CN114838000B - Variable-length tubular type vortex reducer system of aero-engine compressor - Google Patents

Abstract

The invention discloses a variable-length tubular vortex reducer system of an aircraft engine gas compressor, which comprises an upstream gas compressor disk, a downstream gas compressor disk, a drum barrel and a vortex reducer, wherein the drum barrel is arranged between the upstream gas compressor disk and the downstream gas compressor disk, and the drum barrel, the upstream gas compressor disk and the downstream gas compressor disk form a disk cavity of the gas compressor; the vortex reducer comprises a chuck and a plurality of vortex reducing pipes arranged on the side wall of the chuck along the circumferential direction, the chuck is arranged in a disc cavity, two ends of the chuck are respectively connected with an upstream compressor disc and a downstream compressor disc, and the vortex reducing pipes are of pipe body structures which are hollow inside and adjustable in length; according to the invention, the vortex reducing pipe is designed into a structure capable of changing the length of the pipe, so that the position of an inlet and an outlet of the vortex reducing pipe can be adjusted according to the real-time working condition of an engine, the separation vortex formed by airflow at the inlet of the vortex reducing pipe is reduced, and cooling gas meeting the corresponding pressure is provided according to the requirement of a high-temperature component, so that the vortex reducing pipe has good vortex reducing performance in the whole flight envelope.

Description

A variable-length tubular vortex reducer system for an aero-engine compressor

technical field

The invention relates to the technical field of aero-engines, in particular to a variable-length tubular vortex reducer system for an aero-engine compressor.

Background technique

The cold air of the internal flow air cooling system of the aero-engine is drawn from the middle stage of the compressor. On the one hand, this part of cold air is used to cool high-temperature components such as turbine blades, casings, discs, shafts and bearings, so that they can work under the allowable temperature and ensure the pressure of the intake air of relevant parts; on the other hand, the cold air will It is used for sealing to prevent gas backflow; at the same time, the cold air is also used to maintain the pressure of the unloading chamber, so as to meet the requirements of the axial force of the engine. However, a large amount of bleed air will reduce the overall efficiency of the engine and increase fuel consumption, so the flow of cold air from the compressor should be controlled within a certain range. Secondly, when designing the cooling of the hot-end components, in addition to considering the overall average temperature of the hot-end components, their temperature gradient should also be considered. In addition, after the cold air drawn from the compressor flows radially from the drum hole, it must flow through a designed flow path before reaching the final target area. During this process, the pressure of the cold air will gradually decrease and the temperature will gradually increase due to the effect of resistance. High, so it needs to be designed reasonably to reduce the resistance as much as possible to reduce the pressure loss. Therefore, the amount of cooling air used, the overall temperature and temperature gradient, and pressure loss have become the key issues that designers of inflow cooling air systems need to consider.

The most widely used vortex reducer at this stage is the straight tube vortex reducer. This type of straight-tube vortex reducer is installed in the above-mentioned radial air-inducing chamber to guide the airflow. Since the working conditions of the engine will continue to change during the entire working process, the existing straight-tube vortex reducer has an ideal drag reduction effect only in a certain working condition or a small range of working conditions due to its fixed structure , when deviating from the design working condition, because the gas in the chamber has a relatively large circumferential velocity relative to the vortex reducer, when entering the vortex reducer, flow separation will occur at the inlet of the vortex reducer, resulting in a large pressure loss.

Contents of the invention

The purpose of the present invention is to provide a variable-length tubular vortex reducer system for an aero-engine compressor to solve the above-mentioned problems in the prior art. According to the demand, the cooling gas that meets the corresponding pressure is provided, so that the vortex reducer has good vortex reduction performance in the entire flight envelope.

To achieve the above object, the present invention provides the following scheme:

The present invention provides a variable-length tubular vortex reducer system for an aero-engine compressor, which includes an upstream compressor disc, a downstream compressor disc, a drum and a vortex reducer, and the drum is arranged on the upstream compressor disc and the upstream compressor disc. Between the downstream compressor disks, the drum, the upstream compressor disk and the downstream compressor disk constitute the disk chamber of the compressor; Several vortex reducing tubes on the wall, the chuck is arranged in the disc cavity and the two ends are respectively connected to the upstream compressor disc and the downstream compressor disc, the vortex reducing tubes are tubes with a hollow interior and adjustable length body structure.

Preferably, a plurality of drum holes for introducing airflow into the disc cavity are evenly distributed on the outer wall of the drum in the circumferential direction.

Preferably, the drum hole is an oblong hole.

Preferably, the upstream compressor disk, the downstream compressor disk, the drum and the chuck are coaxially arranged.

Preferably, each of the vortex reducers is evenly distributed in the circumferential direction of the chuck.

