CN113864082B

CN113864082B - Aviation jet engine

Info

Publication number: CN113864082B
Application number: CN202111069907.XA
Authority: CN
Inventors: 邓云娣; 梅柯; 王国栋; 邸京京; 陈鹏; 门景龙
Original assignee: Shanghai Xinyuncai Aviation Technology Co ltd
Current assignee: Shanghai Xinyuncai Aviation Technology Co ltd
Priority date: 2021-09-13
Filing date: 2021-09-13
Publication date: 2022-12-06
Anticipated expiration: 2041-09-13
Also published as: CN113864082A

Abstract

The invention relates to the technical field of aero-engines, in particular to an aero-jet engine, which comprises a shell, wherein the shell comprises an air inlet end and an air outlet end, and the interior of the shell is sequentially provided with the following components in the direction from the air inlet end to the air outlet end: the invention has the advantages of ingenious design and novel structure, ensures that an engine can obtain high thrust-weight ratio or low fuel consumption rate, can simultaneously meet the power requirements of high maneuverability, high navigational speed and long navigational range of an airplane, and has remarkable popularization significance.

Description

Aviation jet engine

Technical Field

The invention relates to the technical field of aero-engines, in particular to an aero-jet engine.

Background

At present, an aviation jet engine mainly adopts a turbofan engine, civil airplanes mainly adopt the turbofan engine with a large bypass ratio, military airplanes mainly adopt the booster turbofan engine with a small bypass ratio so as to meet the requirements of economy and high navigational speed, and only a few fighters which emphasize high-speed performance adopt the turbojet engine. Although the power of the turbojet engine is strong, the fuel consumption rate is high, and the requirement of a high-performance warplane for a large combat radius cannot be met. In order to reduce oil consumption, an outer duct is added on the basis of a turbojet engine to form a turbofan engine, and the turbofan engine becomes the mainstream of the current aircraft engine. However, although the oil consumption of the turbofan engine is low, the unit thrust density is low, the thrust weight is low, and although the afterburner is added, the instantaneous thrust is improved so as to meet the requirements of high-performance aircraft takeoff and air high-mobility operation. However, because the oxygen content of the gas entering the afterburner is low, the air pressure is far lower than that of the main combustion chamber, the speed is also as high as 120-180m/s, the speed is far higher than that of the main combustion chamber, the combustion is unstable, the combustion is insufficient, in order to improve the combustion efficiency, the afterburner is usually designed to be longer, the factors increase the complexity of the design of the afterburner, and the structure weight is larger. In addition, the operating temperature of the afterburner is high, so that the engine cannot work for a long time, and therefore, even if the afterburner is installed, the requirement of supersonic cruising of a high-performance fighter cannot be met.

Patent No. CN112483275A, a Propeller and aircraft, discloses a jet engine with double combustion chambers and no shaft turbine drive. The rear part of the high-pressure compressor is divided into an outer duct, wherein a first combustion chamber is arranged in the inner duct and is mainly used for driving a shaftless turbine to drive the shaftless compressor to work; the second combustion chamber is arranged in the outer duct, the gas of the second combustion chamber is consistent with the gas of the first combustion chamber, the gas is fresh high-pressure air, the combustion is stable and sufficient at speed, the combustion efficiency is high, and the supersonic cruise device is mainly used for generating continuous thrust and meeting the requirement of supersonic cruise. When the second combustion chamber does not work, the outer duct is equivalent to the outer duct of the turbofan engine to generate thrust, and at the moment, the oil consumption of the engine is low, so that the low-speed and low-oil-consumption patrol flight requirement is met. The design can obtain continuous high thrust theoretically, has high thrust-weight ratio and low oil consumption rate, and meets the requirements of supersonic cruise and low-speed economical patrol of high-performance fighters. However, the method is in a concept stage, so that the development difficulty is high, the period is long, and the method cannot be put into practical application in a short time.

Therefore, an aviation jet engine with high thrust-weight ratio, low oil consumption rate and relatively mature technology needs to be developed.

