Aeroengine and its compressor transition section

Abstract

The application relates to an aero-engine and a compressor transition section thereof. The compressor transition section comprises a body and a jet device. The body comprises a casing, a hub and a plurality of blades. The casing is sleeved outside the hub and forms a through-flow cavity between the casing and the hub. The plurality of blades are arranged at intervals along the circumference of the hub and are connected between the casing and the hub. The jet device comprises an air guide port, an air guide flow channel and a jet port. The air guide port is arranged on the surface of the blade and communicates with the through-flow cavity. The air guide flow channel is arranged inside the blade. The jet port is arranged on the hub and is located on the side of the blade where the air guide flow channel is located along the circumference of the hub. The jet port communicates with the air guide port through the air guide flow channel. Therefore, the flow loss of the compressor transition section can be effectively reduced.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aero-engines, in particular to an aero-engine and a compressor transition section thereof. BACKGROUND

[0002] In an aero-engine, a compression system usually comprises a booster stage, a compressor transition section and a high-pressure compressor, the compressor transition section being connected between the booster stage and the high-pressure compressor for realizing airflow transition between the booster stage and the high-pressure compressor.

[0003] In the working process, the flow loss of the compressor transition section is large, which affects the aerodynamic performance of the compression system and reduces the efficiency of the compression system. SUMMARY

[0004] One technical problem to be solved by the present application is to reduce the flow loss of the compressor transition section of an aero-engine.

[0005] To solve the above technical problem, the present application provides a compressor transition section, comprising:

[0006] a body comprising a casing, a hub and a plurality of blades, the casing being sleeved outside the hub and forming a through-flow cavity with the hub, the plurality of blades being arranged in a circumferential direction of the hub and connected between the casing and the hub; and

[0007] a jet device comprising a bleed port, a bleed flow channel and a jet port, the bleed port being arranged on a surface of the blade and communicating with the through-flow cavity, the bleed flow channel being arranged inside the blade, and the jet port being arranged on the hub and located on one side of the blade where the bleed flow channel is located in the circumferential direction of the hub, the jet port communicating with the bleed port through the bleed flow channel.

[0008] In some embodiments, in the axial direction of the hub, the jet port is located between the leading edge and the trailing edge of the blade, and the axial distance between the jet port and the leading edge is 45%-65% of the axial distance between the leading edge and the trailing edge.

[0009] In some embodiments, the axial distance between the jet port and the leading edge is 50% of the axial distance between the leading edge and the trailing edge.

[0010] In some embodiments, the included angle γ between the jet direction of the jet port and the tangent of the hub at the jet port is greater than 0° and less than or equal to 20°.

[0011] In some embodiments, the jet device comprises two jet ports, the two jet ports being located on opposite sides of the blade where the bleed flow channel is located in the circumferential direction of the hub and communicating with the same bleed flow channel.

[0012] In some embodiments, the jet device comprises a shunt pipe, and the two jet ports communicate with the bleed flow channel through the shunt pipe.

[0013] In some embodiments, the bleed port is located on the pressure surface of the blade.

[0014] In some embodiments, the bleed port is located at 50% to 60% of the blade height of the blade.

[0015] In some embodiments, one jet device is arranged every other blade among all the blades.

[0016] The application further provides an aero-engine, which comprises a booster stage and a high-pressure compressor, and further comprises a compressor transition section of the embodiments of the application, the compressor transition section connecting the booster stage and the high-pressure compressor.

[0017] In the application, the jet device can effectively reduce the flow loss of the compressor transition section by performing self-bleed jet, improve the aerodynamic performance of the compression system, and improve the efficiency of the compression system.

[0018] Other features and advantages of the application will become apparent from the following detailed description of exemplary embodiments of the application with reference to the drawings. BRIEF DESCRIPTION OF DRAWINGS

[0019] In order to more clearly illustrate the technical solutions in the embodiments of the application or the prior art, the drawings needed to be used in the embodiments or prior art description will be briefly introduced below. Obviously, the drawings in the following description are only some embodiments of the application, and other drawings can be obtained by those skilled in the art without creative labor under the premise of the drawings.

[0020] Figure 1 It is a schematic diagram of an aero-engine.

[0021] Figure 2 It is a three-dimensional schematic diagram of the compressor transition section in the embodiments of the application.

