Variable geometry axisymmetric supersonic inlet, aircraft and control method

Abstract

The present disclosure relates to the field of aviation technology, and particularly provides a variable geometry axisymmetric supersonic inlet, an aircraft and a control method to solve the problem that the variable geometry axisymmetric supersonic profile is prone to inducing flow separation due to the change of the center cone moving distance in the prior art. The variable geometry axisymmetric supersonic inlet is applied to an engine, and the variable geometry axisymmetric supersonic inlet comprises a center cone, an outer casing, a support assembly and a leakage ring. The center cone is located in the outer casing, the leading end of the center cone is in sliding connection with the inner side wall of the air inlet end of the outer casing, the support assembly is fixed to the air outlet end of the outer casing, the trailing end of the center cone is fixed to the inner side wall of the air outlet end of the outer casing through the support assembly, the leading end of the center cone is provided with a plurality of leakage grooves for fluid inflow, and the leakage ring is arranged at the leakage grooves. The variable geometry axisymmetric supersonic inlet, the aircraft and the control method provided by the present disclosure are used in the design of an aero-engine.

Description

TECHNICAL FIELD

[0001] The present disclosure relates to the field of aviation technology, and particularly relates to a variable geometry axisymmetric supersonic inlet, an aircraft and a control method. BACKGROUND

[0002] As a windward component of a power system, the performance of the axisymmetric supersonic inlet directly affects the performance of the core engine of the aircraft. In the low-altitude and low-speed flight state of the aircraft, in order to avoid the inlet from not starting, the inlet is required to have a large throat area. In the high-altitude and high-speed cruise state of the aircraft, the axisymmetric supersonic inlet is required to increase the contraction ratio of the inlet by reducing the throat area, so as to improve the compression capacity of the inlet to the flow.

[0003] At present, the axisymmetric supersonic inlet is divided into fixed geometry and variable geometry. The throat area of the fixed geometry axisymmetric supersonic inlet is not adjustable, and the through-flow capacity of the axisymmetric supersonic inlet is improved by means of a leakage slot to realize the starting in the low-speed state. This case meets the high-speed flight and low-speed flight of the aircraft. The variable geometry axisymmetric supersonic inlet generally takes the high-speed cruise state as a design point, and changes the throat area by moving the center cone. The performance of the full-speed domain is obviously improved relative to the fixed geometry axisymmetric supersonic inlet, but the inlet surface changes sharply and is easy to induce flow separation. For the variable geometry axisymmetric supersonic inlet, if the moving distance of the center cone is small, the working Mach number range is narrow, and the inlet is easy to not start in the low Mach number state or the compression capacity of the inlet to the flow is insufficient in the high Mach number state. If the moving distance of the center cone is large, the inlet surface changes sharply, and although the starting in the low Mach number state can be realized, the sharply changed surface is easy to induce flow separation.

[0004] Therefore, how to solve the problem that the sharply changed surface of the variable geometry axisymmetric supersonic inlet is easy to induce flow separation due to the moving distance of the center cone in the prior art is one of the important problems to be solved in the field. SUMMARY

[0005] Therefore, the embodiments of the present disclosure provide a variable geometry axisymmetric supersonic inlet, an aircraft and a control method to solve the problem that the sharply changed surface of the variable geometry axisymmetric supersonic inlet is easy to induce flow separation due to the moving distance of the center cone in the prior art.

[0006] According to one aspect of the present disclosure, a variable geometry axisymmetric supersonic inlet is provided, which is applied to an engine, and the variable geometry axisymmetric supersonic inlet comprises a center cone, an outer casing, a support assembly and a leakage ring, the center cone is located in the outer casing, a leading end of the center cone is in sliding connection with an inner side wall of an air inlet end of the outer casing, the support assembly is fixed at an air outlet end of the outer casing, a trailing end of the center cone is fixed on the inner side wall of the air outlet end of the outer casing through the support assembly, the leading end of the center cone is provided with a plurality of leakage grooves for fluid inflow, and the leakage ring is arranged at the leakage grooves.

[0007] When the engine is in a first state, the leakage ring is used to close all the leakage grooves.

