Multi-combustion mode combined hydrogen fuel aviation power system for hypersonic aircraft

Abstract

The application relates to the field of aerospace power design and manufacturing, and discloses a multi-combustion mode combined hydrogen fuel aviation power system for a super-speed aircraft, which is based on a single-rotor turbojet engine and is additionally provided with an outer channel detonation combustion chamber and a scramjet engine after the turbine, is compact and simple in structure, is in series combination of various parts, can minimize the volume and weight of the power system, is more suitable for small and medium-sized aircrafts for super-speed, can meet the wide-speed-domain flight of the aircraft, can effectively utilize the high-speed and high-pressure airflow of the outer channel by adjusting the opening and closing mode of the nozzle with the suction hole to adapt to different working modes of the combined engine, and can meet the technical requirements of low carbon, high energy efficiency, high speed and long-time flight when the turbine engine works independently, the turbine engine and the detonation combustion chamber work together, the scramjet engine works independently and the scramjet and the turbine engine work together.

Description

Technical Field

[0001] This invention relates to the field of aerospace propulsion design and manufacturing, and discloses a multi-combustion mode combined hydrogen fuel aero-propulsion system for hypersonic vehicles. Background Technology

[0002] In existing aerospace propulsion systems, most high-speed aircraft propulsion systems employ scramjet combustion, combined with rockets or turbines, to meet the high-speed flight propulsion requirements. Detonation combustion is a novel pressurized combustion technology; its strong coupling between shock waves and the flame surface allows ignition under a wide range of speed and pressure conditions, providing extremely high thrust and enabling the aircraft to reach higher Mach numbers. However, detonation combustion currently faces two problems: firstly, the difficulty in achieving precise long-term control of the operating mode; and secondly, the issue of thermal protection during long-term use. Therefore, detonation combustion is currently limited to short-term unsteady-state combustion. Furthermore, both detonation and scramjet combustion modes used in high-speed aircraft face two challenges: firstly, how to extract electrical energy during high-speed flight; and secondly, the thermal protection of the combustion chamber during high-speed flight. Finally, for continuous-running, multi-chamber, multi-mode propulsion systems, effectively organizing mass transfer and combustion is of paramount importance.

[0003] Application number CN202311803106 discloses a compressor-driven parallel afterburning detonation aero-engine, including the engine body and a bleed air system. By bleeding air from the intermediate stage of the high-pressure compressor of a gas turbine engine, rotating detonation combustion is organized in the afterburner. By controlling the opening and closing of the bleed air valve, baffle, and adjustable nozzle, engine efficiency can be improved and fuel consumption reduced. However, in this configuration, the detonation combustion chamber only serves the afterburning function and cannot operate independently. Due to the limitations of the turbine operating mode, this configuration cannot be designed for hypersonic vehicles, and the thermal management of the afterburner detonation combustion chamber is difficult to control effectively. Summary of the Invention

[0004] The purpose of this invention is to provide a multi-combustion mode combined hydrogen fuel aviation propulsion system for hypersonic aircraft, which can meet the technical requirements of low-carbon, high-efficiency, high-speed, and long-duration flight when the turbine engine operates independently, the turbine engine and the detonation combustion chamber operate together, the scramjet engine operates independently, and the scramjet and the turbine generator operate together, while satisfying the wide speed range flight requirements of the aircraft.

[0005] To achieve the above-mentioned technical effects, the technical solution adopted by the present invention is as follows:

[0006] A multi-combustion mode combined hydrogen fuel aero-propulsion system for hypersonic vehicles, including:

[0007] A turbojet engine, comprising a compressor, a turbine in front of an annular combustion chamber, and a gas turbine arranged coaxially in the direction of airflow;

[0008] The bypass casing is coaxially disposed on the outer periphery of the turbojet engine formed by the compressor, the turbine in front of the full annular combustion chamber and the gas turbine, and the bypass casing forms an bypass flow channel between the turbojet engine and the outer wall of the turbojet engine;

[0009] A detonation combustion chamber is disposed within the outer bypass channel, and the detonation combustion chamber is located downstream of the axial position corresponding to the gas turbine.

[0010] The tail nozzle is coaxially arranged with the outer bypass casing and located downstream of the outer bypass casing. It is used to exhaust the gas in the turbine in front of the full-annular combustion chamber and the detonation combustion chamber to generate thrust.

[0011] The ramjet combustion chamber is located downstream of the gas turbine tail end, near the tail nozzle, and is used to mix and burn the gas flow passing through the tail end of the bypass duct with the fuel to form gas.

