Turbine engine system and aircraft

Abstract

The invention relates to the technical field of aircrafts, in particular to a turbine engine system and an aircraft, and the turbine engine system comprises a propulsion channel, a capillary heat exchanger and a turbine engine; a first micro-channel nozzle is arranged in the propelling channel, and at least part of the first micro-channel nozzle is arranged on the inner wall of a channel section, located between the air inlet and the turbine engine, of the propelling channel; the turbine engine comprises a gas compressor, the gas compressor is provided with at least two stages of guide vanes, each stage of guide vane comprises a movable vane and a stationary vane, and a second micro-channel nozzle is arranged on the stationary vane; the capillary tube heat exchanger comprises a heat exchange tube for heat exchange water to flow so as to cool air entering from the air inlet; and the first micro-channel nozzle and the second micro-channel nozzle are respectively communicated with the hot water outlet of the heat exchange pipe, so that the heated heat exchange water is ejected from the first micro-channel nozzle and the second micro-channel nozzle to generate latent heat phase change, the heat exchange efficiency is improved, and the high-speed and long-endurance operation of the engine is ensured.

Description

Technical Field

[0001] The present disclosure relates to the technical field of aircraft, and in particular to a turbine engine system and an aircraft. Background Art

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

[0003] Turbine engines are a vital component of aircraft, providing propulsion. To increase the speed of turbine engines, conventional technologies use plate-fin heat exchangers to cool the ram air intake, reducing the inlet air temperature and, to a certain extent, offsetting the temperature increase caused by aerodynamic heating during high-speed flight. However, the heat exchange efficiency of these heat exchangers is low, making it difficult for turbine engines to operate at high speeds. Summary of the Invention

[0004] In order to solve the above technical problems or at least partially solve the above technical problems, the present disclosure provides a turbine engine system and an aircraft to improve heat exchange efficiency.

[0005] In a first aspect, the present disclosure provides a turbine engine system comprising a propulsion passage, a capillary tube heat exchanger, and a turbine engine;

[0006] One end of the propulsion channel has an air inlet, and an end of the propulsion channel away from the air inlet is connected to a tail nozzle; the capillary heat exchanger and the turbine engine are located in the propulsion channel and are arranged in sequence along the direction from the air inlet to the tail nozzle;

[0007] A first microchannel nozzle is provided in the propulsion channel, and at least part of the first microchannel nozzle is provided on the inner wall of a channel section of the propulsion channel located between the air inlet and the turbine engine;

[0008] The turbine engine includes a compressor, the compressor having at least two stages of guide vanes, each stage of the guide vanes including moving blades and stationary blades sequentially arranged along the axial direction of the compressor, and the stationary blades are provided with second microchannel nozzles;

[0009] The capillary heat exchanger includes a capillary heat exchange tube for flowing hot water to cool the air entering through the air inlet; the first microchannel nozzle and the second microchannel nozzle are respectively connected to the hot water outlet of the capillary heat exchange tube, so that the heated hot water is ejected from the first microchannel nozzle and the second microchannel nozzle to undergo latent heat phase change.

[0010] Optionally, the second microchannel nozzle is arranged on a side of the stationary blade facing the capillary tube heat exchanger and / or on a side of the stationary blade facing away from the capillary tube heat exchanger.

[0011] Optionally, in a direction from the outer edge of the stationary blade to the inner end of the stationary blade, the second microchannel nozzle is arranged to extend obliquely in a direction away from the stationary blade;

[0012] The included angle between the second microchannel nozzle and the stationary blade is in the range of 10° to 30°.

[0013] Optionally, the second microchannel nozzle includes a nozzle tube, the inlet of the nozzle tube is connected to the hot water outlet of the capillary heat exchange tube, and a guide groove is provided at the outlet of the nozzle tube, and the guide groove extends in a direction from the inlet of the nozzle tube to the outlet of the nozzle tube;

[0014] The groove width of the guide groove at one end close to the inlet of the nozzle pipe is smaller than the groove width of the guide groove at one end close to the outlet of the nozzle pipe.

