An aircraft engine and aircraft

Abstract

The application provides an aero-engine and an aircraft. The aero-engine comprises: a body provided with a compression device; a ring-shaped mounting provided at an end of the body, an inner part of the mounting is provided with a mounting cavity, an inner hole of the mounting forms an air inlet passage, an air outlet side of the air inlet passage is communicated with an air inlet side of the compression device, a side of the mounting facing the air inlet passage is provided with an air injection through hole, and a side of the mounting facing away from the air inlet passage is provided with an auxiliary through hole, the air injection through hole and the auxiliary through hole are communicated through the mounting cavity; and a control device arranged in the mounting cavity. The aero-engine has a simple structure, and when applied to the aircraft, the control device can be well cooled in a ground test mode and a flight task mode.

Description

Technical Field

[0001] This article relates to engine equipment technology, particularly an aero engine and aircraft. Background Technology

[0002] Miniature aircraft engines have two common operating modes: ground testing mode and flight mission mode. Ground testing mode: the aircraft is stationary; flight mission mode: the aircraft flies at high speed in the air.

[0003] When designing micro aero engines, in order to save space, the control unit (ECU) is integrated into the mounting cavity formed by the fairing and the intake casing. However, the heat dissipation problem of the control unit with this integrated layout is more prominent than that of the control unit with an independent layout.

[0004] Independently designed control units typically have a heat-conducting housing with heat dissipation fins, allowing heat generated during operation to be quickly dissipated. However, integrated control units lack the means to connect heat dissipation fins. Therefore, achieving heat dissipation for integrated control units in both operating modes of a micro-air engine is a technical challenge that urgently needs to be addressed by those skilled in the art. Summary of the Invention

[0005] To solve at least one of the aforementioned technical problems, this application provides an aircraft engine that is not only simple in structure but also capable of effectively dissipating heat from the control device in both ground testing and flight mission modes.

[0006] This application also provides an aircraft.

[0007] The aero-engine provided in this embodiment of the invention includes: a body having a compressor; an annular mounting member disposed at the end of the body, the mounting member having a mounting cavity inside, the inner hole of the mounting member forming an air intake passage, the air outlet side of the air intake passage communicating with the air intake side of the compressor, the mounting member having an air ejector through hole on the side facing the air intake passage and an auxiliary through hole on the side facing away from the air intake passage, the air ejector through hole and the auxiliary through hole communicating through the mounting cavity; and a control device disposed within the mounting cavity.

[0008] In some exemplary embodiments, the diameter of the air ejector through-hole is no greater than 1 mm.

[0009] In some exemplary embodiments, the air ejector through-holes include a plurality of holes, which are spaced apart circumferentially along the mounting member.

[0010] In some exemplary embodiments, the auxiliary through holes include a plurality of holes, which are spaced apart circumferentially along the mounting member.

[0011] In some exemplary embodiments, the diameter of the auxiliary through-hole is not smaller than the diameter of the air ejector through-hole.

[0012] In some exemplary embodiments, the inner peripheral wall of the mounting member includes a guide section that tapers from the air intake side of the air intake passage to the air outlet side of the air intake passage, and the air ejector through-hole is located on the guide section.

[0013] In some exemplary embodiments, in the axial direction of the air intake passage: the auxiliary through-hole is located between the air ejector through-hole and the air compressor.

[0014] In some exemplary embodiments, the mounting component includes an annular intake casing and an annular fairing, the intake casing being disposed at the end of the body, and the fairing being fitted onto the intake casing and together with the intake casing forming the mounting cavity.

[0015] In some exemplary embodiments, the mounting cavity is an annular cavity.

[0016] The aircraft proposed in the embodiments of the present invention includes the aero-engine described in any of the above embodiments.

[0017] In some exemplary embodiments, the aircraft further includes: a curved air intake duct, the exhaust side of which is connected to the air intake side of the air intake passage; wherein, an inner peripheral surface of the exhaust side of the air intake duct is formed with a boundary layer separation region, and the axis of the air ejector through-hole is arranged along the extension direction of the boundary layer separation region.

