A dual-drive rotating detonation engine integrated with a fuel cell

Abstract

The present invention proposes a dual-drive rotating detonation engine integrated with a fuel cell, belonging to the field of aircraft propulsion systems. The present invention utilizes the electrochemical reaction within a high-temperature proton membrane fuel cell system to generate electrical energy. When the clutch is disconnected, the fan is directly connected to the motor to drive the fan, and the compressor is directly connected to the motor to drive the compressor. The exhaust gas or incompletely reacted gas can also drive the turbine to rotate. When the clutch is closed, the fan is driven to rotate and the compressor to operate. The two drive modes complement each other and solve the problem of poor matching between the aircraft's thermodynamic performance and propulsion performance. Structurally, the present invention tightly couples the turbine engine and the motor, increasing the aircraft's energy efficiency and payload.

Description

Technical Field

[0001] The present invention belongs to the field of aircraft propulsion systems and relates to a jet turbofan engine, in particular to a dual-drive rotating detonation engine integrated with a fuel cell. Background Art

[0002] Currently, most aircraft propulsion systems are composed of a single or multiple engines. Excessive engines not only increases the aircraft's weight and reduces its payload, but also complicates the internal layout, resulting in uneven force distribution during flight, poor propulsion, and low efficiency. Furthermore, using only aviation kerosene would release large amounts of pollutants into the air, increasing carbon dioxide emissions and incompatible with the development goals of a "green carbon economy." Therefore, a high-power, high-energy-density, low-pollution propulsion system is urgently needed.

[0003] Traditional fuel cell aircraft engines are suitable for takeoff and low-altitude flight, but they require gas turbine propulsion during the transition from low altitude to high altitude. Furthermore, using a gas turbine engine during high-altitude flight, where the turbine drives the compressor, increases the engine's energy consumption. From the perspective of the combustion chamber, traditional engines generate pressure losses during combustion and lack cyclic combustion, reducing energy efficiency. Using an annular combustion chamber allows for constant-volume combustion within the combustion chamber, rather than the constant-pressure combustion of traditional jet engines, which can cause a sudden change in the engine's thermodynamic efficiency. While traditional engines generate pressure losses during combustion, rotating detonation engines using an annular combustion chamber can achieve supercharged cyclic combustion, improving engine efficiency. From an energy perspective, using high-temperature proton membrane fuel cells not only reduces carbon dioxide and nitrogen oxide emissions, but also uses mechanical energy from the turbine to drive the compressor, which is more difficult to capture and store than using electrical energy generated by the fuel cell. Furthermore, directly using hydrogen fuel in conjunction with the fuel cell allows for a more convenient and rapid electrochemical reaction. Summary of the Invention

[0004] In view of this, in order to overcome the problems of low payload, low energy utilization, low propulsion efficiency and high environmental pollution rate of aircraft exhaust gas after combustion in existing aircraft propulsion systems, and in order to obtain a high-efficiency engine with a supercharged cycle, the present invention proposes a dual-drive rotating detonation engine integrated with a fuel cell, which is mainly used in aviation engine systems.

[0005] To achieve the above-mentioned purpose, the present invention adopts the following technical scheme: a dual-drive rotating detonation engine integrated with a fuel cell, comprising a high-temperature proton membrane fuel cell system, a fuel tank, a fuel pump, an annular combustion chamber, an aircraft propulsion system and an air delivery passage, wherein the aircraft propulsion system comprises a fan, a compressor, a compressor direct-connected motor, a fan direct-connected motor, a mode selection valve, a high-pressure turbine, a low-pressure turbine, a low-pressure shaft, a high-pressure shaft and a tail nozzle, wherein the outer periphery of the high-pressure shaft is sequentially provided with a compressor, a compressor direct-connected motor, an annular combustion chamber and a high-pressure turbine from front to back, a low-pressure shaft passes through the high-pressure shaft, a fan direct-connected motor is installed outside the low-pressure shaft, a fan is installed at the front end of the low-pressure shaft, and a low-pressure turbine is connected to the rear end, the fan direct-connected motor drives the fan to rotate, the compressor direct-connected motor drives the compressor to work, the low-pressure turbine drives the low-pressure shaft to rotate and drives the fan to rotate, and the high-pressure turbine drives the high-pressure shaft to rotate and drives the compressor to work;

[0006] The air delivery passage includes an inner duct and an outer duct. External cold air enters the inner duct and the outer duct inside the engine through the fan, and then is compressed by the compressor and sent to the high-temperature proton membrane fuel cell system through the air cavity channel above the high-temperature proton membrane fuel cell system.

