Three-duct variable-pressure-ratio multi-mode aero-engine

Abstract

The invention provides a three-duct variable-pressure-ratio multi-mode aero-engine, and belongs to the technical field of aero-engines. Comprising a front fan, a rear fan and an exhaust nozzle, the front fan and the rear fan are communicated with the exhaust nozzle through an inner duct and a second outer duct, the inner duct comprises a low-pressure compressor, a high-pressure compressor, a front combustion chamber, a high-pressure turbine, a rear combustion chamber and a low-pressure turbine, and the front fan, the rear fan, the low-pressure compressor and the low-pressure turbine are connected through low-pressure rotors. The structure that the inner culvert is axially connected with the double combustion chambers in series is adopted, three combustion chamber working modes are achieved through the double combustion chambers, duct mode conversion and the air flow direction of the second outer culvert are controlled through the mode selection valve, and switching between multiple ducts and a single duct and change of the supercharging ratio under different modes are achieved; therefore, enough thrust and good fuel economy are guaranteed, and the flight speed range is widened.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aero-engines, and particularly relates to a three-cascade variable-pressure-ratio multi-modal aero-engine. BACKGROUND

[0002] In the process of continuous development in the field of aviation, aircraft engine technology has always been the core element of promoting the progress of the aviation industry. As an important representative of early aviation power, the aircraft piston engine has played a key role in the long history of aviation development. With the advantages of relatively simple structure and high technical maturity, it still occupies a place in specific flight scenarios and application fields.

[0003] However, with the rapid development of aviation technology and the continuous improvement of people's requirements for flight performance, the piston engine attempts to break through the speed limit and enter the high-speed flight state. The size limitation of the engine cannot meet the demand for efficient intake and compression of a large amount of air at high speed, which directly leads to a sharp reduction in power output of the engine at high speed, thereby causing a problem of thrust reduction, and cannot realize wide-speed-domain flight and higher cruising speed. Therefore, it is urgent to design an innovative structure of an engine to realize wide-speed-domain flight, take into account economy and large thrust, promote the development of aero-engines, and provide a feasible path for the next generation of aviation power systems. SUMMARY

[0004] (1) Technical problem to be solved In view of the shortcomings of the prior art, the purpose of the present application is to provide a three-cascade variable-pressure-ratio multi-modal aero-engine to solve the problem that the existing aero-engine cannot realize wide-speed-domain flight and higher cruising speed when entering a high-speed flight state due to the increase of intake temperature and the decrease of through-flow capacity, which leads to a sharp reduction in power output of the engine at high speed, thereby causing a problem of thrust reduction.

[0005] (2) Technical scheme In order to solve the above technical problems, the present application provides a three-cascade variable-pressure-ratio multi-modal aero-engine, comprising a front fan, a rear fan and a tail nozzle, the front fan and the rear fan are communicated with the tail nozzle through an inner channel and a second outer channel, the inner channel comprises a low-pressure compressor, a high-pressure compressor, a front combustion chamber, a high-pressure turbine, a rear combustion chamber and a low-pressure turbine, the front fan, the rear fan, the low-pressure compressor and the low-pressure turbine are connected by a low-pressure rotor, and the high-pressure compressor and the high-pressure turbine are connected by a high-pressure rotor. The first mode selection valve is located between the rear fan and the low-pressure compressor, and is used to control opening and closing of the inner bypass channel; the second mode selection valve is located between the low-pressure compressor and the high-pressure compressor, and is used to control opening and closing of the first outer bypass channel; and the third mode selection valve is located between the high-pressure turbine and the low-pressure turbine, and is used to control communication and closing of the second outer bypass channel with the rear combustion chamber.

[0006] Preferably, the engine has a front combustion mode, a double combustion mode and a rear combustion mode.

