CN111102098A - Turbojet propulsion system based on front-mounted compression guide impeller and control method - Google Patents
Turbojet propulsion system based on front-mounted compression guide impeller and control method Download PDFInfo
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- CN111102098A CN111102098A CN202010004512.0A CN202010004512A CN111102098A CN 111102098 A CN111102098 A CN 111102098A CN 202010004512 A CN202010004512 A CN 202010004512A CN 111102098 A CN111102098 A CN 111102098A
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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
- F02K3/075—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 controlling flow ratio between flows
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- 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
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Abstract
The invention provides a turbojet propulsion system based on a front-mounted compression guide impeller and a control method thereof, wherein the front-mounted compression guide impeller is adopted to adjust the pressure ratio of an engine, so that the engine can be circularly changed in the whole envelope range, the bypass ratio can be effectively adjusted, and the internal and external bypass powers can be greatly adjusted and redistributed; through the combined adjustment design of the engine and the air inlet, the area ratio of the throat of the air inlet works within an acceptable efficiency range, and the matching of the air inlet flow and the working flow of a main engine is realized, so that the thrust loss is reduced, and the propulsion efficiency is improved. When the incoming air flow speed is low, the compressed guide impeller is used for guiding flow, the fan is in a small pressure ratio mode, the throat area of the air inlet channel is in a small state, so that the bypass ratio is effectively increased, and the oil consumption rate is reduced; when the incoming flow speed of the inlet air is high, the compression guide vane wheel is used as the air compressor, the throat area of the inlet channel is in a large state, the fan pressure ratio is increased, the bypass ratio is reduced, the gas temperature in front of the turbine is increased, and the unit thrust is improved.
Description
Technical Field
The invention relates to the field of aircraft power design, in particular to a supersonic aircraft propulsion system, and a control method thereof, wherein the turbojet propulsion system is based on a front-mounted compression guide impeller, and the novel turbojet propulsion system can realize the characteristics that an engine has a large unit thrust of a turbojet engine or a turbofan engine with a small bypass ratio in a supersonic speed state and has the characteristics of a smaller unit thrust, lower noise and lower oil consumption rate than the turbofan engine with a large bypass ratio in a subsonic speed in the aircraft flight process.
Background
The appearance of the turbojet engine replaces an aviation piston engine, so that the flight performance of the airplane is greatly improved, people can realize supersonic flight, but the turbojet engine has superior performance at high speed and poor economy at low speed. In order to maintain better performance of the turbojet at high speeds and to improve its economy at low speeds, turbofan engines have been invented by adding an external bypass to the engine. The turbofan engine for the military supersonic fighter adopts a small bypass ratio design, mixes exhaust gas, and is provided with an afterburner, so that the engine has the advantages of large boost ratio, good subacoustic cruise economy and superior performance of supersonic flying. The turbofan engine for the large-scale transport plane/passenger plane adopts a large bypass ratio design, adopts separated exhaust or mixed exhaust, has large take-off thrust and good cruising economy. With the further development of the aircraft engine technology, new requirements of military fighters provide new requirements for the design of new-generation engines, and besides the requirement of higher thrust-weight ratio, the engines are required to have the characteristic of high unit thrust of turbojet engines so as to meet the requirements of supersonic cruise, combat maneuver flight, transonic speed acceleration and the like; and the characteristic of low oil consumption rate of the turbofan engine during subsonic cruising is required to meet requirements of subsonic cruising, standby, air patrol and the like. Obviously, to achieve the above conflicting cycle objectives to some extent, variable cycle turbofan engines are undoubtedly ideal propulsion devices, and therefore, in order to take into account the characteristics of a turbojet engine or a small bypass ratio turbofan engine with a high unit thrust in a supersonic speed state, and the characteristics of a large bypass ratio turbofan engine with a low unit thrust, low noise and low fuel consumption in a subsonic speed state, variable cycle turbofan engines have been developed. However, the variable cycle structure is complex, the quantity of adjustable geometric parameters to be adjusted is large, and the design difficulty is extremely high. For example, a three-duct self-adaptive circulating turbofan engine with an FLADE fan, which is developed in the United states, is additionally provided with a third duct on the basis of the traditional variable circulating turbofan engine so as to realize the functions of greatly adjusting the flow, reducing the overflow resistance and cooling high-temperature components; the second-stage FLADE blade punctures the inner wall of a third duct and extends into the third duct, the front and the rear of the second-stage FLADE blade are respectively provided with a circle of adjustable guide vanes for adjusting the bypass air flow of the third duct, the problem of coordinated adjustment of adjustable guide vanes and other adjustable mechanisms is involved, and a series of sealing problems are caused after the FLADE blade punctures; the core machine is provided with a core machine driving fan, and large-range adjustment of the bypass ratio is realized by adjusting the rotating speed and the guide vane angle. The adjustable geometric mechanism brings great challenges to development, cooperative control and engine flight mode control, and simultaneously increases development cost and risks.
Furthermore, as the flight mach number increases, the engine requires a smaller ratio of inlet throat to inlet area. If the air inlet channel is designed according to a certain high Ma, the ratio of the throat area to the inlet area is relatively small, if the air inlet channel works by being smaller than the designed Ma, the throat area is too small, the throat blockage phenomenon occurs, the air flow entering the air inlet channel is limited, and a normal shock wave occurs in front of the air inlet channel, so that the overflow loss is very large. In order to reduce the total pressure loss, the overflow resistance, the additional resistance and the like on the performance and simplify the adjusting mechanism, the throat area of the air inlet channel needs to be geometrically adjustable along with the flight state of the airplane, so that the overflow resistance is avoided.
