An electro-vortex engine, aircraft and control method

Abstract

The present disclosure relates to the field of aviation technology, and particularly provides an electric turbofan engine, an aircraft and a control method to solve the problem that the existing technology cannot match the required electric energy of the engine, which is one of the important problems to be solved in the field. The electric turbofan engine is applied to the aircraft, and the electric turbofan engine comprises: a supercharging assembly, a shell, a first flow channel, a second flow channel, a flow channel switching mechanism, a plurality of turbine generators and a plurality of guide tubes. The supercharging assembly is arranged in the shell, the supercharging assembly is located in the first flow channel, and the second flow channel is formed between the outer side wall of the supercharging assembly and the inner side wall of the shell. Each guide tube is annularly arranged on the outer side wall of the shell, each guide tube is in communication with the second flow channel, the flow channel switching mechanism is arranged at the inlet of each guide tube, and each turbine generator is arranged at the outlet of each guide tube. The electric turbofan engine, the aircraft and the control method provided by the present disclosure are used to increase the power of the electric turbofan engine.

Description

TECHNICAL FIELD

[0001] The present disclosure relates to the field of aviation technology, in particular to an electric turbofan engine, an aircraft and a control method. BACKGROUND

[0002] The combination of laser weapons and intelligent electric unmanned aircraft is an inevitable trend of future combat platforms, and the comprehensive combat effectiveness cannot be achieved without efficient and high-performance propulsion systems. At the same time, with a substantial power extraction from the power propulsion system, based on the existing mature turbofan engine, it is the most feasible technical approach to develop a multi-electric engine to replace the conventional mechanical system for high-power extraction.

[0003] As a key technology for the performance improvement of next-generation aircraft and engines, multi-electric technology has attracted keen attention from countries around the world. Some countries have carried out verification of the embedded starter-generator dual-shaft power extraction scheme, with output power reaching megawatt level. With the development of multi-electric technology, the output power of multi-electric technology cannot meet the huge demand of new sensors and directional energy weapons on future fighter aircraft. Therefore, through the power-optimized aircraft (POA) technology verification program, some countries have carried out research and verification of the integration of embedded generators and engine components / systems. However, advanced airborne systems and weapon systems require more than megawatt of power.

[0004] Therefore, how to solve the problem that the existing technology cannot match the required electric energy of the engine is one of the important problems to be solved in the field. SUMMARY

[0005] Therefore, the embodiments of the present disclosure provide an electric turbofan engine, an aircraft and a control method to solve the problem that the existing technology cannot match the required electric energy of the engine.

[0006] According to one aspect of the present disclosure, an electric turbofan engine is provided, which is applied to an aircraft, and the electric turbofan engine comprises: a supercharging assembly, a shell, a first flow channel, a second flow channel, a flow channel switching mechanism, a plurality of turbine generators and a plurality of guide tubes, the supercharging assembly is arranged in the shell, the supercharging assembly is located in the first flow channel, the second flow channel is formed between the outer side wall of the supercharging assembly and the inner side wall of the shell, each guide tube is annularly arranged on the outer side wall of the shell, each guide tube is in communication with the second flow channel, the flow channel switching mechanism is arranged at the inlet of each guide tube, and each turbine generator is arranged at the outlet of each guide tube.

[0007] When the aircraft is in a first state, the flow channel switching mechanism is used to close the inlet of each guide tube.

[0008] When the aircraft is in a second state, the flow channel switching mechanism is used to open the inlet of each guide tube.

[0009] In addition, the turbofan engine according to an aspect of the present disclosure, the booster assembly comprises: a fan, a compressor, a combustion chamber and a turbine, the fan is fixedly connected with one end of the compressor, the other end of the compressor is communicated with the combustion chamber, the combustion chamber is connected with the turbine, each guide pipe is arranged at a position close to the compressor.