[0007] Preferably, the front combustion mode is: the first mode selection valve opens the inner bypass channel, the second mode selection valve opens the first outer bypass channel, the third mode selection valve opens the second outer bypass channel, the front combustion chamber is in an ignition state, the rear combustion chamber is in a blowing state, after the airflow passes through the front fan and the rear fan, the airflow is divided into two paths A and B, the B path gas is divided into two paths C and D after passing through the low-pressure compressor, the A path gas is mixed with the C path gas and then directly discharged through the second outer bypass channel, and the D path gas passes through the high-pressure compressor, the front combustion chamber, the high-pressure turbine, the rear combustion chamber and the low-pressure turbine, and is discharged from the tail nozzle to complete a thermodynamic cycle.

[0008] Preferably, the front combustion mode works when Ma < 1.2.

[0009] Preferably, the front combustion mode works when the bypass ratio is 4-5.

[0010] Preferably, the double combustion mode is: the first mode selection valve opens the inner bypass channel, the second mode selection valve closes the first outer bypass channel, the third mode selection valve opens the second outer bypass channel, the front combustion chamber and the rear combustion chamber are in an ignition state, after the airflow passes through the front fan and the rear fan, the airflow is divided into two paths A and B, the A path gas is directly discharged through the second outer bypass channel, and the B path gas passes through the low-pressure compressor, the high-pressure compressor, the front combustion chamber, the high-pressure turbine, the rear combustion chamber and the low-pressure turbine, and is discharged from the tail nozzle to complete a reheating cycle.

[0011] Preferably, the double combustion mode works when 1.2 < Ma < 2.2.

[0012] Preferably, the double combustion mode works when the bypass ratio is 0.5-0.8.

[0013] Preferably, the afterburning mode is that the afterburning chamber is in ignition state, the first mode selection valve closes the inner channel, the second mode selection valve closes the first outer channel, the third mode selection valve closes the second outer channel, the control gas enters the afterburning chamber from the second outer channel, after mode conversion, the airflow passes through the front fan, the rear fan and the second outer channel, ignites and burns in the afterburning chamber through the third mode selection valve, and then completes the heat cycle through the low-pressure turbine to do work and is discharged from the tail nozzle.

[0014] Preferably, when the afterburning mode works, 2.2

[0015] Preferably, the engine changes the pressure ratio through switching of the first mode selection valve, the second mode selection valve and the second mode selection valve.

[0016] Preferably, the engine has high pressure ratio in the front combustion mode and the double combustion mode, and has low pressure ratio in the afterburning mode.

[0017] Preferably, the engine adopts high pressure ratio in low-speed flight and adopts low pressure ratio in high-speed flight.

[0018] (3) Beneficial effects Compared with the prior art, the beneficial effects of the present application are that: In the above scheme, the mode selection valve is used to control the channel mode conversion and the second outer channel air flow direction, the change of the pressure ratio in different modes is realized, the inner channel and the axial double combustion chamber structure are used, the three combustion chamber working modes are realized by the double combustion chamber, the mode selection valve is used to realize the switching of the multi-channel and single-channel modes, the large channel ratio flight at low Mach and the turbojet mode flight at high Mach are realized, and the high efficiency and large thrust in the range of 0-3 Mach are realized.

[0019] In the above scheme, the engine pressure ratio is changed to realize wide-speed flight, the mode selection valve is used to control the inner and outer channel conversion and the first outer channel air flow direction, the change of the pressure ratio in different modes is realized, in low Mach, the first mode selection valve, the second mode selection valve and the third mode selection valve are controlled to open the inner channel, the first outer channel and the second outer channel at the same time, and the total increase ratio is the pressure ratio accumulation of the inlet channel, the fan and the compressor, in medium Mach, the first mode selection valve, the second mode selection valve and the third mode selection valve are controlled to open the inner channel and the second outer channel and close the first outer channel, and the total increase ratio is the pressure ratio accumulation of the inlet channel, the fan and the compressor, in high Mach flight, the first mode selection valve closes the inner channel, the second mode selection valve closes the first outer channel, and the third mode selection valve closes the second outer channel, and the total increase ratio is the pressure ratio accumulation of the inlet channel and the fan, so as to ensure sufficient thrust and widen the flight speed range.

