Splitter ring and aircraft engine and aircraft comprising same

Abstract

The application discloses a flow splitting ring, an aero-engine and an aircraft comprising the same. The flow splitting ring is internally provided with an annular cavity. An air inlet pipe is connected to the flow splitting ring. The air inlet pipe supplies gas into the annular cavity. The flow splitting ring further comprises a flow guide. The flow guide is arranged in the annular cavity. One end of the flow guide along a gas flow direction is arranged towards a circumferential direction of the annular cavity. The other end of the flow guide along the gas flow direction is arranged corresponding to a position where the air inlet pipe is communicated with the annular cavity. The application enhances the circulation flow rate of the gas, especially the high-temperature gas supplied into the annular cavity of the flow splitting ring by the flow guide. The high-temperature gas is distributed in each position of the annular cavity. The inner wall of the annular cavity is heated by the high-temperature gas. The thermal efficiency of the flow splitting ring is increased. The efficiency of heating and removing the ice of the flow splitting ring caused by the cold air is improved. The flight safety of the aircraft is ensured.

Description

Technical Field

[0001] This invention relates to a shunt ring and an aero-engine and aircraft incorporating the same. Background Technology

[0002] Icing has a significant impact on flight safety, especially for aircraft engines. Once icing occurs, it can cause minor issues such as obstructed engine air intake and reduced engine efficiency; in severe cases, it can cause engine shutdown or even physical damage, leading to serious safety accidents. The engine splitter ring, as a physical element used to separate the airflow between the inner and outer bypass shafts after the fan, is a typical anti-icing component, and its most common anti-icing method is hot gas anti-icing.

[0003] As a ring-shaped cavity, the flow divider ring will inevitably experience circumferential temperature drop when hot air enters the flow divider ring cavity due to circumferential flow. In order to ensure the anti-icing effect while minimizing the amount of anti-icing induced air, the uniformity of heat distribution is a key technology in the design of the flow divider ring anti-icing system.

[0004] To address the issue of heat distribution uniformity, the primary technical approach currently used is multi-pipe air supply. The main characteristic of this structure is that, as the high-pressure bleed pipe approaches the splitting ring, a distribution pipe divides the original single airflow into multiple circumferentially distributed airflows. These multiple airflows then enter the splitting ring cavity through corresponding circumferentially distributed splitting pipes, thereby increasing heat uniformity. The main drawback of this structure is that while multiple splitting rings improve heat uniformity, weight limitations restrict the number of splitting pipes, resulting in a significant distance between adjacent pipes. When the airflow exits each splitting ring pipe, it primarily impacts the cavity wall directly opposite the pipe opening, leading to a higher wall temperature there. A temperature difference persists on the wall further away from the pipe opening, resulting in poor airflow uniformity. Furthermore, cold air encountered during flight can ic up on the cooler splitting ring surface, impacting flight safety. Summary of the Invention

[0005] The technical problem to be solved by the present invention is to overcome the defect of poor gas flow in the prior art, and to provide a flow divider ring and an aero-engine and aircraft containing the same.

[0006] The present invention solves the above-mentioned technical problems through the following technical solution:

[0007] A flow divider ring includes an annular cavity within the flow divider ring and an intake pipe connected to the flow divider ring. The intake pipe supplies gas to the annular cavity. The flow divider ring also includes a flow guide, which is disposed within the annular cavity. One end of the flow guide along the gas flow direction is positioned towards the circumferential direction of the annular cavity, and the other end of the flow guide along the gas flow direction is positioned at the connection point between the intake pipe and the annular cavity.

[0008] In this design, the guide vane is positioned corresponding to the air intake pipe, with one end extending axially towards the annular cavity. The gas supplied by the air intake pipe flows through the guide vane, causing the gas to flow circumferentially along the annular cavity. This improves the uniformity of gas flow within the annular cavity. Furthermore, when the gas is high-temperature gas, it can heat the inner wall of the annular cavity through which it flows, thereby achieving uniform heating of the circumference of the flow divider ring. This uniform heating of the flow divider ring removes the ice that forms on the outer periphery of the flow divider ring, ensuring the flight safety of the aircraft.

[0009] Preferably, the flow guide is connected to the inner wall of the annular cavity.

