CN114623019A - Large-bypass-ratio split type variable-circulation turbofan engine - Google Patents

Large-bypass-ratio split type variable-circulation turbofan engine Download PDF

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Publication number
CN114623019A
CN114623019A CN202210525756.2A CN202210525756A CN114623019A CN 114623019 A CN114623019 A CN 114623019A CN 202210525756 A CN202210525756 A CN 202210525756A CN 114623019 A CN114623019 A CN 114623019A
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CN
China
Prior art keywords
fan
combustion chamber
turbojet
adjustable
casing
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Granted
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CN202210525756.2A
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Chinese (zh)
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CN114623019B (en
Inventor
刘子扬
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Xi'an Xingyun Aviation Technology Co ltd
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Xi'an Xingyun Aviation Technology Co ltd
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Publication of CN114623019A publication Critical patent/CN114623019A/en
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/30—Exhaust heads, chambers, or the like
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F02C7/057—Control or regulation

Abstract

The invention discloses a split variable-circulation turbofan engine with a large bypass ratio, which comprises a turbojet part, a fan part, a gas pipeline and a heat insulation layer, wherein the turbojet part and the fan part are independent respectively, the turbojet part and the fan part are connected through the gas pipeline, the turbojet part and the fan part are communicated with the inside of the gas pipeline, and the heat insulation layer is sleeved on the outer side of the gas pipeline. The invention adopts the split arrangement of the turbojet part and the fan part, separates the work of the turbojet part and the fan part, realizes that the working conditions of the turbojet part and the fan part are not influenced mutually, realizes the reduction of resistance through the adjustment of the sectional area of the air inlet, prevents the fan from being broken by shock waves at supersonic speed, and realizes the capacity of enabling the turbofan engine with large bypass ratio to have supersonic speed through the arrangement of the second combustion chamber and the third combustion chamber.

Description

Large-bypass-ratio split type variable-circulation turbofan engine
Technical Field
The invention relates to the technical field of aero-engines, in particular to a split variable-circulation turbofan engine with a large bypass ratio.
Background
Aircraft engines are an important component of aircraft and currently in use include turbojet engines, turbofan engines, turboprop engines and the like, wherein the turbofan engine is a form of engine that uses rotating members to extract kinetic energy from a fluid passing through it, being one of the internal combustion engines. In the prior art, a turbine jet is serially connected behind a fan in a large bypass ratio turbofan engine, the design causes that the windward area of the fan is too large and the windward resistance is too high, if the windward resistance is directly adjusted by sleeving an adjustable section air inlet channel outside the engine, the windward resistance of the fan can be reduced, but the working efficiency of the fan is rapidly reduced or even damaged due to the contraction of an air inlet, and the working condition of the turbojet part is disturbed by turbulent flow, so that the conventional large bypass ratio turbofan engine does not have the capability of supersonic flight.
Disclosure of Invention
Aiming at the technical defects, the invention discloses a split variable-circulation turbofan engine with a large bypass ratio, which aims to solve the problem that supersonic flight cannot be realized due to too large windward resistance of the turbofan engine with the large bypass ratio.
The invention adopts the following technical scheme:
a split variable-circulation turbofan engine with a large bypass ratio comprises a turbojet part, a fan part, a gas pipeline and a heat insulation layer, wherein the turbojet part and the fan part are independent from each other, the turbojet part and the fan part are connected through the gas pipeline, the turbojet part and the fan part are communicated with the inside of the gas pipeline, and the heat insulation layer is sleeved on the outer side of the gas pipeline;
the vortex spraying part is used for providing high-temperature and high-pressure gas to the fan part during low-speed flight and providing partial thrust during high-speed flight by afterburning, the vortex spraying part is provided with a vortex spraying shell, a vortex spraying adjustable tail spraying pipe, a vortex spraying main shaft, an air compressing movable blade wheel set, an air compressing fixed blade wheel set, a first combustion chamber, a turbine fixed blade, a turbine movable blade, an opening and closing spoiler, a second combustion chamber and a bypass pipeline, the vortex spraying adjustable tail spraying pipe is arranged at the right end of the vortex spraying shell, the vortex spraying main shaft is arranged at the center inside of the vortex spraying shell, the air compressing movable blade wheel set, the air compressing fixed blade wheel set, the first combustion chamber, the turbine fixed blade and the turbine movable blade are sequentially sleeved and arranged on the vortex spraying main shaft from left to right, the air compressing movable blade wheel set array is provided with the air compressing movable blade, the air compressing fixed blade wheel set is fixedly arranged at the inner side of the vortex spraying shell, the first combustion chamber is fixedly arranged in the turbojet casing, the turbine stationary blade is fixedly arranged on the inner side of the turbojet casing, the opening and closing spoiler is arranged in the turbojet casing, the opening and closing spoiler is arranged on the right side of the turbojet main shaft, the second combustion chamber is arranged between the right side of the opening and closing spoiler and the left side of the turbojet adjustable tail nozzle, the circumferential array of the bypass pipelines is arranged on the outer side of the turbojet casing, and the bypass pipelines are communicated with the inside of the turbojet casing;
wherein the fan part is used for converting the internal energy of high-temperature and high-pressure gas provided by the turbojet part into mechanical energy during low-speed flight, providing main thrust for low-speed flight and providing partial thrust for stamping combustion during high-speed flight, the fan part is provided with an air inlet, a fan main shaft, a fan rectifying impeller, a fan inner shell, a radial turbine and a third combustion chamber, the air inlet channel is arranged at the left end of the fan shell, the adjustable tail nozzle of the fan is arranged at the right end of the fan shell, the fan main shaft is arranged at the center of the inner part of the fan shell, the fan is arranged at the left end of the fan main shaft, the fan inner shell is arranged in the fan shell, the fan rectifying impeller is fixedly arranged at the left end of the fan inner shell, the radial turbine is arranged in the fan inner shell, the radial turbine is arranged at the right end of the fan main shaft, and the third combustion chamber is arranged between the fan inner shell and the adjustable tail nozzle of the fan;
the gas pipeline is used for conveying high-temperature and high-pressure gas of the turbojet part to the fan part, one end of the gas pipeline is arranged on the turbojet shell between the turbine movable blades and the opening and closing spoiler, and the other end of the gas pipeline is eccentrically arranged on the fan inner shell on the outer side of the radial turbine.