[0015] Optionally, the width of the guide groove gradually increases in the direction from the inlet of the nozzle pipe to the outlet of the nozzle pipe;

[0016] And / or, the depth of the guide groove ranges from 15 μm to 25 μm.

[0017] Optionally, an opening and closing member is provided at the air inlet, and the opening and closing member can rotate relative to the air inlet to adjust the opening and closing state of the air inlet;

[0018] The first micro-channel nozzle is provided on the inner wall of the opening and closing member.

[0019] Optionally, the turbine engine system further includes a water pump;

[0020] The capillary heat exchanger, the first microchannel nozzle and the second microchannel nozzle are respectively connected to the water pump, and the water pump is used to pump the heated water in the capillary heat exchanger from the hot water outlet of the capillary heat exchange tube into the first microchannel nozzle and the second microchannel nozzle.

[0021] Optionally, the turbine engine system further comprises a hot water supply container;

[0022] The hot water supply container is communicated with the water pump, and the water pump is further used to pump the hot water in the hot water supply container into the capillary heat exchange tube through the cold water inlet of the capillary heat exchange tube.

[0023] Optionally, the turbine engine system further includes a control unit;

[0024] The control unit is connected to the water pump to adjust the pressure of the water pump.

[0025] In a second aspect, the present disclosure provides an aircraft comprising the turbine engine system described above.

[0026] The turbine engine system and aircraft provided by the present disclosure are characterized by arranging a capillary heat exchanger on the air intake side of the turbine engine, performing sensible heat exchange with the air entering through the air intake through the capillary heat exchanger, and the heat exchange water in the capillary heat exchanger absorbs the heat in the air to cool the air, while the temperature of the heat exchange water is greatly increased. At the same time, a first microchannel nozzle is arranged in the propulsion channel, so that at least part of the first microchannel nozzle is located on the inner wall of the air intake section of the propulsion channel, and a second microchannel nozzle is arranged on the stator blade of the compressor, so that the first microchannel nozzle and the second microchannel nozzle are respectively connected to the capillary heat exchanger. The hot water outlets of the pipes are connected, so that the heated water in the capillary heat exchanger is ejected from the first microchannel nozzle and the second microchannel nozzle. The ejected high-temperature spray film undergoes a latent heat evaporation phase change, that is, the high-temperature water is used to absorb heat during evaporation to reduce the temperature rise and cool the air again. In other words, the dual synergistic precooling of sensible heat precooling + jet latent heat precooling improves the heat exchange efficiency of the air, thereby ensuring the stable operation of the engine under high-speed conditions. At the same time, since the heat sink of the jet precooling comes from the heat exchange of the capillary heat exchanger, the heat sink of the capillary heat exchanger is effectively utilized, saving costs.

[0027] Furthermore, the high-temperature spray film ejected by the first and second microchannel nozzles provides effective thermal protection for the channel and guide vane walls, reducing the thermal impact of the high-temperature airflow on the propulsion channel and guide vane walls, thereby extending the service life of related engine system components. At the same time, because the second microchannel nozzles are installed on the stators of each level of guide vanes, that is, through staged spray cooling, the compressor's operating environment during high-speed flight is improved, the compressor's compression efficiency is increased, and energy loss during the gas compression process is reduced, which allows the compressor's temperature limit to be raised, thereby expanding the engine's operating range and performance, and providing guarantees for the engine's long-range, high-speed flight.

[0028] It is to be understood that both the foregoing general description and the following detailed description are exemplary, and are intended to provide further explanation of the technology as claimed. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] The above and other purposes, features, and advantages of the present disclosure will become more apparent through a more detailed description of the embodiments of the present disclosure in conjunction with the accompanying drawings. The accompanying drawings are intended to provide a further understanding of the embodiments of the present disclosure and constitute a part of the specification. Together with the embodiments of the present disclosure, they are used to explain the present disclosure and are not intended to limit the present disclosure. In the drawings, the same reference numerals generally represent the same components or steps.