[0010] The turbofan engine according to an aspect of the present disclosure, the flow passage switching mechanism comprises a first baffle and a second baffle, the first baffle has a first hinged side and a first extended side opposite to the first hinged side, the second baffle has a second hinged side and a second extended side opposite to the second hinged side, the thickness of the first hinged side is greater than the thickness of the first extended side;

[0011] The first hinged side is hinged on the outer side wall of the compressor, the second hinged side is hinged on the inlet position of each guide pipe, and the thickness of the second hinged side is greater than the thickness of the second extended side.

[0012] The turbofan engine according to an aspect of the present disclosure, the thickness of the first baffle decreases along the distribution direction from the first hinged side to the first extended side, and the thickness of the second baffle decreases along the distribution direction from the second hinged side to the second extended side.

[0013] The turbofan engine according to an aspect of the present disclosure, the first baffle and the second baffle are both wedge-shaped baffles, the wedge-shaped tip of the first baffle is the extended end of the first baffle, and the wedge-shaped tip of the second baffle is the extended end of the second baffle; and / or,

[0014] The flow passage switching mechanism further comprises a connecting rod, the connecting rod is hinged on the same side of the first baffle and the second baffle, and the connecting rod is arranged on the outer side wall of the compressor.

[0015] The turbofan engine according to an aspect of the present disclosure, the flow guiding surface of the first baffle and the flow guiding surface of the second baffle are planar or curved.

[0016] The turbofan engine according to an aspect of the present disclosure, the fan, the compressor, the combustion chamber and the turbine rotate coaxially.

[0017] According to another aspect of the present disclosure, a kind of aircraft is provided, comprising the above-mentioned turbofan engine.

[0018] According to yet another aspect of the present disclosure, a kind of flight mode control method is provided, the above-mentioned aircraft, method comprises:

[0019] In response to the first state flight instruction, control turbofan engine works;

[0020] Control the flow passage switching mechanism to close the inlet of each guide pipe, and the first state flight instruction is used to indicate that each turbine generator is in non-working state;

[0021] in response to the second state flight instruction, control the electric turbofan engine to work;

[0022] The control flow passage switching mechanism opens the inlet of each guide pipe, and the first state flight instruction is used to instruct each turbine generator to be in a working state.

[0023] According to the flight mode control method of one aspect of the present disclosure, the flight mode control method further comprises:

[0024] determining a target physical parameter of the electric turbofan engine under the condition that the electric turbofan engine works normally;

[0025] determining an actual physical parameter of the turbine generator according to the electric turbofan engine under the condition that the electric turbofan engine works normally;

[0026] determining a geometric parameter of the turbine generator based on the target physical parameter and the actual physical parameter.

[0027] The above at least one technical solution adopted by the embodiments of the present disclosure can achieve the following beneficial effects: The electric turbofan engine is applied to an aircraft, the supercharging assembly is arranged in the shell, the supercharging assembly is located in the first flow passage, the second flow passage is formed between the outer side wall of the supercharging assembly and the inner side wall of the shell, each guide pipe is arranged on the outer side wall of the shell in a ring shape, each guide pipe is in communication with the second flow passage, the flow passage switching mechanism is arranged at the inlet of each guide pipe, and each turbine generator is arranged at the outlet of each guide pipe. When air flows into the supercharging assembly, part of the fluid flows into the first flow passage, and another part of the fluid flows into the second flow passage. When the aircraft is in a first state, the flow passage switching mechanism is used to close the inlet of each guide pipe. At this time, each turbine generator arranged at the outlet of each guide pipe is in a non-working state. When the aircraft is in a second state, the flow passage switching mechanism is used to open the inlet of each guide pipe. Part of the fluid flows into each guide pipe to drive each turbine generator arranged at the outlet of each guide pipe to work. Each turbine generator works to generate electricity to provide electric energy for the electric turbofan engine, effectively solving the problem that the electric energy required by the engine cannot be matched in the prior art. At the same time, each turbine generator is arranged in each guide pipe, which has little influence on the structural layout of the prime mover system and is flexible in installation mode, compact in structure, and can realize modular installation according to the required power of the electric turbofan engine. BRIEF DESCRIPTION OF DRAWINGS

[0028] 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 the prior art description will be briefly introduced below. Obviously, the drawings in the following description are only some embodiments of the present disclosure, and other drawings can be obtained by those skilled in the art without creating any creative labor.