Disclosure of Invention
The technical problem to be solved by the invention is as follows:
aiming at the defects and shortcomings of the prior art, in order to realize that the engine has the characteristic of high unit thrust of a turbojet engine or a turbofan engine with a small bypass ratio in a supersonic speed state and has the characteristics of lower unit thrust, low noise and low oil consumption rate of the turbofan engine with a large bypass ratio in a subsonic speed state in the flying process of the airplane, the invention provides an advance-generation combined regulation turbojet propulsion system based on a front-mounted compression guide impeller, wherein the change of the bypass ratio of the engine and the redistribution of internal and external bypass flow are realized through the change of the compression or guide function working mode of the front-mounted compression guide impeller in the whole flying process of the airplane, so that the working cycle mode of the engine is changed, and the optimal combination effect of high thrust and low oil consumption is realized in different using stages of the airplane; meanwhile, the engine and the air inlet channel are designed in a fused mode and adjusted in a combined mode, the position of the air inlet flow guide cone is adjustable along with the flight state, namely the throat area of the air inlet channel is adjustable, and overflow resistance is effectively reduced.
(II) the technical scheme of the invention is as follows:
the technical scheme adopted by the invention for solving the technical problem is as follows:
a turbojet propulsion system based on a front-mounted compression guide impeller comprises a core machine, an air inlet channel and an afterburner, wherein the air inlet channel is arranged at the upstream of the core machine, the afterburner is arranged at the downstream of the core machine, the core machine comprises a fan, a high-pressure air compressor, a main combustion chamber, a high-pressure turbine and a low-pressure turbine which are sequentially arranged along the flow direction, the high-pressure turbine is connected with the high-pressure air compressor through a high-pressure shaft, the low-pressure turbine is connected with the fan through a low-pressure shaft, an air inlet guide cone is fixedly arranged at the axle center of the front end of the air inlet channel, an initial throat of the air inlet channel is formed at the position of the maximum cross section of the air inlet guide cone, the fan is arranged in the air inlet channel, the part of the air inlet channel, which is positioned at the downstream of the fan, is formed into an, the exhaust port of the outer duct is communicated with the air inlet of the afterburner, and the afterburner is characterized in that:
an outer culvert flow regulating valve is arranged at an exhaust port of the outer culvert and used for controlling the outer culvert to be opened or closed so as to realize variable culvert control;
the air inlet channel is also provided with a compression guide impeller, the compression guide impeller is positioned between the air inlet guide cone and the fan, the front end of a rotating shaft of the compression guide impeller is rotatably supported at the rear end of the air inlet guide cone through a bearing, the rear end of the rotating shaft of the compression guide impeller is connected with the rotating shaft of the fan through a coupler, and the rotating shaft of the compression guide impeller is controlled to be disconnected or combined with the rotating shaft of the fan through the coupler;
a rotor braking mechanism is further arranged on a rotating shaft of the compression guide vane wheel, and when the rotating shaft of the compression guide vane wheel is disconnected with the rotating shaft of the fan, the rotating speed of the compression guide vane wheel is reduced through the rotor braking mechanism until locking and braking are achieved;
the compression guide vane wheel is uniformly provided with a plurality of compression guide vanes with adjustable attack angles along the circumferential direction, when the compression guide vane wheel is adjusted to a large attack angle state, the compression guide vane wheel is used for compressing inlet air, and when the compression guide vane wheel is adjusted to a small attack angle state, the compression guide vane wheel is used for guiding the inlet air;
an annular resistance-reducing through-flow duct is coaxially arranged on the outer side of the air inlet channel, the resistance-reducing through-flow duct extends along the axial direction, an air outlet of the resistance-reducing through-flow duct is communicated with an air inlet of the afterburner, an air inlet of the resistance-reducing through-flow duct is communicated with the air inlet channel, the position of the air inlet of the resistance-reducing through-flow duct is located at the position of an initial throat of the air inlet channel in the axial direction, a resistance-reducing through-flow valve is arranged at the air inlet of the resistance-reducing through-flow duct, when the resistance-reducing through-flow valve is opened, the resistance-reducing through-flow duct is communicated with the air inlet channel, part; when the drag reduction vent valve is closed, the throat position of the inlet is unchanged and remains at the initial throat position.
Preferably, a radial flame stabilizer is arranged at an inlet at the front end of an air inlet of the afterburner, and a corrugated heat insulating material is adopted for a shell of the afterburner.
Preferably, the high-pressure compressor has 5-6 stages and the pressure ratio is 6-8.
Further, a guide vane adjusting mechanism is arranged in the high-pressure air compressor to match the change of the inlet working environment of the high-pressure air compressor.
Preferably, the main combustion chamber is a short annular combustion chamber, and the temperature of the outlet gas of the short annular combustion chamber is about 2150K-2200K, so that the fuel oil and the air are quickly mixed, and a uniform inlet temperature field is provided for the high-pressure turbine.
Preferably, the number of the fan stages is 2-3, and the pressure ratio is about 4. The fan and the compression guide impeller work in a matching mode under different working modes of the engine.
Preferably, the rotor braking mechanism is disposed adjacent the coupling.