[0007] The fuel in the fuel tank is transported to the fuel pump under the action of the electric motor. The fuel pump pumps the fuel to the high-temperature proton membrane fuel cell system for electrochemical reaction in one path, and transports the fuel to the annular combustion chamber for combustion reaction in another path. The transport ratio between the two can be adjusted.

[0008] The exhaust gas and unreacted gas of the high-temperature proton membrane fuel cell are sent to the annular combustion chamber for combustion. The high-temperature combustion gas generated enters the tail nozzle, where it is accelerated to produce an internal and external pressure difference, propelling the aircraft forward. The low-temperature air enters the nozzle and expands to do work, also propelling the aircraft forward. Finally, the exhaust gases of both are discharged into the atmosphere.

[0009] Furthermore, the compressor, fuel tank, fuel pump, electric motor, fan direct-connected motor, compressor direct-connected motor, high-pressure turbine, low-pressure turbine, low-pressure shaft, low-pressure shaft clutch, high-pressure shaft, high-pressure shaft clutch, mode selection valve and annular combustion chamber are all located in the front half of the engine, and the high-temperature proton membrane fuel cell system, the No. 2 diversion gas channel passing through the turbine and the tail nozzle and other devices are all located in the rear half of the engine.

[0010] Furthermore, the high-temperature proton membrane fuel cell system is composed of multiple cells connected in series. Fuel and air are supplied to each cell, and electrochemical reactions occur in each cell to generate direct current that is gathered and supplied to the electric motor, fan direct-connected motor and compressor direct-connected motor.

[0011] Furthermore, the high-temperature proton membrane fuel cell system includes a cathode inlet for air entry, an anode inlet for hydrogen entry, an electrolyte assembly between the cathode and the anode, and an exhaust gas outlet. The high-temperature proton membrane fuel cell system is a plate structure, and substances react in the cathode-electrolyte-anode assembly. The cathode inlet is located on the left side above the high-temperature proton membrane fuel cell system, the anode inlet is located on the right side above the high-temperature proton membrane fuel cell system, and there is an exhaust gas outlet below the high-temperature proton membrane fuel cell system.

[0012] Furthermore, the annular combustion chamber is a tubular structure, including an inlet for air and exhaust gas or unreacted gas, a gas outlet and a fuel inlet. The outer layer of the annular combustion chamber is a circular tube. The inlet for air and exhaust gas or unreacted gas is located on the left side of the annular combustion chamber, the gas outlet is located on the right side of the annular combustion chamber, and the fuel inlet is located on the upper side of the annular combustion chamber.

[0013] Furthermore, the exhaust gas outlet of the high-temperature proton membrane fuel cell system is collected through a pipeline and sent to the inlet of the annular combustion chamber and the exhaust gas or unreacted gas. The exhaust gas generated by the combustion in the annular combustion chamber flows out from the gas outlet and is sent to the tail nozzle through the diversion gas channel.

[0014] Furthermore, the air delivery path is the inner duct and outer duct through which the external cold air enters the engine through the fan, and then is compressed by the compressor and delivered to the cathode inlet of the high-temperature proton membrane fuel cell system through the air cavity channel above the high-temperature proton membrane fuel cell system; the fuel delivery path is the fuel in the fuel tank is delivered to the fuel pump under the action of the electric motor, and the fuel pump pumps the fuel to the anode inlet of the high-temperature proton membrane fuel cell system for electrochemical reaction, and the other path delivers the fuel to the fuel inlet of the annular combustion chamber for combustion reaction.

