Variable geometric axis symmetric supersonic air inlet channel, aircraft and control method

Abstract

The invention relates to the technical field of aviation, in particular to a variable-geometric-axis symmetric supersonic air inlet channel, an aircraft and a control method, and aims to solve the problem that in the prior art, due to the movement distance of a center cone, a molded surface with variable-geometric-axis symmetric supersonic speed dramatic change easily induces flow separation. The variable geometric axis symmetric supersonic air inlet channel is applied to an engine and comprises a center cone, an outer casing, a supporting assembly and a drainage ring, the center cone is located in the outer casing, the head end of the center cone is in sliding connection with the inner side wall of the air inlet end of the outer casing, and the supporting assembly is fixed to the air outlet end of the outer casing; the tail end of the center cone is fixed to the inner side wall of the air outlet end of the outer casing through a supporting assembly, a plurality of drainage grooves allowing fluid to flow in are formed in the head end of the center cone, and the drainage rings are arranged at the drainage grooves. The variable geometric axisymmetric supersonic air inlet channel, the aircraft and the control method are used for design of aero-engines.

Description

Technical Field

[0001] The present disclosure relates to the field of aviation technology, and in particular to a variable geometry axisymmetric supersonic air inlet, an aircraft, and a control method. Background Art

[0002] The performance of the axisymmetric supersonic inlet, as the windward component of the power system, directly affects the performance of the core engine of the aircraft. When the aircraft is flying at low altitude and low speed, the inlet is required to have a larger throat area to avoid the inlet not starting. When the aircraft is cruising at high altitude and high speed, the axisymmetric supersonic inlet is required to increase the contraction ratio of the inlet by reducing the throat area and improve the compression capacity of the incoming flow.

[0003] At present, axisymmetric supersonic inlets are divided into two categories: fixed geometry and variable geometry. The throat area of ​​fixed geometry axisymmetric supersonic inlets cannot be adjusted. The flow capacity of axisymmetric supersonic inlets is improved by means of venting grooves to achieve starting at low speeds. This situation meets the requirements of aircraft in high-speed and low-speed flights. Variable geometry axisymmetric supersonic inlets are generally designed with high-speed cruising states as the design point. The throat area is changed by moving the center cone. The performance in the full-speed range is significantly improved compared with fixed geometry axisymmetric supersonic inlets, but the inlet profile changes dramatically, which is easy to induce flow separation. For variable geometry axisymmetric supersonic inlets, if the center cone moves a small distance, the working Mach number range is narrow, and it is easy to have the phenomenon of the inlet not starting at low Mach numbers or the phenomenon of insufficient airflow compression capacity at high Mach numbers. If the center cone moves a large distance, the inlet profile changes dramatically. Although it can achieve starting at low Mach numbers, the dramatically changing profile is easy to induce flow separation.

[0004] Therefore, how to solve the problem in the prior art that the variable geometry axisymmetric profile with drastic changes in supersonic velocity is prone to induce flow separation due to the moving distance of the central cone is one of the important issues to be solved urgently in this field. Summary of the invention

[0005] In view of this, the embodiments of the present disclosure provide a variable geometry axisymmetric supersonic inlet, an aircraft and a control method to solve the problem in the prior art that the variable geometry axisymmetric supersonic speed profile that changes dramatically due to the moving distance of the central cone easily induces flow separation.

[0006] According to one aspect of the present disclosure, a variable geometry axisymmetric supersonic inlet is provided, which is applied to an engine. The variable geometry axisymmetric supersonic inlet comprises: a central cone, an outer casing, a support assembly and a leakage ring. The central cone is located in the outer casing, the head end of the central cone is slidably connected to 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 tail end of the central cone is fixed to the inner side wall of the air outlet end of the outer casing through the support assembly, the head end of the central cone is provided with a plurality of leakage grooves for fluid to flow in, and the leakage ring is arranged at the leakage grooves;

[0007] When the engine is in the first state, the drain ring is used to close all the drain grooves;

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

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

[0010] According to a variable geometry axisymmetric supersonic inlet according to one aspect of the present disclosure, the central cone is a hollow structure, and each drainage groove opened at the head end of the central cone is evenly spaced and distributed in a matrix.