Disclosure of Invention

In view of the above technical problems, an object of the present invention is to provide an aviation jet engine, which can obtain a high thrust-weight ratio or a low fuel consumption rate by redesigning an engine structure, and can simultaneously meet the power requirements of high maneuverability, high navigational speed and long range of an aircraft, and the technical scheme adopted by the present invention is as follows:

the utility model provides an aviation jet engine, includes the casing, the casing includes inlet end and exhaust end, the inside of casing is installed in proper order along inlet end to exhaust end direction:

the air inlet system comprises an air inlet channel, and air flows into the interior of the engine through the air inlet channel;

a compressor system that decelerates and supercharges air flowing into the engine via an intake;

the bypass system comprises an inner bypass and a first outer bypass, and air pressurized by the compressor system respectively enters the inner bypass and the first outer bypass; a first duct shunting control valve is arranged at the position, close to the compressor system, of the first outer duct;

a combustion chamber system comprising a first combustion chamber disposed inside the inner duct and a second combustion chamber disposed inside the first outer duct;

the turbine system is arranged in the inner duct, and the high-temperature gas ejected from the first combustion chamber drives the turbine system to operate;

the tail spraying system comprises an inner duct tail nozzle and an outer duct tail nozzle, high-temperature gas from the first combustion chamber and passing through the turbine system is sprayed out from the inner duct tail nozzle, high-temperature gas from the second combustion chamber is sprayed out from the outer duct tail nozzle,

a drive shaft system for connecting the turbine system and the compressor system.

Preferably, the second combustion chamber is one of a single-tube combustion chamber, a circular-tube combustion chamber or a circular combustion chamber.

Preferably, the second combustion chamber is a single-tube combustion chamber, and is arranged in the first outer duct along the circumferential direction, and each single-tube combustion chamber works independently.

Preferably, the air inlet system further comprises an air inlet fairing arranged at the front end of the compressor system and used for reducing air resistance generated by the compressor system.

Preferably, the compressor system comprises a low-pressure compressor and a high-pressure compressor, a second outer duct is directly connected between the low-pressure compressor and the exhaust end of the shell, and a second duct flow dividing control valve is arranged on the second outer duct at a position close to the low-pressure compressor.

Preferably, the external engine mixing chamber is communicated with an external engine fuel oil pipeline.

Preferably, still include external engine exhaust nozzle cooling system, external engine exhaust nozzle cooling system set up in the outer wall of external engine combustion chamber and the outer wall of external engine exhaust nozzle, external engine exhaust nozzle cooling system respectively with high pressure outer bleed pipe external engine mixing chamber intercommunication, the high-pressure air that bleed pipe flowed in outside from high pressure it is right in the external engine exhaust nozzle cooling system external engine combustion chamber with the pipe wall of external engine exhaust nozzle cools down, takes away the heat, flows out again external engine mixing chamber.

Preferably, the outboard engine deflection device is further included, the outboard engine is mounted on the outboard engine deflection device, and the direction of the outboard engine exhaust nozzle is adjusted by the outboard engine deflection device, so that the thrust direction of the outboard engine is further controlled.

Preferably, a stamping duct is communicated between the air inlet duct and the first outer duct, a stamping duct air inlet valve is arranged at the joint of the stamping duct and the air inlet duct, and a stamping duct exhaust valve is arranged at the joint of the stamping duct and the first outer duct.

Preferably, the air inlet system further comprises an adjustable shock wave cone of the air inlet and a shock wave cone supporting structural member, the adjustable shock wave cone of the air inlet is installed at the front end of the air compressor system through the shock wave cone supporting structural member, and the shock wave cone supporting structural member is a telescopic structural member.

Compared with the prior art, the invention has the beneficial effects that:

(1) The invention adds the outer duct and the second combustion chamber behind the air compressor of the prior turbojet engine, so that all the combustion chambers of the engine are in the best working state and are fully combusted, thereby obtaining larger additional thrust, lower oil consumption and longer continuous working time than an afterburner, and leading an airplane loaded with the engine to obtain supersonic cruising ability;

(2) The scheme of adding the external engine is equivalent to that one core engine with the compressor drives a plurality of external engines only comprising combustion chambers and a tail nozzle system to work simultaneously, so that the effect that the plurality of engines work simultaneously to generate thrust in different directions is obtained at the cost of smaller structural weight, and the external engine system is suitable for being used for vertical take-off and landing airplanes;

(3) The scheme of adding the ramjet duct enables the engine to have the capability of freely switching between the normal jet mode and the sub-combustion ramjet mode, so that the airplane loaded with the engine can be accelerated to Mach 2 through the normal jet mode in the takeoff stage and then switched to the sub-combustion ramjet mode, and the airplane can fly at supersonic speed and even hypersonic speed.

Drawings

The above features, technical features, advantages and modes of implementing the present invention will be further described in the following detailed description of preferred embodiments in a clearly understandable manner by referring to the accompanying drawings.