Preferably, the shaft coupling adopts an end surface friction type shaft coupling, the rotating shaft of the fan gradually drives the compression guide impeller to increase the rotating speed in a mode that the friction force is gradually increased after the two end surfaces are contacted, and when the rotating speeds are consistent, the position is fixed. The pressure ratio of the compression guide vane wheel is about 1.5.
Preferably, when the coupling is opened, the compression guide vane is adjusted to a large attack angle state, and at the moment, the compression guide vane wheel is in a compression air inlet state; when the coupler is disconnected, the compression guide vane is adjusted to be in a small attack angle state, and at the moment, the compression guide vane wheel is in a flow guiding and air inlet state.
Preferably, when the pressure ratio needs to be increased, the compression guide vane is adjusted to a large attack angle state, and then the coupling is started to enable the compression guide vane wheel to be linked with the fan; when the pressure ratio needs to be reduced, the coupler is disconnected firstly, the compression guide vane wheel is separated from the fan, then the rotor brake mechanism is started, the rotating speed of the compression guide vane wheel is reduced gradually, and finally the compression guide vane is adjusted to be in a small attack angle state.
Preferably, the throat area ratio of the initial throat is greater than the throat area ratio of the extended throat, wherein the throat area ratio is the ratio of the throat area to the inlet area of the inlet.
Further, the throat area ratio of the initial throat ensures that the engine operates at acceptable efficiency in the range of 0-1.6M, and the throat area ratio of the extended throat ensures that the engine operates at acceptable efficiency in the range of 1.6-2.5M.
The invention provides a turbojet propulsion system based on a front-mounted compression guide impeller, which adopts a fused design and combined adjustment of an engine and an air inlet channel, and combines the adjustment of the throat area of the air inlet channel with the combined action control of an air inlet guide cone and the compression guide impeller of the engine, so as to reduce the total pressure loss and the additional resistance in performance.
In addition, in the turbojet propulsion system based on the front-mounted compression guide impeller, in order to simplify a control mechanism and reduce complexity, the air inlet channel adopts two-stage adjustability, and only the areas of two initial throats and two expansion throats are determined. The initial throat area ratio ensures that the engine can operate at acceptable efficiency in the range of 0-1.6M and the extended throat area ratio operates at acceptable efficiency in the range of 1.6-2.5M. The flow of the resistance-reducing through-flow duct is adjusted by opening or closing the resistance-reducing through-flow valve, the problem that the flow of the air inlet channel is incompatible with the flow of the engine is solved, and overflowing air flow enters the afterburner of the engine through the resistance-reducing through-flow duct to be mixed with fuel gas.
The turbojet propulsion system based on the front-mounted compression guide vane wheel adopts the compression guide vane wheel, and the angle of the compression guide vane wheel is adjustable and is controlled integrally with the actuation of the air inlet guide cone. In order to facilitate the realization of engineering, the adjustment of the attack angle of the compressed guide vane is carried out after the rotor is locked. In take-off and middle and low altitudes, the compressed guide vanes are used as guide vanes, and when the pressure ratio needs to be increased, the compressed guide vanes firstly move to a large attack angle state and then are linked with the fan rotor through the coupler. When the pressure ratio needs to be reduced, the compression guide vane wheel is separated from the fan rotor through the coupler, then the braking system is started, the rotating speed of the rotor is gradually reduced, and then the attack angle of the compression guide vane is adjusted. The gradual regulation mode can reduce the requirement of actuating energy to simplify an actuating mechanism and prevent surge caused by sharp change of the working environment of the engine. When the compression guide vane wheel is designed, the position clamping groove of the vane with a large attack angle is fixed at a position where aerodynamic force and mechanical force generated during high-speed rotation can be automatically locked, so that effective position is utilized for keeping.
Another object of the invention is to provide a method for controlling a turbojet propulsion system based on a front-mounted compressed air guiding impeller, characterized in that,
when the incoming air flow speed is not greater than the preset Mach number, the compression guide impeller is used as an inlet guide impeller, at the moment, the compression guide impeller is separated from a rotating shaft of the fan through the coupler, the rotating shaft of the compression guide impeller is locked by the brake mechanism, the inlet attack angle of the compression guide blade is adjusted to be in a small attack angle state to guide incoming air flow, at the moment, the resistance reduction through-flow valve is adjusted to be in a closed state, the throat of the air inlet channel is in an initial throat position, and the outer bypass regulating valve is adjusted to be in a large flow state to increase the flow capacity of the outer bypass, so that the engine works in a large bypass ratio oil-saving mode;
when the incoming air flow speed is larger than the preset Mach number, the compression guide impeller is used as a compression impeller, at the moment, the compression guide impeller is linked with a rotating shaft of the fan through the coupler, the compression guide vanes are adjusted to be in a large attack angle state and used for compressing incoming air flow, at the moment, the resistance reduction through-flow valve is adjusted to be in an open state, the throat of the air inlet channel is in an expansion throat position, the outer duct adjusting valve is adjusted to be in a small flow state, and the engine works in a small duct bypass ratio.
Preferably, when the incoming air flow speed is transited from being greater than the preset mach number to being not greater than the preset mach number, the coupler is disconnected firstly to disconnect the compression guide vane wheel from the fan, then the rotor braking mechanism is started to gradually reduce the rotating speed of the compression guide vane wheel, and finally the compression guide vane is adjusted to be in a small attack angle state.
Preferably, when the incoming airflow speed of the intake air is transited from being not greater than the preset mach number to being greater than the preset mach number, the compression guide vane is firstly adjusted to a large attack angle state, and then the coupling is started to enable the compression guide vane wheel to be linked with the fan.