[0020] The above scheme realizes three combustion chamber working modes by axially connecting the front combustion chamber and the rear combustion chamber in the inner channel, when Ma<1.2, only the front combustion chamber ignites and burns, when 1.2

[0021] The above scheme realizes three channel, double channel and single channel switching through the control of the three channels by the three mode selection valves, adopts three channels at low Mach number, adjusts the three channel airflow distribution, greatly improves the gas flow capacity and ensures sufficient thrust, adopts double channel at medium Mach number, adjusts the double channel airflow distribution, improves the fuel efficiency and ensures sufficient thrust, at high Mach number, closes the inner channel and the first outer channel to become single channel, controls the gas entering the rear combustion chamber from the second outer channel through the third mode selection valve, reduces the total gas pressure ratio, improves the gas flow capacity, greatly reduces the number of control variables and the complexity of combined control, and reduces the cost and difficulty of engine manufacturing. BRIEF DESCRIPTION OF DRAWINGS

[0022] Figure 1 It is a schematic structural diagram of a three-channel variable pressure ratio multi-mode aero-engine.

[0023] Figure 2 It is a schematic structural diagram of a low Mach number mode of a three-channel variable pressure ratio multi-mode aero-engine.

[0024] Figure 3 It is a schematic structural diagram of a medium Mach number mode of a three-channel variable pressure ratio multi-mode aero-engine.

[0025] Figure 4 It is a schematic structural diagram of a high Mach number mode of a three-channel variable pressure ratio multi-mode aero-engine.

[0026] Figure 5 It is a schematic structural diagram of an engine combustion chamber mode of a three-channel variable pressure ratio multi-mode aero-engine.

[0027] The marks in the drawings are: 1, front fan; 2, rear fan; 3, first mode selection valve; 4, first outer channel; 5, second mode selection valve; 6, second outer channel; 7, third mode selection valve; 8, low pressure rotor; 9, low pressure compressor; 10, high pressure compressor; 11, front combustion chamber; 12, high pressure rotor; 13, high pressure turbine; 14, rear combustion chamber; 15, low pressure turbine; 16, tail nozzle; 17, inner channel.

[0028] As shown in the drawings, in order to clearly realize the structure of the embodiments of the present application, specific structures and devices are marked in the drawings, but this is only for the need of illustration, and is not intended to limit the present application in the specific structures, devices and environments, and those skilled in the art can adjust or modify these devices and environments according to specific needs, and the adjustment or modification still includes in the scope of the appended claims. DETAILED DESCRIPTION

[0029] The embodiments of the present application provide a three-cascade variable pressure ratio multi-modal aero-engine, which realizes high efficiency and large thrust in the range of 0-3 Mach of the aero-engine, comprising a front fan 1, a rear fan 2 and a tail nozzle 16, the front fan 1 and the rear fan 2 are communicated with the tail nozzle 16 through an inner channel 17 and a second outer channel 6, the inner channel 17 comprises a low-pressure compressor 9, a high-pressure compressor 10, a front combustion chamber 11, a high-pressure turbine 13, a rear combustion chamber 14 and a low-pressure turbine 15, the front fan 1, the rear fan 2, the low-pressure compressor 9 and the low-pressure turbine 15 are connected by a low-pressure rotor 8, and the high-pressure compressor 10 and the high-pressure turbine 13 are connected by a high-pressure rotor 12. A first modal selection valve 3, a first outer channel 4 and a third modal selection valve 7 are arranged between the inner channel 17 and the second outer channel 6, the first modal selection valve 3 is located between the rear fan 2 and the low-pressure compressor 9, the first modal selection valve 3 is used for controlling the opening and closing of the inner channel 17, the second modal selection valve 5 is arranged at the first outer channel 4, the second modal selection valve 5 is located between the low-pressure compressor 9 and the high-pressure compressor 10, the second modal selection valve 5 is used for controlling the opening and closing of the first outer channel 4, and the third modal selection valve 7 is located between the high-pressure turbine 13 and the low-pressure turbine 15, the third modal selection valve 7 is used for controlling the communication and closing of the second outer channel 6 and the rear combustion chamber 14.