FIG. 1 is a side view of an aircraft jet engine according to a first embodiment of the invention;

FIG. 2 is a schematic view of a second outer duct of an aero-jet engine in a closed state according to a first embodiment of the present invention;

FIG. 3 is a schematic view of an aircraft jet engine with a second outer duct open according to a first embodiment of the invention;

FIG. 4 is a schematic view of a first version of an aero jet engine with an external engine according to a second embodiment of the present invention;

FIG. 5 is a schematic view of a second version of an aircraft jet engine with an external engine according to a second embodiment of the invention;

FIG. 6 is a schematic view of a third version of an aircraft jet engine with an external engine according to a second embodiment of the invention;

FIG. 7 is a schematic view of a third embodiment of the invention showing the normal turbine operating state of an aero jet engine with a ramjet duct;

fig. 8 is a schematic view of the ramjet operation state of the aero-jet engine with the ramjet duct according to the third embodiment of the present invention.

The reference numbers indicate:

10. an air inlet system; 11. an air inlet channel; 12. an inlet duct cowling; 13. a compressor front air guide vane; 14. the air inlet channel can adjust a shock wave cone; 15. a shock cone supporting structure; 16. stamping a duct air inlet valve;

20. a compressor system; 21. a compressor casing; 211. a low pressure compressor case; 212. a high pressure compressor case; 22. compressor rotor blades; 221. a low pressure compressor rotor blade; 222. high pressure compressor rotor blades; 23. compressor stator blades; 231. stator blades of a low-pressure compressor; 232. stator blades of a high-pressure compressor;

30. a ducted system; 31. a high-pressure air guide blade; 32. a bypass shunt control valve; 321. a first duct flow-dividing control valve; 322. a second duct flow-dividing control valve; 33. an inner duct; 34. an outer duct; 341. a first outer duct; 342. a second outer duct; 35. a gas guide vane; 38. stamping a duct; 39. stamping a duct exhaust valve;

40. a combustion chamber system; 41. a first combustion chamber; 42. a first combustion chamber casing; 43. a second combustion chamber;

50. a turbine system; 51. a turbine rotor blade; 511. high pressure turbine rotor blades; 512. low pressure turbine rotor blades; 52. turbine stator blades; 521. high pressure turbine stator vanes; 522. low pressure turbine stator vanes; 53. a turbine case; 531. a high pressure turbine case; 532. a low pressure turbine case;

60. a tail spray system; 61. a gas guide vane; 62. a gas dome; 63. the inner duct tail nozzle; 64. the tail nozzle of the outer duct;

70. a drive shaft system; 71. a drive shaft; 711. a low-pressure transmission shaft; 712. a high-pressure transmission shaft; 72. a drive bearing;

80. a housing; 810. an air inlet end; 820. an exhaust end;

90. an outboard engine; 91. an external engine mixing chamber; 92. an external engine combustion chamber; 93. an external engine tail nozzle; 94. a high pressure external bleed air pipe; 95. a high pressure external bleed air pipe valve; 96. an engine fuel oil pipeline is externally arranged; 97. an external engine tail nozzle cooling system; 98. an external engine deflection device.

Detailed Description

In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following description will be made with reference to the accompanying drawings. It is obvious that the drawings in the following description are only some examples of the invention, and that for a person skilled in the art, other drawings and embodiments can be derived from them without inventive effort.

For the sake of simplicity, only the parts relevant to the invention are schematically shown in the drawings, and they do not represent the actual structure as a product. Moreover, in the interest of brevity and understanding, only one of the components having the same structure or function is illustrated schematically or designated in some of the drawings. In this document, "a" means not only "only one of this but also a case of" more than one ".

It should be further understood that the term "and/or" as used in this specification and the appended claims refers to any and all possible combinations of one or more of the associated listed items and includes such combinations.

In this context, it is to be understood that, unless otherwise explicitly stated or limited, the terms "mounted," "connected," and "connected" are to be construed broadly and may include, for example, a fixed connection, a removable connection, or an integral connection; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.

In addition, in the description of the present application, the terms "first," "second," and the like are used only for distinguishing the description, and are not intended to indicate or imply relative importance.

Example 1

Referring to fig. 1 to 3, the present invention provides an aircraft jet engine, which includes a housing 80, wherein the housing 80 is disposed at the periphery of the engine, provides an installation space for each system, and provides a fixed joint for installing the engine on an aircraft. The casing 80 comprises an air inlet end 810 and an air outlet end 820, and the air inlet system 10, the compressor system 20, the duct system 30, the combustion chamber system 40, the turbine system 50, the tail injection system 60 and the transmission shafting 70 are sequentially installed in the casing along the direction from the air inlet end 810 to the air outlet end 820.