[0022] Figure 3 It is a sectional view when cutting along a surface perpendicular to the axial direction. Figure 2

[0023] It is a sectional view when cutting along a surface perpendicular to the axial direction. Figure 4 Figure 2 It is a sectional view when cutting along a surface perpendicular to the axial direction.

[0024] Figure 5 Figure 2 It is a sectional view when cutting along a surface perpendicular to the axial direction.

[0025] Figure 6 It is a sectional view when cutting along a surface perpendicular to the axial direction. Figure 5

[0026] ​​​Figure 7 Fig. 1 is a structural schematic diagram of a blade in an embodiment of the present application.

[0027] Legend of reference signs:

[0028] 100, an aero-engine; 10, a fan; 20, a booster stage; 30, a compressor transition section; 40, a high-pressure compressor;

[0029] 1, a body; 11, a casing; 12, a hub; 13, a throughflow cavity; 14, a blade; 15, a pressure surface; 16, a suction surface; 17, a leading edge; 18, a trailing edge;

[0030] 2, a fluidic device; 21, an air bleed port; 22, an air bleed passage; 23, a fluidic port; 24, a flow divider; 25, a pipe joint; 26, an expansion nut; 27, a connecting device. DETAILED DESCRIPTION

[0031] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments of the present application. The following description of at least one exemplary embodiment is merely illustrative in nature and in no way should be taken as any limitation of the present application and its applications or uses. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without any creative effort are within the scope of protection of the present application.

[0032] The technologies, methods and devices known to those of ordinary skill in the relevant art can not be discussed in detail, but should be considered as part of the specification when appropriate.

[0033] In the description of the present application, it should be understood that the orientation words such as "front, back, up, down, left, right", "transverse, vertical, perpendicular, horizontal" and "top, bottom" and the like indicate the orientation or position relationship shown in the drawings, which are only for the convenience of describing the present application and simplifying the description, and do not indicate and imply that the indicated devices or elements must have a particular orientation or be constructed and operated in a particular orientation, therefore cannot be understood as a limitation on the scope of protection of the present application; the orientation words "inner, outer" refer to the inner and outer relative to the contour of each component.

[0034] In the description of the present application, it should be understood that the use of the words "first", "second" and the like to define parts is only for the convenience of distinguishing the corresponding parts, and the above words have no special meaning unless otherwise stated, therefore cannot be understood as a limitation on the scope of protection of the present application.

[0035] In addition, the technical features involved in the different embodiments of the present application described below can be combined with each other as long as there is no conflict.

[0036] Figure 1 Part of the structure of the aircraft engine is shown.

[0037] Referring to Figure 1 , the aircraft engine 100 comprises a fan 10, a booster stage 20, a compressor transition section 30 and a high pressure compressor 40 arranged in sequence along the gas flow direction. Among them, the compressor transition section 30 connects the booster stage 20 and the high pressure compressor 40, and is used to realize the flow transition between the booster stage 20 and the high pressure compressor 40. The compressor transition section 30, together with the booster stage 20 and the high pressure compressor 40, constitutes the compression system of the aircraft engine 100.

[0038] Because the rotational speed of the booster stage 20 is the same as that of the fan 10, and the booster stage 20 is "forced" to work at a lower working rotational speed due to the limitation of the tip tangential velocity of the fan 10, in order to obtain a better boosting effect, the booster stage 20 is usually higher than the high pressure compressor 40 in the radial direction to increase the tip tangential velocity and improve the working capacity of the airflow, so as to achieve a better boosting effect.

[0039] Because the booster stage 20 is higher than the high pressure compressor 40 in the radial direction, there is a large radial height difference between the inlet and the outlet of the compressor transition section 30, and in order to reduce the weight, the axial distance between the inlet and the outlet of the compressor transition section 30 is usually short, which means that the airflow needs to flow from a position with a high radial height to a position with a low radial height through a short axial distance when flowing through the compressor transition section 30. In this process, flow separation inevitably occurs, and the corresponding flow separation will cause flow loss, affect the aerodynamic performance of the entire compression system, and reduce the efficiency of the entire compression system.

[0040] And referring to Figure 2 , in order to meet the force bearing requirement and achieve the flow guiding purpose, a plurality of blades 14 (also known as struts) are usually arranged between the casing 11 and the hub 12 of the compressor transition section 30. These blades 14 and the transition area (referred to as the corner area) near the casing 11 and the hub 12 have corner vortices, which will further aggravate the flow separation, increase the flow loss, affect the aerodynamic performance of the entire compression system, and reduce the efficiency of the entire compression system. In particular, the flow separation phenomenon is more serious in the corner area between the root of the blade 14 and the hub 12 due to the large adverse pressure gradient and the influence of the boundary layer. Among them, the corner vortex refers to the vortex structure formed by the airflow in the corner area.