Preferably, the vortex reducer includes a vortex reducer inlet section, a vortex reducer outlet section, a vortex reducer fixed section and a transmission rod, and the two ends of the vortex reducer fixed section are respectively connected to the vortex reducer through the transmission rod. The inlet section of the tube and the outlet section of the vortex reducer. The inlet section of the vortex reducer and the outlet section of the vortex reducer adjust the axial position through the transmission rod.

Preferably, a threaded pipe is arranged inside the fixed section of the vortex reducer, and the two ends of the threaded pipe are respectively threaded with one of the transmission rods, and a mounting shell is arranged outside the fixed section of the vortex reducer, the A servo motor is arranged in the installation shell, and a first gear is arranged on the rotating shaft of the servo motor, and the first gear meshes with the second gear fixed in the middle of the threaded pipe.

Preferably, several mounting holes are evenly distributed on the outer wall of the chuck in the circumferential direction, and the vortex reducer is arranged on the corresponding mounting holes through the vortex reducer fixing joint.

Compared with the prior art, the present invention has achieved the following beneficial technical effects:

The variable-length tubular vortex reducer system of the aero-engine compressor provided by the present invention can adjust the inlet and outlet positions of the vortex reducer according to the real-time working conditions of the engine by designing the vortex reducer into a structure capable of changing the length of the tube. Reduce the separation vortex formed by the airflow at the inlet of the vortex reducer, and provide cooling gas that meets the corresponding pressure according to the needs of high-temperature components, so that the vortex reducer has good vortex reduction performance in the entire flight envelope.

Description of drawings

In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the accompanying drawings required in the embodiments. Obviously, the accompanying drawings in the following description are only some of the present invention. Embodiments, for those of ordinary skill in the art, other drawings can also be obtained based on these drawings without any creative effort.

Fig. 1 is the structural representation of variable-length tubular vortex reducer system of aero-engine compressor in the present invention;

Fig. 2 is the structural representation of vortex reducing tube in the present invention;

Fig. 3 is a schematic diagram of the internal structure of the vortex reducer in the present invention;

In the figure: 1-Upstream compressor disc, 2-Downstream compressor disc, 3-Drum, 4-Vortex reducer, 5-Disc cavity, 6-Chuck, 7-Vortex reducer, 71-Vortex reducer inlet Section, 72-Vortex Reducer Outlet Section, 73-Vortex Reducer Fixed Section, 74-Transmission Rod, 75-Threaded Pipe, 76-Installation Shell, 77-Servo Motor, 78-First Gear, 79-Second Gear, 8-drum hole.

Detailed ways

The following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only some, not all, embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by persons of ordinary skill in the art without making creative efforts belong to the protection scope of the present invention.

The object of the present invention is to provide a variable-length tubular vortex reducer system for an aero-engine compressor to solve the problems in the prior art.

In order to make the above objects, features and advantages of the present invention more comprehensible, the present invention will be further described in detail below in conjunction with the accompanying drawings and specific embodiments.

The aeroengine compressor variable length tubular vortex reducer system in this embodiment, as shown in Figure 1, includes an upstream compressor disc 1, a downstream compressor disc 2, a drum 3 and a vortex reducer 4, and the drum 3 Set between the upstream compressor disk 1 and the downstream compressor disk 2, the drum 3, the upstream compressor disk 1 and the downstream compressor disk 2 constitute the disk cavity 5 of the compressor; the vortex reducer 4 includes a chuck 6 and a peripheral To the number of vortex reducing tubes 7 arranged on the side wall of the chuck 6, the chuck 6 is arranged in the disc cavity 5 and the two ends are respectively connected to the upstream compressor disc 1 and the downstream compressor disc 2, and the vortex reducing tubes 7 are hollow inside And the tube body structure whose length can be adjusted.

In this specific embodiment, the outer wall of the drum 3 is evenly distributed in the circumferential direction with a number of drum holes 8 for introducing airflow into the disk chamber 5, and the drum holes 8 are specifically oblong holes.

In this specific embodiment, the upstream compressor disk 1 , the downstream compressor disk 2 , the drum 3 and the chuck 6 are coaxially arranged.

In this specific embodiment, the structures and sizes of the vortex reducers 7 are the same, and they are evenly distributed in the circumferential direction of the chuck 6 .