[0030] Figure 1 A schematic structural diagram of a turbine engine system provided in one embodiment of the present disclosure;

[0031] Figure 2 A schematic structural diagram of a stationary blade and a second microchannel nozzle of a turbine engine system provided by an embodiment of the present disclosure;

[0032] Figure 3 A schematic structural diagram of a turbine engine system provided in another embodiment of the present disclosure.

[0033] Among them, 1. Propulsion channel; 11. Air inlet; 111. Opening and closing parts; 12. Tail nozzle; 2. Capillary heat exchanger; 21. Capillary heat exchange tube; 3. Turbine engine; 31. Compressor; 311. Moving blades; 312. Stationary blades; 4. First microchannel nozzle; 5. Second microchannel nozzle; 51. Nozzle tube; 52. Guide groove; 6. Water pump; 7. Hot water supply container. DETAILED DESCRIPTION

[0034] In order to make the purpose, technical solutions and advantages of the present disclosure more apparent, the following will describe in detail exemplary embodiments of the present disclosure 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, and it should be understood that the present disclosure is not limited to the exemplary embodiments described herein.

[0035] The term "including" and its variations used in this document are open inclusions, that is, "including but not limited to". The term "based on" means "based at least in part on". The term "one embodiment" means "at least one embodiment"; the term "another embodiment" means "at least one other embodiment"; the term "some embodiments" means "at least some embodiments". The relevant definitions of other terms will be given in the description below. It should be noted that the concepts of "first", "second", etc. mentioned in this disclosure are only used to distinguish different devices, modules or units, and are not used to limit the order or interdependence of the functions performed by these devices, modules or units.

[0036] It should be noted that the modifications of "one" and "multiple" mentioned in the present disclosure are illustrative rather than restrictive, and those skilled in the art should understand that unless otherwise clearly indicated in the context, they should be understood as "one or more".

[0037] Reference Figures 1 to 3 As shown, an embodiment of the present disclosure provides a turbine engine system, which can be specifically applied on an aircraft.

[0038] The turbine engine system provided by the embodiment of the present disclosure includes: a propulsion channel 1, a capillary heat exchanger 2 and a turbine engine 3.

[0039] One end of the propulsion channel 1 has an air inlet 11, and the end of the propulsion channel 1, away from the air inlet 11, is connected to the tail nozzle 12. The capillary tube heat exchanger 2 and the turbine engine 3 are located within the propulsion channel 1 and are arranged sequentially from the air inlet 11 to the tail nozzle 12. It is understood that the capillary tube heat exchanger 2 is located on the air inlet side of the turbine engine 3.

[0040] Capillary heat exchanger 2 includes capillary heat exchange tubes 21 through which hot water flows. Capillary heat exchange tubes 21 are used to cool the air entering through air inlet 11. Specifically, the hot water exchanges sensible heat with the air entering through air inlet 11 through the walls of capillary heat exchange tubes 21, absorbing heat from the air. This heat increases the temperature of the hot water itself, thereby cooling the air.

[0041] For example, the inner diameter of the capillary heat exchange tube 21 can be set between 0.5 mm and 1.2 mm, and the total length of the capillary heat exchange tube 21 can be set between 1.2 m and 1.8 m. The material of the capillary heat exchange tube 21 can be high-temperature alloy GH4169, so as to improve heat exchange efficiency.

[0042] For example, the capillary heat exchange tube 21 may include multiple tube segments arranged sequentially, with the ends of two adjacent tube segments connected by an arc tube, forming a multi-layer capillary tube array structure. This enhances the convective heat exchange system and further achieves sufficient heat exchange between water and high-temperature air. For example, the relationship between the curvature radius R of the arc tube and the inner diameter D of the tube segment can be R = 5D.