[0041] In addition, the unsteady separation vortex existing in the compressor transition section 30 is located at the inlet of the high-pressure compressor 40, and also affects the flow field quality of the high-pressure compressor 40, so that the aerodynamic performance of the entire compression system is poor, and the efficiency is reduced. The unsteady separation vortex refers to an unsteady vortex, which is variable in time and space.

[0042] It can be seen that the flow separation in the compressor transition section 30 is serious, the flow loss is large, the aerodynamic performance is poor, and the efficiency is low, which needs to be further improved.

[0043] In view of the above situation, the structure of the compressor transition section 30 is improved in the application to suppress flow separation, reduce flow loss, improve aerodynamic performance, and improve efficiency.

[0044] Figures 2-7 The structure of the compressor transition section 30 of the application is exemplarily shown.

[0045] Referring to Figures 2-7 In the application, the compressor transition section 30 includes a body 1 and a jet device 2. The body 1 includes a casing 11, a hub 12, and a plurality of blades 14, the casing 11 is sleeved outside the hub 12 and forms a through-flow cavity 13 with the hub 12, and the plurality of blades 14 are arranged in a circumferential direction of the hub 12 and are connected between the casing 11 and the hub 12. The jet device 2 includes an air bleeding port 21, an air bleeding flow passage 22, and a jet port 23, the air bleeding port 21 is arranged on the surface of the blade 14 and communicates with the through-flow cavity 13, the air bleeding flow passage 22 is arranged inside the blade 14, the jet port 23 is arranged on the hub 12 and is located on one side of the blade 14 where the air bleeding flow passage 22 is located in the circumferential direction of the hub 12, and the jet port 23 communicates with the air bleeding port 21 through the air bleeding flow passage 22.

[0046] In the above arrangement, since the air bleeding port 21 of the jet device 2 located on the surface of the blade 14 and the air bleeding flow passage 22 located inside the blade 14 communicate the through-flow cavity 13 located between the casing 11 and the hub 12 with the jet port 23 located on the hub 12, during operation, the airflow entering the through-flow cavity 13 and flowing through the surface of the blade 14 can flow to the jet port 23 through the air bleeding port 21 and the air bleeding flow passage 22, and be ejected from the jet port 23 to the vicinity of the outer surface of the hub 12 to form a jet, which plays a role of an aerodynamic vortex generator, injects kinetic energy to the low-energy gas near the outer surface of the hub 12 which is seriously divided due to the boundary layer and large adverse pressure gradient, and plays a role of pushing the airflow with low kinetic energy near the outer surface of the hub 12, so that the kinetic energy of the corresponding airflow with low kinetic energy becomes large, and the airflow can flow downstream more smoothly. In this way, the flow separation in the corner region (referred to as corner separation) between the blade 14 and the hub 12 can be effectively suppressed, so that the flow loss can be effectively reduced, the aerodynamic performance of the entire compression system can be improved, and the efficiency of the entire compression system can be improved.

[0047] And, since the high pressure gas used by the jet flow device 2 to form the jet flow is from the gas flowing through the compressor transition section 30 itself, rather than from other places such as the high pressure compressor 40, the jet flow device 2 is a self bleed air jet flow device, which uses a self bleed air mode to achieve jet flow and suppress flow separation. Since the gas flowing through the surface of the blade 14 is high pressure gas, the gas flowing through the surface of the blade 14 can smoothly reach the jet flow port 23 through the bleed air port 21 and the bleed air flow channel 22, and be ejected to form a jet flow.

[0048] Since the jet flow is achieved by using a self bleed air mode, there is no need to introduce external high pressure gas, so the problem of high pressure gas source can be cleverly solved. On the one hand, since there is no need to additionally provide a high pressure gas source outside the compressor transition section 30, and there is no need to additionally provide a pipeline to introduce the high pressure gas from the high pressure gas source to the compressor transition section 30, and there is no need to provide a valve or other control mechanism to control whether to introduce bleed air, the structure is relatively simple and the cost is relatively low. On the other hand, since the flow path from the external high pressure gas source to the compressor transition section 30 can be omitted, the gas path of the high pressure gas flowing to the jet flow port is shortened, so the response speed is faster and the flow separation can be suppressed more timely. On the other hand, since the bleed air is not introduced into the compressor transition section 30, the aerodynamic performance of other mechanisms will not be affected. For example, when the high pressure compressor 40 is used as the high pressure gas source, since the high pressure gas in the high pressure compressor 40 needs to be introduced into the compressor transition section 30, the aerodynamic performance of the high pressure compressor 40 will inevitably be affected. However, since the present application does not need to introduce bleed air from the high pressure compressor 40, the aerodynamic performance of the high pressure compressor 40 will not be affected by the introduction of bleed air into the compressor transition section 30.