[0029] Figure 1FIG. 1 is a structural schematic diagram of an electric turbofan engine to which an embodiment of the present disclosure is applied.

[0030] Figure 2 FIG. 1 is a structural schematic diagram of an electric turbofan engine to which an embodiment of the present disclosure is applied.

[0031] Reference Signs:

[0032] 1 - supercharging assembly, 11 - fan, 12 - compressor, 13 - combustion chamber, 14 - turbine, 2 - flow guide pipe, 3 - turbo generator, 4 - first flow passage, 5 - second flow passage, 6 - flow passage switching mechanism, 61 - first damper, 62 - connecting rod, 63 - second damper. DETAILED DESCRIPTION

[0033] Embodiments of the present disclosure will be described in more detail by referring to the drawings. Although certain embodiments of the present disclosure are shown in the drawings, it is understood that the present disclosure can be implemented in various forms and should not be construed as being limited to the embodiments set forth herein, but rather these embodiments are provided so as to more completely and thoroughly understand the present disclosure. It is understood that the drawings and embodiments of the present disclosure are for exemplary purposes only and are not intended to limit the scope of the present disclosure.

[0034] It is understood that each of the steps recited in the method embodiments of the present disclosure can be executed in different orders and / or in parallel. In addition, the method embodiments can include additional steps and / or omit the execution of the steps shown. The scope of the present disclosure is not limited in this respect.

[0035] The term "comprising" and variations thereof as used in the present disclosure are open-ended, that is, "including 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 are given throughout the description. It is noted that the concepts "first", "second", etc. mentioned in the present disclosure are merely used to distinguish different apparatuses, modules or units, and are not intended to limit the order or interdependence of the functions performed by these apparatuses, modules or units.

[0036] It is noted that the modification "one", "multiple" mentioned in the present disclosure is illustrative and not restrictive, and those skilled in the art should understand that, unless otherwise explicitly indicated in the context, it should be understood as "one or more".

[0037] The names of the messages or information exchanged between the plurality of apparatuses in the embodiments of the present disclosure are merely for illustrative purposes, and are not intended to limit the scope of the messages or information.

[0038] The combination of laser weapons and intelligent electric unmanned aircraft is an inevitable trend of future combat platforms, and the comprehensive combat effectiveness cannot be achieved without efficient and high-performance propulsion systems. With a large power extraction from the existing mature turbofan engine, the development of a multi-electric engine to replace the conventional mechanical system is the most feasible technical approach.

[0039] As a key technology for the performance improvement of next-generation aircraft and engines, multi-electric technology has attracted keen attention from countries around the world. Some countries have carried out verification of the embedded starter-generator dual-shaft power extraction scheme, with output power reaching megawatt level. With the development of multi-electric technology, the output power of multi-electric technology cannot meet the huge demand of new sensors and directional energy weapons on future fighter aircraft. Therefore, through the power optimization aircraft (POA) technology verification program, some countries have carried out research and verification of the integration of embedded generators and engine components / systems. However, advanced airborne systems and weapon systems require more than megawatt of power.

[0040] To solve the problem that the existing technology cannot match the required electric energy of the engine, the electric turbofan combined cycle engine, aircraft and control method according to the embodiments of the present disclosure are provided.

[0041] An electric turbofan combined cycle engine according to the embodiments of the present disclosure will be described in detail below with reference to the accompanying drawings.

