Turbine engine system and aircraft

Abstract

The invention relates to the technical field of aircrafts, and particularly provides a turbine engine system and an aircraft, the turbine engine system comprises a main propulsion channel, an auxiliary channel, a turbine engine, a precooler, an air turbine and a radiator; the end, away from the main air inlet, of the main propelling channel is connected with a first exhaust nozzle. The precooler and the turbine engine are located in the main propulsion channel. The end, away from the auxiliary air inlet, of the auxiliary channel is connected with a second exhaust nozzle. The air turbine and the radiator are located in the auxiliary channel. The precooler is provided with a first channel for a heat exchange working medium to flow, the radiator is provided with a second channel for the heat exchange working medium to flow, an outlet of the first channel is communicated with an inlet of the second channel, and an outlet of the second channel is communicated with an inlet of the first channel, so that the heat exchange working medium circularly flows between the first channel and the second channel; therefore, the regeneration and self-sustaining of the heat sink are realized while the air inlet temperature of the engine is reduced, and the long-endurance operation of the engine is ensured.

Description

Technical Field

[0001] The present disclosure relates to the technical field of aircraft, and more particularly to a turbofan engine system and an aircraft. Background Art

[0002] The high-speed development of aircraft has extremely important military and civilian values. However, the aerodynamic heating effect generated during high-speed flight restricts the flight limit of the aircraft.

[0003] The turbofan engine is an important component of the aircraft, which is used to provide flight power for the aircraft. In order to increase the flight speed of the turbofan engine, the related art cools the ram air intake through a pre-cooler structure to reduce the temperature of the inlet air flow, so as to offset the temperature increase caused by the aerodynamic heating effect during high-speed flight to a certain extent. However, in the related art, the heat sink required for flight thermal management is insufficient, resulting in an unsustainable pre-cooling process, making it difficult for the turbofan engine to operate for a long time. Summary of the Invention

[0004] In view of the above problems, the present disclosure provides a turbofan engine system and an aircraft.

[0005] In a first aspect, the present disclosure provides a turbofan engine system, including a main propulsion channel, an auxiliary channel, a turbofan engine, a pre-cooler, an air turbine, and a radiator;

[0006] One end of the main propulsion channel has a main air inlet, and the end of the main propulsion channel far from the main air inlet is connected to a first tail nozzle; the pre-cooler and the turbofan engine are located in the main propulsion channel and are arranged in sequence in the direction from the main air inlet to the first tail nozzle;

[0007] One end of the auxiliary channel has an auxiliary air inlet, and the end of the auxiliary channel far from the auxiliary air inlet is connected to a second tail nozzle; the air turbine and the radiator are located in the auxiliary channel and are arranged in sequence in the direction from the auxiliary air inlet to the second tail nozzle;

[0008] The pre-cooler has a first channel for the heat transfer working medium to flow through to cool the air entering from the main air inlet, and the radiator has a second channel for the heat transfer working medium to flow through to heat the air cooled by the expansion of the air turbine. The outlet of the first channel is communicated with the inlet of the second channel, and the outlet of the second channel is communicated with the inlet of the first channel, so that the heat transfer working medium circulates between the first channel and the second channel.

[0009] Optionally, the turbofan engine system further includes a heat transfer working medium storage container for storing the heat transfer working medium;

[0010] The first channel and the second channel are respectively communicated with the heat exchange medium storage container.

[0011] Optionally, the turbine engine system further includes a working fluid pump;

[0012] The working fluid pump is connected between the first channel and the second channel, and is used to drive the heat exchange working fluid to circulate between the first channel and the second channel;

[0013] And / or, the heat exchange medium storage container is located outside the main propulsion channel and the auxiliary channel.

[0014] Optionally, the heat exchange medium is a liquid-cooled metal medium.

[0015] Optionally, the liquid-cooled metal working fluid includes a gallium-based alloy.

[0016] Optionally, the turbine engine system further includes an auxiliary compressor;

[0017] The auxiliary compressor is arranged in the auxiliary channel and is located on a side of the radiator away from the air turbine. The auxiliary compressor is connected to the air turbine, and the second tail nozzle is communicated with an exhaust port of the auxiliary compressor.

[0018] Optionally, a first opening and closing member is provided at the main air inlet, and the first opening and closing member can rotate relative to the main air inlet to adjust the opening and closing state of the main air inlet;

[0019] And / or, a second opening and closing member is provided at the auxiliary air inlet, and the second opening and closing member can rotate relative to the auxiliary air inlet to adjust the opening and closing state of the auxiliary air inlet.