As a further technical solution of the present invention, the first combustion chamber is an annular combustion chamber, the first combustion chamber is a main combustion chamber, the first combustion chamber is provided with an air inlet, a fixing rod, a first oil nozzle and a first igniter, the air inlet array is disposed on an inner wall and an outer wall of the first combustion chamber, the fixing rod circumferential array is disposed outside the first combustion chamber, one end of the fixing rod is disposed inside the turbojet casing, the first oil nozzle circumferential array is disposed at a left end of the first combustion chamber, and the first igniter is disposed above one of the first oil nozzles.
As a further technical scheme of the present invention, the second combustion chamber is an afterburner and ram combustion chamber, the second combustion chamber is a cavity formed by a turbojet casing, an opening and closing spoiler and a turbojet adjustable tail nozzle, the second combustion chamber is provided with a turbojet flame stabilizer, a second oil nozzle and a second igniter, the turbojet flame stabilizer is arranged on the right side of the opening and closing spoiler, the second oil nozzle is circumferentially arrayed on the turbojet casing, and the second igniter is arranged on the right side of one of the second oil nozzles.
As a further technical scheme, the number of the bypass pipelines is 3-6, bypass valves are arranged on the bypass pipelines, an air inlet at the left end of each bypass pipeline is arranged on a turbojet casing on the outer side of the middle section of the air compression wheel set, and an air outlet at the right end of each bypass pipeline is arranged on a turbojet casing on the left side of a turbojet flame stabilizer in the second combustion chamber.
As a further technical scheme of the invention, the air inlet channel is provided with a variable-section fairing, an adjustable shock wave guide cone bracket and a telescopic motor, the variable cross-section fairing is arranged at the left end of the fan shell, the adjustable shock wave guide cone bracket is arranged inside the variable cross-section fairing, the outer side of the adjustable shock wave diversion cone bracket is circumferentially arrayed with bracket rods, one end of each bracket rod is arranged on the inner side of the variable-section fairing, the adjustable shock wave diversion cone is arranged at the left end of the adjustable shock wave diversion cone bracket, the right side in the adjustable shock wave diversion cone is provided with a screw hole, the adjustable shock wave diversion cone is connected with the adjustable shock wave diversion cone support in a sliding mode, the telescopic motor is fixedly arranged inside the adjustable shock wave diversion cone support, the left end of the telescopic motor is provided with a screw, and the screw is in threaded connection with the screw.
As a further technical scheme of the present invention, the third combustion chamber is a stamping combustion chamber, the third combustion chamber is a cavity formed by a fan housing, a fan inner housing and a fan adjustable tail nozzle, the third combustion chamber is provided with a fan flame stabilizer, a third oil nozzle and a third igniter, the fan flame stabilizer is arranged on the right side of the fan inner housing, the third oil nozzle is circumferentially arranged on the fan housing in an array manner, and the third igniter is arranged on the right side of one of the third oil nozzles.
As a further technical scheme of the invention, the heat-insulating layer material is an aerogel heat-insulating material.
As a further technical scheme of the invention, the adjustable turbojet tail pipe and the adjustable fan tail pipe are adjustable Laval spray pipes.
As a further technical scheme of the invention, the left end of the adjustable shock wave guide cone shape is a cone, the right end of the adjustable shock wave guide cone shape is a cylinder, and the middle of the adjustable shock wave guide cone shape is provided with an arc for transition.
As a further technical scheme of the invention, one vortex spraying part can be connected with 1-3 fan parts according to specific use conditions.