The air inlet system 10 is arranged inside an inner cavity of the housing 80 and comprises an air inlet 11, the air inlet 11 provides an air inlet channel for the engine, and air flows into the engine through the air inlet 11;

as a preferred embodiment, the air intake duct system 10 in this embodiment further includes an air intake duct fairing 12 disposed at the front end of the compressor system 20, and is configured to reduce air resistance generated by the compressor system 20, the air intake duct system 10 further includes a compressor front air guide vane 13, the compressor front air guide vane 13 is disposed at the front end of the compressor system 20, and is connected to the housing 80 and the air intake duct fairing 12, and when the air intake duct fairing 12 is fixedly supported, the air flowing into the compressor system 20 from the air intake duct is guided, the air flow direction is changed, and the air intake duct is decelerated and pressurized, so as to improve the working efficiency of the compressor system 20.

The compressor system 20 is arranged on the inner side of an inner cavity of the shell 80 and used for decelerating and pressurizing air flowing into the engine through the air inlet 11, the compressor system 20 is arranged between the air inlet system 10 and the bypass system 30, and the compressor system 20 comprises a compressor casing 21 arranged on the transmission shaft system 70, compressor rotor blades 22 arranged on the outer side of the compressor casing 21 and compressor stator blades 23 arranged on the inner side of the shell 80. When the transmission shaft system 70 rotates, the compressor casing 21 drives the compressor rotor blades 22 to rotate, so as to push the air entering from the air inlet system 10 to move towards the rear side, increase the air flow rate, and convert the mechanical energy of the transmission system 10 into the kinetic energy of the air; when air accelerated by the compressor rotor blades 22 flows through the compressor stator blades 23, the direction of the air flow is changed, and meanwhile, the flow speed is reduced, and the pressure intensity and the temperature are increased, so that the kinetic energy of the air is converted into the internal energy of the air. The air flowing from the air inlet system 10 is pressurized by the compressor system 20 to form high-pressure air, and then flows into the ducted system 30.

In a preferred embodiment, the compressor system 20 in this embodiment further includes a low-pressure compressor 21 and a high-pressure compressor 22, the low-pressure compressor 21 is disposed near the rear end of the air intake duct 11, and the high-pressure compressor 22 is disposed between the low-pressure compressor 21 and the combustion chamber system 40. The specific connection and function of the low-pressure compressor 21 and the high-pressure compressor 22 are the same as those of the compressor system 20, and will not be described in detail. Preferably, the direction of rotation of the low-pressure compressor 21 is opposite to the direction of rotation of the high-pressure compressor 22.

The ducted system 30 is arranged inside the inner cavity of the casing 80, the ducted system 30 comprises an inner duct 33 and a first outer duct 341, the inner duct 33 is arranged between the compressor system 20 and the turbine system 50, and one side of the inner cavity of the casing 80 close to the transmission shafting 70; the first bypass 341 is disposed between the compressor system 20 and the tail injection system 60, between the inner cavity of the casing 80 and the bypass 33.

In a preferred embodiment, the bypass system 30 in this embodiment further includes a high-pressure air guide vane 31 disposed at the rear end of the compressor system 20, and is used for adjusting the direction of the high-pressure air flow from the compressor system 20, and further decelerating and pressurizing the high-pressure air flow, so as to increase the internal energy of the high-pressure air flow.

In a preferred embodiment, the ducted system 30 in this embodiment further includes a gas guide vane 35 disposed at the front end of the turbine system 50, and is used for changing the direction of the gas flow flowing out from the combustion chamber system 40, reducing the pressure and increasing the speed, converting the internal energy of the high-temperature and high-pressure gas into kinetic energy, and improving the working efficiency of the turbine system 50.

In a preferred embodiment, the bypass system 30 in this embodiment further includes a first bypass split control valve 321 disposed at the rear end of the high pressure air guide vane 31, for adjusting the proportional relationship of the high pressure air flowing into the inner bypass 33 and the first outer bypass 341, so as to make the combustion chamber system 40 of the engine in the target working state.

In a preferred embodiment, the bypass system 30 in this embodiment further includes a second bypass 342 disposed between the compressor system 20 and the tail portion of the engine and outside the first bypass 341, and is configured to bleed air directly from the low-pressure compressor 21 of the compressor system 20 to the tail portion of the engine, and eject the bleed air out of the engine to generate thrust.

In a preferred embodiment, the bypass system 30 in this embodiment further includes a second bypass flow-dividing control valve 322 disposed at the rear end of the low-pressure air guide vane 31, and is used for adjusting the proportional relationship between the air flowing into the high-pressure compressor 22 and the second bypass 341, so as to enable the fuel efficiency of the engine to be in a high state.