As shown in Figure 2, the vortex reducer 7 includes a vortex reducer inlet section 71, a vortex reducer outlet section 72, a vortex reducer fixed section 73 and a transmission rod 74, and the two ends of the vortex reducer fixed section 73 are respectively connected by a transmission rod 74 The vortex reducer inlet section 71 and the vortex reducer outlet section 72, the vortex reducer inlet section 71 and the vortex reducer outlet section 72 adjust the axial position through the transmission rod 74; as shown in Figure 3, the inside of the vortex reducer fixed section 73 A threaded pipe 75 is provided, and the two ends of the threaded pipe 75 are threadedly connected with a transmission rod 74 respectively. An installation shell 76 is arranged on the outside of the vortex reducer fixed joint 73, and a servo motor 77 is arranged inside the installation shell 76. The rotating shaft of the servo motor 77 The first gear 78 is arranged on the top, and the first gear 78 meshes with the fixed second gear 79 in the middle of the threaded pipe 75. The rotation of the servo motor 77 drives the first gear 78 to rotate, and the first gear 78 drives the second gear 79 to rotate through the engagement. The second gear 79 rotates the threaded pipe 75, and the threaded pipe 75 rotates to drive the transmission rods 74 threaded at both ends to move synchronously, thereby driving the axial displacement of the vortex reducer inlet section 71 and the vortex reducer outlet section 72 to realize the entrance of the vortex reducer 7 and adjustment of the outlet position.

In this specific embodiment, the outer wall of the chuck 6 is evenly distributed with several installation holes in the circumferential direction, and the vortex reducer 7 is clamped on the corresponding installation hole by setting a shoulder on the vortex reducer fixing joint 73, and the vortex reducer The inlet section 71 is located outside the chuck 6 , and the outlet section 72 of the vortex reducer is located inside the chuck 6 .

When the compressor is running, several vortex reducers 7 rotate coaxially, at the same speed, and in the same direction as the upstream compressor disk 1 and downstream compressor disk 2 on both sides, and the airflow flows through several drum holes 8 into the compressed air In the disc cavity 5 of the compressor, the flow is led into the axial channel of the compressor through the scroll reducer 7; the length of the scroll reducer 7 can be adjusted, so that the scroll reducer 7 can adjust the inlet and outlet positions according to the real-time working conditions of the engine, reducing The separated vortex formed by the small air flow at the entrance of the vortex reducer 7 provides cooling gas that meets the corresponding pressure according to the requirements of high-temperature components, so that the vortex reducer 7 has good vortex reduction performance in the entire flight envelope.

The present invention uses specific examples to illustrate the principle and implementation of the present invention. The description of the above embodiments is only used to help understand the method of the present invention and its core idea; meanwhile, for those of ordinary skill in the art, according to the present invention Thoughts, there will be changes in specific implementation methods and application ranges. In summary, the content of this specification should not be construed as limiting the present invention.

Claims (6)

1. The utility model provides an aeroengine compressor variable length tubular subtracts whirlpool ware system which characterized in that: the device comprises an upstream compressor disk, a downstream compressor disk, a drum and a vortex reducer, wherein the drum is arranged between the upstream compressor disk and the downstream compressor disk, and the drum, the upstream compressor disk and the downstream compressor disk form a disk cavity of a compressor; the vortex reducer comprises a chuck and a plurality of vortex reducing pipes arranged on the side wall of the chuck along the circumferential direction, the chuck is arranged in the disc cavity, two ends of the chuck are respectively connected with the upstream compressor disc and the downstream compressor disc, and the vortex reducing pipes are of pipe body structures which are hollow inside and adjustable in length;

the vortex reducing pipe comprises a vortex reducing pipe inlet joint, a vortex reducing pipe outlet joint, a vortex reducing pipe fixed joint and a transmission rod, wherein two ends of the vortex reducing pipe fixed joint are respectively connected with the vortex reducing pipe inlet joint and the vortex reducing pipe outlet joint through the transmission rod, and the axial positions of the vortex reducing pipe inlet joint and the vortex reducing pipe outlet joint are adjusted through the transmission rod; the scroll reducer is characterized in that a threaded pipe is arranged inside the scroll reducer fixing joint, two ends of the threaded pipe are respectively in threaded connection with one transmission rod, an installation shell is arranged outside the scroll reducer fixing joint, a servo motor is arranged in the installation shell, a first gear is arranged on a rotating shaft of the servo motor, and the first gear is meshed with a second gear fixed in the middle of the threaded pipe.

2. The aircraft engine compressor variable length tubular deswirler system of claim 1, wherein: and a plurality of drum holes for introducing air flow into the disc cavity are uniformly distributed on the outer wall of the drum in the circumferential direction.

3. The aircraft engine compressor variable length tubular deswirler system of claim 2, wherein: the drum holes are oblong holes.

4. The aero-engine compressor variable length tubular deswirler system of claim 1, wherein: the upstream compressor disk, the downstream compressor disk, the drum and the chuck are arranged coaxially.

5. The aircraft engine compressor variable length tubular deswirler system of claim 1, wherein: the vortex reducing pipes are uniformly distributed on the chuck in the circumferential direction.

6. The aero-engine compressor variable length tubular deswirler system of claim 1, wherein: the circumference equipartition has a plurality of mounting holes on the outer wall of chuck, it passes through to subtract the scroll fixed knot and set up in corresponding on the mounting hole.

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