The invention has the positive beneficial effects that:
the invention adopts the split arrangement of the turbojet part and the fan part, separates the work of the turbojet part and the fan part, realizes that the working conditions of the turbojet part and the fan part are not influenced mutually, realizes the reduction of resistance and the prevention of the fan from being broken by shock waves through the adjustment of the sectional area of the air inlet passage, and realizes the capability of enabling the engine to have supersonic speed through the arrangement of the second combustion chamber and the third combustion chamber.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without inventive exercise, wherein:
FIG. 1 is a schematic perspective view of the present invention;
FIG. 2 is a schematic perspective view of a turbojet part according to the present invention;
FIG. 3 is a schematic structural view of a front view of a turbojet part in accordance with the present invention;
FIG. 4 is a schematic structural diagram of a top view of a turbojet part in accordance with the present invention;
FIG. 5 is a left side view of the turbojet part of the present invention;
FIG. 6 is an enlarged view of a portion of FIG. 3;
FIG. 7 is a perspective view of a fan assembly according to the present invention;
FIG. 8 is a front view of the fan section of the present invention;
FIG. 9 is a schematic top view of a fan assembly according to the present invention;
FIG. 10 is a left side view of the fan assembly of the present invention;
FIG. 11 is a schematic cross-sectional view of an inlet according to the present invention;
FIG. 12 is a cross-sectional view of the inner structure of the fan according to the present invention;
FIG. 13 is a schematic top view of a fan assembly according to the present invention;
the labels in the figure are:
1-a vortex spraying part; 11-a turbojet casing; 12-vortex spraying adjustable tail nozzle; 13-a main shaft of vortex spraying; 14-a pneumatic impeller set; 15-a stationary compressed air vane wheel set; 16-a first combustion chamber; 161-air inlet holes; 162-a fixation rod; 163-first fuel injector; 164-a first igniter; 17-a turbine vane; 18-turbine buckets; 19-an opening and closing spoiler; 110-a turbojet flame stabilizer; 111-a second combustion chamber; 112-a bypass line; 113-a bypass valve; 114-a second fuel injector; 115-a second igniter; 2-a fan section; 20-an air inlet channel; 21-a variable section fairing; 22-a fan housing; 23-a fan variable nozzle; 24-an adjustable shock wave guide cone; 241-screw holes; 25-adjustable shock wave guide cone support; 251-a support bar; 26-a telescopic motor; 261-screw; 27-a fan main shaft; 28-a fan; 29-a fan straightening impeller; 210-a fan inner housing; 211-radial turbine; 212-a fan flame holder; 213-a third combustion chamber; 214-a third fuel injector; 215-third igniter; 3-a gas pipeline; 4-heat preservation layer.
Detailed Description
The preferred embodiments of the present invention will be described below with reference to the accompanying drawings, and it should be understood that the embodiments described herein are merely for the purpose of illustrating and explaining the present invention and are not intended to limit the present invention.
As shown in fig. 1-13, the split variable-circulation turbofan engine with a large bypass ratio comprises a turbojet part 1, a fan part 2, a gas pipeline 3 and an insulating layer 4, wherein the turbojet part 1 and the fan part 2 are independent from each other, the turbojet part 1 and the fan part 2 are connected through the gas pipeline 3, the turbojet part 1 and the fan part 2 are communicated with the inside of the gas pipeline 3, and the insulating layer 4 is sleeved on the outer side of the gas pipeline 3;
in the specific embodiment, when the split variable-circulation turbofan engine with the large bypass ratio works and is cruising at a low speed, high-temperature and high-pressure fuel gas generated by the turbojet part 1 is conveyed to the fan part 2 through the fuel gas pipeline 3 to push the fan part 2 to rotate so as to provide power for flight; in the supersonic speed stage, the gas is accelerated by the turbojet part 1 through the turbojet adjustable tail nozzle 12 and is sprayed to the rear of the engine to provide power for flying; when Mach 3 is about, the vortex spraying part 1 and the fan part 2 work simultaneously to generate hypersonic airflow to provide power for flight.