[0008] When the engine is in a second state, the leakage ring is used to open all the leakage grooves.

[0009] According to the variable geometry axisymmetric supersonic inlet of one aspect of the present disclosure, when the engine is in a third state, the leakage ring is used to open the first two leakage grooves at the leading end of the center cone.

[0010] According to the variable geometry axisymmetric supersonic inlet of one aspect of the present disclosure, the center cone is a hollow structure, and each leakage groove at the leading end of the center cone is distributed in a matrix at equal intervals.

[0011] According to the variable geometry axisymmetric supersonic inlet of one aspect of the present disclosure, an included angle between each leakage groove and the inner side wall of the air inlet end of the outer casing is 45°.

[0012] According to the variable geometry axisymmetric supersonic inlet of one aspect of the present disclosure, a groove width of each leakage groove is 5 mm.

[0013] According to the variable geometry axisymmetric supersonic inlet of one aspect of the present disclosure, a first through hole for fluid outflow is arranged at the trailing end of the center cone.

[0014] According to the variable geometry axisymmetric supersonic inlet of one aspect of the present disclosure, a second through hole for fluid outflow is arranged at the air outlet end of the outer casing.

[0015] According to the variable geometry axisymmetric supersonic inlet of one aspect of the present disclosure, the support assembly comprises a plurality of first support members and a second support member, one end of each first support member is annularly arranged on an outer side wall of the second support member, and the other end of each first support member is annularly arranged on an inner side wall of the air inlet end of the outer casing.

[0016] According to another aspect of the present disclosure, a flying vehicle is provided, which comprises an engine and the above-mentioned variable geometry axisymmetric supersonic inlet, and the variable geometry axisymmetric supersonic inlet is applied to the engine.

[0017] The variable geometry axisymmetric supersonic inlet control method according to one aspect of the present disclosure is applied to the aircraft, and the variable geometry axisymmetric supersonic inlet control method comprises the following steps of:

[0018] controlling the engine to work in response to a first state flight instruction;

[0019] controlling the bleed ring to close all the bleed slots, and the first state flight instruction is used to indicate that the engine is in a low-speed working state;

[0020] controlling the engine to work in response to a second state flight instruction;

[0021] controlling the bleed ring to open all the bleed slots, and the second state flight instruction is used to indicate that the engine is in a medium-speed working state;

[0022] controlling the engine to work in response to a third state flight instruction;

[0023] controlling the bleed ring to open the first two bleed slots at the leading end of the center cone, and the third state flight instruction is used to indicate that the engine is in a high-speed working state.

[0024] The above at least one technical solution adopted by the embodiment of the present disclosure can achieve the following beneficial effects: the variable geometry axisymmetric supersonic inlet is applied to the engine, the center cone is located in the outer casing, the leading end of the center cone is in sliding connection with the inner side wall of the inlet end of the outer casing, the support assembly is fixed to the gas outlet end of the outer casing, the tail end of the center cone is fixed to the inner side wall of the gas outlet end of the outer casing through the support assembly, the leading end of the center cone is provided with a plurality of bleed slots for fluid inflow, and the bleed ring is arranged at the bleed slots. Based on this, when the fluid flows into the variable geometry axisymmetric supersonic inlet, the fluid flow will drive the center cone to move in the outer casing, the internal contraction ratio is increased by moving the center cone, the flow capture capability of the variable geometry axisymmetric supersonic inlet is improved, and the flow compensation is realized. Secondly, the moving of the center cone can change the throat area of the variable geometry axisymmetric supersonic inlet. On this basis, the bleed slots can eliminate the separation zone formed by the interference between the shock wave and the boundary layer in the variable geometry axisymmetric supersonic inlet, the internal contraction ratio of the variable geometry axisymmetric supersonic inlet is increased under the premise of ensuring that the variable geometry axisymmetric supersonic inlet is in a working state, and the compression efficiency of the inlet is further improved. At the same time, when the engine is in a first state, the bleed ring is used to close all the bleed slots, when the engine is in a second state, the bleed ring is used to open all the bleed slots, the excessive flow loss caused by the bleed slots is effectively avoided, and the problem that the variable geometry axisymmetric supersonic inlet is prone to flow separation due to the change of the profile caused by the moving distance of the center cone in the prior art is effectively solved. BRIEF DESCRIPTION OF DRAWINGS

[0025] In order to more clearly illustrate the technical solutions in the embodiments of the present disclosure, 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 only represent some embodiments of the present disclosure, and for those skilled in the art, other drawings can also be obtained without creative labor on the basis of these drawings.

