CN106523187A - Divided flow gas turbine engine - Google Patents
Divided flow gas turbine engine Download PDFInfo
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- CN106523187A CN106523187A CN201510577423.4A CN201510577423A CN106523187A CN 106523187 A CN106523187 A CN 106523187A CN 201510577423 A CN201510577423 A CN 201510577423A CN 106523187 A CN106523187 A CN 106523187A
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- high pressure
- thrust
- shunting
- pressure draught
- gas
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Abstract
The invention discloses an aviation gas turbine engine. The power and the thrust of the engine can be increased by 2.2 times or more under a condition that the structural mass is the same by means of increasing the air flow and additionally arranging a flow division spraying pipe; the non-augmented thrust-weight ratio can be up to 14.08 or more (generally, the non-augmented thrust-weight ratio of a turbine engine is 6.4); the augmented thrust-weight ratio can be up to 28.16 or more (generally, the augmented thrust-weight ratio of an augmented engine is 8); furthermore, under an augmented state, the engine can continuously work until an augmenting working medium is used up; and compared with the fuel consumption of the general turbine engine, the fuel consumption can be at least reduced twice or more. The engine can easily enter a state at the level of 3Ma or above for cruising, and also can enable an airplane to fly in and out of an atmospheric layer or in a near space.
Description
Technical field
The invention belongs to aero-engine field
Background technology
Usual aero gas turbine engine, due to structural principle, the reason for manufacturing and designing with the aspect such as material, makes reinforcing thrust-weight ratio seldom have and reaches 10.Less efficient due to gas-turbine unit compressor wheel mechanism that generally band is reinforced, after-burner can only be worked again at lower pressures, make total oil consumption rate of engine higher.Under the restriction of engine, contemporary aircraft is difficult to be greatly improved in terms of aeroplane performance.
The content of the invention
It is an object of the invention to provide a kind of existing design manufacturing technology using domestic Taihang engine WS10, it becomes possible to which the not accentuated push anharmonic ratio of enforcement is up to more than 14.08；Thrust-weight ratio is reinforced up to more than 28.16；Turbogenerator thrust, power than phase homogenous quantities is big more than 2.2 times；The oil consumption rate Multifunctional navigation gas-turbine unit tactic pattern minimum low more than 1 times than existing aero-turbine.
In order to achieve the above object, the technical scheme is that：Increase high pressure air flow by a relatively large margin and set up separate nozzle.Separate nozzle is cylindrical shape, is individually symmetrically mounted on the both sides of high pressure draught generator.Specific practice is：The appropriate length of blade for lengthening high pressure draught generator；The appropriate series for increasing gas turbine, the available work for producing the burning of high pressure draught generator, compressor is given all in theory, so as to produce the high pressure gas flow of 1.5 times more of the consumption that works more than high pressure draught generator itself, work 4.5 times of fuel mixed combustion more than consumption into shunting combustion room and than high pressure draught generator itself, power and thrust that the available work for producing of burning than high pressure draught generator is directly changed into big more than 4.5 times of thrust is produced, promotes aircraft to advance.Thus make not accentuated push anharmonic ratio increase to more than 14.08；Reinforcing thrust-weight ratio increases to more than 28.16；Power and thrust are more than 2.2 times of phase homogenous quantities turbogenerator.Oil consumption rate is low more than 1 times.Due to no compressor wheel mechanism in separate nozzle, also the loss in efficiency for not having compressor wheel mechanism to bring, limit with ignition temperature, so the operating temperature and operating efficiency of separate nozzle are significantly larger than usual turbogenerator, the saying that more than 1 times lower than usual turbogenerator of oil consumption rate is conservative saying.The big 2.2 times of sayings for reaching 14.08 and 28.16 with thrust-weight ratio of thrust are also very conservative saying, because these sayings, all the factor that separate nozzle improves operating temperature and cancellation compressor wheel loss in efficiency are not taken into account.
Shunting gas-turbine unit is made up of high pressure draught generator (turbojet engine of deformation) and separate nozzle two large divisions.
High pressure draught generator is by compressor, combustion chamber, gas turbine, fuel system and often effect igniter is constituted.
Separate nozzle constitutes (both sides that two separate nozzles are arranged separately in high pressure draught generator) by punching press air inlet pipe, high pressure gas pipeline, air-flow power and energy valve, shunting combustion room, propelling nozzle, adjustable nozzle, long-acting igniter and fuel system.
Due to no rotating mechanism in separate nozzle, operating characteristic is not restricted by the moment of inertia.
