cold air flow area of ​​the high-pressure area (A), thereby preventing the mainstream gas (S1) from invading the high-pressure area (A) and causing an over-temperature risk; An air collecting cavity (3) is located on a side of the stator blade (2) facing the central axis of the turbine, the wall of the air collecting cavity (3) being provided with air holes, the air holes being in communication with the gap; The flow path of the cooling gas (S2) is as follows: the cooling gas (S2) in the gas collecting cavity (3) flows out through the air holes into the disc cavity formed by the rotor blade (1) and the stator blade (2), and the cooling gas in the disc cavity is divided into two paths. The first cooling gas enters the channel between the first edge plate (11) and the second edge plate (21) to serve as sealing gas; the second cooling gas flows to the side of the stator blade (2) away from the rotor blade (1).

2. The turbine according to claim 1, characterized in that The rotor blades (1) and the stator blades (2) are arranged in one-to-one correspondence, and the second edge plate (21) of each stator blade (2) is provided with the inner recess (22).

3. The turbine according to claim 1, characterized in that The second edge plate (21) is provided with one or more inner recesses (22).

4. The turbine according to claim 1, characterized in that Each of the rotor blades (1) is provided with a first edge plate (11); in the radial direction of the turbine, the first edge plate (11) is located on a side of the second edge plate (21) close to the center axis of the turbine; and the inner recess (22) is provided on a side of the second edge plate (21) facing the first edge plate (11).

5. The turbine according to claim 1, characterized in that Each of the rotor blades (1) is provided with two first edge plates (11); in the radial direction of the turbine, the second edge plate (21) is located between the two first edge plates (11).

6. The turbine according to claim 5, characterized in that The inner recess (22) is provided on a side of the second edge plate (21) facing the central axis of the turbine.

7. The turbine according to claim 5, characterized in that The inner recess (22) is provided on the side of the second edge plate (21) facing the central axis of the turbine, and the inner recess (22) is also provided on the side of the second edge plate (21) facing away from the central axis of the turbine.

8. The turbine according to claim 7, characterized in that Also includes: A comb teeth sealing assembly (4) is arranged on a side of the air collecting cavity (3) away from the stator blades (2).

9. An aircraft engine, characterized in that: It comprises the turbine described in any one of claims 1 to 8.

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