Vortex reducer and aircraft engine

Abstract

The application discloses a vortex reducer and an aero-engine, and relates to the technical field of aero-engine cooling systems. The vortex reducer comprises a vortex reduction pipe and a support ring, the vortex reduction pipe is arranged in the support ring, and the vortex reducer is further provided with an airflow channel. The airflow channel sequentially flows through the support ring, the pipe wall of the vortex reduction pipe and the support ring. The cooling airflow introduced by the pressure compressor can be introduced into the airflow channel. The cooling airflow exchanges heat with the radial airflow flowing through the vortex reduction pipe, so that the temperature of the radial airflow can be effectively reduced, the radial airflow can be kept at a lower temperature, and the performance of the aero-engine cooling system is ensured.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aero-engines, in particular to a vortex reducer and an aero-engine. BACKGROUND

[0002] The radial flow guiding vortex reducer is usually located in the disc cavity of the rotor disc of the compressor of the aero-engine, and the radial flow guiding vortex reducer reduces the pressure loss in the air system during air extraction by inhibiting the development of the circumferential velocity of the airflow, which plays a very important role in improving the efficiency of the engine.

[0003] The vortex reducer usually comprises a vortex reduction pipe and a support ring, and one end of the vortex reduction pipe is inserted into the support ring. The radial airflow can directly flow through the vortex reduction pipe and the support ring. The temperature of the radial airflow is usually high, and the vortex reducer cannot effectively cool the radial airflow. SUMMARY

[0004] The technical problem to be solved by the present application is to overcome the above-mentioned defects in the prior art that the vortex reducer cannot cool the radial airflow, and to provide a vortex reducer and an aero-engine.

[0005] The present application solves the above-mentioned technical problems by the following technical solutions:

[0006] A vortex reducer, comprising a vortex reduction pipe and a support ring, the vortex reduction pipe being inserted into the support ring, and the vortex reducer further comprising an airflow passage, the airflow passage sequentially flowing through the support ring, the pipe wall of the vortex reduction pipe and the support ring.

[0007] In the present application, by adopting the above structure, the airflow passage is provided in the vortex reducer, and cooling airflow extracted from the compressor can be introduced into the airflow passage. The cooling airflow exchanges heat with the radial airflow flowing through the vortex reduction pipe, thereby effectively reducing the temperature of the radial airflow, and further maintaining a lower temperature of the radial airflow, and ensuring the performance of the cooling system of the aero-engine.

[0008] Preferably, the airflow passage comprises an inlet section, a cooling section and an outlet section, the inlet section and the outlet section are arranged on the support ring, and the cooling section is arranged on the pipe wall.

[0009] In the present application, by adopting the above structure, the structure and function of the airflow passage are relatively concentrated, which can simplify the design of the airflow passage and improve the cooling effect of the radial airflow.

[0010] Preferably, the cooling section comprises a straight-through pipe and a coiled pipe connected in series, the straight-through pipe extends from the first end of the vortex reduction pipe to the second end of the vortex reduction pipe, and the coiled pipe extends from the second end of the vortex reduction pipe to the first end of the vortex reduction pipe.

[0011] The first end of the vortex-reducing pipe is the end of the vortex-reducing pipe close to the support ring, and the second end of the vortex-reducing pipe is the end of the vortex-reducing pipe away from the support ring.

[0012] In the scheme, by adopting the above structure, the cooling airflow directly flows to the second end of the vortex-reducing pipe through the straight pipe, and the cooling airflow flows into the coiled pipe from the second end, so that the temperature of the cooling airflow in the coiled pipe at the second end is relatively lower, and the temperature difference with the radial airflow is larger, and the cooling airflow can more efficiently and rapidly exchange heat with the radial airflow, and the temperature of the radial airflow can be more effectively reduced. The cooling airflow flows from the second end to the first end, and the cooling airflow and the radial airflow continuously exchange heat, the temperature of the cooling airflow gradually rises, and the temperature of the radial airflow gradually and uniformly decreases, so that the temperature of the radial airflow is more stable, the temperature difference between the radial airflow close to the pipe wall and the radial airflow away from the pipe wall is avoided to be too large, and the stability and reliability of the aircraft engine cooling system can be improved.