[0012] A mode switching valve, hinged to the compressor inlet, is used to switch the opening and closing of the compressor inlet or the bypass duct inlet. Specifically, when the aircraft is at 0 to Mach 1, the mode switching valve switches the compressor inlet to open and the bypass duct inlet to close, with only the turbojet engine operating. When the aircraft is at Mach 1 to Mach 2, the mode switching valve switches the compressor inlet to open and the bypass duct inlet to open, with only the turbojet engine and the detonation combustor operating. When the aircraft is at Mach 2 to Mach 3, the mode switching valve switches the compressor inlet to close and the bypass duct inlet to open, stopping the turbojet engine and the detonation combustor, while the ramjet combustor operates.

[0013] An air intake passage is located between the compressor exhaust end and the stator casing of the gas turbine. A normally closed intake switching valve is installed at the location of the air intake passage. When the aircraft exceeds the third Mach number, the mode switching valve keeps the compressor inlet closed and the bypass duct inlet open. The intake switching valve opens the air intake passage to introduce part of the airflow from the bypass duct into the turbine inlet annular combustion chamber. The turbine inlet annular combustion chamber and the ramjet combustion chamber are both in operation.

[0014] The nozzle with suction port has an adjustable outlet area and is coaxially disposed at the tail of the gas turbine. When the aircraft is above the third Mach number, the nozzle with suction port is closed to the state with the minimum outlet area; when the aircraft is between 0 and the first Mach number, the nozzle with suction port is opened to the state with the maximum outlet area.

[0015] Furthermore, it also includes a controller, a duct flow path, and a first, second, third, fourth, fifth, sixth, seventh, and eighth duct flow holes interconnected by the duct flow path. The duct flow path is equipped with valves to control the interconnection between the duct flow holes. The first duct flow hole is located on the outer bypass casing and is used to introduce airflow from the external environment into the duct flow path. The second duct flow hole is located within the outer bypass channel and is used to introduce airflow from the outer bypass channel into the duct flow path. The third duct flow hole is located on the compressor casing of the upstream stage of the compressor intermediate stage and is used to introduce airflow from the compressor low-pressure stage into the duct flow path. The fourth... The first air intake port is located on the compressor casing corresponding to the intermediate stage of the compressor, and is used to introduce the airflow of the intermediate stage of the compressor into the air intake path; the second air intake port is located on the compressor casing corresponding to the downstream stage of the intermediate stage of the compressor, and is used to introduce the airflow of the high-pressure stage of the compressor into the air intake path; the third air intake port is located in the turbine inlet annular combustion chamber near the head of the turbine inlet annular combustion chamber, and is used to introduce the airflow of the turbine inlet annular combustion chamber into the air intake path; the fourth air intake port is located on the inner wall of the detonation combustion chamber, and is used to introduce the airflow of the detonation combustion chamber into the air intake path; the fifth air intake port is located near the head of the ramjet combustion chamber, and is used to introduce the airflow of the ramjet combustion chamber into the air intake path.

[0016] Furthermore, when the aircraft is at a speed of 0 to the first Mach number, the controller is used to control all valves in the bleed air path to close, and the bleed air holes are not connected to each other;

[0017] When the aircraft is at the first Mach number to the second Mach number, the controller is used to control the valves in the airflow path, so that the first air vent, the second air vent, the fifth air vent, and the seventh air vent are interconnected to form a first airflow path, and the first air vent, the fourth air vent, the sixth air vent, and the seventh air vent are interconnected to form a second airflow path.

[0018] When the aircraft is at the second to third Mach number, the controller is used to control the valves in the airflow path, so that the first air vent, the third air vent, the sixth air vent, and the eighth air vent are interconnected to form a third airflow path, and the second air vent, the fourth air vent, the fifth air vent, and the eighth air vent are interconnected to form a fourth airflow path.

[0019] Furthermore, when the aircraft reaches a speed greater than the third Mach number, the controller is used to control the valves in the airflow path, so that the first air vent, the third air vent, the fifth air vent, and the eighth air vent are interconnected to form a fifth airflow path, and the first air vent, the fourth air vent, the fifth air vent, and the eighth air vent are interconnected to form a sixth airflow path.

[0020] Furthermore, it also includes a generator to provide electrical energy to start the compressor rotor during the takeoff phase when the aircraft is at the first Mach number, and to ignite the turbine in the in-circular combustion chamber after it reaches the ignition speed, or to generate electricity by driving the gas turbine to rotate when the aircraft is at a speed greater than the third Mach number.