[0015] Furthermore, the first working mode of driving the compressor is that the unreacted gas and the exhaust gas produced by combustion are sent to the annular combustion chamber for combustion, and then the mode selection valve is switched to the diverter gas channel 2, so that the gas passes through the high-pressure turbine and the low-pressure turbine to drive the turbine to rotate, and then the low-pressure shaft clutch and the high-pressure shaft clutch are closed, and the low-pressure turbine, high-pressure turbine, compressor and fan rotate coaxially, and the turbine drives the compressor and fan to work.

[0016] Furthermore, the second working mode of driving the compressor is to switch the mode selection valve 9 to the diverter gas channel 1, and the gas generated in the annular combustion chamber does not pass through the high-pressure turbine and the low-pressure turbine, and the low-pressure shaft clutch and the high-pressure shaft clutch are disconnected. The DC power generated by the high-temperature proton membrane fuel cell drives the fan direct-connected motor to drive the fan to work, and drives the compressor direct-connected motor to work to drive the compressor to work.

[0017] Compared with the prior art, the dual-drive rotating detonation engine integrated with a fuel cell according to the present invention has the following beneficial effects:

[0018] (1) The present invention not only retains the turbine of an ordinary engine in terms of structure, but also adds specific electric motors such as an electric motor for pumping fuel, a fan direct-connected motor for driving the fan, and a compressor direct-connected motor for driving the compressor, so that the aircraft can be more freely transformed between high altitude and low altitude, which not only increases the aircraft's payload and propulsion efficiency, but also reduces the pressure loss of the combustion process due to the use of an annular combustion chamber, and the gas can also be burned in a constant volume cycle in the combustion chamber, thereby increasing the energy utilization efficiency and thermodynamic efficiency of the aircraft.

[0019] (2) Compared with traditional gas turbine aircraft engines and fuel cell aircraft engines, the performance of the present invention is a combination of the two. The two working modes complement each other, thereby solving the problem of poor matching between thermodynamic performance and propulsion performance, and increasing the working efficiency and operational flexibility of the aircraft at high altitudes.

[0020] (3) The present invention adopts a mode selection valve and a clutch, which can achieve independent operation of the gas turbine engine and the fuel cell engine, and the two working modes can be switched freely.

[0021] (4) The high-temperature proton membrane fuel cell system used in the present invention has high thermal efficiency and produces low levels of nitrogen oxides and carbon dioxide in the exhaust gas, which can solve the problem of low energy consumption and high pollution during the use of traditional aircraft engines. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] The accompanying drawings, which constitute part of the present invention, are provided to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are provided to explain the present invention and do not constitute an undue limitation of the present invention. In the accompanying drawings:

[0023] Figure 1 This is a schematic structural diagram of a dual-drive rotating detonation engine integrated with a fuel cell according to the present invention;

[0024] Figure 2 This is a schematic diagram of the working principle of the high-temperature proton membrane fuel cell of the present invention;

[0025] Figure 3 This is a schematic diagram of the working principle of the combustion chamber of the present invention;

[0026] In the figure: 1-fan; 2-inner duct; 3-outer duct; 4-compressor; 5-fuel pump; 6-fuel tank; 7-electric motor; 8-annular combustion chamber; 9-mode selection valve; 10-high-temperature proton membrane fuel cell; 11-tail nozzle; 12-low-pressure turbine; 13-high-pressure turbine; 14-high-pressure shaft clutch; 15-compressor direct-connected motor; 16-high-pressure shaft; 17-low-pressure shaft clutch; 18-fan direct-connected motor; 19-low-pressure shaft; 20-cathode inlet; 21-cathode-electrolyte-anode assembly; 22-anode inlet; 23-exhaust gas outlet; 24-inlet for air and exhaust gas or unreacted gas; 25-gas outlet; 26-fuel inlet, 27-nozzle, 28-No. 1 diverter gas channel 1 that does not pass through the turbine, 29-No. 2 diverter gas channel 2 that passes through the turbine. DETAILED DESCRIPTION

[0027] The following will clearly and completely explain the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. It should be noted that the embodiments of the present invention and the features therein can be combined with each other without conflict, and the embodiments described are only part of the embodiments of the present invention, not all of them.

