Structure for regulating axial force of turbine disk by exhaust gas ejection and aeroengine

Abstract

The present invention discloses a structure for adjusting the axial force of a turbine disk by exhaust gas ejection and an aeroengine. The structure for adjusting the axial force of a turbine disk by exhaust gas ejection includes: a turbine disk and turbine rotor blades. The turbine disk is located in a mixing chamber, the turbine rotor blades are located in a main flow chamber, and an ejection hole is provided on one side of the mixing chamber away from the turbine disk; a protective casing, which is arranged at an interval from the turbine disk. A rear cavity of the turbine disk is formed between the protective casing and the turbine disk, and the rear cavity of the turbine disk is communicated with the mixing chamber. The beneficial effect of the present invention is that by opening an ejection hole at the center of the tail nozzle, the gas path behind the turbine disk is changed, and the outflow position is adjusted from a relatively high-pressure position at the turbine rim to a relatively low-pressure position of the exhaust gas. Thus, on the basis of meeting the cooling and sealing requirements of the rear cavity of the turbine disk, the pressure behind the turbine disk is greatly reduced, and the effective regulation of the rotor axial force is realized.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engine air systems, and particularly to a structure and an aero-engine for regulating the axial force of a turbine disk by using exhaust gas ejection. Background Art

[0002] In the aero-engine air system, secondary air needs to be provided to the turbine disk cavity (composed of a turbine rotating disk and a stationary disk), and its functions include cooling the turbine disk and sealing the rim clearance between the rotating and stationary disks. Figure 1 The conventional cooling and sealing gas path and structure of the rear turbine disk cavity are shown. The low-temperature gas led out from the compressor enters the rear turbine disk cavity 5 through the axial vent hole 3. The air flow entering the rear turbine disk cavity first flows over the surface of the turbine rotating disk 2 to cool the rotating disk; then it flows out from the turbine rim rotating-stator clearance 6 between the rotating disk 2 and the stationary disk 4 and mixes with the mainstream, playing a role in sealing the turbine rim rotating-stator clearance 6 to prevent the high-temperature gas in the mainstream from entering the rear compartment of the turbine disk through the turbine rim rotating-stator clearance 6 and causing the overheating of the turbine disk. This structure also includes a mainstream exhaust guide vane 7, which is arranged at intervals with the mainstream vane 1.

[0003] Since all the axial loads on the rotating disks in the aero-engine air system jointly determine the total axial load of the rotor, the pressure in the disk cavity will directly affect the total axial force of the rotor. Although the above flow path and structure layout meet the requirements of cooling and sealing the rear turbine disk cavity, they also result in a relatively high pressure in the rear turbine disk cavity and cannot be adjusted and optimized according to the design requirements of the total axial force of the rotor (the pressure in the rear turbine disk cavity is equivalent to the pressure at the turbine rim rotating-stator clearance 6, and the pressure at the turbine rim rotating-stator clearance 6 is determined by the mainstream vane 1, and once the vane is designed, it cannot be adjusted anymore). How to further optimize the design of the flow path and structure of the rear turbine disk cavity and realize the regulation of the pressure in the rear turbine disk cavity on the basis of meeting the cooling and sealing requirements of the rear turbine disk cavity is of great significance for regulating the axial force of the rotor. [[ID='14']] Summary of the Invention

[0004] In view of this, the present invention discloses a structure and an aero-engine for regulating the axial force of a turbine disk by using exhaust gas ejection, so as to achieve the purpose of regulating the pressure in the rear turbine disk cavity and further changing the axial force of the rotor to the working range of the thrust bearing.

[0005] The embodiments of the present specification provide the following technical solutions: A structure for regulating the axial force of a turbine disk by using exhaust gas ejection, comprising: a turbine disk and turbine rotor blades, the turbine disk is located in the mixing cavity, the turbine rotor blades are located in the mainstream cavity, and an ejection hole is arranged on the side of the mixing cavity far from the turbine disk; a protective casing, which is arranged at intervals with the turbine disk, a rear turbine disk cavity is formed between the protective casing and the turbine disk, and the rear turbine disk cavity is communicated with the mixing cavity.

[0006] Further, a cooling and sealing flow passage is provided inside the turbine disk, and the cooling and sealing flow passage is connected to the rear cavity of the turbine disk through an axial vent hole.

[0007] Further, the structure for adjusting the axial force of the turbine disk by exhaust gas ejection further includes a main exhaust guide vane, which is arranged at intervals with the turbine rotor blades, and the main exhaust guide vane extends from the main flow cavity to the mixing cavity; the protection casing is fixedly connected to the main exhaust guide vane in the mixing cavity.

[0008] Further, the main flow cavity is communicated with the rear cavity of the turbine disk through the turbine rim rotating-stator clearance.