[0042] Figure 1 The structure of the electric turbofan engine according to the embodiments of the present disclosure is shown in the figure. As shown in the figure, Figure 1 An electric turbofan engine is applied to an aircraft, and the electric turbofan engine includes a supercharging assembly 1, a housing, a first flow channel 4, a second flow channel 5, a flow channel switching mechanism 6, a plurality of turbine generators 3 and a plurality of guide tubes 2. The supercharging assembly 1 is arranged in the housing, the supercharging assembly 1 is located in the first flow channel 4, and the second flow channel 5 is formed between the outer side wall of the supercharging assembly 1 and the inner side wall of the housing. Each guide tube 2 is annularly arranged on the outer side wall of the housing, each guide tube 2 is in communication with the second flow channel 5, the flow channel switching mechanism 6 is arranged at the inlet of each guide tube 2, and each turbine generator 3 is arranged at the outlet of each guide tube 2. When the aircraft is in a first state, the flow channel switching mechanism 6 is used to close the inlet of each guide tube 2. When the aircraft is in a second state, the flow channel switching mechanism 6 is used to open the inlet of each guide tube 2.

[0043] In practical application, the electric vortex engine is applied to the aircraft, the pressurizing assembly 1 is arranged in the shell, the pressurizing assembly 1 is located in the first flow channel 4, the outer side wall of the pressurizing assembly 1 and the inner side wall of the shell form the second flow channel 5, each guide pipe 2 is annularly arranged on the outer side wall of the shell, each guide pipe 2 is communicated with the second flow channel 5, the flow channel switching mechanism 6 is arranged at the inlet of each guide pipe 2, and each turbine generator 3 is arranged at the outlet of each guide pipe 2. When the air flows into the pressurizing assembly 1, part of the fluid flows into the first flow channel 4, and the other part of the fluid flows into the second flow channel 5. When the aircraft is in the first state, the flow channel switching mechanism 6 is used to close the inlet of each guide pipe 2, at this time, each turbine generator 3 arranged at the outlet of each guide pipe 2 is in a non-working state. When the aircraft is in the second state, the flow channel switching mechanism 6 is used to open the inlet of each guide pipe 2, part of the fluid flows into each guide pipe 2, and drives each turbine generator 3 located at the outlet of each guide pipe 2 to work. Each turbine generator 3 works to generate electricity to provide electric energy for the electric vortex engine, effectively solving the problem that the electric energy required by the existing technology cannot be matched with the electric energy required by the engine. At the same time, each turbine generator 3 is arranged in each guide pipe 2, which has little influence on the structural layout of the original power system and flexible installation mode, compact structure, and can realize modular installation according to the required power size of the electric vortex engine.

[0044] As shown in Figure 1 , the pressurizing assembly 1 comprises a fan 11, a compressor 12, a combustion chamber 13 and a turbine 14, the fan 11 is fixedly connected with one end of the compressor 12, the other end of the compressor 12 is communicated with the combustion chamber 13, the combustion chamber 13 is connected with the turbine 14, and each guide pipe 2 is arranged at a position close to the compressor 12. It can be understood that the above-mentioned fan 11, compressor 12, combustion chamber 13 and turbine 14 rotate coaxially.

[0045] In another optional mode, Figure 2 Fig. 2 shows a structural schematic diagram of the flow channel switching mechanism applied according to the embodiment of the present disclosure, as shown in Figure 2 , the flow channel switching mechanism comprises a first baffle 61 and a second baffle 63, the first baffle 61 has a first hinged side and a first extended side opposite to the first hinged side, the second baffle 63 has a second hinged side and a second extended side opposite to the second hinged side, and the thickness of the first hinged side is greater than the thickness of the first extended side. The first hinged side is hinged on the outer side wall of the compressor, and the second hinged side is hinged at the inlet position of each guide pipe, and the thickness of the second hinged side is greater than the thickness of the second extended side. Since the thickness of the second hinged side is greater than the thickness of the second extended side, the first baffle 61 and the second baffle 63 are easy to control, so as to facilitate the control of the positions of the first baffle 61 and the second baffle 63, and effectively switch the flight state of the aircraft.