[0020] Optionally, the radiator includes a hollow tube body and radiating fins;

[0021] The inner cavity of the tube body is formed as the second channel, and the heat dissipation fins are arranged on the outer wall of the tube body and are in heat-conducting contact with the outer wall of the tube body;

[0022] And / or, the precooler comprises at least three heat-conducting baffles stacked in sequence, a channel is formed between two adjacent heat-conducting baffles, and each of the channels is provided with a heat-conducting fin;

[0023] Among two adjacent channels, one channel is formed as the first channel, an inlet of the other channel is communicated with the main air inlet, and an outlet of the other channel is communicated with the air inlet of the turbine engine.

[0024] Optionally, the expansion ratio of the air turbine is in the range of 8:1 to 12:1.

[0025] In a second aspect, the present disclosure provides an aircraft, including the turbofan engine system as described above.

[0026] For the turbofan engine system and the aircraft provided by the embodiments of the present disclosure, by providing a main propulsion channel and an auxiliary channel, a precooler is arranged in the main propulsion channel to cool the air entering from the main air inlet through the precooler. By arranging an air turbine and a radiator in the auxiliary channel, the heat transfer medium channels of the precooler and the radiator are connected, so that the heat transfer medium can circulate in the heat transfer medium channels of the precooler and the radiator. That is, the air cooled by the expansion of the air turbine exchanges heat with the heat transfer medium in the radiator to heat the air. At this time, the heat transfer medium is cooled because it transfers heat to the air in the auxiliary channel. The cooled heat transfer medium enters the heat transfer medium channel of the precooler and exchanges heat with the air entering from the main air inlet, thereby reducing the intake air temperature of the turbofan engine.

[0027] That is to say, while effectively reducing the intake air temperature of the turbofan engine, all the heat sinks required for heat exchange in the precooler come from the heat sinks possessed by the auxiliary channel after the air is cooled by the air turbine. The way of obtaining the heat sink is renewable and self-sustaining, ensuring the continuity of the precooling process, without relying on limited heat sink resources such as the fuel of the aircraft, and avoiding the problem of insufficient heat sink caused by using the fuel for combustion of the aircraft as the heat sink resource. Thus, the long-endurance operation of the turbofan engine is ensured, the flight envelope of the turbofan engine is extended, and the flight performance and the scope of use of the aircraft using the turbofan engine system are improved.

[0028] In addition, since the radiator in the auxiliary channel heats the air, while providing a regenerative heat sink for the precooler, the air ejected from the second nozzle also has a certain amount of energy, providing auxiliary thrust for the flight of the aircraft, greatly offsetting the resistance generated by the projected area of the auxiliary channel, thereby ensuring the flight efficiency of the aircraft.

[0029] That is, for the turbofan engine system and the aircraft provided by the embodiments of the present disclosure, through the coordinated operation among the main propulsion channel, the auxiliary channel, the air turbine, the precooler and the radiator, the regeneration and self-sustenance of the heat sink are realized, so that the turbofan engine can effectively reduce the intake air temperature during high-speed flight, improve the engine performance, ensure the long-endurance operation of the engine and the flight efficiency of the aircraft.

[0030] It should be understood that both the foregoing general description and the following detailed description are exemplary and are intended to provide further explanation of the claimed technology. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] The above and other objects, features, and advantages of the present disclosure will become more apparent by describing the embodiments of the present disclosure in more detail with reference to the accompanying drawings. The accompanying drawings are used to provide a further understanding of the embodiments of the present disclosure and constitute a part of the specification. They are used together with the embodiments of the present disclosure to explain the present disclosure and do not constitute a limitation on the present disclosure. In the drawings, the same reference numerals generally represent the same components or steps.

[0032] Figure 1 Structural schematic of a turbo engine system provided by an embodiment of the present disclosure Figure 1 ;

[0033] Figure 2 Structural schematic of a turbo engine system provided by an embodiment of the present disclosure Figure 2 ;

[0034] Figure 3 Partial structural schematic diagram of a pre-cooler in the turbo engine system according to an embodiment of the present disclosure;

[0035] Figure 4 Structural schematic diagram of a radiator in the turbo engine system according to an embodiment of the present disclosure.