[0030] As Figures 1-5 shown, in the present embodiment, the engine has a front combustion modal, a double combustion modal and a rear combustion modal.

[0031] As Figure 2As shown, in this embodiment, when the pre-combustion mode works, Ma < 1.2, the pre-combustion mode is: in this working mode, in order to achieve low fuel consumption, only the pre-combustion chamber 11 is ignited and combusted, the first mode selection valve 3 opens the inner channel 17, the second mode selection valve 5 opens the first outer channel 4, the third mode selection valve 7 opens the second outer channel 6, the pre-combustion chamber 11 is in an ignition state, and the afterburner 14 is in a blowing state. After the airflow passes through the front fan 1 and the rear fan 2, it is divided into two paths A and B. The B path gas passes through the low-pressure compressor 9 and is divided into two paths C and D. The A path gas and the C path gas are mixed and then discharged directly through the second outer channel 6. The D path gas is combusted in the pre-combustion chamber 11, passes through the high-pressure turbine 13, is blown in the afterburner 14, and then passes through the low-pressure turbine 15, and is discharged from the tail nozzle 16 to complete the thermodynamic cycle. At this time, the air is pressurized by the inlet channel, the front fan 1, the rear fan 2, the low-pressure compressor 9, and the high-pressure compressor 10 during low-Mach flight. After the airflow is compressed by the fan and the compressor, it enters the pre-combustion chamber 11. After the gas is blown in the afterburner 14, it enters the low-pressure turbine 15 to do work. In this working mode, the three channels greatly improve the flow capacity of the gas, ensuring sufficient thrust and high efficiency.

[0032] As shown, Figure 3 As shown, in this embodiment, when the pre-combustion mode works, Ma < 1.2, the pre-combustion mode is: in this working mode, in order to achieve low fuel consumption, only the pre-combustion chamber 11 is ignited and combusted, the first mode selection valve 3 opens the inner channel 17, the second mode selection valve 5 opens the first outer channel 4, the third mode selection valve 7 opens the second outer channel 6, the pre-combustion chamber 11 is in an ignition state, and the afterburner 14 is in a blowing state. After the airflow passes through the front fan 1 and the rear fan 2, it is divided into two paths A and B. The B path gas passes through the low-pressure compressor 9 and is divided into two paths C and D. The A path gas and the C path gas are mixed and then discharged directly through the second outer channel 6. The D path gas is combusted in the pre-combustion chamber 11, passes through the high-pressure turbine 13, is blown in the afterburner 14, and then passes through the low-pressure turbine 15, and is discharged from the tail nozzle 16 to complete the thermodynamic cycle. At this time, the air is pressurized by the inlet channel, the front fan 1, the rear fan 2, the low-pressure compressor 9, and the high-pressure compressor 10 during low-Mach flight. After the airflow is compressed by the fan and the compressor, it enters the pre-combustion chamber 11. After the gas is blown in the afterburner 14, it enters the low-pressure turbine 15 to do work. In this working mode, the three channels greatly improve the flow capacity of the gas, ensuring sufficient thrust and high efficiency.

[0033] As shown, Figure 4As shown, in this embodiment, the afterburning mode operates at 2.2 < Ma < 3.0, and the afterburning mode is as follows: in this operating mode, the afterburning chamber 14 is in the ignition state, the first mode selection valve 3 closes the inner duct 17, the second mode selection valve 5 closes the first outer duct 4, the third mode selection valve 7 closes the second outer duct 6, the control gas enters the afterburning chamber 14 from the second outer duct 6, after mode conversion, the gas flow passes through the front fan 1, the rear fan 2 and the second outer duct 6, flows into the afterburning chamber 14 through the third mode selection valve 7 to ignite and burn, and then completes the heat cycle by doing work through the low-pressure turbine 15 and being discharged from the tail nozzle 16; under the control of the mode selection valve, when flying at high Mach number, air is pressurized by the inlet, the front fan 1 and the rear fan 2, without passing through the compressor, and the gas temperature is slightly lower and the cold gas quality is higher, so that the gas can be fully burned after entering the afterburning chamber 14, the gas flow is directly introduced into the afterburning chamber 14 through the second outer duct 6 after being compressed by the fan, the gas does not pass through the high-pressure turbine 13, but directly enters the low-pressure turbine 15 to do work, and in this operating mode, the total pressure ratio can be prevented from being too high, so as to ensure high compression efficiency and flow capacity, and maintain sufficient temperature rise and thrust.