Preferably, when the compression guide vane is in a small attack angle state, the air inlet attack angle of the compression guide vane is about 0 degree.
Preferably, the macrobypass bypass ratio is about 0.8-1.0.
Preferably, the mini-bypass ratio is about 0.2-0.3.
Preferably, the preset mach number is mach 1.6.
Preferably, the compression guide vane wheel is used as a compression vane wheel when the incoming flow velocity of the inlet air is in the range of Mach 1.6-2.5.
(III) compared with the prior art, the invention has the advantages that:
1) the turbojet propulsion system based on the front-mounted compression guide impeller adopts the fusion design and combined adjustment of the engine and the air inlet channel, the position of the air inlet guide cone is adjustable along with the flight state (namely the throat area of the air inlet channel is adjustable), and the overflow resistance is effectively reduced.
2) The turbojet propulsion system based on the front-mounted compression guide impeller adopts the external culvert channel to replace a third culvert of the conventional self-adaptive variable-circulation turbofan engine, and has relatively light weight and simple structure.
3) The turbojet propulsion system based on the front-mounted compression guide impeller simplifies the control logic and the actuating mechanism by integrating the air inlet channel and the engine compression guide multifunctional impeller and adopting a two-stage regulation control concept.
4) Compared with the existing three-duct self-adaptive variable-circulation turbofan engine, the area geometric adjusting variable of the turbojet propulsion system based on the front-mounted compression guide impeller provided by the invention has 5 less control variables, and is more beneficial to actuation coordination and control mode conversion.
5) The turbojet propulsion system based on the front-mounted compression guide impeller has a relatively simple core machine structure, can adopt a core machine of a traditional turbofan engine, does not have a core machine driving fan, and is relatively simple in pneumatic matching.
6) The turbojet propulsion system based on the front-mounted compression guide impeller is free of three-duct Flade fans, simplifies the mechanism, and does not involve the problems of sealing and the like.
7) The turbine jet propulsion system based on the front-mounted compression guide impeller has the advantages of large adjustable parameter adjustable range, simplified structure and large profit ratio.
Drawings
FIG. 1 is a schematic view of a forward compressed inducer based turbojet propulsion system of the present invention;
FIG. 2 is an enlarged partial schematic view of FIG. 1;
fig. 3 is another partially enlarged view of fig. 1.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present invention clearer, the technical solutions in the embodiments of the present invention will be described in more detail below with reference to the accompanying drawings in the embodiments of the present invention. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are only some, but not all embodiments of the invention. The embodiments described below with reference to the drawings are illustrative and intended to be illustrative of the invention and are not to be construed as limiting the invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention. Embodiments of the present invention will be described in detail below with reference to the accompanying drawings.
As shown in fig. 1-3, the turbojet propulsion system based on the front-mounted compression guide vane wheel of the present invention includes a core engine, an air inlet channel disposed at the upstream of the core engine, and an afterburner 14 disposed at the downstream of the core engine, wherein the core engine includes a fan 8, a high pressure compressor 9, a main combustor 10, a high pressure turbine 11, a low pressure turbine 12, and the like, which are sequentially disposed in a flow direction, the high pressure turbine 11 is connected to a rotating shaft of the high pressure compressor 9 through a high pressure shaft, the low pressure turbine 12 is connected to a rotating shaft of the fan 8 through a low pressure shaft, an air inlet guide cone 1 is fixedly disposed at an axial center of a front end of the air inlet channel, an initial throat 15 of the air inlet channel is formed at a position of a maximum cross section of the air inlet guide cone 1, the fan 8 is disposed in the air inlet channel, portions of the air inlet channel located at the downstream, the exhaust port of the outer duct is communicated with the air inlet of the afterburner 14, an outer duct flow regulating valve 13 is arranged at the exhaust port of the outer duct, and the outer duct flow regulating valve 13 is used for controlling the outer duct to be opened or closed so as to realize duct-variable control.
As shown in fig. 2, a compression guide vane wheel 5 is further disposed in the air inlet channel, the compression guide vane wheel 5 is located between the air inlet guide cone 1 and the fan 8, the front end of the rotating shaft of the compression guide vane wheel 5 is rotatably supported at the rear end of the air inlet guide cone 1 through a bearing 4, the rear end of the rotating shaft of the compression guide vane wheel 5 is connected with the rotating shaft of the fan 8 through a coupling 7, and the rotating shaft of the compression guide vane wheel 5 is controlled to be disconnected or combined with the rotating shaft of the fan 8 through the coupling 7; a rotor braking mechanism 6 is also arranged on the rotating shaft of the compression guide vane wheel 5, and when the rotating shaft of the compression guide vane wheel 5 is disconnected with the rotating shaft of the fan 8, the rotating speed of the compression guide vane wheel 5 is reduced through the rotor braking mechanism 6 until locking and braking are carried out; the compression guide vane wheel 5 is uniformly provided with a plurality of compression guide vanes with adjustable attack angles along the circumferential direction, when the compression guide vane wheel is adjusted to a large attack angle state, the compression guide vane wheel 5 is used for compressing inlet air, and when the compression guide vane wheel is adjusted to a small attack angle state, the compression guide vane wheel 5 is used for guiding the inlet air; an annular drag reduction through-flow duct 3 is coaxially arranged at the outer side of the air inlet channel, the drag reduction through-flow duct 3 extends along the axial direction, an air outlet of the drag reduction through-flow duct is communicated with an air inlet of a afterburner 14 (see figure 3), the air inlet of the drag reduction through-flow duct is communicated with the air inlet channel and is positioned at an initial throat 15 position of the air inlet channel in the axial direction, a drag reduction through-flow valve 2 is arranged at the air inlet of the drag reduction through-flow duct 3, when the drag reduction through-flow valve 2 is opened, the drag reduction through-flow duct 3 is communicated with the air inlet channel, part of air in the air inlet channel enters the drag reduction through-flow duct 3; when the drag reduction throughflow valve 2 is closed, the throat position of the inlet is unchanged, still at the initial throat 15 position.