[0036] Wherein, 1, main propulsion channel; 11, main air inlet; 12, first opening and closing member; 2, auxiliary channel; 21, auxiliary air inlet; 22, second opening and closing member; 3, pre-cooler; 30, first channel; 31, inlet; 32, outlet; 33, heat-conducting partition; 34, heat-conducting fins; 4, turbo engine; 5, air turbine; 6, radiator; 60, second channel; 61, inlet; 62, outlet; 63, pipe body; 64, heat-dissipating fins; 7, auxiliary compressor; 8, first tail nozzle; 9, second tail nozzle; 101, heat exchange working fluid storage container; 102, working fluid pump. Detailed implementation manners

[0037] In order to make the objectives, technical solutions, and advantages of the present disclosure more apparent, exemplary embodiments according to the present disclosure will be described in detail below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present disclosure, rather than all the embodiments of the present disclosure. It should be understood that the present disclosure is not limited by the exemplary embodiments described herein.

[0038] It should be understood that the steps recorded in the method embodiments of the present disclosure can be executed 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 regard.

[0039] As used herein, the term "comprising" and its variants are open-ended, that is, "including but not limited to". The term "based on" means "at least partially based 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". Relevant definitions of other terms will be given in the following description. It should be noted that the concepts such as "first", "second", etc. mentioned in the present 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.

[0040] It should be noted that the modifications of "one" and "a plurality" mentioned in the present disclosure are illustrative rather than restrictive. Those skilled in the art should understand that unless otherwise clearly specified in the context, it should be understood as "one or more".

[0041] Referring to Figures 1 to 4 As shown, an embodiment of the present disclosure provides a turbofan engine system, which is specifically applied to an aircraft.

[0042] The turbofan engine system includes: a main propulsion channel 1, an auxiliary channel 2, a turbofan engine 4, a precooler 3, an air turbine 5, and a radiator 6.

[0043] Among them, one end of the main propulsion channel 1 has a main air inlet 11, and the end of the main propulsion channel 1 far from the main air inlet 11 is connected to a first tail nozzle 8. The precooler 3 and the turbofan engine 4 are located in the main propulsion channel 1 and are arranged in sequence in the direction from the main air inlet 11 to the first tail nozzle 8.

[0044] Among them, one end of the auxiliary channel 2 has an auxiliary air inlet 21, and the end of the auxiliary channel 2 far from the auxiliary air inlet 21 is connected to a second tail nozzle 9. The air turbine 5 and the radiator 6 are located in the auxiliary channel 2 and are arranged in sequence in the direction from the auxiliary air inlet 21 to the second tail nozzle 9.

[0045] The precooler 3 has a first channel 30 for the heat exchange working medium to flow through to cool the air entering from the main air inlet 11. The radiator 6 has a second channel 60 for the heat exchange working medium to flow through to heat the air that has been expanded and cooled by the air turbine 5. Among them, the outlet 32 of the first channel 30 is communicated with the inlet 61 of the second channel 60, and the outlet 62 of the second channel 60 is communicated with the inlet 31 of the first channel 30, so that the heat exchange working medium circulates between the first channel 30 and the second channel 60.

[0046] The main function of the main propulsion channel 1 is to generate the thrust required for the aircraft. Specifically, during the flight of the aircraft, the outside air enters the main propulsion channel 1 through the main air inlet 11, first passes through the precooler 3, and the air exchanges heat with the heat transfer working fluid in the first channel 30 of the precooler 3. The heat of the air is transferred to the heat transfer working fluid. At this time, the air is cooled and the heat transfer working fluid is heated. The cooled air enters the turbofan engine 4, improving the performance of the turbofan engine 4 and ensuring the normal operation of the turbofan engine 4. The gas discharged from the exhaust port of the turbofan engine 4 finally sprays out from the first tail nozzle 8, thereby providing power for the aircraft. Among them, the heat transfer working fluid carrying heat flows out from the outlet 32 of the first channel 30, and then flows into the second channel 60 from the inlet 61 of the second channel 60.

[0047] The main function of the auxiliary channel 2 is to dissipate the heat absorbed by the precooler 3 in the main propulsion channel 1 and at the same time provide a certain amount of auxiliary thrust for the aircraft. Specifically, the air turbine 5 indirectly cools the air by improving the air flow and at the same time increases the pressure of the air. The air entering from the auxiliary air inlet 21 first passes through the air turbine 5, that is, the air temperature drops significantly during this process. The cooled air passes through the radiator 6 and exchanges heat with the heat transfer working fluid in the second channel 60 of the radiator 6, absorbing the heat of the heat transfer working fluid. At this time, the heat transfer working fluid is cooled and the air is heated, that is, energy is input into the air. The heated and pressurized air finally sprays out from the second tail nozzle 9, generating auxiliary thrust for the aircraft and reducing the flight resistance.