Wherein the vortex spraying part 1 is used for providing high-temperature and high-pressure gas to the fan part 2 during low-speed flight and providing partial thrust during high-speed flight by afterburning, the vortex spraying part 1 is provided with a vortex spraying shell 11, a vortex spraying adjustable tail spraying pipe 12, a vortex spraying main shaft 13, an air compressing movable vane wheel set 14, an air compressing stationary vane wheel set 15, a first combustion chamber 16, a turbine stationary vane 17, a turbine movable vane 18, an opening and closing spoiler 19, a second combustion chamber 111 and a bypass pipeline 112, the vortex spraying adjustable tail spraying pipe 12 is arranged at the right end of the vortex spraying shell 11, the vortex spraying main shaft 13 is arranged at the center inside of the vortex spraying shell 11, the air compressing movable vane wheel set 14, the air compressing stationary vane wheel set 15, the first combustion chamber 16, the turbine stationary vane 17 and the turbine movable vane 18 are sequentially sleeved and arranged on the vortex spraying main shaft 13 from left to right, wherein the air compressing movable vane wheel set 14 array is provided with air compressing movable vanes, wherein the air compressing movable vane wheel set 15 array is provided with air compressing stationary vanes, the air compression stationary blade wheel set 15 is fixedly arranged on the inner side of the turbojet casing 11, the first combustion chamber 16 is fixedly arranged inside the turbojet casing 11, the turbine stationary blade 17 is fixedly arranged on the inner side of the turbojet casing 11, the opening and closing spoiler 19 is arranged inside the turbojet casing 11, the opening and closing spoiler 19 is arranged on the right side of the turbojet main shaft 13, the second combustion chamber 111 is arranged between the right side of the opening and closing spoiler 19 and the left side of the turbojet adjustable tail nozzle 12, the bypass pipelines 112 are circumferentially arranged on the outer side of the turbojet casing 11, and the bypass pipelines 112 are communicated with the inside of the turbojet casing 11;
in a specific embodiment, during low-speed cruising, the turbojet part 1 generates high-temperature and high-pressure gas through combustion in the first combustion chamber 16, the high-temperature and high-pressure gas is discharged from the first combustion chamber 16 and then drives the turbine movable blades 18 to rotate, the rotation of the turbine movable blades 18 drives the turbojet main shaft 13 to rotate, the rotation of the turbojet main shaft 13 drives the air compression movable blade wheel set 14 to rotate, the air compression movable blade wheel set 14 is matched with the air compression stationary blade wheel set 15 to pressurize airflow entering the turbojet part 1 and convey the airflow to the first combustion chamber 16, and meanwhile, the high-temperature and high-pressure gas generated through combustion is conveyed to the fan part 2 through the gas pipeline 3; in the supersonic speed stage, the opening and closing spoiler 19 is opened, the second combustion chamber 111 works, the second combustion chamber 111 serves as an afterburner to perform combustion work, high-temperature and high-pressure fuel gas is generated, the fuel gas is accelerated through the turbojet adjustable tail pipe 12, supersonic speed gas is generated, and the supersonic speed gas flows to be ejected towards the rear of the engine to provide power for flying; at about Mach 3, a bypass valve 113 of the turbojet part 1 is opened, partial airflow is directly introduced into the second combustion chamber 111 through a bypass pipeline 112, and the second combustion chamber 111 is converted into a ram combustion chamber from an afterburner to generate hypersonic airflow so as to provide power for flight.
Wherein the fan part 2 is used for converting the high-temperature high-pressure gas internal energy provided by the turbojet part into mechanical energy during low-speed flight, and provides main thrust for low-speed flight, and performs stamping combustion during high-speed flight to provide partial thrust, the fan part 2 is provided with an air inlet 20, a fan main shaft 27, a fan 28, a fan rectifying impeller 29, a fan inner housing 210, a radial turbine 211 and a third combustion chamber 213, the air inlet 20 is arranged at the left end of the fan housing 22, the fan adjustable tail nozzle 23 is arranged at the right end of the fan housing 22, the fan main shaft 27 is arranged at the center inside of the fan housing 22, the fan 28 is arranged at the left end of the fan main shaft 27, the fan inner housing 210 is arranged inside the fan housing 22, the fan rectifying impeller 29 is fixedly arranged at the left end of the fan inner housing 210, the radial turbine 211 is arranged inside the fan inner housing 210, the radial turbine 211 is arranged at the right end of the fan main shaft 27, and the third combustion chamber 213 is arranged between the fan inner shell 210 and the fan variable tail pipe 23;
in a specific embodiment, during low-speed cruising, high-temperature and high-pressure fuel gas is conveyed to the fan inner shell 210 through the fuel gas pipeline 3, so that the radial turbine 211 rotates, the rotation of the radial turbine 211 drives the fan main shaft 27 to rotate, and the rotation of the fan main shaft 27 drives the fan 28 to rotate, thereby generating power during the low-speed cruising; during the supersonic phase, the inlet 20 of the fan section 2 is closed, reducing drag and preventing shock waves from breaking the fan 28; at about mach 3, according to the current airspeed, the air inlet 20 of the fan part 2 is opened to a certain opening degree, high-speed incoming flow is decelerated to subsonic speed in the air inlet 20 and is supplied to the rear third combustion chamber 213, and the third combustion chamber 213 starts to work as a stamping combustion chamber to generate high-speed air flow to provide power for flight.
The gas pipeline 3 is used for conveying high-temperature and high-pressure gas of the turbojet part 1 to the fan part 2, one end of the gas pipeline 3 is arranged on the turbojet casing 11 between the turbine rotor blades 18 and the opening and closing spoiler 19, and the other end of the gas pipeline 3 is eccentrically arranged on a fan inner casing 210 outside the radial turbine 211.
In the embodiment, the gas duct 3 delivers the high-temperature and high-pressure gas generated from the turbojet part 1 to the inner fan casing 210 of the fan part 2 at the low-speed cruising.