[0043] A first microchannel nozzle 4 is disposed within the propulsion channel 1, with at least a portion of the first microchannel nozzle 4 disposed on the inner wall of a channel section of the propulsion channel 1 located between the air inlet 11 and the turbine engine 3. Specifically, the channel section of the propulsion channel 1 located between the air inlet 11 and the turbine engine 3 can be referred to as the air inlet section, i.e., the first microchannel nozzle 4 is disposed on the inner wall surface of the air inlet section. It is understood that the spray direction of the first microchannel nozzle 4 is toward the interior of the propulsion channel 1.

[0044] The turbine engine 3 specifically comprises a compressor 31, a combustion chamber, and a turbine. The compressor 31 compresses air, which is then mixed with fuel and combusted in the combustion chamber, producing high-temperature, high-pressure combustion gas. This high-temperature, high-pressure combustion gas drives the turbine, which is then ejected through the tail nozzle 12 to provide thrust for the aircraft.

[0045] Specifically, the compressor 31 has at least two stages of guide vanes, each of which includes a moving blade 311 and a stationary blade 312 sequentially arranged along the axial direction of the compressor 31. The stationary blade 312 is provided with a second microchannel nozzle 5.

[0046] The first microchannel nozzle 4 and the second microchannel nozzle 5 are respectively connected to the hot water outlet of the capillary heat exchange tube 21, so that the heated water is ejected from the first microchannel nozzle 4 and the second microchannel nozzle 5 to undergo latent heat phase change.

[0047] That is, the high-temperature water after heat exchange with the air in the capillary heat exchange tube 21 enters the first microchannel nozzle 4 and the second microchannel nozzle 5, and is ejected from the first microchannel nozzle 4 and the second microchannel nozzle 5 to form a high-temperature spray film. The high-temperature spray film undergoes a latent heat phase change, that is, the high-temperature spray film absorbs the heat in the air and evaporates, thereby achieving further cooling of the air.

[0048] It should be noted that the nozzle apertures of the first microchannel nozzle 4 and the second microchannel nozzle 5 are less than 100 microns. For example, the nozzle aperture can be set between 50 μm and 100 μm to spray high-temperature water mist, such as a water mist with a Sauter mean diameter (SMD) of no more than 20 μm, thereby forming an ultra-fine spray film and penetrating spray effect, increasing the evaporation rate and thus improving the heat exchange efficiency.

[0049] The turbine engine system provided by the embodiment of the present disclosure is provided with a capillary heat exchanger 2 on the air intake side of the turbine engine 3. The capillary heat exchanger 2 performs sensible heat exchange with the air entering through the air intake 11. The heat exchange water in the capillary heat exchanger 2 absorbs the heat in the air, cools the air, and at the same time, the temperature of the heat exchange water is greatly increased. At the same time, a first microchannel nozzle 4 is provided in the propulsion channel 1, so that at least part of the first microchannel nozzle 4 is located on the inner wall of the air intake section of the propulsion channel 1. A second microchannel nozzle 5 is provided on the stator blade 312 of the compressor 31, so that the first microchannel nozzle 4 and the second microchannel nozzle 5 are respectively connected to the turbine engine 31. The hot water outlet of the capillary heat exchange tube 21 is connected, so that the heated water in the capillary heat exchanger 2 is ejected by the first microchannel nozzle 4 and the second microchannel nozzle 5. The ejected high-temperature spray film undergoes a latent heat evaporation phase change, that is, the high-temperature water evaporation absorbs heat to reduce the temperature rise and cool the air again. In other words, through the dual synergistic precooling effect of sensible heat precooling + jet latent heat precooling, the heat exchange efficiency of the air is improved, which provides a guarantee for the stable operation of the engine under high-speed conditions. At the same time, since the heat sink of the jet precooling comes from the heat exchange of the capillary heat exchanger 2, the heat sink of the capillary heat exchanger 2 is effectively utilized, saving costs.