[0013] Preferably, the coiled pipe is helical as a whole from the second end of the vortex-reducing pipe to the first end of the vortex-reducing pipe.

[0014] Alternatively, the coiled pipe comprises a plurality of U-shaped pipes connected end to end, and the U-shaped pipes are arranged along the axis of the vortex-reducing pipe.

[0015] Alternatively, the cooling section is zigzag as a whole, and the zigzag cooling section is reciprocally arranged from the first end of the vortex-reducing pipe to the second end of the vortex-reducing pipe.

[0016] In the scheme, by adopting the above structure, the structure of the coiled pipe can be simplified, the heat exchange area between the cooling airflow and the radial airflow can be increased, and the cooling effect of the radial airflow can be improved.

[0017] Preferably, the inlet section and the outlet section are sequentially arranged along the axis of the support ring.

[0018] In the scheme, by adopting the above structure, the cooling airflow can flow into the inlet section, and the cooling airflow can flow out of the outlet section and then flow to other areas.

[0019] Preferably, the vortex-reducing device comprises a plurality of the vortex-reducing pipes, the plurality of vortex-reducing pipes are arranged at intervals along the circumference of the support ring, and the airflow channel is correspondingly arranged with the vortex-reducing pipe.

[0020] In the scheme, by adopting the above structure, the temperature of the radial airflow can be more uniformly and effectively reduced, and the stability and reliability of the aircraft engine cooling system can be improved.

[0021] Preferably, the airflow channel further comprises a communication ring segment, the communication ring segment is arranged along the circumference of the support ring, and the communication ring segment is in communication with the inlet segment of the airflow channel or the outlet segment of the airflow channel.

[0022] Alternatively, the number of the communication ring segments is two, and the two communication ring segments are in communication with the inlet segment of the airflow channel and the outlet segment of the airflow channel respectively.

[0023] An aero-engine, the aero-engine comprising the vortex reducer as described above, and the cooling airflow introduced from the compressor of the aero-engine flows along the airflow channel.

[0024] In the present scheme, by adopting the above structure, the aero-engine adopts the vortex reducer as above, so that the stability and reliability of the cooling system of the aero-engine can be improved, and thus the performance and reliability of the aero-engine can be improved.

[0025] Preferably, the aero-engine further comprises a sealing device, the sealing device is arranged between the inlet segment of the airflow channel and the outlet segment of the airflow channel.

[0026] In the present scheme, by adopting the above structure, the sealing device makes the cooling airflow flow into the airflow channel more completely, so that the utilization rate of the cooling airflow can be improved, and the cooling airflow can be prevented from flowing away directly.

[0027] Preferably, the aero-engine further comprises an airflow guide pipe, and the gas flowing out of the airflow channel flows into the airflow guide pipe.

[0028] In the present scheme, by adopting the above structure, the airflow guide pipe can better guide the flow of the cooling airflow, and the cooling airflow can be prevented from mixing with the radial airflow.

[0029] On the basis of common general knowledge in the art, the above-mentioned preferred conditions can be combined arbitrarily, and thus each preferred example of the present application is obtained.

[0030] The positive progress effect of the present application is that:

[0031] The present application sets the airflow channel in the vortex reducer, the cooling airflow introduced from the compressor can flow into the airflow channel, the cooling airflow exchanges heat with the radial airflow flowing in the vortex pipe, so that the temperature of the radial airflow can be effectively reduced, and thus the radial airflow can maintain a lower temperature, and the performance of the cooling system of the aero-engine is ensured. BRIEF DESCRIPTION OF DRAWINGS

[0032] Figure 1 FIG. 1 is a structural schematic diagram of an aero-engine according to a preferred embodiment of the present application, in which only the vortex reducer and the rotor disc are shown.

[0033] Figure 2 is a sectional view of the structure of an aircraft engine in Figure 1 .