[0049] At the same time, when the self bleed air mode is used, the bleed air introduced from the through-flow cavity 13 can eventually return to the through-flow cavity 13 and continue to flow downstream, so the corresponding bleed air process has little effect on the flow rate of the compressor transition section 30 itself and will not cause a large flow loss of the compressor transition section 30.

[0050] As can be seen, the jet flow device 2 provided can effectively suppress the flow separation, especially the corner separation, of the compressor transition section 30 based on a relatively simple structure and a relatively low cost, can effectively reduce the flow loss, improve the aerodynamic performance, and improve the efficiency. Moreover, the jet flow device 2 provided has a fast response speed and little adverse effect, especially little adverse effect on the flow rate of the high pressure compressor 40 and the compressor transition section 30, and will not cause a large flow loss of the high pressure compressor 40 and the compressor transition section 30. In general, the jet flow device 2 provided can effectively reduce the flow loss of the compressor transition section 30, improve the aerodynamic performance of the compression system, and improve the efficiency of the compression system at a relatively small cost.

[0051] The arrangement mode of the bleed air port 21 on the blade 14 can be various.

[0052] For example, the position of the bleed port 21 in the thickness direction of the blade 14 is not limited, and the bleed port 21 can be located on the pressure surface 15 of the blade 14 or on the suction surface 16 of the blade 14. As an example, see Figures 2-7 , the bleed port 21 is located on the pressure surface 15 of the blade 14. Since the pressure of the gas flowing through the pressure surface 15 is greater than that of the gas flowing through the suction surface 16, the bleed port 21 is arranged on the pressure surface 15 to introduce the high-pressure gas flowing through the pressure surface 15 to the jet port 23 to form a jet, which can inject more energy to the low-energy gas, thereby more effectively suppressing flow separation, reducing flow loss, improving aerodynamic performance, and improving efficiency.

[0053] For another example, the position of the bleed port 21 in the height direction of the blade 14 (i.e., the blade height, which is also the radial direction) is not limited. As an example, see Figure 7 , the bleed port 21 is located at 50%-60% of the blade height of the blade 14. In the height direction of the blade 14, the flow condition of the gas at 50%-60% of the blade height is better, so that the bleed port 21 is arranged at 50%-60% of the blade height to minimize the impact of the bleed on the normal flow in the compressor transition section 30. At the same time, the pressure of the gas at 50%-60% of the blade height also meets the jet demand, which can effectively suppress the flow separation of the compressor transition section 30.

[0054] In addition, the arrangement of the jet port 23 on the hub 12 can also be various.

[0055] For example, the axial position of the jet port 23 can have various arrangements. As an example, in the axial direction of the hub 12, the jet port 23 is located between the leading edge 17 and the trailing edge 18 of the blade 14, and the axial distance between the jet port 23 and the leading edge 17 is 45%-65% of the axial distance between the leading edge 17 and the trailing edge 18. It can be understood that the leading edge 17 is the intersection corner part of the pressure surface 15 and the suction surface 16 on the front side (upstream in the airflow direction). The trailing edge 18 is the intersection corner part of the pressure surface 15 and the suction surface 16 on the rear side (downstream in the airflow direction). The axial distance between the leading edge 17 and the trailing edge 18 refers to the axial distance between the intersection point of the leading edge 17 and the hub 12 and the intersection point of the trailing edge 18 and the hub 12.