[0010] In this design, the end of the guide extends circumferentially within the annular cavity, while the side of the guide is connected to the inner wall of the annular cavity. That is, the upper and lower inner walls of the annular cavity are connected by the guide. The guide increases the support points within the annular cavity, thereby increasing the structural strength of the diversion ring. Based on this, the wall of the diversion ring can be thinner. Under the premise of ensuring strength, the inner wall of the diversion ring can be heated by high-temperature gas more quickly, increasing the rate of melting ice on the outer periphery of the diversion ring.

[0011] Preferably, multiple intake pipes are provided along the circumference of the flow divider ring, and multiple flow guides are provided in the annular cavity corresponding to the multiple intake pipes.

[0012] In this design, multiple intake pipes and guide components are provided to ensure that the gas flows in the circumferential direction within the annular cavity, thereby improving the gas's fluidity and uniformity. This allows the gas to form an airflow ring that flows in the radial direction of the diversion ring when it flows within the annular cavity, which is a preferred structural configuration of this application.

[0013] Preferably, the plurality of the flow guides are arranged along the circumferential direction of the middle portion of the annular cavity.

[0014] In this design, one end of the guide extends radially along the flow divider ring, allowing the high-temperature gas flowing in from the intake pipe to flow radially along the annular cavity within the flow divider ring. Furthermore, the guide is positioned in the middle of the annular cavity, enabling the high-temperature gas to flow and circulate around the middle of the annular cavity. During circulation, the high-temperature gas can also diffuse to both sides of the annular cavity, thereby increasing its fluidity and allowing it to circulate to all locations within the annular cavity. This enhances the heating and de-icing capabilities of the flow divider ring, ensuring the aircraft's flight safety.

[0015] Preferably, the annular cavity has a wall, which is located away from the air inlet pipe and separates the annular cavity to form a front cavity and a rear cavity. The guide is located in the rear cavity, and a channel is formed on the wall, through which the air inlet pipe supplies gas to the front cavity.

[0016] In this design, the annular cavity is further divided into a front cavity and a rear cavity. The rear cavity is connected to the air intake pipe, while the front cavity and the rear cavity are connected by a channel. The air intake pipe supplies high-temperature gas to the rear cavity, and a portion of the high-temperature gas flows to the front cavity through the channel. The inner wall of the front cavity is closer to the cold air encountered during flight than the rear cavity. In other words, the front cavity is also prone to icing when it encounters cold air, making the supply of high-temperature gas to the front cavity for de-icing very important.

[0017] Preferably, the front cavity has a tapered cross-section, and an exhaust port is provided on the inner wall of the front cavity. The rear cavity is connected to the exhaust port through the channel.

[0018] In this design, the front cavity has a tapered cross-section to better divide the air and reduce air resistance during flight. In addition, the front cavity increases thermal efficiency through impact heat exchange, that is, high-temperature gas impacts the inner wall of the front cavity through the channel to rapidly increase the temperature of the inner wall of the front cavity. The increased temperature of the front cavity removes the ice that has formed on the front edge of the splitting ring from the cold air. The high-temperature gas supplied to the front cavity is discharged through the exhaust port to avoid the high-temperature gas accumulating in the front cavity after cooling, which would cause the outer periphery of the front cavity to re-ice. This is a preferred structural design of this application.

[0019] Preferably, the channel is an air hole arranged along the axial direction of the flow divider ring, the air hole is on the inner wall of the front cavity, and the exhaust port and the air hole are located in different directions.

[0020] In this design, the vent is arranged along the axial direction of the flow divider ring and is connected to the exhaust port. High-temperature gas is supplied through the vent, while the exhaust port and the vent are located in different directions to prevent the high-temperature gas flowing into the front cavity from being discharged through the exhaust port before heating the inner wall of the front cavity, thereby extending the residence time of the high-temperature gas in the front cavity.

[0021] Preferably, there are multiple air holes, and the exhaust port is arranged vertically and located on the bottom inner wall of the front cavity. The exhaust port is provided in multiple ways corresponding to the air holes.

[0022] In this design, multiple vents are provided to enhance the thermal efficiency of heating and de-icing of the front cavity by increasing the impact of high-temperature gas on the inner wall of the front cavity. At the same time, multiple exhaust ports are also provided to facilitate the discharge of cooled gas from the front cavity after the high-temperature gas heats the inner wall of the front cavity and before the thermal efficiency decreases. This ensures that the front cavity can maintain a high temperature to remove the ice that forms on the leading edge of the splitter ring.

[0023] Preferably, the exhaust port is a through groove penetrating the inner wall of the front cavity, and the through groove is an annular through groove circumferentially surrounding the flow divider ring.