[0021] Furthermore, the fuel is liquid hydrogen, which is stored in a liquid hydrogen tank. The liquid hydrogen tank is connected to a vaporizer, which is used to draw in air to vaporize the liquid hydrogen fuel delivered from the liquid hydrogen tank, and then deliver it to the fuel nozzles of the turbine front annular combustion chamber, the detonation combustion chamber, and the ramjet combustion chamber, respectively.

[0022] Furthermore, the outer casing wall is provided with a conveying pipeline, and the outer casing outer wall is provided with a film cooling hole communicating with the conveying pipeline; the conveying pipeline is connected to the air outlet of the vaporizer, so that the air drawn in by the vaporizer enters the vaporizer to exchange heat with liquid hydrogen, and the cooled air is guided through the conveying pipeline to the film cooling hole for film cooling of the outer casing.

[0023] Furthermore, it also includes an intake cone, wherein the mode switching valve is used to contact and cooperate with the outer wall of the intake cone or the inner wall of the bypass casing to switch the opening and closing of the compressor inlet or the bypass channel inlet.

[0024] Compared with the prior art, the beneficial effects of this invention are as follows: This invention adds an outer bypass duct detonation combustor and a scramjet engine after the turbine to the single-rotor turbojet engine. The structure is compact and simple, and the parts are connected in series, which can minimize the size and weight of the power system and is more suitable for small and medium-sized aircraft with ultra-high speed. Moreover, while meeting the wide speed range flight requirements of the aircraft, by adjusting the opening and closing mode of the nozzle with suction port to adapt to different working modes of the combined engine, the high-speed and high-pressure airflow of the outer bypass duct can be effectively utilized to achieve the technical requirements of low carbon, high energy efficiency, high speed and long flight when the turbine engine is working independently, the turbine engine and the detonation combustor work together, and the scramjet engine is working independently. Attached Figure Description

[0025] Figure 1 This is a schematic diagram of the multi-combustion mode combined hydrogen fuel aero-propulsion system structure for hypersonic vehicles in the embodiment;

[0026] Figure 2 This is a schematic diagram showing the positions of the mode switching valve and the nozzle with suction holes when the turbojet engine and the detonation combustion chamber are in operation, as described in the embodiment.

[0027] Figure 3 This is a schematic diagram showing the positions of the mode switching valve and the nozzle with suction holes when only the ramjet combustion chamber is in operation, as illustrated in the embodiment.

[0028] Figure 4 This is a schematic diagram showing the positions of the mode switching valve and the nozzle with suction holes when the ramjet combustion chamber and the gas turbine are in operation, as illustrated in the embodiment.

[0029] Figure 5 This is a schematic diagram showing the opening positions of each air vent in the embodiment;

[0030] Figure 6 This is a schematic diagram of the airflow path in the embodiment;

[0031] Figure 7 This is a schematic diagram of the structure of the outer bypass casing and film membrane vent in the embodiment;

[0032] The components are as follows: 1. Compressor; 2. Turbine-front annular combustion chamber; 3. Gas turbine; 4. Outer bypass casing; 5. Outer bypass flow channel; 6. Detonation combustion chamber; 7. Tail nozzle; 8. Ram combustion chamber; 9. Mode switching valve; 10. Bleed air passage; 11. Intake switching valve; 12. Nozzle with suction port; 13. Bleed air path; 14. First bleed air port; 15. Second bleed air port; 16. Third bleed air port; 17. Fourth bleed air port; 18. Fifth bleed air port; 19. Sixth bleed air port; 20. Seventh bleed air port; 21. Eighth bleed air port; 22. Starter generator; 23. Liquid hydrogen tank; 24. Carburetor; 25. Delivery pipeline; 26. Film gas port; 27. Intake cone. Detailed Implementation

[0033] The present invention will now be described in further detail with reference to the embodiments and accompanying drawings. However, this should not be construed as limiting the scope of the above-described subject matter of the present invention to the following embodiments; all technologies implemented based on the content of the present invention fall within the scope of the present invention.