[0034] Figure 3 is a sectional view of the structure of a vortex reducer in Figure 1 , in which the flow path of the cooling air flow is shown.

[0035] Figure 4 is a sectional view of the structure of a vortex reducer in Figure 1 , in which the coiled tube as a whole is in the shape of a spiral.

[0036] BRIEF DESCRIPTION OF DRAWINGS

[0037] aircraft engine 100

[0038] seal device 11

[0039] air flow guide pipe 12

[0040] rotor disc 13

[0041] bleed air hole 14

[0042] vortex reducer 20

[0043] vortex reduction tube 21

[0044] first end 211

[0045] second end 212

[0046] tube wall 213

[0047] support ring 22

[0048] air flow passage 30

[0049] inlet section 31

[0050] cooling section 32

[0051] straight-through tube 321

[0052] coiled tube 322

[0053] outlet section 33

[0054] radial air flow A

[0055] cooling air flow B

[0056] axis C DETAILED DESCRIPTION

[0057] The application will be more fully understood and appreciated, as the same becomes better understood, by reference to the following exemplary, non-limiting detailed description of the application taken in connection with the accompanying drawings in which:

[0058] As Figures 1 to 2As shown in the embodiments, the present application provides an aero-engine 100, the aero-engine 100 comprises the vortex reducer 20 as described above, the cooling airflow B introduced from the compressor of the aero-engine 100 flows along the airflow channel 30 of the vortex reducer 20. The aero-engine 100 adopts the vortex reducer 20 as described above, so that the stability and reliability of the cooling system of the aero-engine 100 can be improved, and the performance and reliability of the aero-engine 100 can be improved.

[0059] In Figure 1 and Figure 2 , only the vortex reducer 20, the rotor disc 13, the sealing device 11, the airflow guide pipe 12 and the axis of the aero-engine 100 are shown in the figures, and other components of the aero-engine 100 are not shown in the figures. The radial airflow A flows into the vortex pipe 21 of the vortex reducer 20 from the bleed hole 14. The radial airflow A and the cooling airflow B in this article are only used to distinguish the two different airflow directions, and in the actual running aero-engine 100, the radial airflow A and the cooling airflow B can both be air.

[0060] As shown in Figure 3 and Figure 4 , the vortex reducer 20 as described above is shown in the figures, the vortex reducer 20 comprises the vortex pipe 21 and the support ring 22, the vortex pipe 21 is inserted into the support ring 22, and the vortex reducer 20 is also provided with the airflow channel 30, the airflow channel 30 sequentially flows through the support ring 22, the pipe wall 213 of the vortex pipe 21 and the support ring 22. By arranging the airflow channel 30 in the vortex reducer 20, the cooling airflow B introduced from the compressor can flow into the airflow channel 30, the cooling airflow B exchanges heat with the radial airflow A flowing through the vortex pipe 21, so that the temperature of the radial airflow A can be effectively reduced, and the radial airflow A can be kept at a lower temperature, and the performance of the cooling system of the aero-engine 100 is also ensured.

[0061] As an embodiment, the airflow channel 30 comprises an inlet section 31, a cooling section 32 and an outlet section 33, the inlet section 31 and the outlet section 33 are arranged on the support ring 22, and the cooling section 32 is arranged on the pipe wall 213. The structure and function of the airflow channel 30 are relatively concentrated, the design of the airflow channel 30 can be simplified, and the cooling effect of the radial airflow A can be improved. As shown in Figure 3 , the inlet section 31 and the outlet section 33 can both be in the shape of a broken line, and the inlet section 31 and the outlet section 33 are arranged in reference to the shape of the support ring 22.

[0062] In one specific implementation, the wall 213 of the anti-vortex tube 21 can be thickened, and correspondingly, an airflow channel 30 can be provided within the wall 213. The anti-vortex tube 21 can be manufactured by casting, with the airflow channel 30 pre-set within the wall 213. In other embodiments, the wall 213 of the anti-vortex tube 21 can also have a hollow structure, and the cooling section 32 can be composed of pipes housed within the hollow structure.

