Multi-combustion-mode combined hydrogen fuel aviation power system for ultra-high-speed aircraft

Abstract

The invention relates to the field of aerospace power design and manufacturing, and discloses a multi-combustion-mode combined hydrogen fuel aviation power system for an ultra-high-speed aircraft, an outer duct detonation combustion chamber and a scramjet engine behind a turbine are added on the basis of a single-rotor turbojet engine, the structure is compact and simple, each part belongs to series combination, and the structure is compact and compact. The size and the weight of a power system can be reduced to the maximum extent, and the system is more suitable for ultra-high-speed small and medium-sized aircrafts; when the wide-speed-range flight of the aircraft is met, the opening and closing modes of the nozzles with the suction holes are adjusted to adapt to different working modes of the combined engine, and high-speed and high-pressure airflow of the outer duct can be effectively utilized; 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 engine and the turbine generator work together are met.

Description

TECHNICAL FIELD

[0001] 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 hypersonic aircraft. BACKGROUND

[0002] In existing aviation power system configurations, a hypersonic aircraft power system mostly adopts a scramjet combustion mode, which is combined with a rocket or a turbine to meet the high-speed flight power demand of the aircraft. Detonation combustion is one of new types of supercharged combustion technologies, and the strong coupling characteristics of a shock wave and a flame surface can realize ignition under a wide range of speeds and pressures, provide super-strong thrust, and further make the aircraft reach a high flight Mach number. However, the detonation combustion mode currently faces two problems: one is that it is difficult to accurately control the working mode for a long time, and the other is the heat protection problem of long-time use, so the detonation combustion is currently limited to short-time unsteady combustion. Secondly, the detonation or scramjet combustion mode used by the hypersonic aircraft needs to face two problems: one is how to extract electric energy under the high-speed flight state, and the other is the heat protection problem of the combustion chamber under the high-speed flight. Finally, for a continuous working medium series connection type multi-combustion chamber multi-mode power system, how to effectively organize mass transfer and combustion is the most important.

[0003] Application No. CN202311803106 discloses a compressor bleed air external parallel afterburner detonation aviation engine, which comprises an engine body and a bleed air device. The intermediate stage bleed air is introduced from the high-pressure compressor of the gas turbine engine to organize rotating detonation combustion in the afterburner, and the engine efficiency can be improved and the fuel consumption can be reduced by controlling the opening and closing of the bleed air valve, the baffle and the adjustable nozzle. However, the detonation combustion chamber in the configuration is only used for afterburning and cannot work alone. Due to the limitation of the turbine working mode, the configuration cannot be used for hypersonic aircraft, and the heat management of the afterburner detonation combustion chamber cannot be effectively controlled. SUMMARY

[0004] The application aims to provide a multi-combustion-mode combined hydrogen-fuel aviation power system for a hypersonic aircraft, which 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 generator work together.