[0028] See also Figure 1-3 This embodiment describes a dual-drive rotating detonation engine integrated with a fuel cell, including a high-temperature proton membrane fuel cell system 10, a fuel tank 6, a fuel pump 5, an annular combustion chamber 8, an aircraft propulsion system, an air delivery passage, and a fuel delivery passage. The aircraft propulsion system includes a fan 1, a compressor 4, a compressor direct-connected motor 15, a fan direct-connected motor 18, a mode selection valve 9, a high-pressure turbine 13, a low-pressure turbine 12, a low-pressure shaft 19, a low-pressure shaft clutch 17, a high-pressure shaft 16, a high-pressure shaft clutch 14, and a tail nozzle 11. The high-pressure shaft 1 6 The outer periphery is sequentially provided with a compressor 4, a compressor direct-connected motor 15, an annular combustion chamber 8 and a high-pressure turbine 13 from front to back. A low-pressure shaft 19 is passed through the high-pressure shaft 16. A fan direct-connected motor 18 is installed on the outside of the low-pressure shaft 19. A fan 1 is installed at the front end of the low-pressure shaft 19 and a low-pressure turbine 12 is connected to the rear end. The fan direct-connected motor 18 drives the fan 1 to rotate, the compressor direct-connected motor 15 drives the compressor 4 to work, the low-pressure turbine 12 drives the low-pressure shaft 19 to rotate and drives the fan 1 to rotate, and the high-pressure turbine 13 drives the high-pressure shaft 16 to rotate and drives the compressor 4 to work;

[0029] The air delivery passage is used to bring air into the fuel cell propulsion system. The intake duct is divided into an inner duct 2 and an outer duct 3, located at the upper and lower ends of the electric motor 7, respectively, to achieve double compression of the air. External cold air enters the inner duct 2 and outer duct 3 inside the engine through the fan 1. After being compressed by the compressor 4, it is delivered to the high-temperature proton membrane fuel cell system 10 through the air cavity channel above the high-temperature proton membrane fuel cell system 10.

[0030] The fuel delivery path is that the fuel in the fuel tank 6 is delivered to the fuel pump 5 under the action of the electric motor 7. The fuel pump 5 delivers the hydrogen fuel directly to the high-temperature proton membrane fuel cell system 10 in one path. The hydrogen fuel and air react electrochemically inside the high-temperature proton membrane fuel cell system 10 to generate direct current to supply the electric motor 7, the fan direct-connected motor 18 and the compressor direct-connected motor 15. The other path delivers the fuel to the annular combustion chamber 8 for combustion reaction, and the delivery ratio between the two can be adjusted.

[0031] The exhaust gas and unreacted gas of the high-temperature proton membrane fuel cell 5 are sent to the annular combustion chamber 8 for combustion. The high-temperature combustion gas generated enters the tail nozzle 11, is accelerated in the tail nozzle 11 to produce an internal and external pressure difference, and propels the aircraft forward. The low-temperature air enters the nozzle 27 and expands to do work, which also propels the aircraft forward. Finally, the exhaust gases of both are discharged into the atmosphere.

[0032] The compressor 4, fuel tank 6, fuel pump 5, electric motor 7, fan direct-connected motor 18, compressor direct-connected motor 15, high-pressure turbine 13, low-pressure turbine 12, low-pressure shaft 19, low-pressure shaft clutch 17, high-pressure shaft 16, high-pressure shaft clutch 14, mode selection valve 9 and annular combustion chamber 8 are all located in the front half of the engine, and the high-temperature proton membrane fuel cell system 10, the No. 2 diversion gas channel 29 passing through the turbine and the tail nozzle 11 and other devices are all located in the rear half of the engine.

[0033] The high-pressure shaft clutch 14 is mounted on the high-pressure shaft 16 and divides the high-pressure shaft 16 into two sections. The high-pressure shaft clutch 14 can combine the two sections into one, rotating at the same speed, or it can allow only one section to rotate while the other is stationary. The high-pressure shaft 16 is divided into two sections and connected by the high-pressure shaft clutch 14.