[0034] Example

[0035] See Figures 1 to 7 A multi-combustion mode combined hydrogen fuel aero propulsion system for hypersonic vehicles, including:

[0036] A turbojet engine, comprising a compressor 1, a turbine inlet annular combustion chamber 2, and a gas turbine 3 arranged coaxially in the direction of airflow; the turbine inlet annular combustion chamber 2 is used to mix and burn the airflow introduced by the compressor 1 with fuel to form gas, and the gas turbine 3 is used to drive the engine rotor to rotate using the gas generated by the turbine inlet annular combustion chamber 2, and drive the compressor 1 to do work and compress air.

[0037] The outer bypass casing 4 is coaxially disposed on the outer periphery of the turbojet engine formed by the compressor 1, the turbine in front annular combustion chamber 2 and the gas turbine 3, and the outer bypass casing 4 and the outer wall of the turbojet engine form an outer bypass flow channel 5;

[0038] The detonation combustion chamber 6 is disposed within the outer bypass channel 5, and the detonation combustion chamber 6 is located downstream of the axial position corresponding to the gas turbine 3;

[0039] The tail nozzle 7 is coaxially arranged with the outer bypass casing 4 and located downstream of the outer bypass casing 4. It is used to discharge the gas from the turbine in front of the full-annular combustion chamber 2 and the knock combustion chamber 6 to generate thrust.

[0040] The ramjet combustion chamber 8 is located downstream of the tail of the gas turbine 3, near the tail nozzle 7, and is used to mix and burn the airflow flowing through the tail of the bypass duct 5 with the fuel to form gas.

[0041] A mode switching valve 9, hinged to the inlet end of the compressor 1, is used to switch the opening and closing of the compressor 1 inlet or the bypass duct 5 inlet; wherein, for example... Figure 1 As shown, when the aircraft is at speeds from 0 to the first Mach number (e.g., Ma 1.2), the mode switching valve 9 is used to switch the compressor 1 inlet open and the bypass duct 5 inlet closed, with only the turbojet engine in operation; Figure 2 As shown, when the aircraft is at a speed between Mach 1.2 (1.2 Ma) and Mach 2 (e.g., Mach 3), the mode switching valve 9 is used to switch the opening of the compressor 1 inlet and the bypass duct 5 inlet, with only the turbojet engine and the detonation combustion chamber 6 in operation; Figure 3 As shown, when the aircraft is at the second Mach number (3Ma) to the third Mach number (e.g., 5Ma), the mode switching valve 9 is used to switch the compressor 1 inlet closed, the bypass duct 5 inlet open, the turbojet engine and the detonation combustion chamber 6 stop working, and the ramjet combustion chamber 8 is in working state.

[0042] The bleed air passage 10 is located between the exhaust end of the compressor 1 and the stator casing of the gas turbine 3, and a normally closed intake switching valve 11 is installed at the location of the bleed air passage 10; Figure 4 As shown, when the aircraft is above the third Mach number (5Ma), the mode switching valve 9 keeps the compressor 1 inlet closed and the bypass duct 5 inlet open. The intake switching valve 11 opens the bleed air passage 10 to introduce part of the airflow from the bypass duct 5 into the turbine inlet annular combustion chamber 2. The turbine inlet annular combustion chamber 2 and the ramjet combustion chamber 8 are both in working condition.

[0043] The nozzle 12 with suction holes has an adjustable outlet area and is coaxially disposed at the tail of the gas turbine 3. When the aircraft is above the third Mach number (5Ma), the nozzle 12 with suction holes is closed to the state with the smallest outlet area to ensure that the low-pressure gas after the turbine is smoothly discharged under the suction effect of the high-speed airflow outside the nozzle. When the aircraft is between 0 and the first Mach number (1.2Ma), the nozzle 12 with suction holes is opened to the state with the largest outlet area.