[0024] In this design, the exhaust port can also be a through groove, that is, an annular through groove around the circumference of the splitter ring, in order to improve the exhaust efficiency of the front chamber. Of course, other structures can also be selected, the purpose of which is also to remove the cooled gas, so as to ensure the continuous high temperature of the front chamber and effectively remove the ice formed on the front edge of the splitter ring.

[0025] An aircraft engine comprising the flow splitter ring described in any one of the above.

[0026] In this design, the aero-engine using the aforementioned flow divider ring has a higher efficiency in removing ice formed by cold air around the engine. The flow guide enhances the circulation efficiency of high-temperature gas within the engine's front flow divider ring, thereby improving the thermal efficiency of the flow divider ring and resulting in better de-icing performance.

[0027] An aircraft comprising an aircraft engine as described above.

[0028] The positive and progressive effects of this invention are as follows: This invention enhances the circulation flow rate of gas, especially the high-temperature gas supplied to the annular cavity of the splitting ring, through the flow guide, so that the high-temperature gas is distributed throughout the annular cavity. In turn, the high-temperature gas heats the inner wall of the annular cavity, increases the thermal efficiency of the splitting ring, improves the efficiency of heating the splitting ring and removing ice formed when the splitting ring encounters cold air, and ensures the flight safety of the aircraft. Attached Figure Description

[0029] Figure 1 This is a schematic diagram of an anti-icing system for an aircraft engine according to an embodiment of the present invention.

[0030] Figure 2 This is a schematic diagram of the working state of the shunt ring according to an embodiment of the present invention.

[0031] Figure 3 This is a schematic diagram of the structure of a flow divider ring according to an embodiment of the present invention.

[0032] Figure 4 This is a schematic diagram of the working state of a flow guide component according to an embodiment of the present invention.

[0033] Figure 5 This is a perspective view of a flow guide component according to an embodiment of the present invention.

[0034] Figure 6 This is a schematic diagram of gas flow in an annular cavity according to an embodiment of the present invention.

[0035] Figure 7 This is a side view of the gas flow in an annular cavity according to an embodiment of the present invention.

[0036] Explanation of reference numerals in the attached figures:

[0037] Nacelle 1

[0038] shunt ring 2

[0039] Annular cavity 21

[0040] Wall 211

[0041] Channel 2111

[0042] Anterior cavity 212

[0043] Postcavity 213

[0044] Flow guide 22

[0045] Exhaust port 23

[0046] Support plate 3

[0047] High pressure compressor 4

[0048] Combustion Chamber 5

[0049] High-pressure turbine 6

[0050] Air intake device 7

[0051] Control valve 8

[0052] Intake pipe 9 Detailed Implementation

[0053] The present invention will be further illustrated by way of embodiments below, but the present invention is not limited to the scope of the embodiments described herein.

[0054] like Figure 1 As shown, the aero-engine includes a core engine, a nacelle 1, and a flow divider ring 2. The flow divider ring is supported in the nacelle 1 by a support plate 3. The flow divider ring 2 is a physical element used to separate the internal and external bypass airflow after the engine fan. The core engine includes a high-pressure compressor 4, a combustion chamber 5, and a high-pressure turbine 6. The high-pressure compressor 4 further compresses the air before it is discharged into the combustion chamber 5. Fuel and air are mixed and burned, and the combustion gas is then discharged into the high-pressure turbine 6, converting energy into mechanical energy. The air temperature after being compressed by the high-pressure compressor 4 may reach 600°C. The air is drawn out from the high-pressure compressor 4 through the bleed air device 7, and then reaches the intake pipe 9 through the control valve 8, and is injected into the flow divider ring 2 through the intake pipe 9.

[0055] This embodiment provides a flow divider ring, the specific structure of which is as follows: Figure 3 , Figure 6 and Figure 7As shown, the flow divider ring 2 is an annular component with an annular cavity 21 inside. An air inlet pipe 9 is connected to the flow divider ring, supplying gas into the annular cavity 21. The gas entering the annular cavity 21 can fill the annular cavity 21. When the gas is compressed by the high-pressure compressor 4, its temperature can reach 600°C, i.e., high-temperature gas. By filling the annular cavity 21 with high-temperature gas and using the contact between the high-temperature gas and the inner wall of the annular cavity 21, the surface temperature of the flow divider ring 2 is increased. This heats the surface of the flow divider ring 2, thereby removing cold air, such as air containing water droplets or ice crystals, that adheres to the surface of the flow divider ring 2 and freezes. This achieves the effect of hot gas de-icing or heating de-icing, improving the flight safety of the aircraft.