[0034] The low-pressure shaft clutch 17 is installed on the low-pressure shaft 19 . Its principle and structure are the same as those of the high-pressure shaft clutch 14 . The low-pressure shaft 19 is divided into two sections, which are connected by the low-pressure shaft clutch 17 .

[0035] The fuel delivery passage and the gas channel (the No. 1 branch gas channel 28 passing the turbine and the No. 2 branch gas channel 29 passing the turbine) are both located in the upper channel of the air cavity above the high-temperature proton membrane fuel cell system 10. Heat exchange can be carried out between the two to reduce the temperature of the gas and increase the temperature of the fuel, thereby realizing the recycling of waste heat.

[0036] The high-temperature proton membrane fuel cell system 10 is composed of multiple cells connected in series. Fuel and air are supplied to each cell, and electrochemical reactions occur in each cell to generate direct current, which is then gathered and supplied to the electric motor 7, the fan direct-connected motor 18 and the compressor direct-connected motor 15.

[0037] The high-temperature proton membrane fuel cell system 10 includes a cathode inlet 20 for air to enter, an anode inlet 22 for hydrogen to enter, and an electrolyte assembly 21 between the cathode and the anode. The high-temperature proton membrane fuel cell system 10 is a plate structure, and substances react in the cathode-electrolyte-anode assembly 21. The cathode inlet 20 is located on the left side above the high-temperature proton membrane fuel cell system 10, and the anode inlet 22 is located on the right side above the high-temperature proton membrane fuel cell system 10. There is an exhaust gas outlet 23 at the bottom of the high-temperature proton membrane fuel cell system 10.

[0038] The annular combustion chamber 8 is a tubular structure, including an inlet 24 for air and exhaust gas or unreacted gas, a gas outlet 25 and a fuel inlet 26. The outer layer of the annular combustion chamber 8 is a circular tube. The inlet 24 for air and exhaust gas or unreacted gas is located on the left side of the annular combustion chamber 8, the gas outlet 25 is located on the right side of the annular combustion chamber 8, and the fuel inlet 26 is located on the upper side of the annular combustion chamber 8.

[0039] The exhaust gas outlet 23 of the high-temperature proton membrane fuel cell system 10 is collected through a pipeline and sent to the air and exhaust gas or unreacted gas inlet 24 of the annular combustion chamber 8. The exhaust gas generated by the combustion in the annular combustion chamber 8 flows out from the gas outlet 25 and is sent to the tail nozzle 11 through the diverter gas channel.

[0040] The annular combustion chamber 8 is used to burn the unreacted gas, the exhaust gas generated by the reaction and the air. After combustion, the exhaust gas can be sent to the tail nozzle 11 after heat exchange through the No. 1 diversion gas channel 28 that does not pass through the turbine and the No. 2 diversion gas channel 29 that passes through the turbine, so that the engine generates thrust and propels the aircraft forward.

[0041] The air delivery path is the inner duct 2 and the outer duct 3 through which the external cold air enters the engine through the fan 1, and then is compressed by the compressor 4 and sent to the cathode inlet 20 of the high-temperature proton membrane fuel cell system 10 through the air cavity channel above the high-temperature proton membrane fuel cell system 10; the fuel delivery path is the fuel in the fuel tank 6 is delivered to the fuel pump 5 under the action of the electric motor 7, and the fuel pump 5 pumps the fuel to the anode inlet 22 of the high-temperature proton membrane fuel cell system 10 for electrochemical reaction, and the other path delivers the fuel to the fuel inlet 26 of the annular combustion chamber 8 for combustion reaction, and the delivery ratio between the two is adjustable.

[0042] The first operating mode of the compressor drive is to send the unreacted gas and the exhaust gas produced by combustion to the annular combustion chamber 8 for combustion, and then switch the mode selection valve 9 to the second diversion gas channel 29, allowing the gas to pass through the high-pressure turbine 13 and the low-pressure turbine 12 to drive the turbines to rotate. Then, the low-pressure shaft clutch 17 and the high-pressure shaft clutch 14 are both closed, and the low-pressure turbine 12, the high-pressure turbine 13, the compressor 4 and the fan 1 rotate coaxially, and the turbine drives the compressor 4 and the fan 1 to work. The second operating mode is to switch the mode selection valve 9 to the diversion gas channel 1 28, and the gas produced by the annular combustion chamber 8 does not pass through the high-pressure turbine 13 and the low-pressure turbine 12. The low-pressure shaft clutch 17 and the high-pressure shaft clutch 14 are disconnected, and the DC power generated by the high-temperature proton membrane fuel cell 10 drives the fan direct-connected motor 18 to drive the fan 1 to work, and drives the compressor direct-connected motor 15 to work and drive the compressor 4 to work.