[0048] Among them, the heat transfer working fluid cooled due to heat exchange in the second channel 60 enters the first channel 30 of the precooler 3, and then exchanges heat with the air entering from the main air inlet 11, absorbing the heat in the air. At this time, the heat transfer working fluid is heated, and the heat transfer working fluid carrying heat enters the second channel 60 from the outlet 32 of the first channel 30, and exchanges heat with the air passing through the radiator 6 in the auxiliary channel 2 again, transferring the heat to the air passing through the radiator 6. In this way, a continuous heat sink is provided for the precooler 3, making the heat sink renewable and self-sustaining, ensuring the sufficiency of the heat sink, and thus ensuring the long endurance operation of the turbofan engine 4.

[0049] The turbine engine system provided by the embodiments of the present disclosure is provided with a main propulsion channel 1 and an auxiliary channel 2. A precooler 3 is arranged in the main propulsion channel 1. The air entering from the main air inlet 11 is cooled by the precooler 3. By arranging an air turbine 5 and a radiator 6 in the auxiliary channel 2, the heat transfer working medium channels of the precooler 3 and the radiator 6 are connected, so that the heat transfer working medium can circulate in the heat transfer working medium channels of the precooler 3 and the radiator 6. That is, the air cooled by expanding through the air turbine 5 exchanges heat with the heat transfer working medium in the radiator 6 to heat the air. At this time, the heat transfer working medium is cooled because it transfers heat to the air in the auxiliary channel 2. The cooled heat transfer working medium enters the heat transfer working medium channel of the precooler 3 and exchanges heat with the air entering from the main air inlet 11, thereby reducing the intake air temperature of the turbine engine 4.

[0050] That is to say, while effectively reducing the intake air temperature of the turbine engine 4, all the heat sinks required for heat exchange of the precooler 3 come from the heat sinks possessed by the auxiliary channel 2 after being cooled by the air turbine 5. The acquisition method of the heat sink is renewable and self-sustaining, ensuring the continuity of the precooling process. There is no need to rely on limited heat sink resources such as the fuel of the aircraft, avoiding the problem of insufficient heat sink caused by using the fuel for combustion of the aircraft as the heat sink resource, thereby ensuring the long endurance operation of the turbine engine 4, expanding the flight envelope of the turbine engine 4, and improving the flight performance and application range of the aircraft using this turbine engine system.

[0051] In addition, since the radiator 6 in the auxiliary channel 2 heats the air, while providing a regenerative heat sink for the precooler 3, it also makes the air ejected from the second tail nozzle 9 have a certain amount of energy, providing auxiliary thrust for the flight of the aircraft, and greatly offsetting the resistance generated by the projected area of the auxiliary channel 2, thereby ensuring the flight efficiency of the aircraft.

[0052] That is to say, the turbine engine system and the aircraft provided by the embodiments of the present disclosure achieve the regeneration and self-sustenance of the heat sink through the coordinated operation of the main propulsion channel 1, the auxiliary channel 2, the air turbine 5, the precooler 3, and the radiator 6. When the turbine engine 4 is flying at high speed, it can effectively reduce the intake air temperature, improve the engine performance, ensure the long endurance operation of the engine, and the flight efficiency of the aircraft.

[0053] The turbine engine system provided by the embodiments of the present disclosure can be specifically applied to aeroengines, variable cycle engines, hypersonic turbine-based combined cycle engines, etc.

[0054] In some embodiments, the expansion ratio of the air turbine 5 can be set between 8:1 and 12:1, which further ensures the cooling and pressurizing effect of the air turbine 5 on the air, thereby further ensuring the heat exchange efficiency, improving the renewability and sustainability of the heat sink, and increasing the pressure of the gas ejected from the second tail nozzle 9, thereby increasing the thrust.

[0055] In some embodiments, the turbine engine system also includes an auxiliary compressor 7, which is arranged in the auxiliary channel 2 and is located on the side of the radiator 6 facing away from the air turbine 5. The auxiliary compressor 7 is connected to the air turbine 5, and the second tail nozzle 9 is connected to the exhaust port of the auxiliary compressor 7.