[0046] For example, the thickness of the first baffle 61 decreases along the distribution direction from the first hinged side to the first extended side, so that the first baffle 61 is more easily controlled due to the air pressure difference caused by different air flows when the air flows through the first baffle 61. The thickness of the second baffle 63 decreases along the distribution direction from the second hinged side to the second extended side. Here, the beneficial effects of the first baffle 61 are the same as described above, and will not be repeated here.

[0047] As shown in the example, Figure 2 The first baffle 61 and the second baffle 63 are both wedge-shaped baffles, the wedge-shaped tip of the first baffle 61 is the extended end of the first baffle 61, and the wedge-shaped tip of the second baffle 63 is the extended end of the second baffle 63. The wedge-shaped structure has a self-locking feature, so that the flow channel switching mechanism can stably guide the air flow in different flight states of the aircraft, thereby obtaining sufficient thrust in different flight modes of the aircraft.

[0048] In an example, as shown in the example, Figure 2 The flow channel switching mechanism of the example embodiment of the present disclosure further includes a connecting rod 62, which is hinged on the same side of the first baffle 61 and the second baffle 63 and is arranged on the outer side wall of the compressor. The first baffle 61 and the second baffle 63 are connected through the connecting rod 62, and the connecting rod 62 drives the second baffle 63 to open or close the inlet of each flow guide pipe, and also synchronously deflects the first baffle 61 and the second baffle 63. At the same time, the connecting rod 62 arranged on the outer side wall of the compressor can reduce air resistance.

[0049] As shown in the example, Figure 2 The flow guide surface of the first baffle 61 and the flow guide surface of the second baffle 63 are flat or curved.

[0050] The example embodiment of the present disclosure provides an aircraft, which includes the electric turbofan engine of the example embodiment of the present disclosure. It should be understood that the aircraft of the example embodiment of the present disclosure can be a drone or a manned aircraft. For example,

[0051] Compared with the prior art, the beneficial effects of the aircraft provided by the example embodiment of the present disclosure refer to the beneficial effects of the electric turbofan engine, which will not be repeated here.

[0052] It should be noted that the aircraft provided by the example 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 flow channel switching mechanism to open or close the inlet of each flow guide pipe according to the flight state.

[0053] In an optional manner, the flight mode control method is applied to the aircraft described above, and the flight mode control method includes:

[0054] controlling the electric turbofan engine to operate in response to the first flight instruction;

[0055] the control flow passage switching mechanism closes the inlet of each guide pipe, and the first flight instruction is used to indicate that each turbo-generator is in a non-operating state;

[0056] controlling the electric turbofan engine to operate in response to the second flight instruction;

[0057] the control flow passage switching mechanism opens the inlet of each guide pipe, and the first flight instruction is used to indicate that each turbo-generator is in an operating state.

[0058] Exemplarily, the flight mode control method further comprises:

[0059] determining a target physical parameter of the electric turbofan engine under normal operation of the electric turbofan engine, and it can be understood that the target physical parameter is actually the power Pz required for the normal operation of the electric turbofan engine.

[0060] determining an actual physical parameter of the turbo-generator according to the electric turbofan engine under normal operation, and it can be understood that the actual physical parameter is actually the power Px generated by the second flow passage gas passing through the turbo-generator per unit air mass flow, wherein the calculation formula of Px is as follows:

[0061]

[0062] wherein C P is the constant pressure specific heat capacity of air, η T is the turbine efficiency, ηe is the generator efficiency, T t is the turbine inlet temperature, P t is the turbine inlet pressure, P0 is the standard atmospheric pressure, and k is the constant entropy index of air.

[0063] determining the geometric parameters of the turbo-generator based on the target physical parameter and the actual physical parameter, and it can be understood that the required air mass flow and quantity of the turbo-generator are determined according to the values of Pz and Px.