[0043] The working principle of the dual-drive rotating detonation engine integrated with a fuel cell described in the present invention is:

[0044] The present invention is different from traditional turbine engines that use a turbine to drive the compressor; the present invention is also different from using only the electricity from the fuel cell to power the electric motor to drive the compressor. Instead, it combines the two and adopts a dual-drive mode to improve the aircraft's payload and propulsion efficiency.

[0045] And unlike ordinary dual-drive aircraft engines, the present invention adopts an annular combustion chamber 8, which can reduce the pressure loss generated by the combustion process, realize supercharged cycle combustion, and improve the efficiency of the engine. At the same time, if an annular combustion chamber is used, the gas can be burned at a constant volume in the annular combustion chamber 8, and constant volume combustion can greatly improve the thermodynamic efficiency of the engine. During takeoff and low altitude, a fuel cell can be used to generate current to drive the electric motor to drive the compressor. When converting from low altitude to high altitude, a gas turbine is used to drive the compressor 4. When flying stably at high altitude, it can be switched to fuel cell mode. The two modes complement each other and make the engine's conversion between low altitude and high altitude more flexible.

[0046] The specific operation process and working principle of the dual-drive rotating detonation engine integrated with a fuel cell described in the present invention are as follows:

[0047] The external cold air enters the inner duct 2 and the outer duct 3 inside the engine through the fan 1, and then is compressed by the compressor 4 and sent to the cathode inlet 20 of the high-temperature proton membrane fuel cell system 10 through the air cavity channel above the high-temperature proton membrane fuel cell system 10; the fuel in the fuel tank 6 is transported to the fuel pump 5 under the action of the electric motor 7, and the fuel pump 5 pumps the fuel to the anode inlet 22 of the high-temperature proton membrane fuel cell system 10 for electrochemical reaction.

[0048] The direct current generated by the high-temperature proton membrane fuel cell system 10 is supplied to the electric motor 7, the fan direct-connected motor 18, and the compressor direct-connected motor 15. The exhaust gas outlet 23 of the high-temperature proton membrane fuel cell system 10 is collected via a pipeline and supplied to the air and exhaust gas or unreacted gas inlet 24 of the annular combustion chamber 8. After constant-volume cyclic combustion in the annular combustion chamber 8, the mode selector valve 9 can be switched to the diverter gas channel 1 28, disengaging the low-pressure shaft clutch 17 and the high-pressure shaft clutch 14. At this point, the gas generated in the annular combustion chamber 8 does not pass through the high-pressure turbine 13 and the low-pressure turbine 12. The direct current generated by the high-temperature proton membrane fuel cell 10 drives the fan direct-connected motor 18 to operate the fan 1, and drives the compressor direct-connected motor 15 to operate the compressor 4.

[0049] If mode selector valve 9 is switched to diverter gas channel 2 29, both pressure shaft clutch 17 and high-pressure shaft clutch 14 are closed, allowing the gas to pass through high-pressure turbine 13 and low-pressure turbine 12, driving the turbines. Low-pressure turbine 12, high-pressure turbine 13, compressor 4, and fan 1 rotate coaxially. The turbines now drive compressor 4 and fan 1. The gas generated through gas outlet 25 of annular combustion chamber 8 is delivered to tail nozzle 11 through the upper channel of the air cavity above high-temperature proton membrane fuel cell system 10. There, it is accelerated and the high-temperature, high-speed gas is discharged into the atmosphere. Ultimately, the reaction force generated by the airflow realizes the propulsion function of the aircraft.