The combustor system 40 includes a first combustor casing 42 disposed between the compressor system 20 and the turbine system 50, a first combustor 41 disposed within the inner duct 33, and a second combustor 43 disposed within the first outer duct 341. A part of the high-pressure air from the compressor system 20 enters the first combustion chamber 41 through the inner duct 33, and is fully combusted with the fuel to form high-temperature and high-pressure fuel gas, and the high-temperature and high-pressure fuel gas is accelerated in the inner duct 33 to convert the internal energy of the fuel gas into the kinetic energy of the fuel gas to form high-temperature, high-pressure and high-speed fuel gas, and the high-temperature, high-pressure and high-speed fuel gas flows into the turbine system 50; the other part enters the second combustion chamber 43 through the first outer duct 341, and is fully combusted with the fuel to form high-temperature and high-pressure fuel gas, which flows into the tail injection system 60.

As a preferred embodiment, the second combustion chamber 43 in this embodiment is one of a single-tube combustion chamber, a ring-tube combustion chamber and a ring-shaped combustion chamber, preferably a single-tube combustion chamber; the plurality of single-tube combustion chambers are arranged in the first outer duct 341 along the circumferential direction; the adjacent single-pipe combustion chambers can be communicated or not communicated through the flame connectors, the single-pipe combustion chambers are preferably not communicated in the embodiment, each single-pipe combustion chamber can work independently, and the most appropriate number of the single-pipe combustion chambers is selected to participate in the work according to the thrust requirements of the airplane at different speeds and at different heights, so that the effect that the engine can work at the highest efficiency under the condition of different thrust requirements is achieved, and the fuel economy is improved.

The turbine system 50 is disposed inside an inner cavity of the casing 80 and between the first combustion chamber 41 and the tail injection system 60, and the turbine system 50 includes a turbine casing 51 mounted on the drive shaft system 70, turbine rotor blades 52 mounted on an outer side of the turbine casing 51, and turbine stator blades 53 mounted on an inner side of the casing 80. When the high-temperature high-pressure high-speed combustion gas flowing out of the first combustion chamber 41 flows through the turbine rotor blades 52, the turbine rotor blades 52 are driven to rotate, the turbine rotor blades 52 drive the transmission shaft system 70 to rotate through the turbine casing 51, the kinetic energy of the combustion gas is converted into mechanical energy, and the mechanical energy is transmitted to the compressor system 20 through the transmission shaft system 70; when the high-temperature and high-pressure gas flowing out of the turbine rotor blades 52 flows through the turbine stator blades 53, the gas flow direction of the gas is changed, the flow velocity is increased, the temperature and the pressure are reduced, the internal energy of the gas is converted into kinetic energy, the high-temperature, high-pressure and high-speed gas is formed, and the high-temperature, high-pressure and high-speed gas enters the next group of turbine rotor blades 52 or the tail injection system 60.

In a preferred embodiment, the turbine system 50 in the present embodiment further includes a high-pressure turbine 51 and a low-pressure turbine 52, the high-pressure turbine 51 is disposed near the rear end of the first combustion chamber 41, and the low-pressure turbine 52 is disposed between the high-pressure turbine 51 and the tail injection system 60. The specific connection and function of the high-pressure turbine 51 and the low-pressure turbine 52 are identical to those of the turbine system 50, and will not be described in detail. Preferably, the high pressure turbine 51 rotates in the opposite direction to the low pressure turbine 52.

The aft jet system 60 is disposed inside the inner cavity of the casing 80, near the end of the engine, and includes an inner duct aft jet 63 disposed at the rear of the turbine system 50 and an outer duct aft jet 64 disposed at the rear of the second combustion chamber 43 inside the first outer duct 341. The high-temperature high-pressure high-speed gas flowing out of the rear part of the turbine system 50 is cooled, depressurized and accelerated at the tail nozzle 63 of the inner duct, the internal energy of the gas is fully converted into kinetic energy, and the kinetic energy is discharged backwards to generate thrust; the high-temperature, high-pressure and high-speed fuel gas flowing out of the rear part of the second combustion chamber 43 is cooled, depressurized and accelerated at the tail nozzle 64 of the outer duct, so that the internal energy of the fuel gas is fully converted into kinetic energy and then is discharged backwards to generate thrust; if the second combustion chamber 43 does not work, the high-pressure air flowing out from the rear part of the second combustion chamber 43 is subjected to pressure reduction and speed increase at the tail nozzle 64 of the outer duct, the internal energy of the air is fully converted into kinetic energy, and then the kinetic energy is discharged backwards to generate thrust.

