CN109854377B - Novel aviation turbofan engine - Google Patents
Novel aviation turbofan engine Download PDFInfo
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- CN109854377B CN109854377B CN201910274675.8A CN201910274675A CN109854377B CN 109854377 B CN109854377 B CN 109854377B CN 201910274675 A CN201910274675 A CN 201910274675A CN 109854377 B CN109854377 B CN 109854377B
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- 239000003638 chemical reducing agent Substances 0.000 claims abstract description 37
- 238000002485 combustion reaction Methods 0.000 claims abstract description 9
- 238000010586 diagram Methods 0.000 description 4
- 230000004323 axial length Effects 0.000 description 2
- 239000007921 spray Substances 0.000 description 2
- 206010034719 Personality change Diseases 0.000 description 1
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- 230000001141 propulsive effect Effects 0.000 description 1
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Abstract
The invention relates to an aviation power device, in particular to a novel aviation turbofan engine. The technical proposal is as follows: the low-pressure rotor system is formed by the first-stage fan, the second-stage fan, the air compressor, the combustion chamber, the high-pressure turbine, the low-pressure turbine, the first speed reducer and the second speed reducer, one end of the first-stage fan is arranged on a first bearing, the other end of the first-stage fan is connected with the first speed reducer, and the other end of the first speed reducer is connected with the low-pressure turbine through a low-pressure turbine shaft; the second-stage fan, the second speed reducer, the air compressor and the high-pressure turbine form a high-pressure rotor system, the second-stage fan is connected with one end of the second speed reducer, and the other end of the second speed reducer is connected with a front shaft neck of the air compressor. The invention has the advantages of small gyro moment, compact structure, light weight, high efficiency and the like.
Description
Technical Field
The invention relates to an aviation power device, in particular to a novel aviation turbofan engine.
Background
Turbofan engines are an important structural type of aeroengines and are widely used due to their high propulsive efficiency and high thrust. The traditional turbofan engine mainly comprises a fan, a gas compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine, an outer culvert casing, a spray pipe and other unit bodies. The basic working principle is that the high-pressure and low-pressure systems of the engine rotate to suck the air at the inlet into the engine, one part of the air is discharged through the outer duct to generate a part of thrust, and the other part of the air is compressed by the compressor, combusted in the combustion chamber and discharged at a high speed after the high-pressure turbine works to generate another part of thrust. The traditional turbofan engine generally adopts a design concept that a high-pressure turbine drives a compressor, a low-pressure turbine drives a fan, and a high-pressure rotor system and a low-pressure rotor system rotate in the same direction, and the working mode has the advantages of clear working principle, simple pneumatic and structural design, but the disadvantages brought by the working mode are that the gyro moment of the rotor system of the engine is larger, stator transition is needed between rotors at all levels, and the axial length of the engine is long and the weight of the whole engine is large.
Disclosure of Invention
The invention provides a novel aviation turbofan engine, which has the advantages of small gyroscopic moment, compact structure, light weight, high efficiency and the like.
The technical scheme of the invention is as follows:
The novel aviation turbofan engine comprises a first-stage fan, a second-stage fan, a gas compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine, a first speed reducer and a second speed reducer, wherein the first-stage fan, the first speed reducer and the low-pressure turbine form a low-pressure rotor system, the low-pressure rotor system is supported on a first bearing, a third bearing, a fourth bearing and a seventh bearing, one end of the first-stage fan is mounted on the first bearing, the other end of the first-stage fan is connected with the first speed reducer, the first bearing is mounted on a front casing, the other end of the first speed reducer is connected with the low-pressure turbine through a low-pressure turbine shaft, the low-pressure turbine is supported on the seventh bearing, and the seventh bearing is mounted on a rear casing; the third bearing and the fourth bearing are intermediate bearings, the third bearing is arranged between the first speed reducer and the second-stage fan, and the fourth bearing is arranged between the front journal of the compressor and the low-pressure turbine shaft; the second-stage fan, the second speed reducer, the air compressor and the high-pressure turbine form a high-pressure rotor system, the high-pressure rotor system is supported on a second bearing, a fifth bearing and a sixth bearing, the second-stage fan is connected with one end of the second speed reducer, the other end of the second speed reducer is connected with a front journal of the air compressor, the high-pressure turbine is supported on the sixth bearing, and the second bearing and the fifth bearing are mounted on an intermediate casing; the sixth bearing is an intermediate bearing and is mounted on the rear journal of the high-pressure turbine and the low-pressure turbine shaft.