[0050] The embodiments of the present invention disclosed above are intended only to illustrate the present invention. The embodiments do not describe all details in detail, nor do they limit the present invention to the specific embodiments described. Numerous modifications and variations are possible based on the contents of this specification. These embodiments are selected and described in detail in this specification to better explain the principles and practical applications of the present invention, thereby enabling those skilled in the art to better understand and utilize the present invention.

Claims

1. A dual-drive rotating detonation engine integrated with a fuel cell, characterized by: The invention comprises a high-temperature proton membrane fuel cell system (10), a fuel tank (6), a fuel pump (5), an annular combustion chamber (8), an aircraft propulsion system and an air delivery passage, wherein the aircraft propulsion system comprises a fan (1), a compressor (4), a compressor direct-connected motor (15), a fan direct-connected motor (18), a mode selection valve (9), a high-pressure turbine (13), a low-pressure turbine (12), a low-pressure shaft (19), a high-pressure shaft (16) and a tail nozzle (11), wherein the outer periphery of the high-pressure shaft (16) is provided with a compressor (4), a compressor direct-connected motor (15), an annular combustion chamber (8) and an air delivery passage in sequence from front to back. A combustion chamber (8) and a high-pressure turbine (13), a low-pressure shaft (19) passes through the high-pressure shaft (16), a fan direct-connected motor (18) is installed outside the low-pressure shaft (19), a fan (1) is installed at the front end of the low-pressure shaft (19), and a low-pressure turbine (12) is connected to the rear end, the fan direct-connected motor (18) drives the fan (1) to rotate, the compressor direct-connected motor (15) drives the compressor (4) to work, the low-pressure turbine (12) drives the low-pressure shaft (19) to rotate and drives the fan (1) to rotate, and the high-pressure turbine (13) drives the high-pressure shaft (16) to rotate and drives the compressor (4) to work; The air delivery passage comprises an inner duct (2) and an outer duct (3), and external cold air enters the inner duct (2) and the outer duct (3) inside the engine through the fan (1), and then is compressed by the compressor (4) and sent to the high-temperature proton membrane fuel cell system (10) through the air cavity channel above the high-temperature proton membrane fuel cell system (10); The fuel in the fuel tank (6) is transported to the fuel pump (5) under the action of the electric motor (7). The fuel pump (5) pumps the fuel to the high-temperature proton membrane fuel cell system (10) for electrochemical reaction in one path, and transports the fuel to the annular combustion chamber (8) for combustion reaction in another path, and the transport ratio between the two can be adjusted; The tail gas and unreacted gas of the high-temperature proton membrane fuel cell system (10) are sent to the annular combustion chamber (8) for combustion, and the generated high-temperature combustion gas enters the tail nozzle (11), is accelerated in the tail nozzle (11), and generates a pressure difference between the inside and outside, thereby propelling the aircraft forward. The low-temperature air enters the nozzle (27) and expands to do work, thereby also propelling the aircraft forward. Finally, the tail gases of both gases are discharged into the atmosphere. A high-pressure shaft clutch (14) is mounted on the high-pressure shaft (16), wherein the high-pressure shaft clutch (14) divides the high-pressure shaft (16) into two sections, and a low-pressure shaft clutch (17) is mounted on the low-pressure shaft (19); The first working mode of driving the compressor (4) is that the unreacted gas and the exhaust gas generated by the combustion are sent to the annular combustion chamber (8) for combustion, and then the mode selection valve (9) is switched to the second diversion gas channel (29) passing through the turbine, so that the gas passes through the high-pressure turbine (13) and the low-pressure turbine (12), driving the turbine to rotate, and then the low-pressure shaft clutch (17) and the high-pressure shaft clutch (14) are both closed, the low-pressure turbine (12) and the fan (1) rotate coaxially, the high-pressure turbine (13) and the compressor (4) rotate coaxially, and the high-pressure turbine (13) and the low-pressure turbine (12) drive the compressor (4) and the fan (1) to work; The second working mode of driving the compressor (4) is to switch the mode selection valve (9) to the No. 1 diversion gas channel (28) that does not pass through the turbine, so that the gas generated by the annular combustion chamber (8) does not pass through the high-pressure turbine (13) and the low-pressure turbine (12), and the low-pressure shaft clutch (17) and the high-pressure shaft clutch (14) are disconnected. The direct current generated by the high-temperature proton membrane fuel cell system (10) drives the fan direct-connected motor (18) to drive the fan (1) to work, and drives the compressor direct-connected motor (15) to work to drive the compressor (4) to work.