In the turbojet propulsion system based on the front-mounted compression guide impeller, the air inlet channel and the engine adopt a fusion design, the adjustment of the throat area of the air inlet channel is controlled by the joint action of the air inlet guide cone 1 of the engine and the compression guide impeller 5, the flow of the resistance-reducing through-flow duct 3 can be adjusted by opening the resistance-reducing through-flow valve 2, the problem that the air inlet channel is incompatible with the flow of the engine is solved, and the overflowing air flow enters the engine afterburner 14 through the resistance-reducing through-flow duct 3 to be mixed with fuel gas. The air inlet flow guide cone 1 is provided with a resistance reduction flow duct 3, the normal shock wave position is adjusted by subsonic waves, the oblique shock wave position is adjusted by ultrasonic waves, the total pressure loss is reduced, and the additional resistance is reduced. And opening the resistance-reducing through-flow valve 2, adjusting the flow of the resistance-reducing through-flow duct 3, solving the problem that the flow of the air inlet channel is incompatible with the flow of the engine, and enabling the overflowing air flow to enter an engine afterburning chamber 14 through the resistance-reducing through-flow pipeline 3 to be mixed with the fuel gas. When the air inflow is large, the outer bypass adjusting valve 13 is opened, and when the air inflow reaches a certain threshold value, the outer bypass adjusting valve is in a large-flow state, so that the flow capacity of the outer bypass is increased.
In the turbojet propulsion system based on the front-mounted compression guide impeller, the bearing 4 can support the compression guide impeller 5 in a high-speed driving state of the coupler 7, and the compression guide impeller 5 can be stably linked with a rotating shaft of the fan 8 at a high speed on the premise of keeping the air inlet guide cone 1 from rotating.
In the turbojet propulsion system based on the front-mounted compression guide impeller, the compression guide impeller 5 is adjustable in angle and is in actuation integrated control with the air inlet guide cone 1, and when the pressure ratio needs to be increased, the compression guide impeller is firstly actuated to a large attack angle state and then is linked with a fan rotor through the coupler 7. When the pressure ratio needs to be reduced, the compression guide vane wheel 5 is firstly separated from the fan rotor through the coupler 7, then the rotor brake mechanism 6 is started, the rotating speed of the rotor is gradually reduced, and finally the angle of the compression guide vane is adjusted to be in a small attack angle state.
In the turbojet propulsion system based on the front-mounted compression guide impeller, the number of stages of the fan 8 is 2-3 stages, the pressure ratio is about 4, and the fan and the compression guide impeller 5 work in a matching mode under different working modes of an engine. The rotor brake mechanism 6 can lock and brake the compression guide vane wheel 5, so that the compression guide vane wheel 5 is in an air inlet guide state. The coupling 7 can realize the linkage of the compression guide impeller 5 and the fan rotor shaft, and the compression guide impeller 5 is in a compression air inlet flow state. The number of stages of the high-pressure compressor 9 is 5-6, the pressure ratio can be 6-8, the existing high-pressure compressor design technology can be adopted, the increase is not needed, and measures such as guide vane adjustment are needed to match the change of the inlet working environment. The temperature of the gas at the outlet of the main combustion chamber 10 reaches about 2150K-2200K, and the short annular combustion chamber is adopted, so that the fuel oil and the air can be rapidly mixed, and a uniform inlet temperature field is provided for the high-pressure turbine 11. The high-pressure turbine 11 is capable of extracting energy from the high-pressure combustion gas flow discharged from the main combustion chamber 10, and converting a part of the energy into shaft power to drive the high-pressure compressor 9. A low-pressure turbine 12 is located downstream of the high-pressure turbine 11, receiving the flow of hot combustion gases exiting from the core, the low-pressure turbine 12 converting the energy of the high-pressure combustion gases exiting from the high-pressure turbine 11 into shaft power, driving the fan 8.
In the turbojet propulsion system based on the front-mounted compression guide impeller, when the compression guide impeller 5 is used as an inlet guide vane, the outer culvert flow regulating valve 13 releases the impeller from a fan rotor shaft by virtue of the coupler 7, and the rotor braking system 6 locks the compression guide impeller 5, so that the engine with the air inlet guide function works in a large-culvert oil-saving mode; the compression guide impeller 5 is used as a compression impeller, the impeller is linked with a fan rotor shaft through a coupler 7, an outer duct adjusting valve 13 is in a low-flow state, and the engine works in a small duct mode. The afterburner 14 is designed using known technology wherein radial flame holders and corrugated insulation can be used.
The invention provides a turbojet propulsion system based on a front-mounted compression guide impeller, which adopts a fused design and combined adjustment of an engine and an air inlet channel, and combines the adjustment of the throat area of the air inlet channel with the combined action control of an air inlet guide cone and the compression guide impeller of the engine, so as to reduce the total pressure loss and the additional resistance in performance.