[0056] Specifically, the shaft work generated by the air turbine 5 drives the auxiliary compressor 7 to pressurize the air after absorbing heat, and finally the pressurized air is ejected from the second tail nozzle 9. While meeting the heat sink requirement, it further ensures that the air ejected from the second tail nozzle 9 has a suitable temperature, pressure and speed, so as to further reduce the adverse effects on the flight of the aircraft.

[0057] In some embodiments, the turbine engine 4 may specifically be a turbojet engine with a relatively small pressure ratio and a relatively high pre-turbine temperature, or a small bypass ratio turbofan engine.

[0058] Reference Figure 2 As shown, in some embodiments, the turbine engine system further includes a heat exchange medium storage container 101 for storing heat exchange medium. The first channel 30 and the second channel 60 are respectively connected to the heat exchange medium storage container 101.

[0059] By providing the heat exchange medium storage container 101, the heat exchange medium can be stored and buffered, further ensuring the continuity of the heat exchange medium flow in the first channel 30 and the second channel 60, thereby further ensuring the regeneration and sustainability of the heat sink and improving the heat exchange efficiency.

[0060] Exemplarily, the heat exchange medium storage container 101 has a first port and a second port, the first port is communicated with the inlet 61 of the second channel 60 , and the second port is communicated with the outlet 32 of the first channel 30 .

[0061] Continue to refer to Figure 2 As shown, in some embodiments, a working fluid pump 102 may be further provided, and the working fluid pump 102 is connected between the first channel 30 and the second channel 60. The working fluid pump 102 is used to drive the heat exchange working fluid to circulate between the first channel 30 and the second channel 60.

[0062] By setting up the working fluid pump 102, the heat exchange working fluid circulates between the pre-cooler 3 and the radiator 6 under the drive of the working fluid pump 102, further improving the smoothness of the flow of the heat exchange working fluid between the first channel 30 and the second channel 60, and thus further ensuring the heat transfer efficiency and the regeneration of the heat sink.

[0063] Among them, the heat exchange working fluid storage container 101 can specifically be protected by covering with inert gas, thus preventing situations such as the reaction of the heat exchange working fluid with oxygen, ensuring the purity, performance and stability of the heat exchange working fluid, and thus further ensuring the heat exchange efficiency.

[0064] Refer to Figure 2 As shown, in some embodiments, the heat exchange working fluid storage container 101 is located outside the main propulsion channel 1 and the auxiliary channel 2.

[0065] In this way, the heat exchange working fluid storage container does not occupy the internal space of the main propulsion channel 1 and the auxiliary channel 2, ensuring the smoothness of the gas flow and the gas flow rate in the main propulsion channel 1 and the auxiliary channel 2, ensuring the sustainability of the heat sink, and further reducing the flight resistance.

[0066] In addition, the working fluid pump 102 can also be arranged outside the main propulsion channel 1 and the auxiliary channel 2.

[0067] In some embodiments, the heat exchange working fluid can be a liquid-cooled metal working fluid.

[0068] The liquid metal working fluid can specifically include gallium-based alloys. Due to the strong heat exchange ability of gallium-based alloys, the overall heat exchange efficiency is improved, and the heat sink regeneration efficiency and sustainability are further improved.

[0069] Exemplarily, in the potassium-based alloy, the mass percentage of potassium Ga is 68.5%, the mass percentage of indium In is 21.5%, and the mass percentage of tin Sn is 10%.

[0070] Exemplarily, the phase change temperature range of the liquid metal working fluid is between -19°C and 2000°C. For example, the flow rate of the heat exchange working fluid can be controlled to be 2 m / s to 5 m / s, so as to further improve the heat exchange ability.

[0071] In other implementation manners, the liquid metal working fluid can also be sodium-potassium alloy, etc. In addition, the heat exchange working fluid can also be other phase-state heat exchange working fluids such as heat exchange gases.

[0072] During specific implementation, the working fluid pump 102 can, for example, select a magnetohydrodynamic pump according to the characteristics of the liquid metal working fluid. Of course, a mechanical pump driven by an electric motor or mechanical power led out by an engine can also be used.

[0073] In some embodiments, the opening size of the main air inlet 11 can be made adjustable. Continue to refer toFigure 1 and Figure 2 As shown in Figure 2 , specifically, a first opening and closing member 12 is provided at the main air inlet 11. The first opening and closing member 12 can rotate relative to the main air inlet 11 to adjust the opening and closing state of the main air inlet 11.

[0074] The adjustment of the opening and closing state of the main air inlet 11 here includes: adjusting the main air inlet 11 to a fully open state, adjusting the main air inlet 11 to a fully closed state, and adjusting the opening size of the main air inlet 11.