[0063] Preferably, the cooling section 32 includes a straight pipe 321 and a coiled pipe 322 connected to each other. The straight pipe 321 extends from the first end 211 of the vortex reducer 21 to the second end 212 of the vortex reducer 21, and the coiled pipe 322 extends from the second end 212 of the vortex reducer 21 to the first end 211 of the vortex reducer 21. The first end 211 of the vortex reducer 21 is the end of the vortex reducer 21 closer to the support ring 22, and the second end 212 of the vortex reducer 21 is the end of the vortex reducer 21 further away from the support ring 22. The cooling airflow B flows directly through the straight pipe 321 to the second end 212 of the vortex reducer 21, and then flows into the coiled pipe 322 from the second end 212. This results in a relatively lower temperature for the cooling airflow B within the coiled pipe 322 at the second end 212, leading to a larger temperature difference with the radial airflow A. This allows for more efficient and rapid heat exchange with the radial airflow A, and more effectively reduces the temperature of the radial airflow A. Cooling airflow B flows from the second end 212 to the first end 211. Cooling airflow B and radial airflow A continuously exchange heat. The temperature of cooling airflow B gradually rises, while the temperature of radial airflow A gradually and evenly decreases. This makes the temperature of radial airflow A more stable, avoids excessive temperature difference between radial airflow A near pipe wall 213 and radial airflow A far from pipe wall 213, and also improves the stability and reliability of the cooling system of the aero-engine 100.

[0064] In other embodiments, the cooling section 32 may not have a straight pipe 321, or the cooling section 32 may only include a coiled pipe 322.

[0065] As one implementation method, such as Figure 4 As shown, the coiled tube 322 is spiral-shaped overall, extending from the second end 212 to the first end 211 of the vortex reducer 21. This simplifies the structure of the coiled tube 322, increases the heat exchange area between the cooling airflow B and the radial airflow A, and improves the cooling effect of the radial airflow A. In other embodiments, the coiled tube 322 may include several U-shaped tubes connected end-to-end, arranged along the axial direction of the vortex reducer 21. Adjacent U-shaped tubes are inverted and connected to each other. The length of the U-shaped tubes extends from the first end 211 to the second end 212 of the vortex reducer 21. The cooling section 32 may also be zigzag-shaped overall, extending from the first end 211 to the second end 212 of the vortex reducer 21 in a reciprocating manner.

[0066] The inlet section 31 and the outlet section 33 are arranged along the axial direction of the support ring 22, which facilitates the flow of the cooling airflow B into the inlet section 31 and the flow of the cooling airflow B out of the outlet section 33 and then to other areas. The inlet of the inlet section 31 can also be in the shape of a bell mouth, which facilitates the better flow of the cooling airflow B into the airflow channel 30.

[0067] The vortex reducer 20 includes a plurality of vortex reduction pipes 21, which are arranged along the circumferential direction of the support ring 22. The airflow channel 30 is arranged in one-to-one correspondence with the vortex reduction pipes 21, which can more uniformly and effectively reduce the temperature of the radial airflow A and improve the stability and reliability of the cooling system of the aero-engine 100.

[0068] The airflow channel 30 can also include a communication ring section, which is arranged along the circumferential direction of the support ring 22 and is in communication with the inlet section 31 of the airflow channel 30 or the outlet section 33 of the airflow channel 30. The number of the communication ring section can also be two, and the two communication ring sections are in communication with the inlet section 31 of the airflow channel 30 and the outlet section 33 of the airflow channel 30, respectively.

[0069] As shown in Figure 2 , the aero-engine 100 also includes a sealing device 11 arranged between the inlet section 31 of the airflow channel 30 and the outlet section 33 of the airflow channel 30. The sealing device 11 makes the cooling airflow B flow into the airflow channel 30 more thoroughly, which can improve the utilization rate of the cooling airflow B and avoid the direct flow of the cooling airflow B.