[0044] In this embodiment, before takeoff, the mode switching valve 9 switches the compressor 1 inlet to open and the bypass duct 5 inlet to close. During startup, only the turbojet engine operates, accelerating the aircraft to Mach 1 (1.2 Ma). During this process, the nozzle 12 with suction holes opens to its maximum outlet area, meeting the exhaust matching requirements of the turbojet engine. After the aircraft exceeds Mach 1 (1.2 Ma), the mode switching valve 9 switches the compressor 1 inlet to open and the bypass duct 5 inlet to open, with the turbojet engine and detonation combustor 6 operating together, accelerating the aircraft to Mach 2 (3 Ma). After the aircraft exceeds Mach 2 (3 Ma), the mode switching valve 9 switches the compressor 1 inlet to close and the bypass duct 5 inlet to open. At this point, the turbojet engine... The engine and detonation combustion chamber 6 stop working, and the ramjet combustion chamber 8 enters the working state, raising the aircraft to Mach 3 (5Ma) or higher. The nozzle area of ​​the convergent-divergent nozzle mechanism is adjusted according to the gas parameters and flight state to ensure a smooth transition of engine mode switching. When the aircraft exceeds Mach 3 (5Ma), the mode switching valve 9 keeps the compressor 1 inlet closed and the bypass duct 5 inlet open. The intake switching valve 11 opens the bleed air passage 10, introducing part of the airflow from the bypass duct 5 into the turbine inlet annular combustion chamber 2. The turbine inlet annular combustion chamber 2 and the ramjet combustion chamber 8 are both in the working state. During this process, the nozzle 12 with suction holes is closed to the state of minimum outlet area to ensure that the low-pressure gas after the turbine is smoothly discharged under the suction effect of the high-speed airflow outside the nozzle. This embodiment of the multi-combustion mode combined hydrogen fuel aviation propulsion system for high-power electric extraction of hypersonic aircraft not only meets the requirements of wide-speed range flight of the aircraft, but also adapts to different working modes of the combined engine by adjusting the opening and closing mode of the nozzle 12 with suction port. It can effectively utilize the high-speed and high-pressure airflow of the outer bypass duct to achieve the technical requirements of low carbon, high energy efficiency, high speed and long flight when the turbine engine is working independently, the turbine engine and the detonation combustion chamber 6 are working together, and the scramjet engine is working independently.

[0045] like Figure 5The multi-combustion mode combined hydrogen fuel aviation propulsion system for high-power electric extraction of hypersonic aircraft in this embodiment also includes a controller, an airflow path 13, and a first air intake port 14, a second air intake port 15, a third air intake port 16, a fourth air intake port 17, a fifth air intake port 18, a sixth air intake port 19, a seventh air intake port 20, and an eighth air intake port 21 that are interconnected through the airflow path 13. The airflow path 13 is equipped with valves to control the interconnection between the air intake ports. The first air intake port 14 is located on the outer bypass casing 4 and is used to introduce airflow from the external environment into the airflow path 13. The second air intake port 15 is located within the outer bypass channel 5 and is used to introduce airflow from the outer bypass channel 5 into the airflow path 13. The third air intake port 16 is located on the upstream stage of the compressor 1, corresponding to the compressor 1 casing, and is used to introduce compressed air... The airflow from the low-pressure stage of the compressor is introduced into the induced flow path 13; the fourth induced flow port 17 is located on the compressor casing corresponding to the intermediate stage of the compressor 1, and is used to introduce the airflow from the intermediate stage of the compressor into the induced flow path 13; the fifth induced flow port 18 is located on the compressor casing corresponding to the downstream stage of the intermediate stage of the compressor 1, and is used to introduce the airflow from the intermediate stage of the compressor into the induced flow path 13; the sixth induced flow port 19 is located in the turbine inlet annular combustion chamber 2 near the head of the turbine inlet annular combustion chamber 2, and is used to introduce the airflow from the turbine inlet annular combustion chamber 2 into the induced flow path 13; the seventh induced flow port 20 is located on the inner wall of the detonation combustion chamber 6, and is used to introduce the airflow from the detonation combustion chamber 6 into the induced flow path 13; the eighth induced flow port 21 is located near the head of the ram combustion chamber 8, and is used to introduce the airflow from the ram combustion chamber 8 into the induced flow path 13. A multi-channel annular three-dimensional pressure control chamber, controllable by intelligent algorithms, assists in achieving effective combustion organization when the continuous working fluid flows through multiple combustion chambers, wherein:

[0046] When the aircraft is at a speed of 0 to Mach 1 (1.2 Ma), the controller is used to control all valves on the bleed air passage 13 to close, and the bleed air holes are not connected to each other; ensuring that the turbojet engine can operate efficiently and independently under the design conditions.

[0047] When the aircraft is at the first Mach number (1.2 Ma) to the second Mach number (3 Ma), the controller is used to control the valves on the bleed air passage 13, so that the first bleed air port 14, the second bleed air port 15, the fifth bleed air port 18, and the seventh bleed air port 20 are interconnected to form a first airflow passage, and the first bleed air port 14, the fourth bleed air port 17, the sixth bleed air port 19, and the seventh bleed air port 20 are interconnected to form a second airflow passage; by bleeding air from the third-stage compressor 1 of the turbojet engine, and then mixing it with the bypass airflow in terms of pressure and temperature, the working stability of the detonation combustion chamber 6 is improved and the working temperature is reduced, preventing premature combustion of fuel; ultra-precise control of the working fluid flow rate and mixing ratio is achieved, which is not directly affected by upstream pressure fluctuations, greatly reducing the local equivalence ratio fluctuations and mixing inhomogeneity caused by flow pulsation, effectively reducing the entropy growth rate of the combustion system caused by irreversible processes, and making the combustion process closer to the ideal quasi-steady-state process.