[0075] Such a setting enables the rotation of the first opening and closing member 12 to be controlled according to actual usage requirements, and the opening and closing state of the main air inlet 11 can be flexibly adjusted. For example, the main air inlet 11 can be fully opened to ensure the amount of air entering from the main air inlet 11, thereby ensuring the pressure of the air ejected from the first tail nozzle 8, etc., and further ensuring the power of the aircraft.

[0076] In addition, for example, when the aircraft is not in use, the first opening and closing member 12 can be rotated to close the main air inlet 11, thereby preventing dust and other debris from entering the main propulsion channel 1 and affecting the pre-cooler 3, the turbine engine 4, etc., effectively protecting the pre-cooler 3, the turbine engine 4, etc., and further ensuring the normal flight and service life of the aircraft.

[0077] In some embodiments, a second opening and closing member 22 is provided at the auxiliary air inlet 21. The second opening and closing member 22 can rotate relative to the auxiliary air inlet 21 to adjust the opening and closing state of the auxiliary air inlet 21.

[0078] The adjustment of the opening and closing state of the auxiliary air inlet 21 here includes: adjusting the auxiliary air inlet 21 to a fully open state, adjusting the auxiliary air inlet 21 to a fully closed state, and adjusting the opening size of the auxiliary air inlet 21.

[0079] Such a setting enables the rotation of the second opening and closing member 22 to be controlled according to actual usage requirements, and the opening and closing state of the auxiliary air inlet 21 can be flexibly adjusted. For example, the auxiliary air inlet 21 can be fully opened to ensure the amount of air entering from the auxiliary air inlet 21, thereby ensuring the pressure of the air ejected from the second tail nozzle 9, etc.

[0080] In addition, for example, when the aircraft is not in use, the second opening and closing member 22 can be rotated to close the auxiliary air inlet 21, thereby preventing dust and other debris from entering the auxiliary channel 2 and damaging the air turbine 5, the radiator 6, the auxiliary compressor 7, etc., effectively protecting the air turbine 5, the radiator 6, the auxiliary compressor 7, etc., and further ensuring the normal flight and service life of the aircraft.

[0081] For another example, when the aircraft is flying at a low speed, the second opening and closing member 22 can be rotated to close the auxiliary air inlet 21, thereby reducing the flight resistance to a certain extent. When the aircraft is flying at a high speed, the second opening and closing member 22 is rotated to open the auxiliary air inlet 21, providing a heat sink for the precooler 3 and auxiliary thrust for the aircraft.

[0082] In specific implementation, for example, the first opening and closing member 12 and the second opening and closing member 22 can be respectively connected to the flight controller, and the first opening and closing member 12 and the second opening and closing member 22 are controlled to rotate through the flight controller. Alternatively, a controller electrically connected to the first opening and closing member 12 and the second opening and closing member 22 is provided on the engine, and the rotation of the first opening and closing member 12 and the second opening and closing member 22 is controlled through this controller.

[0083] Exemplarily, the first opening and closing member 12 and the second opening and closing member 22 can be, for example, heat-resistant and corrosion-resistant metal plates, etc.

[0084] Among them, the precooler 3 and the radiator 6 can adopt an efficient and compact heat exchanger structure to improve the heat exchange efficiency, such as a plate-fin heat exchanger, so that the precooler 3 can quickly cool the high-temperature air to the temperature range acceptable to the turbine engine 4 under the high Mach number air intake condition. The material of the precooler 3 can be selected as an alloy material with high temperature resistance, corrosion resistance and good thermal conductivity, such as nickel-based superalloy.

[0085] Combined Figures 1 to 3 As shown, in some embodiments, the precooler 3 includes at least three thermally conductive partition plates 33 stacked in sequence, and channels are formed between adjacent two thermally conductive partition plates 33, and thermally conductive fins 34 are arranged in each channel.

[0086] Among them, in two adjacent channels, one of the channels is formed as the first channel 30, the inlet of the other channel is communicated with the main air inlet 11, and the outlet of the other channel is communicated with the air inlet of the turbine engine 4.

[0087] In this way, the air entering from the main air inlet 11 enters the other channel described above, the heat exchange working medium flows in the first channel 30, the heat of the air in the other channel is transferred to the heat exchange working medium through the thermally conductive fins 34 and the thermally conductive partition plates 33, the heat exchange working medium absorbs the heat in the air, and then enters the second channel 60 of the radiator 6 from the outlet 32 of the first channel 30. The heat in the heat exchange working medium is thermally exchanged with the air in the auxiliary channel 2 through the radiator 6, realizing the heating and energy input of the air passing through the radiator 6.