Operating position of the shunting gas-turbine unit by change air-flow power and energy valve, with four kinds of working conditions, shunting gas-turbine unit working condition；Ramjet propulsion working condition；Propulsive thrust working condition；Rocket engine working condition.
Description of the drawings
Below in conjunction with the accompanying drawings the present invention is described in further details.
Fig. 1 is the structure principle chart of the present invention.
Fig. 2 is the working state figure of the shunting gas-turbine unit of shunting gas-turbine unit of the present invention.
Fig. 3 is the ramjet propulsion working state figure of shunting gas-turbine unit of the present invention.
Fig. 4 is the propulsive thrust working state figure of shunting gas-turbine unit of the present invention.
Fig. 5 is the liquid-propellant rocket engine working state figure of shunting gas-turbine unit of the present invention.
Specific embodiment
The structure of shunting gas-turbine unit of the present invention is as shown in figure 1, be made up of high pressure draught generator 1,5 two large divisions of separate nozzle.
High pressure draught generator 1 is by compressor 2, combustion chamber 3, gas turbine 4, fuel system 12 and often effect igniter 13 etc. is constituted.
Separate nozzle 5 is made up of punching press air inlet pipe 6, high pressure gas pipeline 7, air-flow power and energy valve 8, shunting combustion room 9, propelling nozzle 10, adjustable nozzle 11, fuel system 12 and long-acting igniter 13 etc..
The shunting gas-turbine unit working condition of shunting gas-turbine unit.
As shown in Fig. 2 the operating position by changing air-flow power and energy valve 8, makes the port closing of punching press air inlet pipe 6；The import of shunting combustion room 9 is opened；The opening releases of high pressure gas pipeline 7；Air is entered in compressor 2 by 1 entrance of high pressure draught generator, flow out from the outlet of compressor 2 after compression, sub-fraction is entered in combustion chamber 3 and fuel mixed combustion is done work, promote 4 high-speed rotation of turbine, compressor 2 is driven to rotate together by power transmission shaft, high pressure draught is continuously generated, the normal operation of engine is maintained；Most of high pressure draught enters shunting combustion room 9 and more fuel mixed combustions, produce the available work for producing of burning than high pressure draught generator 1 and be directly changed into many times greater power of thrust and thrust, so that power, thrust, thrust-weight ratio are greatly improved.The combustion chamber of shunting combustion room 9 and high pressure draught generator 1 is worked under high-temperature high-pressure state, therefore oil consumption rate is relatively low；Now whole engine enters shunting gas-turbine unit working condition.
The ramjet propulsion working condition of shunting gas-turbine unit.
As shown in figure 3, when the total pressure ratio of the design that the stamping pressure that flying speed is produced is close to shunting gas-turbine unit, opening punching press air inlet pipe 6 by air-flow power and energy valve 8；The import of shunting combustion room 9 is opened；The port closing of high pressure gas pipeline 7；High pressure draught generator 1 quits work；Windstream is directly entered behind shunting combustion room 9 and fuel mixed combustion from punching press air inlet pipe 6, and 10 adjustable nozzle 11 of Jing propelling nozzles sprays, and produces thrust, and whole engine enters punching engine working condition.
The propulsive thrust working condition of shunting gas-turbine unit.
As shown in figure 4, closing the import of shunting combustion room 9 by air-flow power and energy valve 8；Punching press air inlet pipe 6 is opened；The opening releases of high pressure gas pipeline 7；From the high pressure draught that compressor 2 flows out, spray from the import of punching press air inlet pipe 6, produce propulsive thrust, make aircraft down, whole engine enter propulsive thrust working condition.
As shown in figure 5, closing the import of shunting combustion room 9 by air-flow power and energy valve 8；Punching press air inlet pipe 6 is opened；The opening releases of high pressure gas pipeline 7；High pressure draught generator 1 quits work, while spraying into liquid oxidizer and fuel into shunting combustion room 9, whole engine enters liquid-propellant rocket engine working condition；In order to mitigate weight, only in the shunting gas-turbine unit of endoatmosphere flight, this function can be cancelled.