[0056] In addition, such as Figure 3 , Figure 5 As shown, the flow divider ring 2 also includes a flow guide 22, which is disposed within the annular cavity 21. One end of the flow guide 22 along the high-temperature gas flow direction is positioned circumferentially towards the annular cavity 21, and the other end of the flow guide 22 along the high-temperature gas flow direction is positioned at the connection point between the intake pipe 9 and the annular cavity 21. An arc-shaped surface is formed between the two ends of the flow guide 22. The arc-shaped surface is used to change the direction of the airflow into the annular cavity 21. The curvature of the arc-shaped surface is specifically calculated using the formula y = 0.0009898x, where the side closest to the intake pipe 9 is the inner wall surface. 3 -0.06538x 2 +1.798x+0.16, correspondingly, the side furthest from the intake pipe 9 is the outer wall surface, and the curvature of the outer wall surface is calculated using the formula y=0.001195x. 3 -0.08208x 2 +1.984x+3.044, meaning that one end of the guide 22 extends circumferentially within the annular cavity 21. After the high-temperature gas enters the annular cavity 21, it is sprayed onto the arc-shaped surface of the guide 22. The guide 22 extends circumferentially within the annular cavity 21, allowing the high-temperature gas to flow further around the annular cavity 21, thereby increasing the flow rate of the high-temperature gas. Compared to the high-temperature gas being directly injected into the annular cavity 21, the flow distance of the high-temperature gas within the annular cavity 21 is greatly increased. At the same time, the area of ​​the inner wall of the annular cavity 21 that can be heated by the high-temperature gas is larger, improving the thermal efficiency and thermal uniformity of the splitting ring 2, and shortening the time required for the splitting ring 2 to heat and de-ic. For high-speed aircraft, flight safety is more reliably guaranteed, avoiding problems such as engine blade damage caused by ice falling into the engine fan.

[0057] Furthermore, such as Figure 3 , Figure 4As shown, the end of the guide member 22 extends circumferentially into the annular cavity 21, while the side of the guide member 22 is connected to the upper and lower inner walls of the annular cavity 21. One purpose is to improve the structural stability of the guide member 22, thereby preventing the high-temperature gas after compression from being injected too quickly into the annular cavity 21 through the intake pipe 9 and thus dispersing the guide member 22. This ensures that the guide member 22 continuously and stably guides the high-temperature gas to the circumferential direction of the annular cavity 21. Another purpose is that the upper and lower inner walls of the annular cavity 21 are connected by the guide member 22, which can serve as a support point within the annular cavity 21 to support the upper and lower inner walls. Its function is equivalent to a support column or support plate, thereby increasing the structural strength of the diversion ring 2. Based on this, the thickness of the inner wall of the diversion ring 2 can be appropriately reduced. Under the premise of ensuring strength, the inner wall of the diversion ring 2 can be heated by the high-temperature gas more quickly, improving the thermal efficiency and thermal uniformity of the diversion ring 2.

[0058] Preferably, one side of the guide member 22 is welded to the lower inner wall of the annular cavity 21, and the other side of the guide member 22 overlaps with the upper inner wall of the annular cavity 21. This ensures the structural stability of the guide member 22. During the process of changing the flow direction of the high-temperature gas, since the guide member 22 is mostly made of metal, its own temperature rises after being sprayed by the high-temperature gas. In order to ensure the structural stability of the guide member 22, the guide member 22 and the inner wall of the annular cavity 21 of the diverting ring 2 are made of the same material to prevent the guide member 22 from deforming after being heated by the high-temperature gas. This avoids flight safety problems caused by the local structural deformation of the diverting ring 2 when the diverting ring 2 continuously supplies high-temperature gas through the intake pipe 9 during high-speed flight.

[0059] like Figure 4 , Figure 6 As shown, in order to improve the thermal efficiency of the flow divider ring 2, multiple intake pipes 9 can be provided along the circumference of the flow divider ring 2. Multiple guide elements 22 are provided in the annular cavity 21 corresponding to the multiple intake pipes 9. In this embodiment, 10 intake pipes 9 are selected. The intake pipes 9 are spaced apart and simultaneously supply high-temperature gas into the annular cavity 21. The number of intake pipes 9 should not be too large to avoid the flow divider ring 2 structure being too heavy and affecting the flight of the aircraft. One end of the multiple guide elements 22 extends along the circumference of the annular cavity 21. The multiple guide elements 22 have the same guiding direction, and the other end of the guide elements 22 is in the same direction as the axis of the flow divider ring 2, so that the high-temperature gas forms a circumferential flow of high-temperature gas in the annular cavity 21, that is, a clockwise circumferential flow or a counterclockwise circumferential flow. The circumferential flow of high-temperature gas has better fluidity and thermal uniformity.