2. The dual-drive rotating detonation engine integrated with a fuel cell according to claim 1, characterized in that: The compressor (4), fuel tank (6), fuel pump (5), electric motor (7), fan direct-connected motor (18), compressor direct-connected motor (15), high-pressure turbine (13), low-pressure turbine (12), low-pressure shaft (19), low-pressure shaft clutch (17), high-pressure shaft (16), high-pressure shaft clutch (14), mode selection valve (9) and annular combustion chamber (8) are all located in the front half of the engine, and the high-temperature proton membrane fuel cell system (10), the second diversion gas channel (29) passing through the turbine and the tail nozzle (11) are all located in the rear half of the engine.

3. The dual-drive rotating detonation engine integrated with a fuel cell according to claim 1, characterized in that: The high-temperature proton membrane fuel cell system (10) is composed of a plurality of cell stacks connected in parallel. Fuel and air are supplied to each cell stack, and an electrochemical reaction occurs in each cell stack to generate direct current, which is then collected and supplied to a fan direct-connected motor (18), an electric motor (7), and a compressor direct-connected motor (15).

4. The dual-drive rotating detonation engine integrated with a fuel cell according to claim 1, characterized in that: The high-temperature proton membrane fuel cell system (10) includes a cathode inlet (20) for air to enter, an anode inlet (22) for hydrogen to enter, a cathode-electrolyte-anode assembly (21) between the cathode and the anode, and an exhaust gas outlet (23). The high-temperature proton membrane fuel cell system (10) is a plate-type structure, and substances react in the cathode-electrolyte-anode assembly (21). The cathode inlet (20) is located on the left side above the high-temperature proton membrane fuel cell system (10), the anode inlet (22) is located on the right side above the high-temperature proton membrane fuel cell system (10), and the exhaust gas outlet (23) is provided below the high-temperature proton membrane fuel cell system (10).

5. The dual-drive rotating detonation engine integrated with a fuel cell according to claim 4, characterized in that: The annular combustion chamber (8) is a tubular structure, comprising an inlet (24) for air and exhaust gas or unreacted gas, a gas outlet (25) and a fuel inlet (26). The outer layer of the annular combustion chamber (8) is an annular tube. The inlet (24) for air and exhaust gas or unreacted gas is located on the left side of the annular combustion chamber (8), the gas outlet (25) is located on the right side of the annular combustion chamber (8), and the fuel inlet (26) is located on the upper side of the annular combustion chamber (8).

6. The dual-drive rotating detonation engine integrated with a fuel cell according to claim 5, characterized in that: The tail gas outlet (23) of the high-temperature proton membrane fuel cell system (10) is collected through a pipeline and sent to the air and tail gas or unreacted gas inlet (24) of the annular combustion chamber (8). The tail gas generated by the combustion in the annular combustion chamber (8) flows out from the gas outlet (25) and passes through the No. 1 diversion gas channel (28) that does not pass through the turbine or the No. 2 diversion gas channel (29) that passes through the turbine and is sent to the tail nozzle (11).

7. The dual-drive rotating detonation engine integrated with a fuel cell according to claim 6, characterized in that: The air delivery path is an inner duct (2) and an outer duct (3) through which external cold air enters the engine through a fan (1), is compressed by a compressor (4), and is delivered to the cathode inlet (20) of the high-temperature proton membrane fuel cell system (10) through an air cavity channel above the high-temperature proton membrane fuel cell system (10); the fuel delivery path is a path through which fuel in the fuel tank (6) is delivered to the fuel pump (5) under the action of the electric motor (7), and the fuel pump (5) pumps the fuel to the anode inlet (22) of the high-temperature proton membrane fuel cell system (10) for electrochemical reaction, and another path delivers the fuel to the fuel inlet (26) of the annular combustion chamber (8) for combustion reaction.

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