[0005] In order to achieve the above technical effects, the technical scheme adopted by the application is as follows: The multi-combustion-mode combined hydrogen-fuel aviation power system for a hypersonic aircraft comprises: a turbojet engine, the turbojet engine comprising a compressor, a turbine front full-annular combustion chamber and a gas turbine which are coaxially arranged in sequence in the direction of air flow, An outer channel casing is coaxially arranged on the outer periphery of a turbojet engine formed by a compressor, a turbine, a pre-turbine full annular combustion chamber and a gas turbine, and forms an outer channel flow passage between the outer channel casing and the outer wall of the turbojet engine; An explosion combustion chamber is arranged in the outer channel flow passage and is located downstream of the corresponding axial position of the gas turbine; A tail nozzle is coaxially arranged with the outer channel casing and is located downstream of the outer channel casing, and is used to discharge the gas from the pre-turbine full annular combustion chamber and the explosion combustion chamber to generate thrust; A ramjet combustion chamber is located downstream of the tail of the gas turbine and is close to the tail nozzle, and is used to mix and burn the gas flowing through the tail of the outer channel flow passage with fuel to form gas; A mode conversion valve is hinged to the inlet end of the compressor, and is used to switch the opening and closing of the compressor inlet or the outer channel flow passage inlet; when the aircraft is at 0 to the first Mach number, the mode conversion valve is used to switch the opening of the compressor inlet and the closing of the outer channel flow passage inlet, and only the turbojet engine is in working condition; when the aircraft is at the first Mach number to the second Mach number, the mode conversion valve is used to switch the opening of the compressor inlet and the opening of the outer channel flow passage inlet, and only the turbojet engine and the explosion combustion chamber are in working condition; when the aircraft is at the second Mach number to the third Mach number, the mode conversion valve is used to switch the closing of the compressor inlet and the opening of the outer channel flow passage inlet, and the turbojet engine and the explosion combustion chamber stop working, and the ramjet combustion chamber is in working condition; A bleed air passage is arranged between the exhaust end of the compressor and the stator casing of the gas turbine, and a normally closed air inlet conversion valve is arranged at the position of the bleed air passage; when the aircraft is greater than the third Mach number, the mode conversion valve keeps the compressor inlet closed and the outer channel flow passage inlet open, and the air inlet conversion valve opens the bleed air passage to introduce part of the gas flow in the outer channel flow passage into the pre-turbine full annular combustion chamber, and the pre-turbine full annular combustion chamber and the ramjet combustion chamber are in working condition; A nozzle with suction holes is coaxially arranged at the tail of the gas turbine, and the outlet area of the nozzle with suction holes is adjustable; when the aircraft is greater than the third Mach number, the nozzle with suction holes is closed to the state that the outlet area is the smallest; when the aircraft is at 0 to the first Mach number, the nozzle with suction holes is opened to the state that the outlet area is the largest.

[0006] Further, the controller, the bleed air flow path, and the first bleed air hole, the second bleed air hole, the third bleed air hole, the fourth bleed air hole, the fifth bleed air hole, the sixth bleed air hole, the seventh bleed air hole, and the eighth bleed air hole are further included, and the bleed air flow path is provided with valves for controlling the mutual connection and disconnection of the bleed air holes; the first bleed air hole is arranged on the outer casing for introducing the air flow from the outside environment into the bleed air flow path; the second bleed air hole is arranged in the outer channel for introducing the air flow from the outer channel into the bleed air flow path; the third bleed air hole is arranged on the corresponding compressor casing of the front stage upstream of the intermediate stage of the compressor for introducing the air flow from the low-pressure stage of the compressor into the bleed air flow path; the fourth bleed air hole is arranged on the corresponding compressor casing of the intermediate stage of the compressor for introducing the air flow from the intermediate-pressure stage of the compressor into the bleed air flow path; the fifth bleed air hole is arranged on the corresponding compressor casing of the rear stage downstream of the intermediate stage of the compressor for introducing the air flow from the high-pressure stage of the compressor into the bleed air flow path; the sixth bleed air hole is arranged in the pre-turbine full annular combustion chamber close to the head of the pre-turbine full annular combustion chamber for introducing the air flow from the pre-turbine full annular combustion chamber into the bleed air flow path; the seventh bleed air hole is arranged on the inner wall of the detonation combustion chamber for introducing the air flow from the detonation combustion chamber into the bleed air flow path; and the eighth bleed air hole is arranged close to the head of the ramjet combustion chamber for introducing the air flow from the ramjet combustion chamber into the bleed air flow path.

[0007] Further, when the aircraft is at 0 to the first Mach number, the controller is configured to control all the valves on the bleed air flow path to be closed, and the bleed air holes are not connected to each other. When the aircraft is at the first Mach number to the second Mach number, the controller is configured to control the valves on the bleed air flow path, so that the first bleed air hole, the second bleed air hole, the fifth bleed air hole, and the seventh bleed air hole are connected to each other to form a first air flow path, and the first bleed air hole, the fourth bleed air hole, the sixth bleed air hole, and the seventh bleed air hole are connected to each other to form a second air flow path. When the aircraft is at the second Mach number to the third Mach number, the controller is configured to control the valves on the bleed air flow path, so that the first bleed air hole, the third bleed air hole, the sixth bleed air hole, and the eighth bleed air hole are connected to each other to form a third air flow path, and the second bleed air hole, the fourth bleed air hole, the fifth bleed air hole, and the eighth bleed air hole are connected to each other to form a fourth air flow path.