In the present invention, the first combustion chamber 16 is an annular combustion chamber, the first combustion chamber 16 is a main combustion chamber, the first combustion chamber 16 is provided with an air inlet hole 161, a fixing rod 162, a first oil nozzle 163 and a first igniter 164, the air inlet hole 161 is arranged on the inner wall and the outer wall of the first combustion chamber 16 in an array, the fixing rod 162 is arranged outside the first combustion chamber 16 in a circumferential array, one end of the fixing rod 162 is arranged inside the turbojet casing 11, the first oil nozzle 163 is arranged at the left end of the first combustion chamber 16 in a circumferential array, and the first igniter 164 is arranged above one of the first oil nozzles 163.
In a specific embodiment, the first combustion chamber 16 is fixed inside the turbojet casing 11 through a fixing rod 162 to complete the fixing of the first combustion chamber 16, the airflow pressurized by the air compressing movable vane wheel set 14 and the air compressing stationary vane wheel set 15 enters the first combustion chamber 16 through air inlet holes 161 on the inner wall and the outer wall of the first combustion chamber 16, the atomized fuel oil is sprayed into the first combustion chamber 16 through a first oil nozzle 163, the atomized fuel oil is mixed with the airflow entering the first combustion chamber 16, the atomized fuel oil is ignited through a first igniter 164, the fuel oil is combusted in the first combustion chamber 16 to generate high-temperature and high-pressure gas, and the high-temperature and high-pressure gas is discharged backwards to continuously generate the high-temperature and high-pressure gas.
In the present invention, the second combustion chamber 111 is an afterburner and ram combustion chamber, the second combustion chamber 111 is a cavity formed by a turbojet casing 11, an opening and closing spoiler 19 and a turbojet adjustable tail nozzle 12, the second combustion chamber 111 is provided with a turbojet flame stabilizer 110, second oil nozzles 114 and second igniters 115, the turbojet flame stabilizer 110 is arranged on the right side of the opening and closing spoiler 19, the second oil nozzles 114 are circumferentially arranged on the turbojet casing 11 in an array, and the second igniters 115 are arranged on the right side of one of the second oil nozzles 114.
In the specific embodiment, the opening and closing spoiler 19 is opened, the airflow enters the cavity of the second combustion chamber 111, the atomized fuel is injected into the cavity of the second combustion chamber 111 through the second fuel injection nozzle 114, the atomized fuel is mixed with the airflow entering the second combustion chamber 111, the atomized fuel is ignited through the second igniter 115, the stability of fuel combustion is ensured through the vortex flame stabilizer 110, the fuel is combusted in the cavity of the second combustion chamber 111 to generate high-temperature and high-pressure fuel, and is discharged backwards to continuously generate the high-temperature and high-pressure fuel.
In the present invention, the number of the bypass pipelines 112 is 3 to 6, the bypass pipeline 112 is provided with a bypass valve 113, an air inlet at the left end of the bypass pipeline 112 is arranged on the turbojet casing 11 at the outer side of the middle section of the air compressor set, and an air outlet at the right end of the bypass pipeline 112 is arranged on the turbojet casing 11 at the left side of the turbojet flame stabilizer 110 in the second combustion chamber 111.
In a specific embodiment, a bypass valve 113 is disposed on the bypass line 112 to control opening and closing of the bypass line 112, and when the bypass line 112 is opened, an airflow is introduced from an air inlet of the turbojet casing 11 outside the middle section of the compressor wheel set and discharged from an air outlet of the turbojet casing 11 on the left side of the turbojet flame stabilizer 110 in the second combustion chamber 111, so as to realize delivery of the airflow to the second combustion chamber 111.
In the invention, the air inlet channel 20 is provided with a variable cross-section fairing 21, an adjustable shock wave diversion cone 24, an adjustable shock wave diversion cone support 25 and a telescopic motor 26, the variable cross-section fairing 21 is arranged at the left end of a fan shell 22, the adjustable shock wave diversion cone support 25 is arranged inside the variable cross-section fairing 21, support rods 251 are arranged on the outer side of the adjustable shock wave diversion cone support 25 in a circumferential array manner, one end of each support rod 251 is arranged inside the variable cross-section fairing 21, the adjustable shock wave diversion cone 24 is arranged at the left end of the adjustable shock wave diversion cone support 25, screw holes 241 are arranged on the right side inside the adjustable shock wave diversion cone 24, the adjustable shock wave diversion cone 24 is in sliding connection with the adjustable shock wave diversion cone support 25, the telescopic motor 26 is fixedly arranged inside the adjustable shock wave diversion cone support 25, and the left end of the telescopic motor 26 is provided with a screw 261, the screw hole 241 is in threaded connection with the screw 261.
In a specific embodiment, the fixing of the air inlet channel 20 is realized by fixedly connecting the variable-section fairing 21 with the fan shell 22, the adjustable shock wave guide cone support 25 is fixed inside the variable-section fairing 21 through the support rod 251, the adjustable shock wave guide cone support 25 supports the adjustable shock wave guide cone 24, the adjustable shock wave guide cone support 25 fixes the telescopic motor 26, the telescopic motor 26 drives the screw 261 to rotate, the screw 241 is in threaded connection with the screw 261, the rotation of the screw 261 drives the adjustable shock wave guide cone 24 to slide in the adjustable shock wave guide cone support 25, the slide of the adjustable shock wave guide cone 24 is matched with the variable-section fairing 21, and the size of the air inlet in the air inlet channel 20 is controlled.