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

[0012] According to the variable geometry axisymmetric supersonic inlet of one aspect of the present disclosure, the 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 to flow out is provided at the tail end of the central cone.

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

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

[0016] According to another aspect of the present disclosure, an aircraft is provided, comprising an engine and the above-mentioned variable geometry axisymmetric supersonic air inlet, wherein the variable geometry axisymmetric supersonic air inlet is applied to the engine.

[0017] According to an aspect of the present disclosure, a variable geometry axisymmetric supersonic inlet control method is applied to the above-mentioned aircraft. The variable geometry axisymmetric supersonic inlet control method includes:

[0018] In response to the first state flight instruction, controlling the engine to operate;

[0019] The control leakage ring is used 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;

[0020] In response to the second state flight instruction, controlling the engine to operate;

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

[0022] <24P08521CN>

[0023] In response to a third state flight instruction, controlling the engine to operate;

[0024] The control leakage ring is used to open the first two leakage grooves at the head 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.

[0025] At least one of the above technical solutions adopted in the embodiments of the present disclosure can achieve the following beneficial effects: a variable geometry axisymmetric supersonic inlet is applied to an engine, the center cone is located in an outer casing, the head end of the center cone is slidably connected to the inner side wall of the inlet end of the outer casing, the support assembly is fixed to the outlet end of the outer casing, the tail end of the center cone is fixed to the inner side wall of the outlet end of the outer casing through the support assembly, the head end of the center cone is provided with a plurality of drainage grooves for fluid inflow, and a drainage ring is provided at the drainage groove. 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, and 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 flow compensation is achieved. Secondly, the throat area of ​​the variable geometry axisymmetric supersonic inlet can be changed by moving the center cone. 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, and increase the internal contraction ratio of the variable geometry axisymmetric supersonic inlet under the premise of ensuring that the variable geometry axisymmetric supersonic inlet is in a working state, thereby improving the compression efficiency of the inlet. At the same time, when the engine is in the first state, the leakage ring is used to close all the leakage grooves, and when the engine is in the second state, the leakage ring is used to open all the leakage grooves, effectively avoiding excessive flow loss caused by the leakage groove, and effectively solving the problem in the prior art that the variable geometry axisymmetric supersonic profile that changes dramatically due to the moving distance of the center cone is easy to induce flow separation. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] In order to more clearly illustrate the technical solutions in the embodiments of the present disclosure, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present disclosure. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0027] Figure 1 The figure shows a structure diagram of a variable geometry axisymmetric supersonic inlet according to an embodiment of the present disclosure;

[0028] Figure 2 FIG. 1 is a schematic diagram of displacement distance of a variable geometry axisymmetric supersonic inlet with or without a leakage groove according to an embodiment of the present disclosure;

[0029] Figure 3 The figure shows a schematic diagram of Mach number of a variable geometry axisymmetric supersonic inlet with or without a bleed slot according to an embodiment of the present disclosure;

[0030] Figure 4 The diagram shows a structure diagram of a center cone according to an embodiment of the present disclosure;

[0031] Figure 5 is a structural diagram further illustrating a drain trough according to an embodiment of the present disclosure;

[0032] Figure 6 is a structural diagram further illustrating a support assembly according to an embodiment of the present disclosure;

[0033] <24P08521CN>

[0034] Figure 7 is a structural diagram further illustrating an outer casing according to an embodiment of the present disclosure;

[0035] Figure 8 It is a structural diagram further illustrating the closure of the spillway according to an embodiment of the present disclosure.

[0036] Reference numerals:

[0037] 1-center cone, 2-outer casing, 3-support assembly, 31-first support member, 32-second support member, 4-leakage ring, 5-leakage groove, 6-first through hole, 7-second through hole. DETAILED DESCRIPTION

[0038] Embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although certain embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be construed as being limited to the embodiments described herein, which are instead provided for a more thorough and complete understanding of 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.

[0039] It should be understood that the various steps described in the method embodiments of the present disclosure may be performed in different orders and / or in parallel. In addition, the method embodiments may include additional steps and / or omit the steps shown. The scope of the present disclosure is not limited in this respect.