[0048] When the aircraft is at the second Mach number (3Ma) to the third Mach number (5Ma), the controller is used to control the valves on the bleed airflow path 13, so that the first bleed air hole 14, the third bleed air hole 16, the sixth bleed air hole 19, and the eighth bleed air hole 21 are interconnected to form a third airflow path, and the second bleed air hole 15, the fourth bleed air hole 17, the fifth bleed air hole 18, and the eighth bleed air hole 21 are interconnected to form a fourth airflow path; the broadband disturbance energy dissipation mechanism that can be actively combined with high efficiency can be used to specifically strengthen the suppression of specific dangerous modes, realize partial decoupling of the core oscillation mode of the combustion chamber from the disturbance source of the upstream system, thereby weakening the energy source that triggers combustion instability.

[0049] When the aircraft speed exceeds the third Mach number (5Ma), the controller controls the valves on the bleed airflow path 13, so that the first bleed air port 14, the third bleed air port 16, the fifth bleed air port 18, and the eighth bleed air port 21 are interconnected to form a fifth airflow path, and the first bleed air port 14, the fourth bleed air port 17, the fifth bleed air port 18, and the eighth bleed air port 21 are interconnected to form a sixth airflow path. This can significantly shorten the response time of the aero-engine system to combustion instability, realize the active organization of the dynamic characteristics of the combustion chamber pressure field, exert influence within the most critical phase window, thereby suppressing the disturbance energy before it accumulates to a destructive level, and greatly improving the robustness of the system to transient disturbances.

[0050] The multi-combustion mode combined hydrogen fuel aviation propulsion system for high-power electricity extraction for hypersonic aircraft in this embodiment also includes a generator 22, which provides electrical energy to start the compressor 1 rotor to rotate during the takeoff phase when the aircraft is at 0 to the first Mach number (1.2 Ma). After rotating to the ignition speed, the turbine in front of the full-annular combustion chamber 2 is ignited. Alternatively, when the aircraft is at a speed greater than the third Mach number (5 Ma), the generator 22 is driven to rotate and generate electricity through the gas turbine 3, thereby realizing high-power electricity extraction during high-speed flight.

[0051] In this embodiment, the fuel is liquid hydrogen, which is stored in a liquid hydrogen tank 23. The liquid hydrogen tank 23 is connected to a vaporizer 24. The vaporizer 24 is used to draw in air to vaporize the liquid hydrogen fuel supplied by the liquid hydrogen tank 23, and then delivers it to the fuel nozzles of the turbine inlet annular combustion chamber 2, the detonation combustion chamber 6, and the ramjet combustion chamber 8, respectively. The power system in this invention uses liquid hydrogen as fuel to achieve zero carbon emissions.

[0052] In this embodiment, the outer bypass casing 4 is provided with a conveying pipe 25 on its wall surface, and a film cooling hole 26 communicating with the conveying pipe 25 is provided on the outer wall surface of the outer bypass casing 4. The conveying pipe 25 is connected to the air outlet of the vaporizer 24, so that the air drawn into the vaporizer 24 enters the vaporizer 24 to exchange heat with liquid hydrogen, and the cooled air is then guided through the conveying pipe 25 to the film cooling hole 26 for film cooling of the outer bypass casing 4. As the main device for hydrogen fuel phase conversion, the vaporizer 24 needs to achieve stable vaporization of liquid hydrogen. In addition, the low-temperature air cooled by heat exchange is also guided through the conveying pipe 25 to the wall surface of the outer bypass casing and discharged through the film cooling hole 26 to form a low-temperature air film, thereby reducing the base temperature of components such as the engine casing and extending its service life.