[0050] Furthermore, the high-temperature spray film ejected by the first microchannel nozzle 4 and the second microchannel nozzle 5 provides effective thermal protection for the channel and guide vane walls, reducing the thermal impact of the high-temperature airflow on the channel and guide vane walls of the propulsion channel 1, thereby extending the service life of the relevant components of the engine system. At the same time, because the second microchannel nozzle 5 is provided on the stator blades 312 of each level of guide vanes, that is, through staged spray cooling, the operating environment of the compressor 31 during high-speed flight is improved, the compression efficiency of the compressor 31 is increased, and energy loss during the gas compression process is reduced, thereby improving the temperature limit of the compressor 31, thereby expanding the engine's operating range and performance, and providing guarantees for the engine's long-range and high-speed flight.

[0051] Tests have shown that air entering the propulsion channel 1 through the air inlet 11, after undergoing heat exchange in the capillary heat exchanger 2, can reach a temperature exceeding 90°C. This water is then ejected through the first and second microchannel nozzles 4 and 5, forming a spray film that significantly increases the water's evaporation rate. The high-temperature water, after exiting the microchannel nozzles, forms an ultrafine spray film that covers the wall surface and creates a penetrating spray effect. This significantly increases the evaporation rate through the thin film evaporation + turbulent perturbation mechanism, further improving heat exchange efficiency and boosting the flight speed of the turbine engine 3. This also provides thermal protection for the walls, enhancing the compression efficiency and temperature limit of the compressor 31, protecting the safety of the engine and aircraft, and ensuring the engine's rotational speed, thereby maintaining thrust at high speeds.

[0052] For example, the first microchannel nozzle 4 and the second microchannel nozzle 5 can be made of silicon nitride ceramic nozzles to ensure the high temperature resistance and structural strength of the nozzles. Of course, other high temperature resistant and corrosion resistant materials can also be used, and the embodiments of the present disclosure are not limited thereto.

[0053] Continue to refer to Figures 1 to 3 As shown, in some embodiments, the second microchannel nozzle 5 is provided on a surface of the stationary blade 312 facing the capillary heat exchanger 2 and / or a surface of the stationary blade 312 facing away from the capillary heat exchanger 2. Figure 1 and Figure 3 As shown, the side of the stationary blade 312 facing the capillary tube heat exchanger 2 can specifically be the left side of the stationary blade 312 shown in the figure, and the side of the stationary blade 312 facing away from the capillary tube heat exchanger 2 can specifically be the right side of the stationary blade 312 shown in the figure.

[0054] In this way, a high-temperature spray film can be formed on the wall surface of the stator blade 312. The evaporation phase change of the high-temperature spray film further absorbs heat in the air, thereby further improving the heat exchange efficiency. In addition, the high-temperature spray film can effectively thermally protect the wall surface of the stator blade 312, further reducing the thermal shock of the high-temperature airflow on the wall surface of the stator blade 312, extending the service life of the guide vane, and further extending the service life of the compressor 31 and the engine, providing a guarantee for the effective operation of the engine.

[0055] In addition, since the spray film ejected from the microchannel nozzle comes from the high-temperature hot water after heat exchange in the capillary heat exchanger 2, the ejected high-temperature spray film quickly exchanges heat with the air, and the evaporation rate is greatly improved, which to a certain extent avoids the presence of droplets from causing damage to the guide vanes of the compressor 31, thereby protecting the safety of the engine and the aircraft.

[0056] Specifically, the second micro-channel nozzle 5 can be provided on both sides of the stationary blade 312, which makes the heat exchange efficiency higher and the thermal protection effect of the stationary blade better. Of course, the second micro-channel nozzle 5 can also be provided on only one side.