[0056] The jet flow port 23 is arranged between the leading edge 17 and the trailing edge 18 of the blade 14, and the axial distance between the jet flow port 23 and the leading edge 17 is 45%-65% of the axial distance between the leading edge 17 and the trailing edge 18, which can more effectively suppress the flow separation. It is found that the flow separation of the compressor transition section 30 mainly occurs at the position of 70%-75% of the axial distance between the leading edge 17 and the trailing edge 18 (referred to as the main separation zone), and therefore, the jet flow port 23 is arranged at the position of 45%-65% of the axial distance between the leading edge 17 and the trailing edge 18, so that the jet flow port 23 is located upstream of the main separation zone, and the jet flow emitted through the jet flow port 23 can flow for a distance before reaching the main separation zone, and in this process, the jet flow can itself form a streamwise vortex in advance, so that when reaching the main separation zone, the jet flow can more fully mix with the low-energy flow and more effectively push the low-energy flow to flow, thereby more effectively suppressing the flow separation and reducing the flow loss. It can be understood that the streamwise vortex refers to a vortex with a flow direction along the main flow direction.

[0057] For another example, the jet flow port 23 can have various arrangement modes of the jet direction. As an example, referring to Figure 5 and Figure 6 , the included angle γ between the jet direction of the jet flow port 23 and the tangent of the hub 12 at the jet flow port 23 is greater than 0° and less than or equal to 20°. At this time, the jet flow is roughly fan-shaped, and there is a shearing effect, which can not only induce the flow near the wall surface, but also induce the flow slightly far from the wall surface, so that the action between the jet flow and the low-energy flow can be increased, and the flow separation can be more effectively reduced. The jet flow can play a role of injecting kinetic energy into the low-energy fluid, and the shearing between the jet flow and the low-energy fluid on the wall surface can generate a streamwise vortex. By using the interaction between the streamwise vortex and the main flow and the low-energy fluid on the wall surface, the flow separation can be effectively suppressed, and the flow loss can be reduced.

[0058] In the foregoing embodiments, the number of jet flow ports 23 in the jet flow device 2 can be one, two or more.

[0059] For example, referring to Figures 2-5 , in some embodiments, the jet flow device 2 includes two jet flow ports 23, which are located on opposite sides of the blade 14 where the bleed flow channel 22 is located along the circumference of the hub 12 and communicate with the same bleed flow channel 22. Based on this, it is not necessary to open the bleed port 21 and the bleed flow channel 22 on each blade 14, but to open the bleed port 21 and the bleed flow channel 22 every other blade 14, so that the structure is simpler and the cost is lower.

[0060] In order to realize the communication between the two jet flow ports 23 and the bleed flow channel 22, referring to Figures 2-5In some embodiments, the fluidic device 2 comprises a shunt pipe 24, and the two fluidic ports 23 are communicated with the bleed flow channel 22 through the shunt pipe 24. In this way, the two fluidic ports 23 located on the two sides of the blade 14 in the circumferential direction can be conveniently communicated with the bleed flow channel 22 located inside the blade 14.

[0061] Next, further introduce Figures 2-6 the illustrated embodiment.

[0062] As Figures 2-6 shown, in this embodiment, the compressor transition section 30 comprises a body 1 and a fluidic device 2.

[0063] The body 1 is used to connect the booster stage 20 and the high-pressure compressor 40, and comprises a casing 11, a hub 12 and a plurality of blades 14. The casing 11 is sleeved outside the hub 12, and forms a through-flow cavity 13 between the casing 11 and the hub 12. All the blades 14 are arranged in the through-flow cavity 13 and are uniformly arranged along the circumferential direction of the hub 12. Each blade 14 is connected with the casing 11 and the hub 12. The lower end of each blade 14 is inserted into the hub 12.

[0064] Among all the blades 14, one fluidic device 2 is arranged every other blade 14. Each fluidic device 2 has the same structure and comprises a bleed port 21, a bleed flow channel 22, two fluidic ports 23, a shunt pipe 24 and a connecting device 27.

[0065] The bleed port 21 is arranged on the pressure surface 15 of the blade 14 and is located at 55% of the blade height of the pressure surface 15. The bleed port 21 is communicated with the through-flow cavity 13.

[0066] The bleed flow channel 22 is arranged at the middle of the thickness direction of the blade 14, the top end of the bleed flow channel 22 is communicated with the bleed port 21, and the lower end of the bleed flow channel 22 penetrates the lower end of the blade 14 to play a flow guiding role. The cross section of the bleed flow channel 22 is circular or elliptical. Since the bleed flow channel 22 is arranged inside the blade 14, the aerodynamic shape of the blade 14 is not damaged.

[0067] The two fluidic ports 23 are both arranged on the hub 12 and are located on the two sides of the blade 14 along the circumferential direction of the hub 12. Each fluidic port 23 is located between the leading edge 17 and the trailing edge 18 of the blade 14, and the axial distance between each fluidic port 23 and the leading edge 17 is 50% of the axial distance between the leading edge 17 and the trailing edge 18.