[0026] Figure 1 The structural diagram of the variable geometry axisymmetric supersonic inlet applied according to the embodiments of the present disclosure is shown in the figure;

[0027] Figure 2 The displacement distance diagram of the variable geometry axisymmetric supersonic inlet applied according to the embodiments of the present disclosure with or without the bleed slot is shown in the figure;

[0028] Figure 3 The Mach number diagram of the variable geometry axisymmetric supersonic inlet applied according to the embodiments of the present disclosure with or without the bleed slot is shown in the figure;

[0029] Figure 4 The structural diagram of the center cone applied according to the embodiments of the present disclosure is shown in the figure;

[0030] Figure 5 The structural diagram of the bleed slot applied according to the embodiments of the present disclosure is further shown in the figure;

[0031] Figure 6 The structural diagram of the support assembly applied according to the embodiments of the present disclosure is further shown in the figure;

[0032] Figure 7 The structural diagram of the outer casing applied according to the embodiments of the present disclosure is further shown in the figure;

[0033] Figure 8 The structural diagram of the bleed slot closure applied according to the embodiments of the present disclosure is further shown in the figure.

[0034] Reference signs:

[0035] 1 - center cone, 2 - outer casing, 3 - support assembly, 31 - first support piece, 32 - second support piece, 4 - bleed ring, 5 - bleed slot, 6 - first through hole, 7 - second through hole. DETAILED DESCRIPTION

[0036] The embodiments of the present disclosure will be described in more detail below with reference to the drawings. Although some embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure can be implemented in various forms, and should not be interpreted as being limited to the embodiments set forth herein, but rather these embodiments are provided to more thoroughly and completely understand the present disclosure. It should be understood that the drawings and embodiments of the present disclosure are only for exemplary purposes, and are not intended to limit the scope of protection of the present disclosure.

[0037] It should be understood that each of the steps recited in the method embodiments of the present disclosure can be performed in different orders and / or in parallel. In addition, the method embodiments can include additional steps and / or omit performing the steps shown. The scope of the present disclosure is not limited in this respect.

[0038] The term "comprises" and variations thereof used in the present disclosure are open-ended, that is, "comprising but not limited to". The term "based on" is "based, at least in part, on". The term "one embodiment" means "at least one embodiment"; the term "another embodiment" means "at least one additional embodiment"; the term "some embodiments" means "at least some embodiments". Related definitions of other terms will be given in the description below. It should be noted that the concepts of "first", "second" and the like mentioned in the present disclosure are only used to distinguish different devices, modules or units, and are not intended to limit the order or interdependence of the functions performed by these devices, modules or units.

[0039] It should be noted that the modification of "one", "multiple" mentioned in the present disclosure is illustrative and not limiting, and those skilled in the art should understand that unless otherwise explicitly indicated in the context, it should be understood as "one or more".

[0040] The names of the messages or information exchanged between the devices in the embodiments of the present disclosure are only for illustrative purposes, and are not intended to limit the scope of the messages or information.

[0041] As the windward component of the power system, the performance of the component of the axisymmetric supersonic inlet directly affects the performance of the core engine of the aircraft. In the low-altitude and low-speed flight state of the aircraft, in order to avoid the inlet from not starting, the inlet is required to have a larger throat area. In the high-altitude and high-speed cruise state of the aircraft, the axisymmetric supersonic inlet is required to increase the contraction ratio of the inlet by reducing the throat area, so as to improve the compression ability to the flow.