[0057] In some embodiments, the second microchannel nozzle 5 extends obliquely away from the vane 312 from the outer edge of the vane 312 to the inner end of the vane 312. The included angle a between the second microchannel nozzle 5 and the wall of the vane 312 in which it is located ranges from 10° to 30°. Examples of the included angle a include 10°, 13°, 15°, 18°, 20°, 25°, and 30°.

[0058] This arrangement further ensures that the high-temperature water sprayed from the second microchannel nozzle 5 can form a spray film on the wall surface of the stationary blade 312, thereby further improving the evaporation efficiency, and further improving the heat exchange efficiency and the protection effect on the wall surface of the stationary blade 312.

[0059] In some embodiments, the second microchannel nozzle 5 includes a nozzle tube 51, the inlet of the nozzle tube 51 is connected to the hot water outlet of the capillary heat exchange tube 21, and a guide groove 52 is provided at the outlet of the nozzle tube 51. The guide groove 52 extends from the inlet of the nozzle tube 51 to the outlet of the nozzle tube 51. The width of the guide groove 52 at the end near the inlet of the nozzle tube 51 is smaller than the width of the guide groove 52 at the end near the outlet of the nozzle tube 51.

[0060] In this way, the high-temperature water is effectively guided through the guide groove 52, so that the high-temperature water sprayed from the second microchannel nozzle 5 forms a liquid film covering the surface of the guide vane as much as possible. The liquid film evaporates to produce a local low-temperature zone, thereby suppressing the separation of the boundary layer, improving the surge margin of the compressor 31, and thus improving the stability and reliability of the system operation.

[0061] Further, continue to refer to Figure 2 As shown, in some embodiments, the width of the guide groove 52 gradually increases from the inlet to the outlet of the nozzle pipe 51. This further improves the diversion effect of the high-temperature hot water, increases the liquid film coverage area, and further improves the surge margin of the compressor 31 and the system operation stability.

[0062] Reference Figure 2 As shown, for example, the guide groove 52 can be a V-shaped groove. Of course, the guide groove can also be a trapezoidal groove, etc.

[0063] For example, the depth of the guide groove 52 can be set between 15 μm and 25 μm, such as 15 μm, 20 μm, 25 μm, etc., and the thickness of the formed liquid film can be less than or equal to 200 μm.

[0064] For example, the first microchannel nozzle 4 may adopt a cavitation atomization method, such as a cavitation number δ>0.5, and the second microchannel nozzle 5 may adopt a shear atomization method, such as a Weber number We>500, to ensure that the particle sizes in different regions are matched.

[0065] In specific implementation, laser equipment such as femtosecond / picosecond lasers and a five-axis linkage precision motion platform can be used to achieve laser drilling, trimming of the inner wall of the propulsion channel 1, etching of the guide groove 52, etc. by adjusting the wavelength, pulse energy, and number of pulses.

[0066] Continue to refer to Figure 1 and Figure 3 As shown, in some embodiments, an opening and closing member 111 is provided at the air inlet 11 , and the opening and closing member 111 can rotate relative to the air inlet 11 to adjust the opening and closing state of the air inlet 11 .

[0067] Adjusting the opening and closing state of the air inlet 11 herein includes: adjusting the air inlet 11 to a fully open state, adjusting the air inlet 11 to a fully closed state, and adjusting the opening size of the air inlet 11 .

[0068] This arrangement allows the rotation of the opening and closing member 111 to be controlled according to actual usage requirements, and the opening and closing state of the air inlet 11 to be flexibly adjusted. For example, the air inlet 11 can be fully opened to ensure the amount of air entering from the air inlet 11, thereby ensuring the pressure of the air ejected from the tail nozzle 12, and thus ensuring the power of the aircraft.

[0069] In addition, for example, when the aircraft is not in use, the opening and closing member 111 can be rotated to close the air inlet 11, thereby preventing dust and other debris from entering the propulsion channel 1 and affecting the capillary heat exchanger 2, the turbine engine 3, etc., thereby effectively protecting the capillary heat exchanger 2, the turbine engine 3, etc., thereby ensuring the normal flight and life of the aircraft.