[0070] In Figure 2 , the aero-engine 100 also includes an airflow guide pipe 12, into which the gas flowing out of the airflow channel 30 flows. The airflow guide pipe 12 can better guide the flow of the cooling airflow B and avoid the mixing of the cooling airflow B and the radial airflow A. As shown in Figure 2 , the outer side of one end of the airflow guide pipe 12 is outwardly protruded, and the protruded outer side is in abutment with the inner side of the support ring 22, thereby dividing the support ring 22 into two spaces that are not in communication with each other. After the cooling airflow B flows out of the outlet section 33, the cooling airflow B then flows into the interior of the airflow guide pipe 12. After the radial airflow A flows out of the vortex reduction pipe 21, the radial airflow A is blocked by the airflow guide pipe 12 and flows outside the airflow guide pipe 12. The cooling airflow B and the radial airflow A cannot mix through the airflow guide pipe 12.

[0071] The air flow passage 30 is formed on the support ring 22 of the vortex reducer 20 and the wall 213 of the vortex reduction pipe 21 to introduce the air, i.e. the cooling air flow B, from the inner cavity of the compressor, the cooling air flow B flows into the wall 213 of the vortex reduction pipe 21 through the support ring 22, the cooling air flow B exchanges heat with the vortex reduction pipe 21, and the radial air flow A flowing into the vortex reduction pipe 21 is secondarily cooled, so that the radial air flow A maintains a lower temperature, the performance of the cooling system of the aero-engine 100 is ensured, and the utilization rate of the air in the inner cavity of the compressor is improved.

[0072] The vortex reducer 20 improves the utilization rate of the air introduced from the front end of the compressor, so that the front end introduced air can not only ensure the sealing pressure of the bearing at the rear end, but also can secondarily cool the air in the vortex reduction pipe 21.

[0073] The vortex reduction device cools the vortex reduction pipe 21 by using the cooling air in the inner cavity of the compressor, so as to reduce the temperature of the air flow introduced by the vortex reduction pipe 21 to the subsequent high-temperature components, ensure the cooling effect of the hot end components, and improve the utilization rate of the air introduced by the front end of the compressor.

[0074] In the Figure 1 , the vortex reducer 20 is installed between the two rotor discs 13. The vortex reducer 20 includes a vortex reduction pipe 21 and a support ring 22. The support ring 22 of the vortex reducer 20 is provided with an inlet section 31 and an outlet section 33, the wall 213 of the vortex reduction pipe 21 has a certain thickness, a cooling section 32 is formed in the wall 213 of the vortex reduction pipe 21, a sealing device 11 is arranged between the inlet section 31 and the outlet section 33, and the sealing device 11 can be a sealing ring. The cooling hole flowing out of the outlet section 33 of the vortex reducer 20 enters the air flow guide pipe 12. The rotor disc 13 is provided with an air introduction hole 14, and the radial air flow A flows into the vortex reduction pipe 21 through the air introduction hole 14.

[0075] In the Figure 2 and Figure 3 , the radial air flow A is introduced into the vortex reduction pipe 21 from the drum cavity outside through the air introduction hole 14, flows through the vortex reduction pipe 21, and is introduced into the high-temperature components at the rear end from the outside of the air flow guide pipe 12. The cooling air flow B flows into the air flow passage 30 from the inlet section 31 of the support ring 22, then flows into the cooling section 32 in the wall 213 of the vortex reduction pipe 21, circulates in the wall 213 for one cycle, and then flows to the outlet section 33, and flows out of the vortex reducer 20 from the outlet section 33. Since the sealing device 11 is arranged between the inlet section 31 and the outlet section 33, the cooling air flow B introduced into the inner cavity from the front end of the aero-engine 100 can only flow into the inlet section 31, circulate in the vortex reducer 20 for one cycle, and then flow out of the outlet section 33. The cooling air B flowing out of the outlet section 33 enters the air flow guide pipe 12, and flows out along the air flow guide pipe 12. Part of the cooling air seals the bearing cavity at the rear end, and the other part of the excess cooling air is discharged to the outside of the aero-engine 100.