In the turbojet propulsion system based on the front-mounted compression guide impeller, in order to simplify a control mechanism and reduce complexity, the air inlet channel adopts two-stage adjustability, and only the areas of two initial throats and two expansion throats are determined. The initial throat area ratio ensures that the engine can operate at acceptable efficiency in the range of 0-1.6M and the extended throat area ratio operates at acceptable efficiency in the range of 1.6-2.5M. The flow of the resistance-reducing through-flow duct is adjusted by opening or closing the resistance-reducing through-flow valve, the problem that the flow of the air inlet channel is incompatible with the flow of the engine is solved, and overflowing air flow enters the afterburner of the engine through the resistance-reducing through-flow duct to be mixed with fuel gas.
The turbojet propulsion system based on the front-mounted compression guide vane wheel adopts the compression guide vane wheel, and the angle of the compression guide vane wheel is adjustable and is controlled integrally with the actuation of the air inlet guide cone. In order to facilitate the realization of engineering, the adjustment of the attack angle of the compressed guide vane is carried out after the rotor is locked. In take-off and middle and low altitudes, the compressed guide vanes are used as guide vanes, and when the pressure ratio needs to be increased, the compressed guide vanes firstly move to a large attack angle state and then are linked with the fan rotor through the coupler. When the pressure ratio needs to be reduced, the compression guide vane wheel is separated from the fan rotor through the coupler, then the braking system is started, the rotating speed of the rotor is gradually reduced, and then the attack angle of the compression guide vane is adjusted. The gradual regulation mode can reduce the requirement of actuating energy to simplify an actuating mechanism and prevent surge caused by sharp change of the working environment of the engine. When the compression guide vane wheel is designed, the position clamping groove of the vane with a large attack angle is fixed at a position where aerodynamic force and mechanical force generated during high-speed rotation can be automatically locked, so that effective position is utilized for keeping.
When the turbine jet propulsion system based on the front-mounted compression guide impeller is used, the following control strategy is adopted:
when the incoming air flow speed is not greater than the preset mach number, for example, when the incoming air flow speed is not greater than mach 1.6, the compression guide vane wheel 4 is used as an inlet guide vane wheel, at this time, the compression guide vane wheel 5 is separated from a rotating shaft of the fan 8 through the coupler 7, the rotating shaft of the compression guide vane wheel 5 is locked by the rotor brake mechanism 6, the inlet attack angle of the compression guide vane wheel is adjusted to be in a small attack angle state of about 0 degree so as to guide the incoming air flow, at this time, the resistance reduction through-flow valve 2 is adjusted to be in a closed state, the throat of the air inlet is in the position of the initial throat 15, the outer bypass adjusting valve 13 is adjusted to be in a large flow state so as to increase the flow capacity of the outer bypass, and the engine works in an oil saving mode with a large bypass ratio (the bypass ratio is about.
When the incoming air flow speed is greater than the preset mach number, for example, at mach 1.6-2.5, the compression guide vane wheel 5 is used as a compression vane wheel, at this time, the compression guide vane wheel 5 is linked with a rotating shaft of the fan 8 through the coupler 7, the compression guide vane wheel is adjusted to a large attack angle state to compress the incoming air flow, at this time, the drag reduction through-flow valve 2 is adjusted to an open state, the throat of the air inlet channel is at the position of the expansion throat 16, the outer duct adjusting valve 13 is adjusted to a small flow state, and the engine works with a small duct bypass ratio (the duct ratio is about 0.2-0.3).
When the incoming air flow speed is transited from being greater than the preset Mach number to being not greater than the preset Mach number, the coupler 7 is disconnected firstly, the compression guide vane wheel 5 is separated from the fan 8, then the rotor brake mechanism 6 is started, the rotating speed of the compression guide vane wheel 5 is gradually reduced, and finally the compression guide vane is adjusted to be in a small attack angle state. When the incoming air flow speed is changed from being not more than the preset Mach number to being more than the preset Mach number, the compression guide vane is firstly adjusted to be in a large attack angle state, and then the coupler 7 is started, so that the compression guide vane wheel 5 is linked with the fan 8. The invention relates to an inlet-outlet joint adjusting turbojet propulsion system based on a front-mounted compression guide impeller, which adopts a front-mounted compression guide adjustable impeller to adjust the pressure ratio of an engine, thereby realizing the cyclic change of the engine in the whole envelope range, the effective adjustment of the bypass ratio, the large-amplitude adjustment and redistribution of the internal and external bypass power; through the combined adjustment design of the engine and the air inlet, the area ratio of the throat of the air inlet works within an acceptable efficiency range, and the matching of the air inlet flow and the working flow of a main engine is realized, so that the thrust loss is reduced, and the propulsion efficiency is improved. When the incoming air flow speed is below 1.6Ma, the compressed guide vane wheel is used as a guide vane, the fan is in a small pressure ratio mode, and the area of the throat of the air inlet channel is in a small state by moving the position of the air inlet cone of the engine, so that the bypass ratio is effectively increased, and the oil consumption rate is reduced; when the incoming flow speed of the inlet air is 1.6-2.5Ma, the compression guide vane wheel is used as an air compressor, the throat area of an air inlet channel is in a larger state by moving the position of an air inlet cone of the engine, the fan pressure ratio is increased, the bypass ratio is reduced, the temperature of the gas in front of the turbine is increased, and the unit thrust is improved. The multi-working-mode turbojet propulsion system can realize effective regulation of thrust and oil consumption in the whole envelope range through matching regulation of air inflow and working flow of a host and bypass ratio regulation, and effectively reduces installation oil consumption rate. Compared with the traditional fixed cycle engine, the non-installation oil consumption rate is reduced by about 11% in the takeoff and low-altitude climb sections; in the high-altitude climbing and subsonic cruising period, the non-installation oil consumption rate is reduced by about 7 percent; if the supersonic cruise is operated at a large bypass ratio, the non-installation oil consumption rate is reduced by 5 percent, otherwise, the thrust is increased by about 25 percent in a mode with a large bypass ratio. Compared with the self-adaptive circulation turbofan engine in the United states, the multi-working-mode turbojet propulsion system has the advantages that the task function is equivalent, and the structure of the engine is simplified.