In the present invention, the third combustion chamber 213 is a ramjet combustion chamber, the third combustion chamber 213 is a cavity formed by the fan housing 22, the fan inner housing 210, and the fan adjustable tail nozzle 23, the third combustion chamber 213 is provided with a fan flame stabilizer 212, a third oil nozzle 214, and a third igniter 215, the fan flame stabilizer 212 is disposed on the right side of the fan inner housing 210, the third oil nozzle 214 is circumferentially arranged on the fan housing 22 in an array, and the third igniter 215 is disposed on the right side of one of the third oil nozzles 214.
In a specific embodiment, at about mach 3, the air inlet 20 is opened by a certain opening degree to decelerate hypersonic incoming flow to subsonic speed, the hypersonic incoming flow is provided to the third combustion chamber 213, the air flow enters the cavity of the third combustion chamber 213 to be mixed with atomized fuel oil sprayed by the third oil nozzle 214, and then the atomized fuel oil is ignited by the third igniter 215, the fan flame stabilizer 212 ensures the stability of fuel oil combustion, and the fuel oil is combusted in the cavity of the third combustion chamber 213 to generate high-temperature and high-pressure fuel gas, and is discharged backwards to continuously generate the high-temperature and high-pressure fuel gas.
In a particular embodiment, the first igniter 164, the second igniter 115, and the third igniter 215 are glow sticks.
In the invention, the material of the heat-insulating layer 4 is aerogel heat-insulating material.
In the specific embodiment, the heat-insulating layer 4 is coated on the outer side of the gas pipeline 3, so that the energy loss of high-temperature and high-pressure gas in the conveying process is prevented.
In the present invention, the turbojet adjustable exhaust nozzle 12 and the fan adjustable exhaust nozzle 23 are adjustable laval nozzles.
In a particular embodiment, the exit velocity of the gas stream is regulated by an adjustable laval nozzle.
In the invention, the adjustable shock wave guide cone 24 is in the shape of a cone at the left end and a cylinder at the right end, and an arc is arranged in the middle for transition.
In the invention, one turbojet part 1 can be connected with 1-3 fan parts 2 according to specific use requirements.
In a specific embodiment, the number of the fan parts 2 which are matched with one turbojet part 1 can be adjusted within 1-3 according to different use requirements.
The present invention will be further described below in conjunction with specific examples to more clearly understand the present invention.
When the low-speed cruise ship is in operation, the turbojet part 1 supplies high-temperature and high-pressure fuel gas to the fan part 2 during low-speed cruise, the fuel gas is conveyed through the fuel gas pipeline 3, high-temperature maintenance is realized through the heat insulation layer 4 during the conveying process, and the fan part 2 generates power during low-speed cruise; in the stage from transonic speed to supersonic speed, a variable cross-section fairing 21 in an air inlet channel 20 in the fan part 2 and an adjustable shock wave guide cone 24 work cooperatively to close the air inlet channel 20 in the fan part 2, reduce windward resistance and prevent shock waves from breaking a fan 28, meanwhile, a vortex jet part 1 opens an opening and closing spoiler 19 and starts a second combustion chamber 111 of the vortex jet part 1, the second combustion chamber 111 is an afterburner, generated fuel gas is accelerated through a vortex jet adjustable tail jet pipe 12 to generate supersonic speed air to be jetted towards the rear part of an engine, and the vortex jet part 1 provides main power for flying; when the airplane is pushed to about Mach 3, a bypass valve 113 of the turbojet part 1 is opened, part of airflow directly enters a second combustion chamber 111 through a bypass pipeline 112, at the moment, the second combustion chamber 111 is converted into a stamping combustion chamber from an afterburner to generate hypersonic airflow, meanwhile, a variable cross-section fairing 21 and an adjustable shock wave guide cone 24 are opened to a certain opening degree according to the current airspeed to decelerate hypersonic incoming flow to subsonic speed in an air inlet channel 20 and provide the subsonic speed to a third combustion chamber 213 located behind, at the moment, the third combustion chamber 213 is a stamping combustion chamber to generate hypersonic airflow and provide power for hypersonic flight.
Although specific embodiments of the present invention have been described above, it will be understood by those skilled in the art that these specific embodiments are merely illustrative and that various omissions, substitutions and changes in the form of the detail of the methods and systems described above may be made by those skilled in the art without departing from the spirit and scope of the invention. For example, it is within the scope of the present invention to combine the steps of the above-described methods to perform substantially the same function in substantially the same way to achieve substantially the same result. Accordingly, the scope of the invention is to be limited only by the following claims.