[0053] In this embodiment, an intake cone 27 is also included. The mode switching valve 9 is used to contact and cooperate with the outer wall of the intake cone 27 or the inner wall of the bypass casing 4 to switch the opening and closing of the compressor 1 inlet or the bypass duct 5 inlet. The hollow structure of the intake cone 27 can be used to integrate related components such as the carburetor 24 and the starter generator 22. The liquid hydrogen carburetor 24 draws air from the engine intake port through the intake cone 27. During the starting phase, air is drawn in by the booster pump, and during the high-speed phase, air is drawn in naturally. The drawn-in air enters the carburetor 24 to exchange heat with the liquid hydrogen. The cooled low-temperature air is transported to the bypass casing through pipelines for film cooling. This can effectively reduce the various complex pipeline networks that draw air from the hot end of the engine, reduce the engine weight, and the integration of the carburetor 24 into the intake cone 27 improves the compactness of the overall structure.

[0054] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A multi-combustion mode combined hydrogen fuel aero-propulsion system for hypersonic vehicles, characterized in that, include: A turbojet engine, comprising a compressor (1), a turbine inlet annular combustion chamber (2), and a gas turbine (3) arranged coaxially in the direction of airflow. The bypass casing (4) is coaxially disposed on the outer periphery of the turbojet engine formed by the compressor (1), the turbine in front annular combustion chamber (2) and the gas turbine (3), and an bypass flow channel (5) is formed between the bypass casing (4) and the outer wall of the turbojet engine. The detonation combustion chamber (6) is disposed in the outer bypass channel (5), and the detonation combustion chamber (6) is located downstream of the axial position corresponding to the gas turbine (3); The tail nozzle (7) is coaxially arranged with the outer bypass casing (4) and located downstream of the outer bypass casing (4). It is used to discharge the gas from the turbine front annular combustion chamber (2) and the detonation combustion chamber (6) to generate thrust. The ramjet combustion chamber (8) is located downstream of the tail of the gas turbine (3) and close to the tail nozzle (7). It is used to mix and burn the gas flow passing through the tail of the bypass channel (5) with the fuel to form gas. A mode switching valve (9) is hinged to the air inlet of the compressor (1) and is used to switch the opening and closing of the compressor (1) inlet or the bypass duct (5) inlet. When the aircraft is at 0 to the first Mach number, the mode switching valve (9) is used to switch the compressor (1) inlet to open and the bypass duct (5) inlet to close, with only the turbojet engine in operation. When the aircraft is at the first Mach number to the second Mach number, the mode switching valve (9) is used to switch the compressor (1) inlet to open and the bypass duct (5) inlet to open, with only the turbojet engine and the detonation combustor (6) in operation. When the aircraft is at the second Mach number to the third Mach number, the mode switching valve (9) is used to switch the compressor (1) inlet to close and the bypass duct (5) inlet to open, with the turbojet engine and the detonation combustor (6) stopping operation and the ramjet combustor (8) in operation. The bleed air passage (10) is located between the exhaust end of the compressor (1) and the stator casing of the gas turbine (3). A normally closed intake switching valve (11) is installed at the location of the bleed air passage (10). When the aircraft is above the third Mach number, the mode switching valve (9) keeps the compressor (1) inlet closed and the bypass duct (5) inlet open. The intake switching valve (11) opens the bleed air passage (10) to introduce part of the airflow from the bypass duct (5) into the turbine front annular combustion chamber (2). The turbine front annular combustion chamber (2) and the ramjet combustion chamber (8) are both in working condition. The nozzle (12) with suction hole has an adjustable outlet area and is coaxially arranged at the tail of the gas turbine (3). When the aircraft is above the third Mach number, the nozzle (12) with suction hole is closed to the state with the smallest outlet area; when the aircraft is between 0 and the first Mach number, the nozzle (12) with suction hole is opened to the state with the largest outlet area.