[0008] Further, when the aircraft is greater than the third Mach number, the controller is configured to control the valves on the bleed air flow path, so that the first bleed air hole, the third bleed air hole, the fifth bleed air hole, and the eighth bleed air hole are connected to each other to form a fifth air flow path, and the first bleed air hole, the fourth bleed air hole, the fifth bleed air hole, and the eighth bleed air hole are connected to each other to form a sixth air flow path.

[0009] Further, a starter generator is further included to provide electric energy to start the compressor rotor at the take-off stage when the aircraft is at the first Mach number, and to ignite the pre-turbine full annular combustion chamber after the rotor rotates to the ignition speed, or to drive the rotating generator through the gas turbine when the aircraft is greater than the third Mach number.

[0010] Further, the fuel is liquid hydrogen, the liquid hydrogen is stored in a liquid hydrogen tank, the liquid hydrogen tank is in communication with a vaporizer, the vaporizer is used to vaporize the liquid hydrogen fuel delivered by the liquid hydrogen tank after the air is inhaled, and then the vaporized liquid hydrogen fuel is delivered to the fuel nozzles of the pre-turbine full annular combustion chamber, the detonation combustion chamber and the ramjet combustion chamber, respectively.

[0011] Further, the outer bypass casing wall is provided with a delivery pipeline, and the outer wall of the outer bypass casing is provided with air film holes in communication with the delivery pipeline; the delivery pipeline is in communication with the air outlet of the vaporizer, so that the air inhaled by the vaporizer exchanges heat with the liquid hydrogen, and then the cooled air is guided to the air film holes through the delivery pipeline to cool the outer bypass casing by air film cooling.

[0012] Further, an air inlet cone is further included, and the mode conversion valve is used to contact and cooperate with the outer wall of the air inlet cone or the inner wall of the outer bypass casing to switch the opening and closing of the compressor inlet or the outer bypass channel inlet.

[0013] Compared with the prior art, the present application has the beneficial effects that: the present application adds an outer bypass detonation combustion chamber and a scramjet engine after the turbine on the basis of a single-rotor turbojet engine, the structure is compact and simple, each part is in series combination, the volume and weight of the power system can be reduced to the maximum, and the present application is more suitable for small and medium-sized aircrafts facing super-high speed; while meeting the wide-speed-domain flight of the aircraft, the opening and closing mode of the nozzle with suction holes is adjusted to adapt to different working modes of the combined engine, the high-speed and high-pressure airflow of the outer bypass channel can be effectively utilized, and the technical requirements of low carbon, high energy efficiency, high speed and long time flight can be met when the turbine engine works independently, the turbine engine and the detonation combustion chamber work together, and the scramjet engine works independently. BRIEF DESCRIPTION OF DRAWINGS

[0014] Figure 1 It is a structure schematic diagram of a multi-combustion-mode combined hydrogen-fuel aviation power system in the embodiment facing super-high-speed aircrafts; Figure 2 It is a position schematic diagram of a mode conversion valve and a nozzle with suction holes when a turbojet engine and a detonation combustion chamber are in a working state in the embodiment; Figure 3 It is a position schematic diagram of a mode conversion valve and a nozzle with suction holes when only a ramjet combustion chamber is in a working state in the embodiment; Figure 4Fig. 1 is a schematic view of the position of the mode conversion valve and the nozzle with suction hole in the ramjet combustion chamber and the gas turbine in the working state in the embodiment; Figure 5 Fig. 2 is a schematic view of the opening position of each bleed hole in the embodiment; Figure 6 Fig. 3 is a schematic view of the structure of the bleed air flow path in the embodiment; Figure 7 Fig. 4 is a schematic view of the structure of the outer casing and the film hole in the embodiment; In the figure, 1 is a compressor, 2 is a turbine front full annular combustion chamber, 3 is a gas turbine, 4 is an outer casing, 5 is an outer flow passage, 6 is a detonation combustion chamber, 7 is a tail nozzle, 8 is a ramjet combustion chamber, 9 is a mode conversion valve, 10 is a bleed air passage, 11 is an air inlet conversion valve, 12 is a nozzle with a suction hole, 13 is a bleed air flow path, 14 is a first bleed hole, 15 is a second bleed hole, 16 is a third bleed hole, 17 is a fourth bleed hole, 18 is a fifth bleed hole, 19 is a sixth bleed hole, 20 is a seventh bleed hole, 21 is an eighth bleed hole, 22 is a starter generator, 23 is a liquid hydrogen tank, 24 is a vaporizer, 25 is a delivery pipeline, 26 is a film hole, and 27 is an air inlet cone. DETAILED DESCRIPTION