[0034] Wherein at low Mach number Ma < 1.2, the duct ratio is 4-5, which improves the engine propulsion efficiency and ensures large thrust output at low Mach number; at medium Mach number 1.2 < Ma < 2.2, the duct ratio is 0.5-0.8, which improves the fuel efficiency and ensures thrust output at medium Mach number, and at high Mach number 2.2 < Ma < 3.0, the inner duct 17 and the first outer duct 4 are closed and converted into a single duct, so as to work in the mode of a non-afterburning turbojet, improve the gas flow capacity, ensure sufficient thrust and realize high Mach number cruise; Three-mode thermodynamic cycle mode: at Ma < 1.2, the front combustion chamber 11 mode is adopted, the thermodynamic cycle is a lower pressure ratio Brayton cycle, at 1.2 < Ma < 2.2, the double combustion chamber mode is adopted, the reheating cycle is completed through secondary combustion, and the thermodynamic cycle is a higher pressure ratio reheated Brayton cycle, and at 2.2 < Ma < 3.0, the afterburning chamber 14 mode is adopted, the pressure ratio is changed to improve the efficiency, and the thermodynamic cycle is a lower pressure ratio Brayton cycle.

[0035] As Figure 5As shown in Figure a, the thermodynamic cycle of the pre-combustion mode is as follows: air is compressed through the intake duct, fan, low-pressure compressor 9, and high-pressure compressor 10 before entering the pre-combustion chamber 11. During the entire compression process, the air pressure increases, the temperature rises, and the specific volume decreases (0→2→3). Then, the air-fuel mixture is burned in the pre-combustion chamber 11, heating the airflow and raising its temperature (3→4). After passing through the pre-combustion chamber 11, the gas expands in the low-pressure turbine 15. After passing through the turbine, the air pressure and temperature decrease. Then, it expands further through the tail nozzle 16 and is discharged from the engine at a certain speed, generating a reaction thrust (4→5→9). Finally, the engine draws in fresh air and returns to the initial state (9→0), completing the cycle.

[0036] like Figure 5 As shown in b, this is a thermodynamic cycle with dual combustion modes. Air is compressed through the intake duct, fan, low-pressure compressor 9, and high-pressure compressor 10 before entering the pre-combustion chamber 11. During the entire compression process, the air pressure increases, the temperature rises, and the specific volume decreases (0→2→3). After the air-fuel mixture is burned in the pre-combustion chamber 11, it heats the airflow, and the temperature rises (3→4). After passing through the pre-combustion chamber 11, the gas expands in the high-pressure turbine 13. After passing through the turbine, the air pressure and temperature decrease, and it enters the rear combustion chamber 14 (4→3′). After the air-fuel mixture is burned in the rear combustion chamber 14, it heats the airflow again and does work, and the temperature rises again (3′→4′). Then the gas expands in the low-pressure turbine 15. After passing through the turbine, the air pressure and temperature decrease significantly. Then it expands further through the tail nozzle 16 and is discharged from the engine at a certain speed. It generates a reaction thrust by relying on the exhaust speed (4′→5→9). Finally, the engine draws in fresh air and returns to the initial state (9→0), completing the reheat cycle.