[0042] At present, axisymmetric supersonic inlets are divided into fixed geometry and variable geometry. The throat area of the fixed geometry axisymmetric supersonic inlet is not adjustable. The through-flow capacity of the axisymmetric supersonic inlet is improved through the way of leakage slot to realize the starting of low speed state. This case meets the aircraft in high speed flight and low speed flight. The variable geometry axisymmetric supersonic inlet is generally designed with high speed cruise state as the design point. The throat area is changed through the way of moving the center cone. The performance of full speed range is obviously improved compared with the fixed geometry axisymmetric supersonic inlet. However, the inlet surface changes sharply and is easy to induce flow separation. For the variable geometry axisymmetric supersonic inlet, if the moving distance of the center cone is small, the working Mach number range is narrow, and the phenomenon of inlet not starting at low Mach number state or the phenomenon of insufficient air flow compression at high Mach number state is easy to occur. If the moving distance of the center cone is large, the inlet surface changes sharply, and although the starting at low Mach number can be realized, the sharply changed surface is easy to induce flow separation.

[0043] In view of the above problems, the example embodiments of the present disclosure provide a variable geometry axisymmetric supersonic inlet, an aircraft and a control method to solve the problem that the sharply changed surface of the variable geometry axisymmetric supersonic inlet is easy to induce flow separation due to the moving distance of the center cone in the prior art.

[0044] Hereinafter, an electric turbofan combined cycle engine according to an embodiment of the present disclosure will be described in detail with reference to the accompanying drawings.

[0045] Figure 1 The structure diagram of the variable geometry axisymmetric supersonic inlet according to the example embodiments of the present disclosure is shown. As shown in Figure 1 The variable geometry axisymmetric supersonic inlet is applied to an engine. The variable geometry axisymmetric supersonic inlet comprises a center cone 1, an outer casing 2, a support assembly 3 and a leakage ring 4. The center cone 1 is located in the outer casing 2. The leading end of the center cone 1 is in sliding connection with the inner side wall of the inlet end of the outer casing 2. The support assembly 3 is fixed at the outlet end of the outer casing 2. The tail end of the center cone 1 is fixed on the inner side wall of the outlet end of the outer casing 2 through the support assembly 3. The leading end of the center cone 1 is provided with a plurality of leakage slots 5 for fluid inflow. The leakage ring 4 is arranged at the leakage slots 5. When the engine is in a first state, the leakage ring 4 is used to close all the leakage slots 5. When the engine is in a second state, the leakage ring 4 is used to open all the leakage slots 5.

[0046] In actual application, as Figure 1As shown, the variable geometry axisymmetric supersonic inlet is applied to an engine, a center cone 1 is located in an outer casing 2, a leading end of the center cone 1 is in sliding connection with an inner side wall of an air inlet end of the outer casing 2, a support assembly 3 is fixed at an air outlet end of the outer casing 2, a trailing end of the center cone 1 is fixed on the inner side wall of the air outlet end of the outer casing 2 through the support assembly 3, a plurality of flow leakage grooves 5 are arranged at the leading end of the center cone 1 for fluid flowing in, and a flow leakage ring 4 is arranged at the flow leakage grooves 5.

[0047] Figure 2 As shown in the displacement distance schematic diagram of the variable geometry axisymmetric supersonic inlet with or without flow leakage grooves applied according to the embodiments of the present disclosure, Figure 2 wherein 1 represents the variable geometry axisymmetric supersonic inlet without flow leakage grooves, 2 represents the variable geometry axisymmetric supersonic inlet with flow leakage grooves, the horizontal coordinate represents the flight speed Mach number of the engine, and the vertical coordinate represents the displacement distance X and the inlet radius R of the variable geometry axisymmetric supersonic inlet, when the fluid flows into the variable geometry axisymmetric supersonic inlet, the fluid flow will drive the center cone to move in the outer casing, and the variable geometry axisymmetric supersonic inlet with flow leakage grooves can reduce the displacement distance of the center cone by about 8.98%, effectively alleviating the too large change of the inlet surface caused by the relative displacement, increasing the internal contraction ratio by moving the center cone, improving the flow capture capacity of the variable geometry axisymmetric supersonic inlet, and realizing the compensation of the flow. Secondly, the variable geometry axisymmetric supersonic inlet can change the throat area of the variable geometry axisymmetric supersonic inlet by moving the center cone.