[0018] In the technical solution of this invention embodiment, the aero-engine is applied to an aircraft. In ground test mode, the compressor operates, and outside air is drawn into the compressor through the intake passage. During this process, since the aircraft is stationary relative to the ground (the air inside the intake passage is depressurized air), the air pressure inside the intake passage is lower than the air pressure inside the mounting cavity. According to the ejection principle, the negative pressure airflow formed inside the intake passage will eject the air inside the mounting cavity through the air ejection port to the intake passage. Air radially outside the mounting cavity will be replenished into the mounting cavity through the auxiliary port. In this way, a first air jet is formed inside the mounting cavity, flowing from the auxiliary port through the mounting cavity to the air ejection port. The first air jet carries away the heat of the control device during its flow, realizing heat dissipation of the control device. No additional fan is required, and the structure is simpler. Flight mission mode In this configuration, the compressor operates. Due to the high-speed flight of the aircraft, and because the air intake side of the air intake passage faces the front of the aircraft's flight direction, outside air is pumped into the compressor through the air intake passage (the air inside the air intake passage is pressurized air). This causes the air pressure inside the air intake passage to be greater than the air pressure inside the mounting cavity. According to the principle of pressurized air intake, a portion of the outside air pumped into the air intake passage flows into the mounting cavity through the air ejector hole. The air inside the mounting cavity then flows radially outward from the auxiliary through-hole. This creates a second airflow within the mounting cavity, flowing from the air ejector hole through the mounting cavity to the auxiliary through-hole. During this flow, the second airflow carries away the heat from the control device, thus dissipating heat from the control device. The remaining portion of the outside air pumped into the air intake passage is pumped into the compressor.

[0019] Other features and advantages of this application will be set forth in the following description, and will be apparent in part from the description, or may be learned by practicing the application. Other advantages of this application can be realized and obtained by means of the solutions described in the description and the accompanying drawings. Attached Figure Description

[0020] The accompanying drawings are used to provide an understanding of the technical solutions of this application and constitute a part of the specification. They are used together with the embodiments of this application to explain the technical solutions of this application and do not constitute a limitation on the technical solutions of this application.

[0021] Figure 1 This is a three-dimensional structural diagram of an aero-engine according to an embodiment of the present invention;

[0022] Figure 2 for Figure 1 The diagram shows a partial structural schematic of the aero engine; the cowling is not shown.

[0023] Figure 3 for Figure 1 A partial radial cross-sectional view of the centrally mounted component in ground test mode; the arrow indicates the first air duct.

[0024] Figure 4 for Figure 1 A partial schematic diagram of the radial cross-sectional structure of the mid-mounted component in flight mission mode; the arrow indicates the second wind beam.

[0025] Figure 5 This is a three-dimensional structural diagram of an aircraft according to an embodiment of the present invention.

[0026] in, Figures 1 to 5 The correspondence between the reference numerals and component names in the attached drawings is as follows:

[0027] 100 Body, 200 Mounting component, 210 Mounting cavity, 220 Intake passage, 230 Intake casing, 231 Air ejector port, 232 Guide section, 240 Frost, 241 Auxiliary passage, 300 Control device, 400 Intake duct, 410 Boundary layer separation zone. Detailed Implementation

[0028] This application describes several embodiments, but these descriptions are exemplary and not restrictive. To make the objectives, technical solutions, and advantages of the present invention clearer, the embodiments of the present invention will be described in detail below with reference to the accompanying drawings. It should be noted that, unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

[0029] The aero-engine provided in the embodiments of the present invention, such as Figures 1 to 4 As shown, it includes: a body 100 with a compressor; an annular mounting member 200 disposed at the front end of the body 100, the mounting member 200 having a mounting cavity 210 inside, the inner hole of the mounting member 200 forming an air intake passage 220, the air outlet side of the air intake passage 220 communicating with the air intake side of the compressor, an air ejector hole 231 provided on the side of the mounting member 200 facing the air intake passage 220, and an auxiliary through hole 241 provided on the side facing away from the air intake passage 220, the air ejector hole 231 and the auxiliary through hole 241 being connected through the mounting cavity 210, the air intake passage 220 being connected to the mounting cavity 210 through the air ejector hole 231, and the outer side of the mounting cavity 210 being connected to the mounting cavity 210 through the auxiliary through hole 241; and a control device 300 disposed within the mounting cavity 210. Figure 3 As shown, in ground test mode, the air compressor operates, and outside air is drawn into the air compressor through the air intake passage 220. The air pressure inside the air intake passage 220 is lower than the air pressure inside the mounting cavity 210. According to the ejection principle, the air inside the mounting cavity 210 flows to the air intake passage 220 through the air ejection through-hole 231. Air radially outside the mounting cavity 210 is replenished into the mounting cavity 210 through the auxiliary through-hole 241. Figure 4As shown, in flight mission mode, the compressor is in operation. Outside air is pumped into the compressor through the air intake passage 220. The air pressure inside the air intake passage 220 is greater than the air pressure inside the mounting cavity 210. According to the principle of pressurized air intake, part of the outside air pumped into the air intake passage 220 will flow into the mounting cavity 210 through the air ejector hole 231. The air in the mounting cavity 210 will flow radially outward from the auxiliary through hole 241. The other part of the outside air pumped into the air intake passage 220 is pumped into the compressor through the air intake passage 220.