The above-mentioned embodiments are only specific embodiments of the present invention, which are used for illustrating the technical solutions of the present invention and not for limiting the same, and the protection scope of the present invention is not limited thereto, although the present invention is described in detail with reference to the foregoing embodiments, those skilled in the art should understand that: any person skilled in the art can modify or easily conceive the technical solutions described in the foregoing embodiments or equivalent substitutes for some technical features within the technical scope of the present disclosure; such modifications, changes or substitutions do not depart from the spirit and scope of the embodiments of the present invention, and they should be construed as being included therein. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.
Claims (10)
1. A turbojet propulsion system based on a front-mounted compression guide impeller comprises a core machine, an air inlet channel and an afterburner, wherein the air inlet channel is arranged at the upstream of the core machine, the afterburner is arranged at the downstream of the core machine, the core machine comprises a fan, a high-pressure air compressor, a main combustion chamber, a high-pressure turbine and a low-pressure turbine which are sequentially arranged along the flow direction, the high-pressure turbine is connected with the high-pressure air compressor through a high-pressure shaft, the low-pressure turbine is connected with the fan through a low-pressure shaft, an air inlet guide cone is fixedly arranged at the axle center of the front end of the air inlet channel, an initial throat of the air inlet channel is formed at the position of the maximum cross section of the air inlet guide cone, the fan is arranged in the air inlet channel, the part of the air inlet channel, which is positioned at the downstream of the fan, is formed into an, the exhaust port of the outer duct is communicated with the air inlet of the afterburner,
the method is characterized in that:
an outer culvert flow regulating valve is arranged at an exhaust port of the outer culvert and used for controlling the outer culvert to be opened or closed so as to realize variable culvert control;
the air inlet channel is also provided with a compression guide impeller, the compression guide impeller is positioned between the air inlet guide cone and the fan, the front end of a rotating shaft of the compression guide impeller is rotatably supported at the rear end of the air inlet guide cone through a bearing, the rear end of the rotating shaft of the compression guide impeller is connected with the rotating shaft of the fan through a coupler, and the rotating shaft of the compression guide impeller is controlled to be disconnected or combined with the rotating shaft of the fan through the coupler;
a rotor braking mechanism is further arranged on a rotating shaft of the compression guide vane wheel, and when the rotating shaft of the compression guide vane wheel is disconnected with the rotating shaft of the fan, the rotating speed of the compression guide vane wheel is reduced through the rotor braking mechanism until locking and braking are achieved;
the compression guide vane wheel is uniformly provided with a plurality of compression guide vanes with adjustable attack angles along the circumferential direction, when the compression guide vane wheel is adjusted to a large attack angle state, the compression guide vane wheel is used for compressing inlet air, and when the compression guide vane wheel is adjusted to a small attack angle state, the compression guide vane wheel is used for guiding the inlet air;
an annular resistance-reducing through-flow duct is coaxially arranged on the outer side of the air inlet channel, the resistance-reducing through-flow duct extends along the axial direction, an air outlet of the resistance-reducing through-flow duct is communicated with an air inlet of the afterburner, an air inlet of the resistance-reducing through-flow duct is communicated with the air inlet channel, the position of the air inlet of the resistance-reducing through-flow duct is located at the position of an initial throat of the air inlet channel in the axial direction, a resistance-reducing through-flow valve is arranged at the air inlet of the resistance-reducing through-flow duct, when the resistance-reducing through-flow valve is opened, the resistance-reducing through-flow duct is communicated with the air inlet channel, part; when the drag reduction vent valve is closed, the throat position of the inlet is unchanged and remains at the initial throat position.
2. Turbojet propulsion system according to the preceding claim, characterized in that: the afterburner is characterized in that a radial flame stabilizer is arranged at an inlet at the front end of an air inlet of the afterburner, and a shell of the afterburner is made of a corrugated heat-insulating material.
3. Turbojet propulsion system according to the preceding claim, characterized in that: the high-pressure compressor has 5-6 stages and a pressure ratio of 6-8.
4. A turbojet propulsion system according to claim 3, characterized in that: the high-pressure compressor is provided with a guide vane adjusting mechanism to match the change of the working environment of the inlet of the high-pressure compressor.
5. Turbojet propulsion system according to the preceding claim, characterized in that: the main combustion chamber adopts a short annular combustion chamber, the temperature of the outlet gas of the short annular combustion chamber is about 2150K-2200K, so that the fuel oil and the air are rapidly mixed, and a uniform inlet temperature field is provided for the high-pressure turbine.