[0009] Further, a transition cavity is formed between the turbine rim rotating-stator clearance and the rear cavity of the turbine disk, and a labyrinth seal for reducing the intake air volume from the turbine rim rotating-stator clearance is arranged in the transition cavity.

[0010] Further, the labyrinth seal is a straight tooth or a stepped tooth.

[0011] Further, the diameter range of the ejection hole is 10 to 50 mm.

[0012] Further, the clearance range of the rear cavity of the turbine disk is 2 to 20 mm.

[0013] The present invention also provides an aeroengine, including the structure for adjusting the axial force of the turbine disk by exhaust gas ejection as described above.

[0014] Compared with the prior art, the beneficial effects that at least one of the above technical solutions adopted in the embodiments of the present specification can achieve at least include:

[0015] By opening the ejection hole, the gas path behind the turbine disk is changed, and its outflow position is adjusted from the relatively high-pressure part of the turbine rim to the relatively low-pressure part of the exhaust (caused by the high-speed ejection of the exhaust gas), so as to greatly reduce the pressure behind the turbine disk on the basis of meeting the cooling and sealing requirements of the rear cavity of the turbine disk, and effectively control the axial force of the rotor.

[0016] The size and position of the ejection hole can be designed according to the regulation requirements of the axial force, and in engineering practice, rapid scheme adjustment, part processing, and test verification can be carried out.

[0017] By arranging a protection casing closely attached to the disk in the rear cavity of the turbine disk, the cooling effect of the airflow on the disk can be strengthened. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required to be used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application, and those of ordinary skill in the art can also obtain other drawings according to these drawings without creative efforts.

[0019] Figure 1 It is a schematic diagram of the flow path and structure behind the turbine disk in the prior art;

[0020] Figure 2 This is a schematic diagram of the flow path and structure behind the turbine disk using exhaust injection in the present invention.

[0021] The reference numerals in the figure are: 1. turbine rotor blade; 2. turbine disc; 3. axial air vent; 4. tail nozzle inner casing; 5. turbine disc rear cavity; 6. turbine rim rotor-stator gap; 7. mainstream exhaust guide vane; 8. protective casing; 9. grate seal; 10. ejection hole; 11. mixing chamber. DETAILED DESCRIPTION

[0022] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0023] It should be noted that, in the absence of conflict, the embodiments and features of the embodiments in this application can be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0024] like Figure 2 As shown, an embodiment of the present invention provides a flow path and structure for regulating the axial force of an aircraft engine turbine disk, comprising a turbine disk 2, turbine rotor blades 1, and a protective casing 8. The turbine disk 2 is located in a mixing chamber 11, while the turbine rotor blades 1 are located in a main flow chamber. An ejection hole 10 is provided on the side of the mixing chamber 11 facing away from the turbine disk 2. The protective casing 8 is spaced apart from the turbine disk 2, forming a turbine disk rear chamber 5 between the protective casing 8 and the turbine disk 2, and the turbine disk rear chamber 5 is in communication with the mixing chamber 11.

[0025] Since the pressure in the rear cavity 5 of the turbine disk will be greatly reduced after the introduction hole 10 is set, the flow at the turbine rim-stator gap 6 will be reversed, and the mainstream high-temperature combustion gas will enter the rear cavity 5 of the turbine disk. In order to prevent the high-temperature combustion gas from causing harm to the turbine disk, a protective casing 8 is set, which is used to surround and protect the turbine disk. After adding the protective casing 8, the rear cavity 5 of the turbine disk becomes a small cavity adjacent to the turbine disk, and the cavity is filled with cooling sealing air from the compressor introduced from the axial vent 3. After cooling the turbine disk, this airflow will be mixed with the mainstream high-temperature combustion gas from the turbine rim-stator gap 6, and will be merged into the mainstream again from the introduction hole 10.

[0026] Preferably, the ejection hole 10 is arranged at the center of the casing 4 inside the tail nozzle. Since the mainstream flow velocity is very high and the static pressure is very low here, the cooling sealing airflow entering the rear cavity of the turbine disk from the axial vent 3 will no longer flow out from the turbine wheel rim stator gap 6, but will flow out from the ejection hole 10.

[0027] Preferably, the protective casing 8 can be optimized according to the surface shape of the turbine disk. On the premise of ensuring no rubbing in the hot state, reducing the distance between the disk and the casing can further enhance the cooling effect of the cooling and sealing air on the turbine disk.

[0028] A cooling and sealing flow passage is arranged inside the turbine disk 2, and the cooling and sealing flow passage is connected to the rear cavity 5 of the turbine disk through the axial vent hole 3. The rear cavity 5 of the turbine disk is cooled by the cooling and sealing air flow passing through the axial vent hole 3.

[0029] The structure for adjusting the axial force of the turbine disk by exhaust gas ejection further includes a main exhaust guide vane 7, which is arranged at intervals with the turbine rotor blade 1, and the main exhaust guide vane 7 extends from the main cavity to the mixing cavity 11; the protective casing 8 is fixedly connected to the main exhaust guide vane 7 in the mixing cavity 11.