[0048] Figure 3 As shown in the Mach number schematic diagram of the variable geometry axisymmetric supersonic inlet with or without flow leakage grooves applied according to the embodiments of the present disclosure, Figure 3 wherein 1 represents the variable geometry axisymmetric supersonic inlet without flow leakage grooves, 2 represents the variable geometry axisymmetric supersonic inlet with flow leakage grooves, Figure 3 When the Mach number of the fluid entering the variable geometry axisymmetric supersonic inlet is 1.5Ma, the variable geometry axisymmetric supersonic inlet can establish a stable shock wave system in the internal flow passage, realizes the starting of the inlet, and the interference point of the shock wave and the boundary layer of the variable geometry axisymmetric supersonic inlet is located in the flow leakage groove area, effectively alleviating the flow separation caused by the shock wave interference in the variable geometry axisymmetric supersonic inlet. For the variable geometry axisymmetric supersonic inlet with flow leakage grooves, the reduction of the relative displacement distance of the center cone reduces the dramatic change of the inlet surface, and no separation zone is generated behind the terminal shock wave.

[0049] On this basis, the leakage groove can eliminate the separation zone formed by the interference of the shock wave and the boundary layer in the variable-geometry axisymmetric supersonic inlet, increase the retraction ratio of the variable-geometry axisymmetric supersonic inlet under the premise of ensuring the working state of the variable-geometry axisymmetric supersonic inlet, and further improve the compression efficiency of the inlet. Meanwhile, when the engine is in the first state, it should be understood that the first state is actually the flight speed Ma < 1.5 of the aircraft, and the leakage ring is used to close all the leakage grooves. When the engine is in the second state, it should be understood that the second state is actually the flight speed 1.5 ≤ Ma ≤ 3.5 of the aircraft, and the leakage ring is used to open all the leakage grooves, effectively avoiding excessive flow loss caused by the leakage grooves, and effectively solving the problem that the variable-geometry axisymmetric supersonic profile is prone to flow separation due to the movement distance of the center cone in the prior art.

[0050] For example, when the engine is in the third state, it should be understood that the second state is actually the flight speed 3.5 < Ma ≤ 4 of the aircraft, and the leakage ring is used to open the first two leakage grooves at the leading end of the center cone. The leakage flow is determined by the pressure difference between the inlet and outlet of the leakage groove and the area of the leakage groove. When the flight speed of the aircraft is Ma = 4, only the first two leakage grooves are opened to ensure that the separation zone formed by the interference of the shock wave and the boundary layer is eliminated, and the flow area of the leakage groove is also reduced, thereby reducing the leakage flow.

[0051] Figure 4 The structure diagram of the center cone according to the embodiment of the present disclosure is shown in FIG. 1. Figure 4 As shown in FIG. 1, the center cone 1 is a hollow structure, and each leakage groove 5 at the leading end of the center cone 1 is distributed in a matrix at equal intervals.

[0052] In actual application, the center cone 1 is a single-stage cone structure, and in order to meet the requirements of high-temperature corrosion resistance, the center cone 1 is made of lightweight aluminum alloy, high-temperature nickel-based alloy or high-temperature alloy steel as a manufacturing material. The center cone 1 is a hollow structure, which can effectively reduce the weight and make the performance of the engine better. Meanwhile, each leakage groove 5 at the leading end of the center cone 1 is distributed in a matrix at equal intervals. The shock wave in the variable-geometry axisymmetric supersonic inlet is prone to interfere with the boundary layer on the center cone 1 to form a separation zone, and the development of the separation zone is prone to cause the inlet to be not started. The leakage groove 5 can effectively suck away the boundary layer, relieve the interference effect of the shock wave and the boundary layer, and multiple rows of leakage grooves 5 can ensure that the interference point is always located in the area of the leakage groove 5 under a wide Mach number, thereby realizing adaptive adjustment.

[0053] Figure 5 The structure diagram of the leakage groove 5 according to the embodiment of the present disclosure is further shown in FIG. 2. Figure 5As shown, the angle between each bleed groove 5 and the inner wall of the air inlet end of the outer casing is represented by α, where α = 45°. The bleed groove 5 first folds over and then vertically connects to the cavity of the central cone to ensure that the boundary layer adhering to the central cone can quickly flow into the central cone when the bleed ring 4 is opened. The groove width of each bleed groove 5 is 5 mm. Figure 5 The term 'd' indicates that it can reduce the large inflow of fluid and avoid a large loss of fluid.