[0030] This type of aero engine is used in aircraft, such as... Figure 3 As shown, in ground test mode, the compressor operates, and outside air is drawn into the compressor through the air intake passage 220. During this process, since the aircraft is stationary relative to the ground (the air inside the air intake passage 220 is depressurized air), the air pressure inside the air intake passage 220 is lower than the air pressure inside the mounting cavity 210. According to the ejection principle, the negative pressure airflow formed inside the air intake passage 220 will eject the air inside the mounting cavity 210 through the air ejection port 231 to the air intake passage 220. Air radially outside the mounting cavity 210 will be replenished into the mounting cavity 210 through the auxiliary port 241. This creates a first airflow stream within the mounting cavity 210, flowing from the auxiliary port 241 through the mounting cavity 210 to the air ejection port 231. This first airflow stream carries away heat from the control device 300 during its flow, achieving heat dissipation for the control device 300 without the need for an additional fan, resulting in a simpler structure. Figure 4 As shown, in flight mission mode, the compressor operates. During this process, due to the high-speed flight of the aircraft, and because the intake side of the air intake passage 220 faces forward of the aircraft's flight direction, outside air is forced into the compressor through the air intake passage 220 (the air inside the air intake passage 220 is pressurized air). This causes the air pressure inside the air intake passage 220 to be greater than the air pressure inside the mounting cavity 210. According to the principle of pressurized air intake, a portion of the outside air injected into the air intake passage 220 will pass through the air ejector port 2. Air 31 flows into the mounting cavity 210, while the air inside the mounting cavity 210 flows radially outward from the auxiliary through-hole 241. This creates a second airflow within the mounting cavity 210, originating from the air ejection through-hole 231 and flowing through the mounting cavity 210 to the auxiliary through-hole 241. This second airflow carries away heat from the control device 300 during its flow, thus dissipating heat from the control device 300. Another portion of the outside air entering the intake passage 220 is then pumped to the compressor through the intake passage 220. The first and second airflows flow in opposite directions.

[0031] During the acceleration of the aircraft, there is inevitably a speed range where the pressure difference between the air ejector port 231 and the auxiliary port 241 is insufficient to form a second air jet. This speed range is approximately between 0.05 Ma and 0.1 Ma (Ma represents Mach number). Since the aircraft spends a short time in this speed range, the negative impact of this speed range's inability to effectively dissipate heat from the control device 300 can be completely ignored.

[0032] In some embodiments, such as Figures 1 to 4 As shown, the mounting cavity 210 is configured as an annular cavity.

[0033] To ensure that the air ejector through-hole 231 can achieve air ejection flow in ground test mode, the diameter of the air ejector through-hole 231 is set to be no greater than 1 mm. Alternatively, the diameter of the air ejector through-hole 231 can be set to 0.5 mm; or it can be set to 1 mm, etc.; all of the above can achieve the purpose of this application, and their intent does not depart from the design concept of this invention, and will not be elaborated further here, all of which should fall within the protection scope of this application.

[0034] To better ensure the heat dissipation effect of the control device 300, such as Figures 1 to 4 As shown, the air ejector through-holes 231 include a plurality of holes, which are spaced apart circumferentially along the mounting member 200. The number of air ejector through-holes 231 can be 10, 15, or 20, etc. Those skilled in the art can make reasonable settings as needed, and all of the above can achieve the purpose of this application. Their intent does not depart from the design concept of this invention, and will not be elaborated further here; all should fall within the protection scope of this application.

[0035] It could be that the auxiliary through-hole 241 is one; or it could be, as... Figure 1 , Figure 3 and Figure 4 As shown, there are multiple auxiliary through holes 241, and the multiple auxiliary through holes 241 are arranged at intervals along the circumference of the mounting member 200; all of the above can achieve the purpose of this application, and their purpose has not departed from the design concept of this invention, so they will not be repeated here, and should all fall within the protection scope of this application.