[0070] In a specific implementation, for example, the opening and closing member 111 can be connected to a flight controller, and the flight controller can be used to control the rotation of the opening and closing member 111. Alternatively, a controller electrically connected to the opening and closing member 111 can be provided on the engine, and the rotation of the opening and closing member 111 can be controlled by the controller.

[0071] Illustratively, the opening and closing member 111 may be a high-temperature-resistant and corrosion-resistant metal plate.

[0072] Reference Figure 3 As shown, in some embodiments, a first micro-channel nozzle 4 is provided on the inner wall of the opening and closing member 111 .

[0073] That is to say, by arranging the first microchannel nozzle 4 on the inner wall of the air intake section of the propulsion channel 1 and the inner wall of the opening and closing member 111, a high-temperature spray film is sprayed into the propulsion channel 1 through the first microchannel nozzle 4, thereby further improving the heat exchange efficiency of the air, thereby further reducing the intake temperature, and providing further guarantee for the high-speed and long-duration flight of the engine.

[0074] Continue to refer to Figure 1 and Figure 3 As shown, in some embodiments, the turbine engine system further includes a water pump 6. The capillary heat exchanger 2, the first microchannel nozzle 4, and the second microchannel nozzle 5 are respectively connected to the water pump 6. The water pump 6 is used to pump the heated water in the capillary heat exchanger 2 from the hot water outlet of the capillary heat exchange tube 21 to the first microchannel nozzle 4 and the second microchannel nozzle 5.

[0075] The heated water is pumped from the hot water outlet of the capillary heat exchange tube 21 into the microchannel nozzle by the water pump 6, thereby ensuring the continuity of high-temperature water in the microchannel nozzle and the controllability of water volume and water speed, thereby being able to flexibly control the amount of water mist sprayed from the microchannel nozzle according to actual needs.

[0076] In some embodiments, the turbine engine system also includes a hot water supply container 7, which is connected to a water pump 6. The water pump 6 is also used to pump the hot water in the hot water supply container 7 into the capillary heat exchange tube 21 from the cold water inlet of the capillary heat exchange tube 21, thereby ensuring the continuity of the hot water supply, so as to ensure the heat exchange efficiency of the capillary heat exchanger 2, and further ensure the continuous supply of high-temperature water to the first microchannel nozzle 4 and the second microchannel nozzle 5, thereby ensuring the heat exchange efficiency.

[0077] In specific implementation, the water pump 6 can adopt a high-pressure water pump, for example, the pressure of the high-pressure water pump is 20MPa~35MPa. The high-pressure water pump drives the hot water through the capillary heat exchanger 2, and uses forced convection sensible heat exchange to raise the water temperature to 85%~95% of the boiling temperature (sensible heat absorption ≥350kJ / kg).

[0078] In some embodiments, the turbine engine system further includes a control unit connected to the water pump 6 to adjust the pressure of the water pump 6 .

[0079] Specifically, the pressure of the water pump 6 can be adjusted according to the flight Mach number and the intake state to avoid the air lock effect caused by excessive evaporation.

[0080] The control unit can be integrated into the flight controller or set up separately.

[0081] The present disclosure also provides an aircraft including a turbine engine system. For example, the aircraft may be a drone or a manned aircraft.

[0082] The turbine engine system in the embodiment of the present disclosure has the same specific structure and implementation principle as the turbine engine system provided in the above embodiment, and can bring the same or similar technical effects. They will not be described one by one here, and please refer to the description of the above embodiment for details.

[0083] The above descriptions are merely some embodiments of the present disclosure and an illustration of the technical principles employed. 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 a specific combination of the above-mentioned technical features, but also encompasses other technical solutions formed by any combination of the above-mentioned technical features or their equivalents without departing from the above-mentioned disclosed concepts. For example, a technical solution formed by replacing the above-mentioned features with (but not limited to) technical features with similar functions disclosed in the present disclosure.