2. The multi-combustion mode combined hydrogen fuel aero-propulsion system according to claim 1, characterized in that, It also includes a controller, an airflow path (13), and a first air intake hole (14), a second air intake hole (15), a third air intake hole (16), a fourth air intake hole (17), a fifth air intake hole (18), a sixth air intake hole (19), a seventh air intake hole (20), and an eighth air intake hole (21) that are interconnected through the airflow path (13). The airflow path (13) is provided with a valve to control the interconnection between the air intake holes. The first air intake hole (14) is located on the outer bypass casing (4) and is used to introduce the airflow from the external environment into the airflow path (13). The second air intake hole (15) is located in the outer bypass channel (5) and is used to introduce the airflow from the outer bypass channel (5) into the airflow path (13). The third air intake hole (16) is located on the upstream front stage of the compressor (1) corresponding to the compressor (1) casing and is used to introduce the airflow from the low-pressure stage of the compressor into the airflow path (13). The fourth bleed port (17) is located on the compressor casing corresponding to the intermediate stage of the compressor (1) and is used to introduce the airflow of the intermediate stage of the compressor into the bleed airflow path (13); the fifth bleed port (18) is located on the compressor casing corresponding to the downstream stage of the intermediate stage of the compressor (1) and is used to introduce the airflow of the high-pressure stage of the compressor into the bleed airflow path (13); the sixth bleed port (19) is located in the turbine inlet annular combustion chamber (2) near the head of the turbine inlet annular combustion chamber (2) and is used to introduce the airflow of the turbine inlet annular combustion chamber (2) into the bleed airflow path (13); the seventh bleed port (20) is located on the inner wall of the detonation combustion chamber (6) and is used to introduce the airflow of the detonation combustion chamber (6) into the bleed airflow path (13); the eighth bleed port (21) is located near the head of the ramjet combustion chamber (8) and is used to introduce the airflow of the ramjet combustion chamber into the bleed airflow path (13).

3. The multi-combustion mode combined hydrogen fuel aero-propulsion system according to claim 2, characterized in that, When the aircraft is at a speed of 0 to the first Mach number, the controller is used to control all valves on the dredging air path (13) to close, and the dredging holes are not connected to each other; When the aircraft is at the first Mach number to the second Mach number, the controller is used to control the valve on the airflow path (13) so that the first air vent (14), the second air vent (15), the fifth air vent (18), and the seventh air vent (20) are interconnected to form a first airflow path, and the first air vent (14), the fourth air vent (17), the sixth air vent (19), and the seventh air vent (20) are interconnected to form a second airflow path; When the aircraft is at the second to third Mach number, the controller is used to control the valve on the airflow path (13) so that the first air vent (14), the third air vent (16), the sixth air vent (19), and the eighth air vent (21) are interconnected to form a third airflow path, and the second air vent (15), the fourth air vent (17), the fifth air vent (18), and the eighth air vent (21) are interconnected to form a fourth airflow path.

4. The multi-combustion mode combined hydrogen fuel aero-propulsion system according to claim 2, characterized in that, When the aircraft reaches a speed greater than the third Mach number, the controller is used to control the valve on the airflow path (13) so that the first air intake (14), the third air intake (16), the fifth air intake (18), and the eighth air intake (21) are interconnected to form a fifth airflow path, and the first air intake (14), the fourth air intake (17), the fifth air intake (18), and the eighth air intake (21) are interconnected to form a sixth airflow path.

5. The multi-combustion mode combined hydrogen fuel aero-propulsion system according to claim 1, characterized in that, It also includes a generator (22) to provide electrical energy to start the compressor (1) rotor to rotate during the takeoff phase when the aircraft is at 0 to the first Mach number, and to ignite the turbine in front of the full-ring combustion chamber (2) after rotating to the ignition speed, or to generate electricity by rotating the gas turbine (3) when the aircraft is at a speed greater than the third Mach number.

6. The multi-combustion mode combined hydrogen fuel aero-propulsion system according to claim 1, characterized in that, The fuel is liquid hydrogen, which is stored in a liquid hydrogen tank (23). The liquid hydrogen tank (23) is connected to a vaporizer (24). The vaporizer (24) is used to draw in air to vaporize the liquid hydrogen fuel delivered by the liquid hydrogen tank (23) and then deliver it to the fuel nozzles of the turbine front annular combustion chamber (2), the detonation combustion chamber (6), and the ramjet combustion chamber (8).

7. The multi-combustion mode combined hydrogen fuel aero-propulsion system according to claim 6, characterized in that, The outer casing (4) is provided with a conveying pipe (25) on its wall surface, and the outer wall surface of the outer casing (4) is provided with a film cooling hole (26) that communicates with the conveying pipe (25). The conveying pipe (25) is connected to the air outlet of the vaporizer (24) so ​​that the air drawn in by the vaporizer (24) enters the vaporizer (24) and exchanges heat with the liquid hydrogen. After cooling, the air is guided through the conveying pipe (25) to the film cooling hole (26) to perform film cooling on the outer casing (4).

8. The multi-combustion mode combined hydrogen fuel aero-propulsion system according to any one of claims 1-7, characterized in that, It also includes an intake cone (27), and the mode switching valve (9) is used to contact and cooperate with the outer wall of the intake cone (27) or the inner wall of the bypass casing (4) to switch the opening and closing of the compressor (1) inlet or the bypass channel (5) inlet.

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