[0054] like Figure 5 As shown, a bleed ring 4 is provided on the inner wall of the central cone. The bleed ring 4 is moved sequentially along the axial direction of the central cone to control the closing or opening of the bleed groove 5.

[0055] Figure 6 This is a further illustration of the structural diagram of the support component according to an embodiment of the present disclosure, such as... Figure 6 As shown, the support assembly includes multiple first support members 31 and second support members 32. It can be understood that the second support members 32 are nested on the shaft of the central cone. One end of each first support member 31 is encircled on the outer side wall of the second support member 32, and the other end of each first support member 31 is encircled on the inner side wall of the air intake end of the outer casing. Due to the axial sliding between the shaft of the central cone and the second support member 32, it should be understood that since the end face of the shaft of the central cone is in contact with the end face of the second support member 32, the flight speed of the aircraft is Ma = 4. At this time, the radial dimension d1 of the central cone is equal to the radial dimension d2 of the inner side wall of the second support member 32.

[0056] Figure 7 This is a further schematic diagram of the structure of the outer casing according to an embodiment of the present disclosure. Figure 8 This is a further schematic diagram illustrating the structure of the drain channel being closed according to an embodiment of the present disclosure, such as... Figures 7-8 As shown, the tail end of the central cone is provided with a first through hole 6 for fluid to flow out, and the air outlet end of the outer casing 2 is provided with a second through hole 7 for fluid to flow out. It should be understood that the above-mentioned venting groove is connected to the first through hole 6 and the second through hole 7 in sequence. The fluid flowing in from the venting groove flows into the variable geometric axisymmetric supersonic air intake through the first through hole 6 and the second through hole 7.

[0057] like Figure 8 As shown, the overflow channel 5 is closed, which is understandable. Figure 8 This is a structural diagram of the engine in its first state. At this time, the flight speed of the aircraft is Ma<1.5, the throat area of ​​the variable geometry axisymmetric supersonic inlet is the largest, and by moving the central cone 1, the cone apex of the central cone 1 is located at the position furthest from the air intake end of the variable geometry axisymmetric supersonic inlet, the throat area of ​​the inlet reaches its maximum, and the bleed ring 4 is moved to make the bleed groove 5 completely closed.

[0058] The aircraft provided by the embodiment of the present disclosure can be a drone or a manned aircraft.

[0059] Compared with the prior art, the aircraft provided by the embodiment of the present disclosure has the beneficial effects of the electric turbofan engine, which will not be repeated here.

[0060] It should be noted that the aircraft provided by the embodiment of the present disclosure can further include a fuselage, a flight controller, etc. The engine and the flight controller are arranged in the fuselage, and the flight controller can control the bleed ring to open or close the bleed slots according to the flight state.

[0061] The embodiment of the present disclosure provides a variable geometry axisymmetric supersonic inlet control method, which is applied to the above-mentioned aircraft. The variable geometry axisymmetric supersonic inlet control method comprises the following steps:

[0062] In response to a first state flight instruction, the engine is controlled to work. The bleed ring is controlled to close all the bleed slots. The first state flight instruction is used to indicate that the engine is in a low-speed working state. At this time, the flight speed of the aircraft is Ma<1.5. At this time, the inlet pressure of the bleed slot is smaller than the outlet pressure of the bleed slot, and the bleed slot cannot achieve bleed. Therefore, the bleed slot is closed.

[0063] In response to a second state flight instruction, the engine is controlled to work. The bleed ring is controlled to open all the bleed slots. The second state flight instruction is used to indicate that the engine is in a medium-speed working state. At this time, the flight speed of the aircraft is 1.5≤Ma≤3.5. The center cone of the variable geometry axisymmetric supersonic inlet moves along the axial direction. The cone apex gradually approaches the direction of the air inlet end of the outer casing. The bleed ring moves backward, and all the bleed slots are opened.