[0015] The present application will be further described in conjunction with the embodiments and the accompanying drawings. However, it should not be understood that the above-mentioned subject matter of the present application is limited to the following embodiments, and any technology realized based on the content of the present application falls within the scope of the present application.

[0016] Embodiment Reference Figures 1 to 7 , a multi-combustion mode combined hydrogen fuel aviation power system for a hypersonic aircraft, comprising: a turbojet engine, which comprises a compressor 1, a turbine front full annular combustion chamber 2, and a gas turbine 3 arranged coaxially in sequence in the direction of airflow flow; the turbine front full annular combustion chamber 2 is used for mixing and burning the airflow introduced by the compressor 1 with fuel to form gas, and the gas turbine 3 is used for rotating the engine rotor to drive the compressor 1 to compress air by using the gas generated by the turbine front full annular combustion chamber 2; an outer casing 4, which is coaxially arranged on the outer periphery of the turbojet engine formed by the compressor 1, the turbine front full annular combustion chamber 2, and the gas turbine 3, and forms an outer flow passage 5 between the outer casing 4 and the outer wall of the turbojet engine; a detonation combustion chamber 6, which is arranged in the outer flow passage 5 and located downstream of the corresponding axial position of the gas turbine 3; a tail nozzle 7, which is coaxially arranged with the outer casing 4 and located downstream of the outer casing 4, and is used for discharging the gas of the turbine front full 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, 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. 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. 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. 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.

[0017] In the present embodiment, the aircraft control mode conversion valve 9 switches the compressor 1 inlet open and the outer duct 5 inlet closed before the aircraft takes off, and only the turbojet engine works when the aircraft starts, and the aircraft is lifted to the first Mach number (1.2Ma), during which the said suction hole nozzle 12 is opened to the state of the largest outlet area, which can meet the exhaust matching requirements of the turbojet engine; when the aircraft is higher than the first Mach number (1.2Ma), the said mode conversion valve 9 switches the compressor 1 inlet open, and the outer duct 5 inlet open, and the turbojet engine and the detonation combustion chamber 6 are in a common working state, and the aircraft is lifted to the second Mach number (3Ma); when the aircraft is higher than the second Mach number (3Ma), the said mode conversion valve 9 switches the compressor 1 inlet closed, and the outer duct 5 inlet open, at this time the turbojet engine and the detonation combustion chamber 6 stop working, and the ramjet combustion chamber 8 enters the working state, and the aircraft is lifted to the third Mach number (5Ma) or higher, and the converging-diverging nozzle mechanism is adjusted according to the gas parameters and the flight state to make the engine mode conversion transition smoothly; when the aircraft is greater than the third Mach number (5Ma), the said mode conversion valve 9 keeps the compressor 1 inlet closed, and the outer duct 5 inlet open, and the said air inlet conversion valve 11 opens the bleed air passage 10 to introduce part of the airflow in the outer duct 5 into the turbine front full annular combustion chamber 2, and the turbine front full annular combustion chamber 2 and the ramjet combustion chamber 8 are in a common working state, during which the said suction hole nozzle 12 is closed to the state of the smallest outlet area, which ensures that the low-pressure area gas after the turbine is smoothly discharged under the suction effect of the high-speed airflow outside the nozzle. The multi-combustion mode combined hydrogen fuel aviation power system of the present embodiment facing the super-high-speed aircraft high-power electricity extraction meets the wide-speed-domain flight of the aircraft, and adjusts the opening and closing mode of the said suction hole nozzle 12 to adapt to different working modes of the combined engine, which can effectively utilize the high-speed and high-pressure airflow in the outer duct to realize the low-carbon, high-energy efficiency, high-speed and long-time flight technical requirements in the independent working of the turbine engine, the common working of the turbine engine and the detonation combustion chamber 6, and the independent working of the scramjet engine.