[0036] In some examples, the diameter of the auxiliary through-hole 241 is not smaller than the diameter of the air ejection through-hole 231, which facilitates the passage of air through the auxiliary through-hole 241. Further, the diameter of the auxiliary through-hole 241 is set to be larger than the diameter of the air ejection through-hole 231.

[0037] In some examples, such as Figure 1 , Figure 3 and Figure 4As shown, in the axial direction of the air intake passage 220: the auxiliary through hole 241 is located between the air ejector through hole 231 and the air compressor, which can effectively increase the flow range of the first air jet and the second air jet in the mounting cavity 210, and better improve the heat dissipation effect of the control device 300.

[0038] In some examples, such as Figures 1 to 4 As shown, the inner peripheral wall of the mounting component 200 includes a guide section 232 that gradually narrows from the air intake side of the air intake passage 220 to the air outlet side of the air intake passage 220. An air ejector hole 231 is located on the guide section 232. This design is more conducive to the flow of outside air into the air intake passage 220 through the air ejector hole 231 into the mounting cavity 210 in flight mission mode.

[0039] In some examples, such as Figures 1 to 4 As shown, the mounting component 200 includes an annular intake casing 230 and an annular fairing 240. The intake casing 230 is located at the front end of the main body 100. The fairing 240 is fitted onto the intake casing 230 and together with the intake casing 230 forms an annular mounting cavity 210. This solution has a simple structure and mature technology.

[0040] In some embodiments, such as Figure 2 As shown, the control device 300 is configured as a circuit board.

[0041] The aircraft proposed in the embodiments of the present invention, such as Figure 5 As shown, it includes the aircraft engine described in any of the above embodiments.

[0042] This aircraft possesses all the advantages of the aero-engine proposed in any of the above embodiments, which will not be elaborated here.

[0043] In some exemplary embodiments, such as Figure 5 As shown, the aircraft also includes: a curved and extended air intake duct 400, the exhaust side of which is connected to the intake side of the air intake passage 220 to prevent the blades of the compressor from being directly exposed to the outside; wherein, a boundary layer separation zone 410 is formed on the inner circumferential surface of the exhaust side of the air intake duct 400, and the axis of the air ejector through-hole 231 is arranged along the extension direction of the boundary layer separation zone 410. In this way, in flight mission mode, it is more conducive for the outside air entering the air intake passage 220 to flow from the air ejector through-hole 231 into the mounting cavity 210. Correspondingly, the amount of airflow in the boundary layer separation zone 410 that is sucked into the aero engine will be less, which can further improve the intake quality of the aero engine.

[0044] In some embodiments, the air intake 400 is configured as an S-shaped air intake.

[0045] It may be that the aircraft is a drone, an airplane, a rocket, etc. All of the above can achieve the purpose of this application. The gist thereof does not depart from the design concept of the present invention and will not be elaborated herein. All of them shall fall within the protection scope of this application.

[0046] In summary, in the technical solution of the embodiment of the present invention, an aeroengine is applied to an aircraft. In the ground test run mode, the compression device operates, and external air will be sucked into the compression device through the air intake passage. Since the aircraft is stationary relative to the ground during this process (the air inside the air intake passage is decompressed air), the air pressure inside the air intake passage is less than the air pressure inside the installation cavity. According to the ejector principle, the negative pressure air flow formed inside the air intake passage will eject the air inside the installation cavity to flow through the air ejector through hole into the air intake passage, and the air on the radial outer side of the installation cavity will be replenished into the installation cavity through the auxiliary through hole. In this way, a first air flow is formed in the installation cavity that flows from the auxiliary through hole through the installation cavity to the air ejector through hole. The first air flow带走 the heat of the control device during the flowing process, realizing heat dissipation of the control device, and there is no need to additionally equip a fan, and the structure is simpler; In the flight mission mode, the compression device operates. Since the aircraft is flying at a high speed in the air during this process and the air intake side of the air intake passage faces the front of the flying direction of the aircraft, external air will be sent into the compression device through the air intake passage (the air inside the air intake passage is pressurized air), making the air pressure inside the air intake passage greater than the air pressure inside the installation cavity. According to the pressurized air intake principle, a part of the external air sent into the air intake passage will flow through the air ejector through hole into the installation cavity, and the air inside the installation cavity will flow from the auxiliary through hole to the radial outer side of the installation cavity. In this way, a second air flow is formed in the installation cavity that flows from the air ejector through hole through the installation cavity to the auxiliary through hole. The second air flow带走 the heat of the control device during the flowing process, realizing heat dissipation of the control device, and another part of the external air sent into the air intake passage is sent into the compression device through the air intake passage.