Claims (9)

1. The utility model provides a split type variable cycle turbofan engine of big bypass ratio which characterized in that: the vortex spraying device comprises a vortex spraying part (1), a fan part (2), a gas pipeline (3) and an insulating layer (4), wherein the vortex spraying part (1) and the fan part (2) are independent respectively, the vortex spraying part (1) and the fan part (2) are connected through the gas pipeline (3), the vortex spraying part (1) and the fan part (2) are communicated with the inside of the gas pipeline (3), and the insulating layer (4) is sleeved on the outer side of the gas pipeline (3); wherein the vortex spraying part (1) is used for providing high-temperature and high-pressure gas to the fan part (2) during low-speed flight and providing partial thrust during afterburning during high-speed flight, the vortex spraying part (1) is provided with a vortex spraying shell (11), a vortex spraying adjustable tail spraying pipe (12), a vortex spraying main shaft (13), an air compressing movable blade wheel set (14), an air compressing stationary blade wheel set (15), a first combustion chamber (16), a turbine stationary blade (17), a turbine movable blade (18), an opening and closing spoiler (19), a second combustion chamber (111) and a bypass pipeline (112), the vortex spraying adjustable tail spraying pipe (12) is arranged at the right end of the vortex spraying shell (11), the vortex spraying main shaft (13) is arranged at the inner center of the vortex spraying shell (11), the air compressing movable blade wheel set (14), the air compressing stationary blade wheel set (15), the first combustion chamber (16), the turbine stationary blade (17) and the turbine movable blade (18) are sequentially sleeved on the vortex spraying main shaft (13) from left to right, the compressed air movable blade wheel set (14) is provided with compressed air movable blades in an array mode, the compressed air fixed blade wheel set (15) is provided with compressed air fixed blades in an array mode, the compressed air fixed blade wheel set (15) is fixedly arranged on the inner side of a turbojet casing (11), the first combustion chamber (16) is fixedly arranged inside the turbojet casing (11), the turbine fixed blades (17) are fixedly arranged on the inner side of the turbojet casing (11), the opening and closing spoiler (19) is arranged inside the turbojet casing (11), the opening and closing spoiler (19) is arranged on the right side of the turbojet main shaft (13), the second combustion chamber (111) is arranged between the right side of the opening and closing spoiler (19) and the left side of the turbojet adjustable tail nozzle (12), the circumferential array of the bypass pipeline (112) is arranged on the outer side of the turbojet casing (11), and the bypass pipeline (112) is communicated with the inside of the turbojet casing (11);
wherein the fan part (2) is used for converting the high-temperature high-pressure gas internal energy provided by the turbojet part (1) into mechanical energy when flying at a low speed, and provides main thrust when flying at a low speed, and the high-temperature high-pressure gas internal energy provided by the turbojet part (1) is punched and burned to provide partial thrust when flying at a high speed, the fan part (2) is provided with an air inlet (20), a fan main shaft (27), a fan (28), a fan rectifying impeller (29), a fan inner shell (210), a radial turbine (211) and a third combustion chamber (213), the air inlet (20) is arranged at the left end of the fan shell (22), the fan adjustable tail nozzle (23) is arranged at the right end of the fan shell (22), the fan main shaft (27) is arranged at the inner center of the fan shell (22), the fan (28) is arranged at the left end of the fan main shaft (27), the fan inner shell (210) is arranged inside the fan shell (22), the fan rectifying impeller (29) is fixedly arranged at the left end of the fan inner shell (210), the radial turbine (211) is arranged inside the fan inner shell (210), the radial turbine (211) is arranged at the right end of the fan main shaft (27), and the third combustion chamber (213) is arranged between the fan inner shell (210) and the adjustable tail pipe (23) of the fan;
the gas pipeline (3) is used for conveying high-temperature and high-pressure gas of the turbojet part (1) to the fan part (2), one end of the gas pipeline (3) is arranged on the turbojet casing (11) between the turbine movable blades (18) and the opening and closing spoiler (19), and the other end of the gas pipeline (3) is eccentrically arranged on a fan inner casing (210) on the outer side of the radial turbine (211);
one of the vortex spraying parts (1) can be connected with 1-3 fan parts (2) according to specific use requirements.
2. The split variable cycle turbofan engine of claim 1 wherein: first combustion chamber (16) is annular combustion chamber, first combustion chamber (16) is main combustion chamber, first combustion chamber (16) is provided with inlet port (161), dead lever (162), first oil nozzle (163) and first igniter (164), inlet port (161) array sets up on first combustion chamber (16) inner wall and outer wall, dead lever (162) circumference array sets up in first combustion chamber (16) outside, dead lever (162) one end sets up in turbojet casing (11) inboard, first oil nozzle (163) circumference array sets up in first combustion chamber (16) left end, first igniter (164) sets up in the top of one of them first oil nozzle (163).