[0064] For example, taking a medium-thrust turbofan engine as an example, the air turbine inlet temperature Tt is 450K, the air turbine inlet pressure Pt is 2.7 atmospheres, the air turbine efficiency ηT is 0.9, and the generator efficiency ηe is 0.88, then Px = 88.05kw / kg, so that 88.05kw of electric energy can be generated per kg of flow of air in the second flow passage. Two turbo-generators are arranged along the outer side wall of the supercharging assembly, and 3kg of air in the second flow passage is introduced into each turbo-generator, so as to meet the power demand of general airborne laser weapons.

[0065] 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.

[0066] 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 illustrative of some embodiments of the present disclosure and of the principles thereof. It is to be understood that modifications can be made to the above embodiments 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. An electric turbofan engine, characterized in that, For use in aircraft, the turbofan engine includes: a pressurization assembly, a housing, a first flow channel, a second flow channel, a flow channel switching mechanism, multiple turbine generators, and multiple guide pipes. The pressurization assembly is disposed within the housing and located within the first flow channel. The second flow channel is formed between the outer sidewall of the pressurization assembly and the inner sidewall of the housing. Each guide pipe is arranged around the outer sidewall of the housing and communicates with the second flow channel. The flow channel switching mechanism is disposed at the inlet of each guide pipe, and each turbine generator is disposed at the outlet of each guide pipe. When the aircraft is in the first state, the flow channel switching mechanism is used to close the inlet of each of the flow guide pipes; When the aircraft is in the second state, the flow channel switching mechanism is used to open the inlet of each of the flow guide pipes; The supercharging assembly includes a fan, a compressor, a combustion chamber, and a turbine. The fan is fixedly connected to one end of the compressor, the other end of the compressor is connected to the combustion chamber, the combustion chamber is connected to the turbine, and each of the guide pipes is located near the compressor. The flow channel switching mechanism includes a first baffle and a second baffle. The first baffle has a first hinge side and a first extended side opposite to the first hinge side. The second baffle has a second hinge side and a second extended side opposite to the second hinge side. The thickness of the first hinge side is greater than the thickness of the first extended side. The first hinge side is hinged to the outer wall of the compressor, and the second hinge side is hinged to the inlet of each of the guide tubes. The thickness of the second hinge side is greater than the thickness of the second extension side. Both the first baffle and the second baffle are wedge-shaped baffles, with the wedge-shaped tip of the first baffle being the protruding end of the first baffle, and the wedge-shaped tip of the second baffle being the protruding end of the second baffle; and / or, The flow channel switching mechanism also includes a connecting rod, which is hinged to the same side of the first baffle and the second baffle, and the connecting rod is located on the outer wall of the compressor.

2. The electric turbofan engine according to claim 1, characterized in that, The thickness of the first baffle decreases along the distribution direction from the first hinge side to the first extension side, and the thickness of the second baffle decreases along the distribution direction from the second hinge side to the second extension side.

3. The electric turbofan engine according to claim 1, characterized in that, The guiding surface of the first baffle and the guiding surface of the second baffle are either planes or curved surfaces.

4. The electric turbofan engine according to claim 1, characterized in that, The fan, the compressor, the combustion chamber, and the turbine rotate coaxially.

5. An aircraft, characterized in that, Includes the electric turbofan engine as described in any one of claims 1-4.

6. A flight mode control method, characterized in that, For use with the aircraft of claim 5, the method includes: In response to the first-state flight command, control the operation of the turbofan engine; The flow channel switching mechanism closes the inlet of each of the flow channels, and the first state flight command is used to indicate that each turbine generator is in a non-operating state; In response to the second-state flight command, control the operation of the turbofan engine; The flow channel switching mechanism opens the inlet of each of the flow channels, and the first state flight command is used to indicate that each turbine generator is in the working state.

7. The flight mode control method according to claim 6, characterized in that, The flight mode control method also includes: Determine the target physical parameters of the electric turbofan engine under normal operating conditions; Determine the actual physical parameters of the turbine generator based on the normal operating conditions of the electric turbofan engine; The geometric parameters of the turbine generator are determined based on the target physical parameters and the actual physical parameters.

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