[0040] The term "including" and its variations used in this document are open inclusions, that is, "including but not limited to". The term "based on" means "based at least in part on". The term "one embodiment" means "at least one embodiment"; the term "another embodiment" means "at least one other embodiment"; the term "some embodiments" means "at least some embodiments". Relevant definitions of other terms will be given in the description below. It should be noted that the concepts of "first", "second", etc. mentioned in this disclosure are only used to distinguish different devices, modules or units, and are not used to limit the order or interdependence of the functions performed by these devices, modules or units.

[0041] It should be noted that the modifications of "one" and "plurality" mentioned in the present disclosure are illustrative rather than restrictive, and those skilled in the art should understand that unless otherwise clearly indicated in the context, it should be understood as "one or more".

[0042] The names of the messages or information exchanged between multiple devices in the embodiments of the present disclosure are only used for illustrative purposes and are not used to limit the scope of these messages or information.

[0043] The performance of the axisymmetric supersonic inlet, as the windward component of the power system, directly affects the performance of the core engine of the aircraft. When the aircraft is flying at low altitude and low speed, the inlet is required to have a larger throat area to avoid the inlet not starting. When the aircraft is cruising at high altitude and high speed, the axisymmetric supersonic inlet is required to increase the contraction ratio of the inlet by reducing the throat area and improve the compression capacity of the incoming flow.

[0044] At present, axisymmetric supersonic inlets are divided into two categories: fixed geometry and variable geometry. The throat area of ​​fixed geometry axisymmetric supersonic inlets cannot be adjusted. The flow capacity of axisymmetric supersonic inlets is improved by means of venting grooves to achieve starting at low speeds. This situation meets the requirements of aircraft in high-speed and low-speed flights. Variable geometry axisymmetric supersonic inlets are generally designed with high-speed cruising states as the design point. The throat area is changed by moving the center cone. The performance in the full-speed range is significantly improved compared with fixed geometry axisymmetric supersonic inlets, but the inlet profile changes dramatically, which is easy to induce flow separation. For variable geometry axisymmetric supersonic inlets, if the center cone moves a small distance, the working Mach number range is narrow, and it is easy to have the phenomenon of the inlet not starting at low Mach numbers or the phenomenon of insufficient airflow compression capacity at high Mach numbers. If the center cone moves a large distance, the inlet profile changes dramatically. Although it can achieve starting at low Mach numbers, the dramatically changing profile is easy to induce flow separation.

[0045] In response to the above problems, an exemplary embodiment of the present disclosure provides a variable geometry axisymmetric supersonic inlet, an aircraft and a control method to solve the problem in the prior art that the variable geometry axisymmetric supersonic speed profile that changes dramatically due to the movement distance of the central cone easily induces flow separation.

[0046] An electric turbofan combined cycle engine according to an embodiment of the present disclosure will be described in detail below with reference to the accompanying drawings.

[0047] Figure 1 The figure shows a structure diagram of a variable geometry axisymmetric supersonic inlet according to an embodiment of the present disclosure. Figure 1 As shown, a variable geometry axisymmetric supersonic inlet is applied to an engine. The variable geometry axisymmetric supersonic inlet comprises: a central cone 1, an outer casing 2, a support assembly 3 and a leakage ring 4. The central cone 1 is located in the outer casing 2, the head end of the central cone 1 is slidingly connected to the inner wall of the inlet end of the outer casing 2, the support assembly 3 is fixed to the outlet end of the outer casing 2, the tail end of the central cone 1 is fixed to the inner wall of the outlet end of the outer casing 2 through the support assembly 3, a plurality of leakage grooves 5 for fluid to flow into are opened at the head end of the central cone 1, the leakage ring 4 is arranged at the leakage groove 5, when the engine is in the first state, the leakage ring 4 is used to close all the leakage grooves 5, and when the engine is in the second state, the leakage ring 4 is used to open all the leakage grooves 5.