[0030] The main exhaust guide vane 7 is extended inward, and the protective casing 8 is connected to it by welding.

[0031] The main cavity is communicated with the rear cavity 5 of the turbine disk through the turbine rim stator clearance 6.

[0032] Furthermore, a transition cavity is formed between the turbine rim stator clearance 6 and the rear cavity 5 of the turbine disk. Since the main air flow entering the mixing cavity 11 from the turbine rim stator clearance 6 will rejoin the main flow from the ejection hole 10 again, there will be relatively large irreversible losses in the extraction, mixing with the cooling and sealing air flow, and rejoining of this air flow into the main flow, resulting in a reduction in the engine thrust. To reduce this adverse effect, a labyrinth seal 9 is arranged at the high radius of the turbine disk to reduce the flow rate of this air flow. When adopting this kind of flow path layout, the purpose of arranging the sealing structure at the high radius at the rear of the turbine disk is to reduce unnecessary air leakage and reduce the influence of ejection on the engine performance.

[0033] Preferably, the labyrinth seal 9 can adopt straight teeth with a simple structure or stepped teeth with better sealing effect. And the leakage can be reduced by increasing the number of teeth and grooves and reducing the clearance to achieve a better sealing effect.

[0034] Optionally, the diameter range of the ejection hole 10 is 10 to 50 mm. The axial position and the aperture size of the ejection hole 10 can be designed according to the pressure required in the rear cavity of the turbine disk.

[0035] Optionally, the clearance range of the rear cavity 5 of the turbine disk is 2 to 20 mm.

[0036] The embodiment of the present invention also provides an aeroengine, including the structure for adjusting the axial force of the turbine disk by exhaust gas ejection as described above.

[0037] The above are only specific embodiments of the present invention, and the scope of the invention cannot be limited thereby. Therefore, the replacement of equivalent components or equivalent changes and modifications made in accordance with the scope of protection of the present invention should still fall within the scope covered by this patent. In addition, the technical features in the present invention, between technical features, between technical features and technical solutions, and between technical solutions can be freely combined and used.

Claims

1. A structure for adjusting the axial force of a turbine disk by exhaust gas ejection, characterized in that, include: A turbine disc (2) and a turbine rotor blade (1), wherein the turbine disc (2) is located in a mixing chamber (11), the turbine rotor blade (1) is located in a main flow chamber, and an ejection hole (10) is provided on a side of the mixing chamber (11) away from the turbine disc (2); A cooling sealing flow channel is provided on the inner side of the turbine disc (2), and the cooling sealing flow channel is connected to the turbine disc rear cavity (5) through the axial vent hole (3); A protective casing (8) is spaced apart from the turbine disc (2), a turbine disc rear cavity (5) is formed between the protective casing (8) and the turbine disc (2), and the turbine disc rear cavity (5) is communicated with the mixing cavity (11); The structure for regulating the axial force of the turbine disk by exhaust gas injection further comprises a mainstream exhaust guide vane (7) spaced apart from the turbine rotor blade (1), the mainstream exhaust guide vane (7) extending from the mainstream cavity to the mixing cavity (11); a protective casing (8) fixedly connected to the mainstream exhaust guide vane (7) in the mixing cavity (11); The main flow cavity is connected to the turbine disc rear cavity (5) via the turbine rim rotor-stator gap (6); The cooling seal gas from the compressor coming from the axial vent hole (3) is mixed with the mainstream high-temperature combustion gas coming from the turbine wheel rim rotor-stator gap (6) and then merges into the mainstream again from the ejection hole (10).

2. The structure for adjusting the axial force of a turbine disk by exhaust gas ejection according to claim 1, wherein, A transition cavity is formed between the turbine wheel rim rotor-stator gap (6) and the turbine disc rear cavity (5), and a grate seal (9) for reducing the amount of air entering through the turbine wheel rim rotor-stator gap (6) is provided in the transition cavity.

3. The structure for adjusting the axial force of the turbine disk by exhaust gas ejection according to claim 2, characterized in that, The grate teeth seal (9) is a straight tooth or a stepped tooth.

4. The structure for adjusting the axial force of the turbine disk by exhaust gas ejection according to claim 1, characterized in that, The diameter of the ejection hole (10) ranges from 10 to 50 mm.

5. The structure for adjusting the axial force of a turbine disk by exhaust gas ejection according to claim 1, characterized in that The clearance of the turbine disc rear cavity (5) ranges from 2 to 20 mm.

6. An aeroengine, comprising a structure for regulating the axial force of a turbine disk by means of exhaust gas ejection, characterized in that, The structure for adjusting the axial force of the turbine disk by using exhaust gas injection is the structure for adjusting the axial force of the turbine disk by using exhaust gas injection as claimed in any one of claims 1 to 5.

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