[0088] The heat exchange between the heat exchange working medium and the air is realized through the adjacent channels, the structure is simple and the heat exchange efficiency is improved. Moreover, by arranging the thermally conductive fins 34, the heat exchange efficiency can be further improved.

[0089] Exemplarily, the heat-conducting partition 33 and the heat-conducting fins 34 are both made of nickel-based alloy, for example.

[0090] Combined Figure 1 、 Figure 2 and Figure 4 As shown in FIGS., in some embodiments, the radiator 6 may specifically include a hollow tube body 63 and heat-dissipating fins 64. Among them, the inner cavity of the tube body 63 forms a second channel 60, and the heat-dissipating fins 64 are arranged on the outer wall of the tube body 63 and are in heat-conducting contact with the outer wall of the tube body 63.

[0091] The heat-exchanging working medium that has absorbed heat and flows out from the first channel 30 enters the second channel 60, and the heat carried by the heat-exchanging working medium is transferred to the air passing through the radiator 6 in the auxiliary channel 2 through the tube body 63 and the heat-dissipating fins 64, heating the air and inputting energy.

[0092] By arranging the radiator 6 as above, the heat exchange efficiency between the air and the heat-exchanging working medium is further improved, and further, the regeneration efficiency and sustainability of the heat sink are improved.

[0093] In specific implementation, the tube body 63 and the heat-dissipating fins 64 are both made of nickel-based alloy, for example.

[0094] The embodiments of the present disclosure further provide an aircraft, and the aircraft includes a turbo engine system. Exemplarily, the aircraft may be a drone or a manned aircraft, etc.

[0095] The turbo engine system in this embodiment has the same structure as the turbo engine system provided in the above embodiment and can bring the same or similar technical effects, which will not be elaborated here one by one. Specifically, reference may be made to the description of the above embodiment.

[0096] In specific implementation, the aircraft may further include a fuselage, a flight controller, etc. Among them, the turbo engine system, the flight controller, etc. are arranged on the fuselage.

[0097] For example, the flight controller may control the opening and closing states of the first opening and closing member 12 and the second opening and closing member 22 according to the flight speed.

[0098] The turbo engine system and the aircraft provided by the embodiments of the present disclosure will be further described below through specific examples:

[0099] When the aircraft is in the low-speed flight stage with a Mach number Ma < 2:

[0100] Keep the auxiliary air inlet 21 closed to reduce the flight resistance to a certain extent. At this time, since there is no cold source, the pre-cooler 3 does not cool the air entering from the main air inlet 11.

[0101] Among them, the heat exchange working fluid maintains a basic cycle in the first channel 30 and the second channel 60 (for example, the flow rate of the heat exchange working fluid is 0.1 m / s) to prevent the heat exchange working fluid from solidifying.

[0102] In this low-speed stage, the aircraft climbs and accelerates relying on the thrust of the turbofan engine 4.

[0103] When the flight speed reaches Ma2:

[0104] Both the main air inlet 11 and the auxiliary air inlet 21 are opened, and the precooler 3 in the main propulsion channel 1 starts to work. The outside air is cooled by the precooler 3 (for example, the intake air temperature can be controlled below 100 °C through the adjustment of the working fluid pump 102) and then enters the turbofan engine 4, and finally is ejected from the first tail nozzle 8. The reduced intake air temperature helps to improve the efficiency and performance of the engine, which is beneficial to the long-endurance operation of the engine and expands the flight envelope.

[0105] At this time, a part of the outside air enters the auxiliary channel 2. The air entering the auxiliary channel 2 first passes through the air turbine 5. Due to the expansion work of the air turbine 5, the air temperature drops significantly. The low-temperature air absorbs the heat transferred from the precooler 3 through the heat exchange working fluid in the radiator 6 to complete the heat exchange. For example, the liquid metal working fluid transfers heat to the auxiliary air flow that is expanded and cooled to 0 °C by the air turbine 5 in the radiator 6.

[0106] The shaft work generated by the air turbine 5 drives the auxiliary compressor 7 to pressurize the air after absorbing heat and increasing temperature, so that it has enough energy to be ejected from the second tail nozzle 9 to generate a certain thrust to offset the resistance generated by the projected area of the auxiliary channel 2. That is, due to the heating (energy input) of the radiator 6 in the auxiliary channel 2, it will not cause obvious resistance to flight.