[0037] like Figure 5 As shown in Figure c, this is the thermodynamic cycle of the rear-combustion mode. After the air is compressed by the intake duct and the fan does work, it directly enters the rear combustion chamber 14 through the second bypass duct 6. During the entire compression process, the air pressure increases, the temperature rises, and the specific volume decreases (0→2→3′). Then, the air-fuel mixture burns in the rear combustion chamber 14, heating the airflow and doing work, causing the temperature to rise (3′→4′). After passing through the rear combustion chamber 14, the gas expands in the low-pressure turbine 15. After passing through the low-pressure turbine 15, the air pressure and temperature decrease. Then, it expands further through the tail nozzle 16 and is discharged from the engine at a certain speed, generating a reaction thrust by relying on the exhaust speed (4′→5→9). Finally, the engine draws in fresh air and returns to the initial state (9→0), completing the cycle.

[0038] The technical scheme provided by the application adopts the mode selection valve and the internal content axial series double combustion chamber structure, under the medium Mach number, the front combustion chamber 11 and the rear combustion chamber 14 work, the rear combustion chamber 14 completes the reheating cycle by burning oil gas again, improves the work capacity, and guarantees sufficient thrust, under the high Mach number, the gas directly enters the rear combustion chamber 14 from the second outer channel 6 after passing through the fan and is burned, the process reduces the supercharging ratio and improves the through-flow capacity, so that the wide speed range flight is realized, the problems of insufficient thrust and the decrease of thermal cycle efficiency caused by the excessively high supercharging ratio and temperature under the high Mach number are solved, the working load of the engine parts is reduced, the deficiencies of insufficient thrust and the inability to consider economy under the high Mach number flight are compensated, and the three-channel configuration is adopted, the three-channel, double-channel and single-channel switching can be realized by using the three guide mode selection valves, so that the flow regulation and matching work in 0-3Ma are realized by the control of the mode selection valve and the opening and closing of the double combustion chamber, the engine structure is simple, the wide speed range flight is realized, the high mobility and economy are considered, and the application is suitable for civil aviation passenger planes and military planes.

[0039] The present application encompasses any substitutions, modifications, equivalent methods and solutions made on the essence and scope of the present application. In order to make the public have a thorough understanding of the present application, the specific details are described in detail in the above preferred embodiments of the present application, and the present application can also be fully understood without the description of these details for those skilled in the art. In addition, in order to avoid unnecessary confusion to the essence of the present application, well-known methods, processes, procedures, elements and circuits are not described in detail.

[0040] Those skilled in the art can understand that all or part of the steps in the above-mentioned embodiment methods can be completed by programs instructing related hardware, and the programs can be stored in computer readable storage medium, such as ROM / RAM, magnetic disc, optical disc, etc.

[0041] The above is only the preferred embodiment of the present application, and it should be pointed out that, for those skilled in the art, without departing from the principle of the present application, a number of improvements and refinements can be made, and these improvements and refinements should also be regarded as the protection scope of the present application.

Claims

1. A triple spool variable pressure ratio multi-modal aero-engine, characterized in that, The engine comprises a front fan (1), a rear fan (2) and a tail nozzle (16), the front fan (1) and the rear fan (2) are communicated with the tail nozzle (16) through an inner channel (17) and a second outer channel (6), the inner channel (17) comprises a low-pressure compressor (9), a high-pressure compressor (10), a front combustion chamber (11), a high-pressure turbine (13), a rear combustion chamber (14) and a low-pressure turbine (15), the front fan (1), the rear fan (2), the low-pressure compressor (9) and the low-pressure turbine (15) are connected by a low-pressure rotor (8), the high-pressure compressor (10) and the high-pressure turbine (13) are connected by a high-pressure rotor (12); The first modal selection valve (3) is arranged between the rear fan (2) and the low-pressure compressor (9), the first modal selection valve (3) is used for controlling the opening and closing of the inner channel (17), the second modal selection valve (5) is arranged at the first outer channel (4), the second modal selection valve (5) is arranged between the low-pressure compressor (9) and the high-pressure compressor (10), the second modal selection valve (5) is used for controlling the opening and closing of the first outer channel (4), the third modal selection valve (7) is arranged between the high-pressure turbine (13) and the low-pressure turbine (15), the third modal selection valve (7) is used for controlling the communication and closing of the second outer channel (6) and the rear combustion chamber (14).