[0060] like Figure 3 , Figure 4As shown, in this embodiment, the annular cavity 21 is provided with a wall 211. The wall 211 is arranged vertically away from the air inlet pipe 9 and divides the annular cavity 21 to form a front cavity 212 and a rear cavity 213. The guide member 22 is located in the rear cavity 213. The air inlet pipe 9 communicates with the rear cavity 213. A channel 2111 is opened on the wall 211 in the horizontal direction. The front cavity 212 and the rear cavity 213 are connected through the channel 2111. The air inlet pipe 9 can supply high-temperature gas to the front cavity 212 through the channel 2111. Figure 2 As shown, the front cavity 212 corresponds to the leading edge of the outer surface of the flow divider ring 2, while the rear cavity 213 corresponds to the trailing edge of the outer surface of the flow divider ring 2. The flow rate of the high-temperature gas is improved by the guide 22, which mainly affects the heating of the trailing edge. The connection between the front cavity 212 and the rear cavity 213 is achieved through the channel 2111, which increases the flow rate of the high-temperature gas into the front cavity 212. Some of the high-temperature gas flows into the front cavity 212 through the channel 2111, thereby improving the thermal efficiency of the front cavity 212. The inner wall of the front cavity 212 is closer to the cold air encountered during flight than the rear cavity 213. In other words, the front cavity 212 is also prone to icing when it encounters cold air. De-icing the high-temperature gas supplied to the front cavity 212 is also a way to improve the thermal efficiency of the flow divider ring 2.

[0061] Meanwhile, since the front cavity 212 encounters cold air before the rear cavity 213, the cross-section of the front cavity 212 is processed into a conical shape. This is to better divide the air, thereby reducing air resistance during flight. Additionally, an exhaust port 23 is provided on the inner wall of the front cavity 212, and the rear cavity 213 is connected to the exhaust port 23 via a channel 2111. This means that the high-temperature gas supplied from the intake pipe 9 to the rear cavity 213 flows into the front cavity 212 through the channel 2111, and the high-temperature gas flowing into the front cavity 212 is sprayed onto the inner wall of the front cavity 212, heating it. In other words, the front cavity 212 increases heat through impact heat exchange. To improve the thermal efficiency of the flow divider ring 2, the high-temperature gas flowing into the front cavity 212 heats the inner wall of the front cavity 212, but the temperature of the high-temperature gas itself will decrease. The thermal efficiency of the high-temperature gas on the inner wall of the front cavity 212 becomes slower after the temperature decreases, which is not conducive to the heating and de-icing of the flow divider ring 2. At this time, the cooled high-temperature gas is discharged through the exhaust port 23 to accelerate the efficiency of the high-temperature gas circulating into the front cavity 212. The high-temperature gas is reintroduced into the front cavity 212 through the channel 2111 to achieve continuous heating and de-icing of the front cavity 212 and avoid the situation where the leading edge of the flow divider ring 2 re-ices after the high-temperature gas cools down.

[0062] Furthermore, the channel 2111 is a vent provided along the axial direction of the diversion ring 2. The vent penetrates the wall 211 and corresponds to the inner wall of the front cavity 212, so that the high-temperature gas enters the front cavity 212 and is directly sprayed onto the inner wall of the front cavity 212. The exhaust port 23 is provided along the radial direction of the diversion ring 2 and is located on the bottom inner wall of the front cavity 212. The exhaust port 23 is not provided to correspond to the vent, so as to avoid the high-temperature gas flowing into the front cavity 212 being discharged through the exhaust port 23 before heating the inner wall of the front cavity 212. This prolongs the residence time of the high-temperature gas in the front cavity 212 and improves the thermal efficiency of the diversion ring 2.

[0063] Of course, there are multiple air holes, and multiple exhaust ports 23 are provided corresponding to the air holes. By increasing the number of air holes, the intake volume of high-temperature gas impacting the inner wall of the front cavity 212 is increased, thereby increasing the heating and de-icing thermal efficiency of the front cavity 212. At the same time, multiple exhaust ports 23 are also provided to facilitate the cooling of the front cavity 212 after the high-temperature gas heats the inner wall of the front cavity 212 until the thermal efficiency decreases. The cooled gas is discharged from the front cavity 212 through the exhaust ports 23 to ensure that the front cavity 212 can maintain a high temperature to remove the ice formed on the front edge of the diversion ring 2.