[0107] The above description is only some embodiments of the present disclosure and the description of the applied technical principles. Those skilled in the art should understand that the scope of disclosure involved in the present disclosure is not limited to the technical solutions formed by the specific combination of the above technical features, and should also cover other technical solutions formed by any combination of the above technical features or their equivalent features without departing from the above disclosure concept. For example, the technical solutions formed by mutually replacing the above features with the technical features (but not limited to) having similar functions disclosed in the present disclosure.

[0108] Although some specific embodiments of the present disclosure have been described in detail by way of examples, those skilled in the art should understand that the above examples are only for illustration and not for limiting the scope of the present disclosure. Those skilled in the art should understand that the above embodiments can be modified without departing from the scope and spirit of the present disclosure. The scope of the present disclosure is defined by the appended claims.

Claims

1. A turbine engine system, characterized in that, It includes a main propulsion channel, an auxiliary channel, a turbine engine, a precooler, an air turbine, and a radiator; One end of the main propulsion channel has a main air inlet, and the end of the main propulsion channel away from the main air inlet is connected to a first nozzle; the precooler and the turbine engine are located in the main propulsion channel and are arranged in sequence in the direction from the main air inlet to the first nozzle; One end of the auxiliary channel has an auxiliary air inlet, and the end of the auxiliary channel away from the auxiliary air inlet is connected to a second nozzle; the air turbine and the radiator are located in the auxiliary channel and are arranged in sequence in the direction from the auxiliary air inlet to the second nozzle; The precooler has a first channel for the heat exchange working medium to flow through to cool the air entering from the main air inlet, and the radiator has a second channel for the heat exchange working medium to flow through to heat the air that has been cooled by the expansion of the air turbine. The outlet of the first channel is communicated with the inlet of the second channel, and the outlet of the second channel is communicated with the inlet of the first channel, so that the heat exchange working medium circulates between the first channel and the second channel.

2. The turbine engine system according to claim 1, characterized in that, The turbine engine system further includes a heat exchange working medium storage container for storing the heat exchange working medium; The first channel and the second channel are respectively communicated with the heat exchange working medium storage container.

3. The turbine engine system according to claim 2, wherein, The turbine engine system further includes a working medium pump; The working medium pump is connected between the first channel and the second channel, and the working medium pump is used to drive the heat exchange working medium to circulate between the first channel and the second channel; And / or, the heat exchange working medium storage container is located outside the main propulsion channel and the auxiliary channel.

4. The turbine engine system according to claim 1, wherein, The heat exchange working medium is a liquid-cooled metal working medium.

5. The turbine engine system according to claim 4, wherein, The liquid-cooled metal working medium includes a gallium-based alloy.

6. The turbine engine system according to claim 1, characterized in that, The turbine engine system further includes an auxiliary compressor; The auxiliary compressor is arranged in the auxiliary channel and is located on the side of the radiator away from the air turbine. The auxiliary compressor is connected to the air turbine, and the second nozzle is communicated with the exhaust port of the auxiliary compressor.

7. The turbine engine system according to any one of claims 1 to 6, characterized in that A first opening and closing member is provided at the main air inlet, and the first opening and closing member can rotate relative to the main air inlet to adjust the opening and closing state of the main air inlet; And / or, a second opening and closing member is provided at the auxiliary air inlet, and the second opening and closing member can rotate relative to the auxiliary air inlet to adjust the opening and closing state of the auxiliary air inlet.

8. The turbine engine system according to any one of claims 1 to 6, characterized in that, The radiator includes a hollow tube body and heat dissipation fins; The inner cavity of the tube body forms the second channel, and the heat dissipation fins are arranged on the outer wall of the tube body and are in thermal contact with the outer wall of the tube body; And / or, the precooler includes at least three heat conduction partition plates stacked in sequence, and channels are formed between adjacent two heat conduction partition plates, and heat conduction fins are arranged in each channel; In two adjacent channels, one of the channels forms the first channel, the inlet of the other channel is communicated with the main air inlet, and the outlet of the other channel is communicated with the inlet of the turbine engine.

9. The turbine engine system according to any one of claims 1 to 6, characterized in that, The expansion ratio of the air turbine ranges from 8:1 to 12:

1.

10. An aircraft, characterized in that, Comprising a turbine engine system according to any one of claims 1 to 9.

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