[0048] In practical applications, such as Figure 1As shown, a variable geometry axisymmetric supersonic inlet is applied to an engine, wherein a center cone 1 is located inside an outer casing 2, a head end of the center cone 1 is slidingly connected to an inner side wall of an air inlet end of the outer casing 2, a support assembly 3 is fixed to an air outlet end of the outer casing 2, a tail end of the center cone 1 is fixed to an inner side wall of an air outlet end of the outer casing 2 through the support assembly 3, a plurality of drainage grooves 5 for fluid to flow into are provided at the head end of the center cone 1, and a drainage ring 4 is provided at the drainage groove 5.

[0049] Figure 2 FIG. 1 is a schematic diagram of the displacement distance of a variable geometry axisymmetric supersonic inlet with or without a leakage groove according to an embodiment of the present disclosure. Figure 2 As shown, 1 represents a variable geometry axisymmetric supersonic inlet without a drain groove, 2 represents a variable geometry axisymmetric supersonic inlet with a drain groove, the horizontal axis represents the flight speed Mach number of the engine, and the vertical axis represents the displacement distance X and the variable geometry axisymmetric supersonic inlet inlet radius R. 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 drain groove can reduce the displacement distance of the center cone by about 8.98%, effectively alleviating the excessive change of the inlet profile caused by the relative displacement, increasing the internal contraction ratio by moving the center cone, improving the flow capture capability of the variable geometry axisymmetric supersonic inlet, and realizing flow compensation. Secondly, the throat area of ​​the variable geometry axisymmetric supersonic inlet can be changed by moving the center cone.

[0050] Figure 3 The figure shows a schematic diagram of the Mach number of a variable geometry axisymmetric supersonic inlet with or without a bleed slot according to an embodiment of the present disclosure. Figure 3 As shown, 1 indicates that the variable geometry axisymmetric supersonic inlet without a discharge groove has a separation zone, and 2 indicates that the variable geometry axisymmetric supersonic inlet with a discharge groove has no separation zone. 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 inner flow channel to realize the starting of the inlet. The interference point between the shock wave and the boundary layer of the variable geometry axisymmetric supersonic inlet is located in the discharge groove area, which effectively alleviates the flow separation caused by the shock wave interference in the variable geometry axisymmetric supersonic inlet. For the variable geometry axisymmetric supersonic inlet with discharge grooves, the relative displacement distance of the central cone is reduced, which reduces the drastic change of the inlet surface and no separation zone is generated behind the ending shock wave.

[0051] 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, and on the premise of ensuring that the variable geometry axisymmetric supersonic inlet is in a working state, the inward contraction ratio of the variable geometry axisymmetric supersonic inlet is increased, thereby improving the compression efficiency of the inlet. At the same time, when the engine is in the first state, it should be understood that the first state is actually the flight speed of the aircraft Ma<1.5, 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 of the aircraft 1.5≤Ma≤3.5, and the leakage ring is used to open all the leakage grooves, effectively avoiding excessive flow loss caused by the leakage groove, and effectively solving the problem in the prior art that the variable geometry axisymmetric supersonic profile with drastic changes due to the moving distance of the center cone is easy to induce flow separation.

[0052] Exemplarily, when the engine is in the third state, it should be understood that the second state is actually the flight speed of the aircraft 3.5<Ma≤4, and the leakage ring is used to open the first two leakage grooves at the head 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 Ma=4, only opening the first two leakage grooves can ensure the elimination of the separation zone formed by the interference of the shock wave boundary layer, and can also reduce the flow area of ​​the leakage groove, thereby reducing the leakage flow.

[0053] Figure 4 The structure diagram of the central cone according to the embodiment of the present disclosure is shown in FIG. Figure 4 As shown, the central cone 1 is a hollow structure, and each of the discharge grooves 5 opened at the head end of the central cone 1 is evenly spaced and distributed in a matrix.