[0068] The shunt pipe 24 is connected with the blade 14 through the connecting device 27, and the two jet ports 23 are communicated with the bleed air flow channel 22. The two branches of the shunt pipe 24 are inserted into the two jet ports 23 respectively, and the head ends of the two branches of the shunt pipe 24 are gathered together and connected with the blade 14 through the connecting device 27. By adjusting the connection direction of the shunt pipe 24 in the two jet ports 23, the angle γ between the exit direction of the two jet ports 23 and the tangent of the hub 12 at the corresponding position is formed as 5°, 10°, 15° or 18°.

[0069] The connecting device 27 comprises a pipe joint 25 and a flared nut 26. The flared nut 26 is connected with the blade 14 through the pipe joint 25, and is connected with the head end of the shunt pipe 24 to realize the connection between the shunt pipe 24 and the blade 14.

[0070] Based on the above setting, the embodiment can realize the self-bleed jet flow based on the simpler structure, thereby effectively inhibiting the flow separation in the compressor transition section 30, reducing the flow loss in the compressor transition section 30, improving the aerodynamic performance of the compression system, and improving the efficiency of the compression system at a smaller cost.

[0071] The application further provides an aero-engine 100 comprising a booster stage 20 and a high-pressure compressor 40, and further comprising the compressor transition section 30 of the embodiment, which is connected between the booster stage 20 and the high-pressure compressor 40.

[0072] The above only describes the exemplary embodiments of the application, and is not used to limit the application. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the application shall be included in the protection scope of the application.

Claims

1. A compressor transition section (30) characterized by, Comprise: a body (1) comprising a casing (11), a hub (12) and a plurality of blades (14), the casing (11) being arranged outside the hub (12) and forming a through-flow cavity (13) with the hub (12), and the plurality of blades (14) being arranged along the circumference of the hub (12) and being connected between the casing (11) and the hub (12); and a fluidic device (2) comprising a bleed port (21), a bleed channel (22) and a fluidic port (23), the bleed port (21) being arranged on the surface of the blade (14) and being in communication with the through-flow cavity (13), the bleed channel (22) being arranged inside the blade (14), and the fluidic port (23) being arranged on the hub (12) and being located on one side of the blade (14) where the bleed channel (22) is arranged along the circumference of the hub (12), the fluidic port (23) being in communication with the bleed port (21) through the bleed channel (22), the fluidic device (2) comprising two fluidic ports (23), the two fluidic ports (23) being located on opposite sides of the blade (14) where the bleed channel (22) is arranged along the circumference of the hub (12) and being in communication with the same bleed channel (22). In the axial direction of the hub (12), the fluidic port (23) is located between the leading edge (17) and the trailing edge (18) of the blade (14), and the axial distance between the fluidic port (23) and the leading edge (17) is 45%-65% of the axial distance between the leading edge (17) and the trailing edge (18).

2. The compressor transition duct (30) of claim 1, characterized in that, The axial distance between the fluidic port (23) and the leading edge (17) is 50% of the axial distance between the leading edge (17) and the trailing edge (18).

3. The compressor transition duct (30) of claim 2, characterized by The fluidic device (2) comprises a shunt pipe (24), and the two fluidic ports (23) are in communication with the bleed channel (22) through the shunt pipe (24).

4. The compressor transition duct (30) of claim 1, wherein, an angle between the direction of exit of the jet (23) and a tangent to the hub (12) at the jet (23) greater than 0° and less than or equal to 20°.

5. The compressor transition duct (30) according to any one of claims 1-4, characterized in that, The bleed port (21) is located on the pressure surface (15) of the blade (14).

6. The compressor transition duct (30) according to any one of claims 1-4, characterized in that The bleed port (21) is located at 50%-60% of the height of the blade (14).

7. The compressor transition duct (30) according to any one of claims 1-4, characterized in that In all blades (14), one fluidic device (2) is arranged every other blade (14).

8. The compressor transition duct (30) of any one of claims 1-4, wherein, Further comprising a compressor transition section (30) as claimed in any one of claims 1-8, the compressor transition section (30) connecting the booster stage (20) and the high-pressure compressor (40).

9. An aeroengine (100) comprising a booster stage (20) and a high pressure compressor (40), characterised in that, ​

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