As a preferred embodiment, the tail injection system 60 in this embodiment further includes a gas fairing 62 disposed at the rear end of the turbine system 50, and is used for guiding the high-temperature and high-speed gas flowing out of the turbine system 50, so as to reduce resistance, and also perform further pressure reduction and speed increase on the gas, increase kinetic energy of the gas, and improve engine efficiency.

In a preferred embodiment, the tail injection system 60 in this embodiment further includes a gas guide vane 61, the gas guide vane 61 is disposed at the rear end of the turbine system 50, and is connected to the casing 80 and the gas fairing 62, so as to guide the high-temperature and high-speed gas flowing out of the turbine system 50 while fixing and supporting the gas fairing 62, change the direction of the gas flow, reduce the pressure and increase the speed, and improve the efficiency of the tail injection system.

The transmission shaft system 70 is arranged at a central axis of an inner cavity of the engine and comprises a transmission shaft 71 connected with the compressor casing 23 and the turbine casing 53 and a transmission bearing 72 supporting the transmission shaft 71 on the first combustion chamber casing 42, and the transmission shaft 71 can rotate freely through the transmission bearing 72.

Preferably, the transmission shaft system 71 further includes a low-pressure transmission shaft 711 and a high-pressure transmission shaft 712, the low-pressure transmission shaft 711 is used for connecting the low-pressure compressor casing 211 and the low-pressure turbine casing 512, the high-pressure transmission shaft 712 is used for connecting the high-pressure compressor casing 212 and the high-pressure turbine casing 511, and the high-pressure transmission shaft 712 and the low-pressure transmission shaft 711 are coaxially installed and are disposed on an outer ring of the low-pressure transmission shaft 711 and movably connected through a transmission bearing 72.

In this embodiment, on the basis of the conventional turbojet engine, the first outer duct 341 and the second combustion chamber 43 are added, and part of the high-pressure fresh air after the compressor is introduced into the second combustion chamber 43 for sufficient combustion, so that the generated high-temperature and high-pressure fuel gas is accelerated by the tail injection system 60 and then is ejected out of the engine, and strong thrust can be continuously generated. Compared with the common turbofan engine outer duct, the mode only sprays relatively low-speed air, so that the energy density is higher, and the thrust-weight ratio is higher; compared with the existing afterburner, the afterburner has the advantages of good working conditions, high pressure, sufficient oxygen, low speed, sustainable and stable combustion and high combustion efficiency, and the length of the required combustor is shorter than that of the afterburner during full combustion, so that the structural weight is relatively small. Meanwhile, the second combustion chamber 43 can be closed, so that the high-pressure fresh air in the first outer duct 341 is directly accelerated by the tail nozzle and then is ejected out of the engine backwards, and at the moment, the working mode is similar to that of a turbofan engine, although the thrust is reduced, the fuel economy is higher.

Example II

Referring to fig. 4 to 6, an outboard engine 90 is added to the embodiment i, and the same parts as those in the embodiment i are not repeated herein. As shown in fig. 4, the engine of the present embodiment further includes an external engine 90, where the external engine 90 includes an external engine mixing chamber 91, an external engine combustion chamber 92, an external engine exhaust nozzle 93, a high-pressure external bleed air pipe 94, a high-pressure external bleed air pipe valve 95, and an external engine fuel pipeline 96. The external engine mixing chamber 91 is arranged at the front section of the external engine 90 and provides a pre-mixing space for high-pressure air and fuel; the external engine combustion chamber 92 is arranged at the rear section of the external engine mixing chamber 91 and is communicated with the external engine combustion chamber to provide a sufficient combustion space for mixed gas of high-pressure air and fuel; the outboard engine tail nozzle 93 is arranged at the rear section of the outboard engine 90, is communicated with the outboard engine combustion chamber 92, and is used for cooling, depressurizing and accelerating the high-temperature and high-pressure gas flowing out of the outboard engine combustion chamber 92, fully converting the internal energy of the gas into kinetic energy, and discharging the kinetic energy backwards to generate thrust; one end of the high-pressure outer bleed air pipe 94 is communicated with the first outer duct 341, the other end is communicated with the external engine mixing chamber 91, and high-pressure air in the first outer duct 341 is introduced into the external engine mixing chamber 91 to be mixed with fuel; the high-pressure outer bleed air pipe valve 95 is arranged at the connection between the high-pressure outer bleed air pipe 94 and the first outer duct 341, and is used for controlling the proportion of high-pressure air in the first outer duct 341 flowing into the external engine 90; the outboard engine fuel line 96 communicates with the outboard engine mixing chamber 91 to direct fuel into the outboard engine mixing chamber 91 for mixing with the high pressure air.