[0064] In a preferred embodiment of this invention, the exhaust port 23 is a through groove that penetrates the inner wall of the front cavity 212. The through groove is an annular through groove that surrounds the circumference of the diversion ring 2, thereby improving the exhaust efficiency of the front cavity 212. The purpose is to remove the high-temperature gas after cooling, so as to ensure the continuous high temperature of the front cavity 212 and effectively remove the ice formed on the front edge of the cold air diversion ring 2. Of course, the exhaust port 23 can also be a round hole or other structures, which is the prior art and will not be described in detail here.

[0065] This embodiment also provides an aero-engine, including the aforementioned splitter ring 2. The aero-engine using the aforementioned splitter ring 2 has a higher efficiency in removing ice formed by cold air around the engine. The flow guide 22 enhances the circulation efficiency of high-temperature gas in the front splitter ring 2 of the engine, thereby improving the thermal efficiency and thermal uniformity of the splitter ring 2 and resulting in a better de-icing effect.

[0066] This embodiment also provides an aircraft, which includes the aforementioned aero-engine. The aero-engine can quickly remove the ice that forms on the flow divider ring 2 after encountering cold air. The guide 22 is used to improve the flow of high-temperature gas, thereby improving the heating and de-icing capacity of each position of the flow divider ring 2, and thus shortening the de-icing time to ensure the flight safety of the aircraft.

[0067] While specific embodiments of the present invention have been described above, those skilled in the art should understand that these are merely illustrative examples, and the scope of protection of the present invention 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 invention, but all such changes and modifications fall within the scope of protection of the present invention.

Claims

1. A flow divider ring, wherein an annular cavity is provided within the flow divider ring, and an air inlet pipe is connected to the flow divider ring, the air inlet pipe supplying gas to the annular cavity, characterized in that, The flow divider ring also includes a flow guide, which is disposed in the annular cavity. One end of the flow guide is disposed at the connection between the air inlet pipe and the annular cavity and extends in the circumferential direction of the annular cavity. An arc-shaped surface is formed between the two ends of the flow guide. The jet direction of the air inlet pipe is aligned with the arc-shaped surface. The arc-shaped surface enables the high-temperature gas air to flow around the annular cavity. The annular cavity has a wall, which is located away from the air inlet pipe and separates the annular cavity to form a front cavity and a rear cavity. The guide is located in the rear cavity, and a channel is opened on the wall. The air inlet pipe supplies gas to the front cavity through the channel. One side of the flow guide is welded to the lower inner wall of the annular cavity, and the other side of the flow guide overlaps with the upper inner wall of the annular cavity.

2. The shunt ring as described in claim 1, characterized in that, The flow guide is connected to the inner wall of the annular cavity.

3. The shunt ring as described in claim 1, characterized in that, The air intake pipes are provided in multiple ways along the circumference of the flow divider ring, and the annular cavity is provided with multiple flow guides corresponding to the multiple air intake pipes.

4. The shunt ring as described in claim 3, characterized in that, The plurality of the flow guides are arranged along the circumferential direction of the middle portion of the annular cavity.

5. The shunt ring as described in claim 1, characterized in that, The front cavity has a tapered cross-section, and an exhaust port is provided on the inner wall of the front cavity. The rear cavity is connected to the exhaust port through the channel.

6. The shunt ring as described in claim 5, characterized in that, The channel is an air hole arranged along the axial direction of the flow divider ring. The air hole is located on the inner wall of the front cavity, and the exhaust port is located in a different direction from the air hole.

7. The shunt ring as described in claim 6, characterized in that, The air vents are provided in multiple ways, and the exhaust port is arranged vertically and located on the bottom inner wall of the front cavity. The exhaust port is provided in multiple ways corresponding to the air vents.

8. The shunt ring as described in claim 6, characterized in that, The exhaust port is a through groove that penetrates the inner wall of the front cavity, and the through groove is an annular through groove that surrounds the circumference of the flow divider ring.

9. An aircraft engine, characterized in that, Includes the shunt ring as described in any one of claims 1-8.

10. An aircraft, characterized in that, It includes the aircraft engine as described in claim 9.

Images