[0054] In practical applications, the central cone 1 is a single-stage cone structure. In order to meet the requirements of high temperature corrosion resistance, <24P08521CN>

[0055] The center cone 1 is made of lightweight aluminum alloy, high-temperature nickel-based alloy or high-temperature alloy steel. The center cone 1 is a hollow structure, which can effectively reduce weight and make the engine performance better. At the same time, each discharge groove 5 opened at the head end of the center cone 1 is evenly spaced and distributed in a matrix. In the variable geometry axisymmetric supersonic inlet, the shock wave easily interferes with the boundary layer on the center cone 1 to form a separation zone. The development of the separation zone easily causes the inlet to fail to start. The discharge groove 5 can effectively suck away the boundary layer and alleviate the interference between the shock wave and the boundary layer. The multiple discharge grooves 5 can ensure that the interference point is always located in the discharge groove 5 area under a wide Mach number, realizing adaptive adjustment.

[0056] Figure 5 is a structural diagram further illustrating the discharge trough 5 according to an embodiment of the present disclosure, such as Figure 5As shown, the angle between each drainage groove 5 and the inner side wall of the air inlet end of the outer casing is represented by α, α = 45°, the drainage groove 5 is first folded, and then vertically connected with the cavity of the central cone to ensure that when the drainage ring 4 is opened, the boundary layer attached to the central cone can quickly flow into the central cone, and the groove width of each drainage groove 5 is 5mm. Figure 5 Indicated by d, it can reduce the large inflow of fluid and avoid the loss of a large amount of fluid.

[0057] like Figure 5 As shown, a leakage ring 4 is provided on the inner side wall of the central cone, and the leakage ring 4 is moved in sequence along the axial direction of the central cone, so as to control the closing or opening of the leakage groove 5.

[0058] Figure 6 is a structural diagram further illustrating a support assembly according to an embodiment of the present disclosure, such as Figure 6 As shown, the support assembly includes: a plurality of first support members 31 and a second support member 32. It can be understood that the second support member 32 is nested on the axis of the center cone, one end of each first support member 31 is arranged on the outer wall of the second support member 32, and the other end of each first support member 31 is arranged on the inner wall of the air intake end of the outer casing. Since the axis of the center cone and the second support member 32 produce axial sliding, it should be understood that since the axial end face of the center cone fits with the end face of the second support member 32, the flight speed of the aircraft is Ma=4, and at this time, the radial dimension d1 of the center cone is equal to the radial dimension d2 of the inner wall of the second support member 32.

[0059] Figure 7 is a structural diagram further illustrating an outer casing according to an embodiment of the present disclosure, Figure 8 is a structural diagram further illustrating the closing of the discharge trough according to an embodiment of the present disclosure, such as Figure 7 - Figure 8 As shown, a first through hole 6 for fluid to flow out is provided at the tail end of the center cone, and a second through hole 7 for fluid to flow out is provided at the air outlet end of the outer casing 2. It should be understood that the above-mentioned drainage groove is connected with the first through hole 6 and the second through hole 7 in sequence, and the fluid flowing in from the drainage groove flows into the variable geometry axisymmetric supersonic inlet again through the first through hole 6 and the second through hole 7.

[0060] like Figure 8 As shown, the discharge trough 5 is closed. It can be understood that Figure 8 The structure diagram of the engine in the 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 the center cone 1 is moved. The cone vertex of the center cone 1 is located at the position farthest from the inlet end of the variable geometry axisymmetric supersonic inlet, and the throat area of ​​the inlet reaches the maximum. The discharge ring 4 is moved.

[0061] The discharge trough 5 is completely closed.

[0062] The exemplary embodiment of the present disclosure provides an aircraft, including the variable geometry axisymmetric supersonic inlet and an engine of the exemplary embodiment of the present disclosure, wherein the variable geometry axisymmetric supersonic inlet is applied to the engine. It should be understood that the aircraft of the embodiment of the present disclosure can be a drone or a manned aircraft. For example:

[0063] Compared with the prior art, the beneficial effects of the aircraft provided by the embodiments of the present disclosure refer to the beneficial effects of the electric turbofan engine, which will not be elaborated here.

[0064] It should be noted that the aircraft provided in the embodiment of the present disclosure may also include a fuselage, a flight controller, etc. The engine, the flight controller, etc. are arranged in the fuselage, and the flight controller can control the leakage ring to open or close the leakage groove according to the flight status.