[0064] In response to a third state flight instruction, the engine is controlled to work. The bleed ring 4 is controlled to open the first two bleed slots at the front end of the center cone. The third state flight instruction is used to indicate that the engine is in a high-speed working state. At this time, the flight speed of the aircraft is 3.5<Ma≤4. The axial end surface of the center cone is in contact with the end surface of the second support. The bleed ring moves forward, and the two bleed slots close to the cone apex are opened.

[0065] Compared with the prior art, the variable geometry axisymmetric supersonic inlet control method provided by the embodiment of the present disclosure has the same beneficial effects as the above-mentioned variable geometry axisymmetric supersonic inlet, which will not be repeated here.

[0066] The above description is merely exemplary of some embodiments of the present disclosure and of the principles thereof. It is to be understood that the disclosure is not limited in scope to the particular embodiments described herein, which are intended as examples only, and that the scope of the disclosure is, instead, defined by the appended claims, along with the full range of equivalents to which such claims are entitled. For example, the features of the various embodiments described above can be combined with each other, unless expressly prohibited by the above description.

[0067] While some specific embodiments of the present disclosure have been described in detail, those skilled in the art should understand that the above examples are merely exemplary and are not intended to limit the scope of the present disclosure. Those skilled in the art should understand that the above embodiments can be modified without departing from the scope and spirit of the present disclosure. The scope of the present disclosure is defined by the appended claims.

Claims

1. A variable geometry axisymmetric supersonic inlet, characterized in that, The variable geometry axisymmetric supersonic inlet applied to an engine comprises a center cone, an outer casing, a support assembly and a leakage ring, the center cone is located in the outer casing, the leading end of the center cone is in sliding connection with the inner side wall of the air inlet end of the outer casing, the support assembly is fixed at the air outlet end of the outer casing, the trailing end of the center cone is fixed on the inner side wall of the air outlet end of the outer casing through the support assembly, the leading end of the center cone is provided with a plurality of leakage grooves for fluid inflow, and the leakage ring is arranged at the leakage grooves. When the engine is in a first state, the leakage ring is used to close all the leakage grooves. When the engine is in a second state, the leakage ring is used to open all the leakage grooves. When the engine is in a third state, the leakage ring is used to open the first two leakage grooves at the leading end of the center cone. The support assembly comprises a plurality of first support members and a second support member, one end of each first support member is annularly arranged on the outer side wall of the second support member, and the other end of each first support member is annularly arranged on the inner side wall of the air inlet end of the outer casing.

2. The variable geometry axisymmetric supersonic inlet according to claim 1, characterized in that, The center cone is a hollow structure, and each leakage groove at the leading end of the center cone is distributed in a matrix at equal intervals.

3. The variable geometry axisymmetric supersonic inlet according to claim 2, characterized in that, The included angle between each leakage groove and the inner side wall of the air inlet end of the outer casing is 45°.

4. The variable geometry axisymmetric supersonic inlet of claim 2, wherein, The groove width of each leakage groove is 5mm.

5. The variable geometry axisymmetric supersonic inlet of claim 2, wherein, The trailing end of the center cone is provided with a first through hole for fluid outflow.

6. The variable geometry axisymmetric supersonic inlet of claim 2, wherein, The air outlet end of the outer casing is provided with a second through hole for fluid outflow.

7. An aircraft, characterized in that An engine and the variable geometry axisymmetric supersonic inlet of any one of claims 1-6 are included, and the variable geometry axisymmetric supersonic inlet is located in the engine.

8. A method of controlling a variable geometry axisymmetric supersonic inlet, characterized in that, The variable geometry axisymmetric supersonic inlet control method applied to the aircraft of claim 7 comprises: In response to a first state flight instruction, the engine is controlled to work; The leakage ring is controlled to close all the leakage grooves, and the first state flight instruction is used to indicate that the engine is in a low-speed working state; In response to a second state flight instruction, the engine is controlled to work; The leakage ring is controlled to open all the leakage grooves, and the second state flight instruction is used to indicate that the engine is in a medium-speed working state; In response to a third state flight instruction, the engine is controlled to work; The leakage ring is controlled to open the first two leakage grooves at the leading end of the center cone, and the third state flight instruction is used to indicate that the engine is in a high-speed working state.

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