3. The split variable cycle turbofan engine of claim 1 wherein: the second combustion chamber (111) is an afterburner and stamping combustion chamber, the second combustion chamber (111) is a cavity formed by a turbojet casing (11), an opening-closing spoiler (19) and a turbojet adjustable tail nozzle (12), the second combustion chamber (111) is provided with a turbojet flame stabilizer (110), a second oil nozzle (114) and a second igniter (115), the turbojet flame stabilizer (110) is arranged on the right side of the opening-closing spoiler (19), the second oil nozzle (114) is circumferentially arrayed on the turbojet casing (11), and the second igniter (115) is arranged on the right side of one of the second oil nozzles (114).
4. The split variable cycle turbofan engine of claim 1 wherein: the quantity of bypass pipeline (112) is 3~6, be provided with bypass valve (113) on bypass pipeline (112), bypass pipeline (112) left end air inlet sets up on turbojet casing (11) in the compressor wheel group middle section outside, bypass pipeline (112) right-hand member gas outlet sets up in second combustion chamber (111) on left turbojet casing (11) of turbojet flame stabilizer (110).
5. The split variable cycle turbofan engine of claim 1 wherein: the air inlet duct (20) is provided with a variable-section fairing (21), an adjustable shock wave diversion cone (24), an adjustable shock wave diversion cone support (25) and a telescopic motor (26), the variable-section fairing (21) is arranged at the left end of a fan shell (22), the adjustable shock wave diversion cone support (25) is arranged inside the variable-section fairing (21), a support rod (251) is arranged on the outer side of the adjustable shock wave diversion cone support (25) in a circumferential array mode, one end of the support rod (251) is arranged on the inner side of the variable-section fairing (21), the adjustable shock wave diversion cone (24) is arranged at the left end of the adjustable shock wave diversion cone support (25), a screw hole (241) is arranged on the right side inside the adjustable shock wave diversion cone (24), the adjustable shock wave diversion cone (24) is in sliding connection with the adjustable shock wave diversion cone support (25), and the telescopic motor (26) is fixedly arranged inside the adjustable shock wave diversion cone support (25), the left end of the telescopic motor (26) is provided with a screw rod (261), and the screw hole (241) is in threaded connection with the screw rod (261).
6. The split variable cycle turbofan engine of claim 1 wherein: third combustion chamber (213) are the punching press combustion chamber, third combustion chamber (213) constitute the cavity by fan casing (22), fan inner casing (210) and the adjustable tail-nozzle (23) of fan, third combustion chamber (213) are provided with fan flame stabilizer (212), third oil nozzle (214) and third igniter (215), fan flame stabilizer (212) set up in fan inner casing (210) right side, third oil nozzle (214) circumference array sets up on fan casing (22), third igniter (215) set up in one of them third oil nozzle (214) right side.
7. The split variable cycle turbofan engine of claim 1 wherein: the heat-insulating layer (4) is made of aerogel heat-insulating material.
8. The split variable cycle turbofan engine of claim 1 wherein: the adjustable turbojet tail pipe (12) and the adjustable fan tail pipe (23) are adjustable Laval nozzles.
9. The split variable cycle turbofan engine of claim 5 wherein: the adjustable shock wave diversion cone (24) is in the shape of a cone at the left end, a cylinder at the right end and an arc arranged in the middle for transition.
CN202210525756.2A 2022-05-16 2022-05-16 Large-bypass-ratio split type variable-circulation turbofan engine Active CN114623019B (en)

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Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1546722A (en) * 1975-06-16 1979-05-31 Gen Electric Gas turbine engine
CN103967649A (en) * 2013-01-24 2014-08-06 李吉光 Turbojet engine
CN108005812A (en) * 2017-12-04 2018-05-08 中国航空发动机研究院 Using adaptive casing and the intelligent engine of adaptive fan
EP3597895A1 (en) * 2018-07-20 2020-01-22 Rolls-Royce plc Supersonic aircraft turbofan engine
CN111102098A (en) * 2020-01-03 2020-05-05 中国科学院工程热物理研究所 Turbojet propulsion system based on front-mounted compression guide impeller and control method
US20200400081A1 (en) * 2019-06-24 2020-12-24 Rolls-Royce Plc Gas turbine engine transfer efficiency
CN112443423A (en) * 2020-11-24 2021-03-05 南京航空航天大学 Jet propulsion power system of air-driven ducted fan

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1546722A (en) * 1975-06-16 1979-05-31 Gen Electric Gas turbine engine
CN103967649A (en) * 2013-01-24 2014-08-06 李吉光 Turbojet engine
CN108005812A (en) * 2017-12-04 2018-05-08 中国航空发动机研究院 Using adaptive casing and the intelligent engine of adaptive fan
EP3597895A1 (en) * 2018-07-20 2020-01-22 Rolls-Royce plc Supersonic aircraft turbofan engine
US20200400081A1 (en) * 2019-06-24 2020-12-24 Rolls-Royce Plc Gas turbine engine transfer efficiency
CN111102098A (en) * 2020-01-03 2020-05-05 中国科学院工程热物理研究所 Turbojet propulsion system based on front-mounted compression guide impeller and control method
CN112443423A (en) * 2020-11-24 2021-03-05 南京航空航天大学 Jet propulsion power system of air-driven ducted fan

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