[0076] In Figure 3 In the embodiment shown in Fig. 1, the cooling air flow B introduced from the front end of the compressor flows into the drum interior through the inlet section 31 of the support ring 22, flows into the cooling section 32 of the vortex-reducing tube 21 in the direction of the arrow, circulates one round along the tube wall 213 of the vortex-reducing tube 21, and then flows out through the outlet section 33 of the support ring 22. The cooling air flow B introduced from the front end of the compressor has a very low temperature, and the temperature of the cooling air flow B is much lower than the temperature of the radial air flow A. The cooling air flow B cools the vortex-reducing tube 21, and the cooled vortex-reducing tube 21 exchanges heat with the radial air flow A, so that the temperature of the radial air flow A is subsequently reduced, and the radial air flow A subsequently flows into the rear end high-temperature component to be cooled.

[0077] Although the specific embodiments of the present application are described above, it should be understood by those skilled in the art that this is only an example, and the protection scope of the present application is defined by the appended claims. Those skilled in the art can make various changes or modifications to these embodiments without departing from the principles and essence of the present application, and these changes and modifications all fall within the protection scope of the present application.

Claims

1. A vortex suppressor, comprising a vortex suppressor tube and a support ring, wherein the vortex suppressor tube is inserted into the support ring, and radial airflow flows into the vortex suppressor tube, characterized in that, The vortex reducer is also provided with an airflow channel, which flows sequentially through the support ring, the tube wall of the vortex reducer tube, and the support ring. The airflow channel is used to introduce cooling airflow from the compressor. The cooling airflow flows along the airflow channel and exchanges heat with the radial airflow flowing inside the vortex reducer tube, thereby cooling the radial airflow. The airflow channel includes an inlet section, a cooling section, and an outlet section. The inlet section and the outlet section are both located on the support ring, and the cooling section is located on the pipe wall.

2. The vortex suppressor as described in claim 1, characterized in that, The cooling section includes a straight pipe and a coiled pipe connected together. The straight pipe extends from the first end of the vortex reducer to the second end of the vortex reducer, and the coiled pipe extends from the second end of the vortex reducer to the first end of the vortex reducer. The first end of the vortex reducer is the end of the vortex reducer that is close to the support ring, and the second end of the vortex reducer is the end of the vortex reducer that is away from the support ring.

3. The vortex suppressor as described in claim 2, characterized in that, The coiled tube is spiral in shape overall; Alternatively, the coiled tube may include a plurality of U-shaped tubes connected end to end, the U-shaped tubes being arranged along the axial direction of the vortex-reducing tube; Alternatively, the cooling section may be zigzag-shaped as a whole.

4. The vortex suppressor as described in claim 1, characterized in that, The inlet section and the outlet section are arranged sequentially along the axial direction of the support ring.

5. The vortex suppressor as described in any one of claims 1-4, characterized in that, The vortex reducer includes a plurality of vortex reducer tubes, which are arranged at circumferential intervals along the support ring, and the airflow channel is arranged correspondingly to each of the vortex reducer tubes.

6. The vortex suppressor as described in claim 5, characterized in that, The airflow channel further includes a connecting loop segment, which is arranged circumferentially along the support ring; The connecting loop segment is connected to either the inlet segment or the outlet segment of the airflow channel; Alternatively, there may be two connecting loop segments, which respectively connect the inlet segment of the airflow channel and the outlet segment of the airflow channel.

7. An aircraft engine, characterized in that, The aero-engine includes a vortex reducer as described in any one of claims 1-6, wherein a cooling airflow introduced from the compressor of the aero-engine flows along the airflow passage.

8. The aero-engine as described in claim 7, characterized in that, The aircraft engine also includes a sealing device, which is located between the inlet section and the outlet section of the airflow passage.

9. The aero-engine as described in claim 7, characterized in that, The aircraft engine also includes an airflow duct, through which gas flowing out of the airflow channel flows into the airflow duct.

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