[0018] As Figure 5The multi-combustion mode combined hydrogen fuel aviation power system for high-power electricity extraction of the hypersonic aircraft in the embodiment further comprises a controller, an air bleed flow path 13, and a first air bleed hole 14, a second air bleed hole 15, a third air bleed hole 16, a fourth air bleed hole 17, a fifth air bleed hole 18, a sixth air bleed hole 19, a seventh air bleed hole 20, and an eighth air bleed hole 21 which are in communication with each other through the air bleed flow path 13, wherein a valve for controlling the mutual on-off of the air bleed holes is arranged on the air bleed flow path 13; the first air bleed hole 14 is arranged on the outer bypass casing 4 and is used to introduce the airflow of the external environment into the air bleed flow path 13; the second air bleed hole 15 is located in the outer bypass duct 5 and is used to introduce the airflow of the outer bypass duct 5 into the air bleed flow path 13; the third air bleed hole 16 is located on the corresponding compressor 1 casing of the front stage of the compressor 1 upstream of the intermediate stage of the compressor 1 and is used to introduce the airflow of the low-pressure stage of the compressor 1 into the air bleed flow path 13; the fourth air bleed hole 17 is located on the corresponding compressor 1 casing of the intermediate stage of the compressor 1 and is used to introduce the airflow of the intermediate-pressure stage of the compressor 1 into the air bleed flow path 13; the fifth air bleed hole 18 is located on the corresponding compressor 1 casing of the rear stage of the compressor 1 downstream of the intermediate stage of the compressor 1 and is used to introduce the airflow of the intermediate-pressure stage of the compressor 1 into the air bleed flow path 13; the sixth air bleed hole 19 is located in the turbine front full annular combustion chamber 2 close to the head of the turbine front full annular combustion chamber 2 and is used to introduce the airflow of the turbine front full annular combustion chamber 2 into the air bleed flow path 13; the seventh air bleed hole 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 air bleed flow path 13; and the eighth air bleed hole 21 is located close to the head of the ramjet combustion chamber 8 and is used to introduce the airflow of the ramjet combustion chamber 8 into the air bleed flow path 13. The multi-channel annular three-dimensional structure pressure control chamber which can be controlled by an intelligent algorithm assists in realizing effective combustion organization when the continuous working medium flows through the multi-combustion chamber, wherein: When the aircraft is at 0 to the first Mach number (1.2Ma), the controller is used to control all the valves on the air bleed flow path 13 to be closed, and the air bleed holes are not in communication with each other; and the turbojet engine can be efficiently and independently operated under the design condition.

[0019] When the aircraft is at the first Mach number (1.2Ma) to the second Mach number (3Ma), the controller is used to control the valve on the bleed air flow path 13, so that the first bleed hole 14, the second bleed hole 15, the fifth bleed hole 18 and the seventh bleed hole 20 are communicated to form a first air flow path, and the first bleed hole 14, the fourth bleed hole 17, the sixth bleed hole 19 and the seventh bleed hole 20 are communicated to form a second air flow path; through the bleed air from the turbine jet engine three-stage compressor 1, and then mixed with the outer duct air flow in pressure and temperature, the working stability of the detonation combustion chamber 6 is improved, and the working temperature is reduced to prevent the fuel from burning too early; the super-precision control of the working fluid flow rate and the mixing ratio is realized, which is not directly affected by the upstream pressure fluctuation, greatly reduces the local equivalence ratio fluctuation and mixing non-uniformity caused by flow pulsation, effectively reduces the entropy increase rate of the combustion system due to irreversible process, and makes the combustion process closer to the ideal quasi-steady process.

[0020] When the aircraft is at the second Mach number (3Ma) to the third Mach number (5Ma), the controller is used to control the valve on the bleed air flow path 13, so that the first bleed hole 14, the third bleed hole 16, the sixth bleed hole 19 and the eighth bleed hole 21 are communicated to form a third air flow path, and the second bleed hole 15, the fourth bleed hole 17, the fifth bleed hole 18 and the eighth bleed hole 21 are communicated to form a fourth air flow path; the high-efficiency active combined broadband disturbance energy dissipation mechanism can specifically suppress the specific dangerous mode, realize the partial decoupling of the combustion chamber core oscillation mode and the upstream system disturbance source, and thus weaken the energy source of exciting combustion instability.