As a preferred embodiment, the outboard engine 90 in this embodiment further includes an outboard engine exhaust nozzle cooling system 97, the outboard engine exhaust nozzle cooling system 97 is disposed at the periphery of the pipe walls of the outboard engine combustion chamber 92 and the outboard engine exhaust nozzle 93, and is communicated with the high-pressure outer bleed air pipe 94 and the outboard engine mixing chamber 91, and the high-pressure air flowing in from the high-pressure outer bleed air pipe 94 cools the pipe walls of the outboard engine combustion chamber 92 and the outboard engine exhaust nozzle 93 in the outboard engine exhaust nozzle cooling system 97, takes away heat, and then flows out to the outboard engine mixing chamber 91;

the external engine exhaust nozzle cooling system 97 in this embodiment may also be respectively communicated with the high-pressure external bleed air pipe 94 and the external atmosphere, and a part of the high-pressure air flowing from the high-pressure external bleed air pipe 94 directly enters the external engine mixing chamber 91 to be mixed with the fuel oil, and the other part of the high-pressure air enters the external engine exhaust nozzle cooling system 97 to cool the pipe walls of the external engine combustion chamber 92 and the external engine exhaust nozzle 93, take away heat, and then flow out to the external atmosphere.

As a preferred embodiment, the outboard engine 90 in this embodiment further includes an outboard engine deflection device 98, the outboard engine deflection device 98 is mounted on the aircraft and connected to the outboard engine 90, and the thrust direction of the outboard engine 90 is controlled by adjusting the direction of the exhaust nozzle of the outboard engine 90.

The scheme of adding the external engine in the embodiment is equivalent to that one core engine with the air compressor drives a plurality of external engines only comprising combustion chambers and tail nozzle systems to work simultaneously, so that the effect that a plurality of engines work simultaneously to generate thrust in different directions is obtained at the cost of lower structural weight, and the external engine system is suitable for vertical take-off and landing airplanes;

example III

Referring to fig. 7 to 8, in this embodiment, a punching duct is added on the basis of the embodiments i and ii, and the same points in this embodiment as those in the embodiments i and ii are not described again. Taking the example of adding the ramjet duct on the basis of the embodiment i, as shown in fig. 7 and 8, the engine of the embodiment further includes the ramjet duct 38, the ramjet duct 38 is disposed in the inner cavity of the casing 80 and the outer ring of the compressor system 20, one end of the ramjet duct 38 is communicated with the air inlet 11, and the other end of the ramjet duct is communicated with the first outer duct 341. When the flying speed of the airplane reaches Mach 2, the engine enters a sub-combustion stamping working mode, incoming air forms high-pressure air after being subjected to speed reduction and pressure boost at the air inlet channel 11, the high-pressure air flows into the first outer duct 341 through the stamping duct 38, the high-pressure air is fully combusted after being mixed with fuel at the second combustion chamber 43, the formed high-temperature high-pressure fuel gas is subjected to temperature reduction, pressure reduction and speed increase at the tail nozzle 64 of the outer duct, the internal energy of the fuel gas is converted into kinetic energy, the kinetic energy is discharged backwards, and thrust is generated.

As a preferred embodiment, the air intake duct system 10 in this embodiment further comprises a ram duct intake valve 16, the ram duct intake valve 16 being provided at the junction of the ram duct 38 and the air intake duct 11 for adjusting the proportion of air flowing into the ram duct 38, or for closing the ram duct 38.

As a preferred embodiment, the air intake duct system 10 in this embodiment further includes an air intake duct adjustable shock cone 14 and a shock cone supporting structure 15, where the air intake duct adjustable shock cone 14 is disposed in the air intake duct 11, and is connected to the shock cone supporting structure 15 at the front end of the compressor system 20. Preferably, the adjustable shock wave cone 14 of the air inlet channel can move back and forth in the shock wave cone supporting structure 15, as shown in fig. 7, the adjustable shock wave cone 14 of the air inlet channel is in a retraction state, at this time, the air inlet area of the air inlet channel 11 is the largest, and is suitable for the airplane in the take-off and subsonic flight states, and the engine is in a normal turbine working state; as shown in figure 8, the adjustable shock cone 14 of the air inlet channel is in an extending state, the air inlet area of the air inlet channel 11 is the smallest, and the adjustable shock cone is suitable for the aircraft in a supersonic flight state, and the engine is switched to a sub-combustion ram working state. The cross section area of the air inlet channel 11 is changed through the adjustable shock wave cone 14 of the air inlet channel, so that the shock wave intensity of incoming air with different Mach numbers entering the air inlet channel 11 is adjusted, and the incoming air is effectively decelerated and pressurized.