6. Turbojet propulsion system according to the preceding claim, characterized in that: the number of stages of the fan is 2-3 stages, and the pressure ratio is about 4.
7. Turbojet propulsion system according to the preceding claim, characterized in that: the rotor brake mechanism is arranged close to the coupler.
8. Turbojet propulsion system according to the preceding claim, characterized in that: the shaft coupling adopts an end face friction type shaft coupling, the rotating shaft of the fan gradually drives the compression guide impeller to increase the rotating speed in a mode that the friction force is gradually increased after the two end faces are contacted, and when the rotating speeds are consistent, the position is fixed.
9. Turbojet propulsion system according to the preceding claim, characterized in that: the pressure ratio of the compression guide vane wheel is about 1.5.
10. Turbojet propulsion system according to the preceding claim, characterized in that: when the coupler is started, the compression guide vane is adjusted to a large attack angle state, and the compression guide vane wheel is in a compression air inlet state; when the coupler is disconnected, the compression guide vane is adjusted to be in a small attack angle state, and at the moment, the compression guide vane wheel is in a flow guiding and air inlet state.
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Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112814949A (en) * | 2021-01-13 | 2021-05-18 | 南京航空航天大学 | Oblique-section type double-freedom-degree inlet adjustable guide vane suitable for wide bypass ratio variation range |
| CN113565649A (en) * | 2021-08-16 | 2021-10-29 | 中国科学院工程热物理研究所 | Three-duct ACE engine with Flade fan and core machine driving fan stages |
| WO2021249184A1 (en) * | 2020-06-08 | 2021-12-16 | 清华大学 | Two-duct turbojet engine with bearing cooling function |
| CN113864082A (en) * | 2021-09-13 | 2021-12-31 | 上海新云彩航空科技有限责任公司 | Aviation jet engine |
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4446696A (en) * | 1981-06-29 | 1984-05-08 | General Electric Company | Compound propulsor |
| CN101021181A (en) * | 2006-02-13 | 2007-08-22 | 通用电气公司 | Double bypass turbofan |
| EP2177735A2 (en) * | 2008-10-20 | 2010-04-21 | Rolls-Royce North American Technologies, Inc. | Turbofan |
| CN102155331A (en) * | 2011-05-05 | 2011-08-17 | 西北工业大学 | Turboramjet combined engine based on knocking combustion |
| US9605557B1 (en) * | 2013-04-30 | 2017-03-28 | United States Of America As Represented By The Secretary Of The Air Force | Variable bypass turbofan engine |
| CN109854378A (en) * | 2019-04-08 | 2019-06-07 | 沈阳建筑大学 | A kind of turbofan with mode tuning ability |
-
2020
- 2020-01-03 CN CN202010004512.0A patent/CN111102098B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4446696A (en) * | 1981-06-29 | 1984-05-08 | General Electric Company | Compound propulsor |
| CN101021181A (en) * | 2006-02-13 | 2007-08-22 | 通用电气公司 | Double bypass turbofan |
| EP2177735A2 (en) * | 2008-10-20 | 2010-04-21 | Rolls-Royce North American Technologies, Inc. | Turbofan |
| CN102155331A (en) * | 2011-05-05 | 2011-08-17 | 西北工业大学 | Turboramjet combined engine based on knocking combustion |
| US9605557B1 (en) * | 2013-04-30 | 2017-03-28 | United States Of America As Represented By The Secretary Of The Air Force | Variable bypass turbofan engine |
| CN109854378A (en) * | 2019-04-08 | 2019-06-07 | 沈阳建筑大学 | A kind of turbofan with mode tuning ability |
Cited By (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2021249184A1 (en) * | 2020-06-08 | 2021-12-16 | 清华大学 | Two-duct turbojet engine with bearing cooling function |
| CN112814949B (en) * | 2021-01-13 | 2022-01-21 | 南京航空航天大学 | Oblique-section type double-freedom-degree inlet adjustable guide vane suitable for wide bypass ratio variation range |
| CN112814949A (en) * | 2021-01-13 | 2021-05-18 | 南京航空航天大学 | Oblique-section type double-freedom-degree inlet adjustable guide vane suitable for wide bypass ratio variation range |
| CN113565649A (en) * | 2021-08-16 | 2021-10-29 | 中国科学院工程热物理研究所 | Three-duct ACE engine with Flade fan and core machine driving fan stages |
| CN113864082B (en) * | 2021-09-13 | 2022-12-06 | 上海新云彩航空科技有限责任公司 | Aviation jet engine |
| CN113864082A (en) * | 2021-09-13 | 2021-12-31 | 上海新云彩航空科技有限责任公司 | Aviation jet engine |
| CN114278617A (en) * | 2021-11-30 | 2022-04-05 | 清华大学 | Compression device and runner system thereof |
| CN114607525A (en) * | 2022-04-02 | 2022-06-10 | 中北大学 | Hydrogen fuel rotary ramjet turbofan engine |
| CN114623019A (en) * | 2022-05-16 | 2022-06-14 | 西安星云航空科技有限公司 | Large-bypass-ratio split type variable-circulation turbofan engine |
| CN114623019B (en) * | 2022-05-16 | 2022-07-19 | 西安星云航空科技有限公司 | Large-bypass-ratio split type variable-circulation turbofan engine |
| CN117662324A (en) * | 2023-12-07 | 2024-03-08 | 中国航空发动机研究院 | Vortex shaft engine, aircraft and method |
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