2. The three- spool variable pressure ratio multimode aeroengine, according to claim 1, characterized in that, The engine has a front combustion mode, a double combustion mode and a rear combustion mode.

3. The three-channel variable pressure ratio multi-modal aero-engine according to claim 2, wherein the front combustion mode is that the first modal selection valve (3) opens the inner channel (17), the second modal selection valve (5) opens the first outer channel (4), the third modal selection valve (7) opens the second outer channel (6), the front combustion chamber (11) is in an ignition state, the rear combustion chamber (14) is in a blowing state, after the airflow passes through the front fan (1) and the rear fan (2), the airflow is divided into two paths A and B, the B-path airflow passes through the low-pressure compressor (9) and is divided into two paths C and D, the A-path airflow and the C-path airflow are mixed and then pass through the second outer channel (6) and are directly discharged, the D-path airflow passes through the high-pressure compressor (10), the front combustion chamber (11), the high-pressure turbine (13), the rear combustion chamber (14) and the low-pressure turbine (15) and is discharged from the tail nozzle (16) to complete a thermodynamic cycle.

4. The three-duct variable pressure ratio multi-modal aeroengine, according to claim 3, wherein, The channel ratio of the front combustion mode is 4-5.

5. The three- spool variable pressure ratio multimode aeroengine, according to claim 4, characterized in that, ​ 6. The three-duct variable pressure ratio multi-modal aeroengine, as recited in claim 2, wherein, The double combustion mode is that the first mode selection valve (3) opens the inner channel (17), the second mode selection valve (5) closes the first outer channel (4), the third mode selection valve (7) opens the second outer channel (6), the front combustion chamber (11) and the rear combustion chamber (14) are in the ignition state, the airflow is divided into two paths A and B after passing through the front fan (1) and the rear fan (2), the A path gas passes through the second outer channel (6) and is directly discharged, the B path gas passes through the low-pressure compressor (9), the high-pressure compressor (10), the front combustion chamber (11), the high-pressure turbine (13), the rear combustion chamber (14) and the low-pressure turbine (15), and is discharged from the tail nozzle (16) to complete the reheating cycle.

7. The three-duct variable pressure ratio multi-modal aeroengine, according to claim 6, wherein, The double combustion mode works when 1.2 8. The three-duct variable pressure ratio multi-modal aeroengine, according to claim 7, wherein, The double combustion mode works when the channel ratio is 0.5-0.

8.

9. The three-duct variable pressure ratio multi-modal aeroengine, as recited in claim 2, wherein, The rear combustion mode is that the rear combustion chamber (14) is in the ignition state, the first mode selection valve (3) closes the inner channel (17), the second mode selection valve (5) closes the first outer channel (4), the third mode selection valve (7) closes the second outer channel (6), the control gas enters the rear combustion chamber (14) from the second outer channel (6), after mode conversion, the airflow passes through the front fan (1), the rear fan (2) and the second outer channel (6), flows into the rear combustion chamber (14) to ignite and burn, and then passes through the low-pressure turbine (15) to do work and is discharged from the tail nozzle (16) to complete the heating cycle.

10. The three- spool variable pressure ratio multimode aeroengine, according to claim 9, characterized by the fact that, The rear combustion mode works when 2.2 11. The three-duct variable pressure ratio multi-modal aeroengine according to any one of claims 3, 6, 9, characterized in that, The engine changes the pressure ratio by switching the first mode selection valve (3), the second mode selection valve (5) and the second mode selection valve (7).

12. The three-duct variable pressure ratio multi-modal aeroengine according to any one of claims 3, 6, 9, characterized in that, The engine has high pressure ratio in the front combustion mode and the double combustion mode, and has low pressure ratio in the rear combustion mode.

13. The three- spool variable pressure ratio multimode aeroengine according to claim 1, characterized in that, The engine adopts high pressure ratio when flying at low speed and adopts low pressure ratio when flying at high speed.

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