Preferably, the ducted system 30 further includes a ram ducted exhaust valve 39, the ram ducted exhaust valve 39 being provided at a connection of the ram ducted 38 and the first outer duct 341 for opening or closing the ram ducted 38.

The same situation is the same when the punching duct is added on the basis of the embodiment II, and the description is omitted.

The scheme of adding the ramjet duct enables the engine to have the capability of freely switching between the normal jet mode and the sub-combustion ramjet mode, so that the airplane loaded with the engine can be accelerated to Mach 2 through the normal jet mode in the takeoff stage and then switched to the sub-combustion ramjet mode, and the airplane can fly at supersonic speed and even hypersonic speed.

It should be noted that the above embodiments can be freely combined as necessary. The foregoing is only a preferred embodiment of the present invention, and it should be noted that it is obvious to those skilled in the art that various modifications and improvements can be made without departing from the principle of the present invention, and these modifications and improvements should also be considered as the protection scope of the present invention.

Claims

1. The utility model provides an aviation jet engine, its characterized in that includes the casing, the casing includes inlet end and exhaust end, the inside of casing is installed along inlet end to exhaust end direction in proper order:

the air inlet system comprises an air inlet, and air flows into the interior of the engine through the air inlet;

a compressor system that decelerates and supercharges air flowing into the engine via an intake duct;

the bypass system comprises an inner bypass and a first outer bypass, and air pressurized by the compressor system respectively enters the inner bypass and the first outer bypass; a first bypass flow dividing control valve is arranged at the position, close to the compressor system, of the first outer bypass;

the transmission shafting is used for connecting the turbine system and the compressor system;

the external engine comprises an external engine mixing chamber, an external engine combustion chamber and an external engine tail jet pipe, wherein the external engine mixing chamber, the external engine combustion chamber and the external engine tail jet pipe are communicated with each other through a high-pressure external air-entraining pipe, one end of the high-pressure external air-entraining pipe is communicated with a first external duct, the other end of the high-pressure external air-entraining pipe is communicated with the external engine mixing chamber, and the external engine mixing chamber is communicated with an external engine fuel oil pipeline.

2. An aircraft jet engine according to claim 1, wherein: the second combustion chamber is one of a single-tube combustion chamber, a circular-tube combustion chamber or a circular combustion chamber.

3. An aircraft jet engine according to claim 2, wherein: the second combustion chamber is a single-pipe combustion chamber, is arranged in the first outer duct along the circumferential direction, and each single-pipe combustion chamber works independently.

4. An aircraft jet engine according to claim 1, wherein: the air inlet system also comprises an air inlet fairing arranged at the front end of the air compressor system and used for reducing air resistance generated by the air compressor system.

5. An aircraft jet engine according to claim 1, wherein: the compressor system comprises a low-pressure compressor and a high-pressure compressor, a second outer duct is directly connected between the low-pressure compressor and the exhaust end of the shell, and a second duct shunting control valve is arranged on the second outer duct at a position close to the low-pressure compressor.

6. An aircraft jet engine according to claim 1, wherein: still include external engine jet-tail pipe cooling system, external engine jet-tail pipe cooling system set up in the outer wall of external engine combustion chamber and the outer wall of external engine jet-tail pipe, external engine jet-tail pipe cooling system respectively with high pressure outer bleed pipe external engine mixing chamber intercommunication, the high-pressure air of bleed pipe inflow outside from high pressure it is right in the external engine jet-tail pipe cooling system external engine combustion chamber with the pipe wall of external engine jet-tail pipe is cooled down, takes away the heat, flows out again external engine mixing chamber.

7. An aircraft jet engine according to claim 1, wherein: the external engine deflection device is arranged on the external engine deflection device, the direction of the tail nozzle of the external engine is adjusted by the external engine deflection device, and the thrust direction of the external engine is further controlled.

8. An aircraft jet engine according to any one of claims 1 to 7, wherein: the punching duct is communicated between the air inlet duct and the first outer duct, a punching duct air inlet valve is arranged at the joint of the punching duct and the air inlet duct, and a punching duct exhaust valve is arranged at the joint of the punching duct and the first outer duct.

9. An aircraft jet engine according to claim 8, wherein: the air inlet system further comprises an adjustable shock wave cone of an air inlet channel and a shock wave cone supporting structural member, the adjustable shock wave cone of the air inlet channel is installed at the front end of the air compressor system through the shock wave cone supporting structural member, and the shock wave cone supporting structural member is a telescopic structural member.