[0065] An exemplary 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 includes:

[0066] In response to the first state flight instruction, the engine is controlled to work. The control leakage ring is used to close all the leakage grooves. 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 leakage groove is lower than the outlet pressure of the leakage groove, and the leakage cannot be achieved, and the leakage groove is closed.

[0067] In response to the second state flight instruction, the engine is controlled to work. The control leakage ring is used to open all the leakage grooves. 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, and the cone vertex gradually approaches the direction of the inlet end of the outer casing, the leakage ring moves backward, and the leakage grooves are all opened.

[0068] In response to the third state flight instruction, the engine is controlled to work. The control leakage ring 4 is used to open the first two leakage grooves at the head 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, when the flight speed of the aircraft is 3.5<Ma≤4, the axial end face of the center cone fits with the end face of the second support member, the leakage ring moves forward, and the two leakage grooves near the cone vertex are opened.

[0069] Compared with the prior art, an exemplary embodiment of the present disclosure provides a variable geometry axisymmetric supersonic inlet control method having the same beneficial effects as the above-mentioned variable geometry axisymmetric supersonic inlet, which will not be elaborated here.

[0070] The above description is only an illustration of some embodiments of the present disclosure and the technical principles used. Those skilled in the art should understand that the scope of disclosure involved in the present disclosure is not limited to the above technical features.

[0071] The technical solutions formed by any combination of the above technical features or their equivalent features shall also be included without departing from the above disclosed concept. For example, the technical solutions formed by replacing the above features with the technical features with similar functions disclosed in this disclosure (but not limited to) are mutually replaced.

[0072] Although some specific embodiments of the present disclosure have been described in detail by way of example, it should be understood by those skilled in the art that the above examples are for illustration only and are not intended to limit the scope of the present disclosure. It should be understood by those skilled in the art that the above embodiments may 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: Applied to an engine, the variable geometry axisymmetric supersonic inlet comprises: a central cone, an outer casing, a support assembly and a leakage ring, wherein the central cone is located in the outer casing, the head end of the central cone is slidably connected to 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 tail end of the central cone is fixed to the inner side wall of the air outlet end of the outer casing through the support assembly, the head end of the central cone is provided with a plurality of leakage grooves for fluid to flow in, and the leakage ring is arranged at the leakage grooves; When the engine is in the first state, the leakage ring is used to close all the leakage grooves; When the engine is in the second state, the leakage ring is used to open all the leakage grooves.

2. The variable geometry axisymmetric supersonic inlet according to claim 1, characterized in that: When the engine is in the third state, the leakage ring is used to open the first two leakage grooves at the head end of the central cone.

3. The variable geometry axisymmetric supersonic inlet according to claim 1, characterized in that: The central cone is a hollow structure, and each drainage groove opened at the head end of the central cone is evenly spaced and distributed in a matrix.

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

5. The variable geometry axisymmetric supersonic inlet according to claim 3, characterized in that: The width of each of the drain grooves is 5 mm.

6. The variable geometry axisymmetric supersonic inlet according to claim 3, characterized in that: The tail end of the central cone is provided with a first through hole for fluid to flow out.

7. The variable geometry axisymmetric supersonic inlet according to claim 3, characterized in that: The air outlet end of the outer casing is provided with a second through hole for fluid to flow out.

8. The variable geometry axisymmetric supersonic inlet according to claim 1, characterized in that: The support assembly includes: a plurality of first support members and a second support member, one end of each of the first support members is arranged on the outer side wall of the second support member, and the other end of each of the first support members is arranged on the inner side wall of the air intake end of the outer casing.

9. An aircraft, characterized in that: It comprises an engine and the variable geometry axisymmetric supersonic inlet as claimed in any one of claims 1 to 8, wherein the variable geometry axisymmetric supersonic inlet is located in the engine.

10. A variable geometry axisymmetric supersonic inlet control method, characterized in that: Applied to the aircraft of claim 9, the variable geometry axisymmetric supersonic inlet control method comprises: In response to the first state flight instruction, controlling the engine to operate; <24P08521CN> The control leakage ring is used 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 the second state flight instruction, controlling the engine to operate; The control leakage ring is used 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, controlling the engine to operate; The control leakage ring is used to open the first two leakage grooves at the head end of the central 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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