[0047] In the description of the present invention, it should be noted that the orientation or positional relationship indicated by terms such as "upper", "lower", "one side", "the other side", "one end", "the other end", "side", "opposite", "four corners", "perimeter", "the "mouth" character structure", etc. is based on the orientation or positional relationship shown in the drawings. It is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the structure referred to has a specific orientation, is constructed and operated in a specific orientation, and therefore cannot be understood as a limitation to the present invention.

[0048] It should be noted that there is an incorrect expression "带走" in the translation of item , which should be "carries away". The corrected translation is as follows: It may be that the aircraft is a drone, an airplane, a rocket, etc. All of the above can achieve the purpose of this application. The gist thereof does not depart from the design concept of the present invention and will not be elaborated herein. All of them shall fall within the protection scope of this application. <00In the description of the embodiments of the present invention, unless otherwise expressly specified and limited, the terms "connection," "direct connection," "indirect connection," "fixed connection," "installation," and "assembly" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection. The terms "installation," "connection," and "fixed connection" can refer to a direct connection or an indirect connection through an intermediate medium, or they can refer to the internal communication between two components. Those skilled in the art can understand the specific meaning of the above terms in the present invention based on the specific circumstances.

[0049] While the embodiments disclosed in this invention are as described above, the content is merely for the purpose of facilitating understanding of the invention and is not intended to limit the invention. Any person skilled in the art to which this invention pertains may make any modifications and changes to the form and details of the implementation without departing from the spirit and scope disclosed herein; however, the scope of patent protection of this invention shall still be defined by the appended claims.

Claims

1. An aircraft, characterized in that The aircraft comprises: a body with a pressure device; a ring-shaped mounting member provided at an end of the body, an inner cavity of the mounting member having an air inlet passage formed therein, an air outlet side of the air inlet passage being in communication with an air inlet side of the pressure device, the mounting member having an air injection through hole on a side facing the air inlet passage and an auxiliary through hole on a side facing away from the air inlet passage, the air injection through hole and the auxiliary through hole being in communication through the inner cavity; and a control device provided in the inner cavity; based on the operation of the pressure device, external air is sucked into the pressure device through the air inlet passage, according to the principle of air injection, the negative pressure flow formed in the air inlet passage injects air in the inner cavity of the mounting member to flow to the air inlet passage through the air injection through hole, and air outside the mounting member is supplemented into the inner cavity through the auxiliary through hole; the aircraft further comprises: a curved air inlet passage, an air outlet side of the air inlet passage being in communication with the air inlet side of the air inlet passage; wherein an inner circumferential surface of the air outlet side of the air inlet passage is formed with a boundary layer separation zone, and an axis of the air injection through hole is arranged along an extension direction of the boundary layer separation zone.

2. The aircraft of claim 1, wherein, The air injection through hole has a diameter of not more than 1 mm.

3. The aircraft of claim 1, wherein, The air injection through hole comprises a plurality of air injection through holes, and the plurality of air injection through holes are arranged at intervals in a circumferential direction of the mounting member.

4. The aircraft of any one of claims 1 to 3, wherein, The auxiliary through hole comprises a plurality of auxiliary through holes, and the plurality of auxiliary through holes are arranged at intervals in the circumferential direction of the mounting member, and the auxiliary through hole has a diameter not less than that of the air injection through hole.

5. The aircraft of any one of claims 1 to 3, wherein, An inner circumferential wall of the mounting member comprises a flow guide section tapering from the air inlet side of the air inlet passage to the air outlet side of the air inlet passage, and the air injection through hole is located on the flow guide section.

6. The aircraft of any one of claims 1 to 3, wherein, In the axial direction of the air inlet passage: the auxiliary through hole is located between the air injection through hole and the pressure device.

7. The aircraft of any one of claims 1 to 3, wherein, The mounting member comprises a ring-shaped air inlet casing provided at an end of the body and a ring-shaped fairing, the fairing being sleeved on the air inlet casing and forming the inner cavity together with the air inlet casing.

8. The aircraft of any one of claims 1-3, wherein, The inner cavity is a ring-shaped cavity.

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