Claims (4)
1. a kind of aviation shunting gas-turbine unit, is made up of high pressure draught generator and separate nozzle two large divisions,
It is characterized in that two separate nozzles are cylindrical shape, the both sides of high pressure draught generator are individually symmetrically mounted on, high pressure draught is sent out
Turbojet engine of the raw device for modification；From the least significant end of compressor, while to shunting combustion room and combustion chamber for same pressure
High pressure draught；High pressure draught generator does not directly produce thrust, but all gives compressor the available work that all burning is produced,
Work many more than the 1.5 times high pressure gas flows of consumption to producing ratio high pressure draught generator itself, in the shunting combustion room with
The fuel mixed combustion many 4.5 times more than high pressure draught generator own combustion consumption, producing can than high pressure draught generator
Diligent power and the thrust for being directly changed into big more than 4.5 times of thrust；Separate nozzle is by punching press air inlet pipe, high pressure gas pipeline, gas
Stream power and energy valve, shunting combustion room, propelling nozzle, adjustable nozzle, long-acting igniter and fuel system etc. are main
Partly to constitute.
2. shunting gas-turbine unit according to claim 1, it is characterised in that no rotating machine in separate nozzle
Structure, operating characteristic are not affected by the moment of inertia.
3. shunting gas-turbine unit according to claim 1, by changing the working position of air-flow power and energy valve
Put, with four kinds of working conditions, it is characterised in that shunting gas-turbine unit working condition；Punching engine work shape
State；Propulsive thrust working condition；Liquid-propellant rocket engine working condition.
4. shunting gas-turbine unit according to claim 1, it is characterised in that the rotating speed of high pressure draught generator,
Do not change with the thrust variation of shunting gas-turbine unit, in the whole thrust variation of shunting gas-turbine unit
In the range of, no matter thrust how acute variation, the compressor of high pressure draught generator is all without there are surge problems.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201510577423.4A CN106523187A (en) | 2015-09-14 | 2015-09-14 | Divided flow gas turbine engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201510577423.4A CN106523187A (en) | 2015-09-14 | 2015-09-14 | Divided flow gas turbine engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN106523187A true CN106523187A (en) | 2017-03-22 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201510577423.4A Pending CN106523187A (en) | 2015-09-14 | 2015-09-14 | Divided flow gas turbine engine |
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| CN (1) | CN106523187A (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108301940A (en) * | 2018-04-23 | 2018-07-20 | 孟金来 | Three rows mouthful turbojet |
| CN110905689A (en) * | 2018-09-17 | 2020-03-24 | 高荣江 | Open water cooling split-flow engine |
| CN114572407A (en) * | 2022-03-10 | 2022-06-03 | 中国航空发动机研究院 | Distributed variable cycle engine and aircraft |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2745247A (en) * | 1952-01-23 | 1956-05-15 | Bell Aircraft Corp | Rocket assisted composite engine arrangement |
| JPH02301653A (en) * | 1989-05-17 | 1990-12-13 | Mitsubishi Heavy Ind Ltd | Air liquefying cycle engine |
| CN2620100Y (en) * | 2003-06-06 | 2004-06-09 | 孔德昌 | Combined fanjet |
| CN201105808Y (en) * | 2007-09-28 | 2008-08-27 | 大连海事大学 | Air breathing type rocket combined cycle engine |
| CN103726952A (en) * | 2012-10-11 | 2014-04-16 | 高荣江 | Divided-flow type gas turbine engine |
-
2015
- 2015-09-14 CN CN201510577423.4A patent/CN106523187A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2745247A (en) * | 1952-01-23 | 1956-05-15 | Bell Aircraft Corp | Rocket assisted composite engine arrangement |
| JPH02301653A (en) * | 1989-05-17 | 1990-12-13 | Mitsubishi Heavy Ind Ltd | Air liquefying cycle engine |
| CN2620100Y (en) * | 2003-06-06 | 2004-06-09 | 孔德昌 | Combined fanjet |
| CN201105808Y (en) * | 2007-09-28 | 2008-08-27 | 大连海事大学 | Air breathing type rocket combined cycle engine |
| CN103726952A (en) * | 2012-10-11 | 2014-04-16 | 高荣江 | Divided-flow type gas turbine engine |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108301940A (en) * | 2018-04-23 | 2018-07-20 | 孟金来 | Three rows mouthful turbojet |
| CN110905689A (en) * | 2018-09-17 | 2020-03-24 | 高荣江 | Open water cooling split-flow engine |
| CN114572407A (en) * | 2022-03-10 | 2022-06-03 | 中国航空发动机研究院 | Distributed variable cycle engine and aircraft |
| CN114572407B (en) * | 2022-03-10 | 2023-12-29 | 中国航空发动机研究院 | Distributed variable-cycle engine and aircraft |
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Application publication date: 20170322 |