[0084] Although some specific embodiments of the present disclosure have been described in detail by way of examples, those skilled in the art will appreciate that the above examples are for illustrative purposes only and are not intended to limit the scope of the present disclosure. Those skilled in the art will appreciate that modifications may 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. A turbine engine system, characterized in that: including propulsion channels, capillary heat exchangers and turbine engines; One end of the propulsion channel has an air inlet, and an end of the propulsion channel away from the air inlet is connected to a tail nozzle; the capillary heat exchanger and the turbine engine are located in the propulsion channel and are arranged in sequence along the direction from the air inlet to the tail nozzle; A first microchannel nozzle is provided in the propulsion channel, and at least part of the first microchannel nozzle is provided on the inner wall of a channel section of the propulsion channel located between the air inlet and the turbine engine; The turbine engine includes a compressor, the compressor having at least two stages of guide vanes, each stage of the guide vanes including moving blades and stationary blades sequentially arranged along the axial direction of the compressor, and the stationary blades are provided with second microchannel nozzles; The capillary heat exchanger includes a capillary heat exchange tube for flowing hot water to cool the air entering through the air inlet; the first microchannel nozzle and the second microchannel nozzle are respectively connected to the hot water outlet of the capillary heat exchange tube, so that the heated hot water is ejected from the first microchannel nozzle and the second microchannel nozzle to undergo latent heat phase change.

2. The turbine engine system according to claim 1, wherein: The second microchannel nozzle is arranged on a surface of the stationary blade facing the capillary tube heat exchanger and / or on a surface of the stationary blade facing away from the capillary tube heat exchanger.

3. The turbine engine system according to claim 2, wherein: In a direction from the outer edge of the stationary blade to the inner end of the stationary blade, the second microchannel nozzle is arranged to extend obliquely in a direction away from the stationary blade; The included angle between the second microchannel nozzle and the stationary blade is in the range of 10° to 30°.

4. The turbine engine system according to claim 2, wherein: The second microchannel nozzle includes a nozzle tube, the inlet of the nozzle tube is connected to the hot water outlet of the capillary heat exchange tube, and a guide groove is provided at the outlet of the nozzle tube, and the guide groove extends in a direction from the inlet of the nozzle tube to the outlet of the nozzle tube; The groove width of the guide groove at one end close to the inlet of the nozzle pipe is smaller than the groove width of the guide groove at one end close to the outlet of the nozzle pipe.

5. The turbine engine system according to claim 4, wherein: The width of the guide groove gradually increases in the direction from the inlet of the nozzle pipe to the outlet of the nozzle pipe; And / or, the depth of the guide groove ranges from 15 μm to 25 μm.

6. The turbine engine system according to claim 1, wherein: An opening and closing member is provided at the air inlet, and the opening and closing member can rotate relative to the air inlet to adjust the opening and closing state of the air inlet; The first micro-channel nozzle is provided on the inner wall of the opening and closing member.

7. The turbine engine system according to any one of claims 1 to 6, characterized in that: The turbine engine system further includes a water pump; The capillary heat exchanger, the first microchannel nozzle and the second microchannel nozzle are respectively connected to the water pump, and the water pump is used to pump the heated water in the capillary heat exchanger from the hot water outlet of the capillary heat exchange tube into the first microchannel nozzle and the second microchannel nozzle.

8. The turbine engine system according to claim 7, wherein: The turbine engine system further includes a hot water supply container; The hot water supply container is communicated with the water pump, and the water pump is further used to pump the hot water in the hot water supply container into the capillary heat exchange tube through the cold water inlet of the capillary heat exchange tube.

9. The turbine engine system according to claim 7, wherein: The turbine engine system further includes a control unit; The control unit is connected to the water pump to adjust the pressure of the water pump.

10. An aircraft, characterized in that: A turbine engine system comprising the turbine engine system according to any one of claims 1 to 9.

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