[0021] When the aircraft is greater than the third Mach number (5Ma), the controller is used to control the valve on the bleed air flow path 13, so that the first bleed hole 14, the third bleed hole 16, the fifth bleed hole 18 and the eighth bleed hole 21 are communicated to form a fifth air flow path, and the first bleed hole 14, the fourth bleed hole 17, the fifth bleed hole 18 and the eighth bleed hole 21 are communicated to form a sixth air flow path; the response time of the aviation power system to combustion instability can be significantly shortened, the active organization of the combustion chamber pressure field dynamic characteristics is realized, the influence is exerted within the most critical phase window, so that the disturbance energy is suppressed before it accumulates to a destructive level, and the robustness of the system to transient disturbance is greatly improved.

[0022] The multi-combustion mode combined hydrogen fuel aviation power system for high-speed aircraft power extraction in the embodiment further comprises a starter generator 22, which is used to provide electric energy to start the rotation of the compressor 1 rotor during the take-off stage when the aircraft is at 0 to the first Mach number (1.2Ma), and ignite the turbine front full-ring combustion chamber 2 after rotating to the ignition speed, or generate electricity by rotating the starter generator 22 through the gas turbine 3 when the aircraft is greater than the third Mach number (5Ma), so as to realize high-power electricity extraction during high-speed flight.

[0023] In the embodiment, the fuel is liquid hydrogen, the liquid hydrogen is stored in a liquid hydrogen tank 23, the liquid hydrogen tank 23 is communicated with a vaporizer 24, the vaporizer 24 is used for vaporizing the liquid hydrogen fuel delivered by the liquid hydrogen tank 23 after the liquid hydrogen fuel is inhaled air, and then delivering the liquid hydrogen fuel to the fuel nozzles of the pre-turbine full annular combustion chamber 2, the knock combustion chamber 6 and the ramjet combustion chamber 8 respectively. The power system in the application uses liquid hydrogen as fuel to achieve the purpose of zero carbon emission.

[0024] In the embodiment, the outer casing 4 wall surface is provided with a delivery pipeline 25, the outer wall surface of the outer casing 4 is provided with a film hole 26 communicated with the delivery pipeline 25; the delivery pipeline 25 is communicated with the air outlet of the vaporizer 24, so that the air inhaled by the vaporizer 24 is introduced into the vaporizer 24 to exchange heat with the liquid hydrogen, and then the cooled air is introduced into the film hole 26 through the delivery pipeline 25 to cool the outer casing 4. The vaporizer 24 is used as the main device for phase conversion of hydrogen fuel, which needs to realize stable vaporization of liquid hydrogen, and the low-temperature air cooled by heat exchange is also introduced into the wall surface of the outer casing through the delivery pipeline 25 and discharged through the film hole 26 to form a low-temperature air film, thereby reducing the base temperature of the engine casing and other components and prolonging the service life.

[0025] In the embodiment, an air inlet cone 27 is further included, and the modal conversion valve 9 is used to contact and cooperate with the outer wall of the air inlet cone 27 or the inner wall of the outer casing 4 to switch the opening and closing of the inlet of the compressor 1 or the inlet of the outer flow passage 5. The hollow structure of the air inlet cone 27 can be used to integrate the vaporizer 24, the starter generator 22 and other related components. The liquid hydrogen vaporizer 24 introduces air from the engine air inlet through the air inlet cone 27, and the air is inhaled by the booster pump in the starting stage and naturally inhaled in the high-speed stage. The inhaled air enters the vaporizer 24 to exchange heat with the liquid hydrogen, and the cooled low-temperature air is delivered to the outer casing through the pipeline for film cooling, which can effectively reduce the complex pipeline network for air introduction from the hot end of the engine, reduce the weight of the engine, and integrate the vaporizer 24 in the air inlet cone 27 to improve the compactness of the overall structure.

[0026] The above is only a preferred embodiment of the application and is not used to limit the application, and any modification, equivalent replacement and improvement made within the spirit and principle of the application shall be included in the protection scope of the application.

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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