Further, the novel aviation turbofan engine comprises a stator system of the engine, wherein the front casing, the intermediate casing, the compressor casing, the combustion chamber, the outer casing, the low-pressure turbine casing, the rear casing and the spray pipe, and the front casing, the intermediate casing and the rear casing are main bearing casings.
The beneficial effects of the invention are as follows: 1. the novel aviation turbofan engine adopts the design idea of contra-rotating a high-pressure rotor system and a low-pressure rotor system, so that the problem of gyro moment increase caused by aircraft attitude change in the engine operation is reduced, and meanwhile, stator blades with partial stages are cancelled in the high-pressure rotor system and the low-pressure rotor system, so that the length of the engine is shortened, and the weight of the whole machine is reduced; 2. according to the novel aviation turbofan engine, the speed reducer is designed in the high-pressure rotor system and the low-pressure rotor system, so that the working capacities of the air compressor, the low-pressure turbine and the two-stage fan are fully exerted, the problem that the high rotating speed of the air compressor and the low-pressure turbine is difficult to match with the low rotating speed of the fan is solved, and meanwhile, the efficiency and the thrust of the engine are improved; 3. the concept of the opposite rotation and deceleration design of the aero-engine can be applied to the technical fields of aerospace, navigation, automobiles and the like with similar use requirements.
Drawings
FIG. 1 is a schematic diagram of the overall structure of a novel aviation turbofan engine;
FIG. 2 is a schematic diagram of a high pressure rotor system;
FIG. 3 is a schematic diagram of a low pressure rotor system;
FIG. 4 is a schematic diagram of an engine stator system.
Detailed Description
As shown in fig. 1, a novel aviation turbofan engine comprises a first-stage fan 2, a second-stage fan 3, a compressor 6, a combustion chamber 7, a high-pressure turbine 9, a low-pressure turbine 10, a first speed reducer 13 and a second speed reducer 16; the second stage fan 3, the second speed reducer 16, the compressor 6, the high-pressure turbine 9 and the like form a high-pressure rotor system, the high-pressure rotor system is supported on a second bearing 14, a fifth bearing 18 and a sixth bearing 20, the second stage fan 3 is connected with one end of the second speed reducer 16, the other end of the second speed reducer 16 is connected with a front journal 5 of the compressor 6, the high-pressure turbine 9 is supported on the sixth bearing 20, the second bearing 14 and the fifth bearing 18 are mounted on an intermediate casing 4, and the sixth bearing 20 is an intermediate bearing and is mounted on a rear journal 8 and a low-pressure turbine shaft 19 of the high-pressure turbine 9; the fifth bearing 18 is a ball bearing, which is used for bearing the axial force and a small part of the radial force of the high-pressure rotor system, and the second bearing 14 and the sixth bearing 20 are roller bearings, which are mainly used for bearing the radial load generated by the system;
The first stage fan 2, the first speed reducer 13, the low pressure turbine 10 and the like form a low pressure rotor system, the low pressure rotor system is supported on a first bearing 12, a third bearing 15, a fourth bearing 17 and a seventh bearing 21, one end of the first stage fan 2 is mounted on the first bearing 12, the other end of the first stage fan is connected with the first speed reducer 13, the other end of the first bearing 12 is mounted on the front casing 1, the other end of the first speed reducer 13 is connected with the low pressure turbine 10 through a low pressure turbine shaft 19, the low pressure turbine 10 is supported on the seventh bearing 21, the seventh bearing 21 is mounted on the rear casing 11, the third bearing 15 is an intermediate bearing and is mounted between the first speed reducer 13 and the second stage fan 3, the fourth bearing 17 is an intermediate bearing and is mounted between the front journal 5 and the low pressure turbine shaft 19 of the compressor 6; the first bearing 12 is a ball bearing and is used for bearing the axial force and a small part of the radial force of the low-pressure rotor system, and the third bearing 15, the fourth bearing 17 and the seventh bearing 21 are roller bearings and are mainly used for bearing the radial load generated by the system;
As shown in fig. 4, the front casing 1, the intermediate casing 4, the compressor casing 22, the combustion chamber 7, the outer culvert casing 23, the low-pressure turbine casing 24, the rear casing 11, the nozzle 25 and the like form a stator system of the engine, wherein the front casing 1, the intermediate casing 4 and the rear casing 11 are main bearing casings, and the stator system bears loads generated by various bearings and gas inside the engine to maintain stable operation of the whole engine system.
Embodiment one: with reference to fig. 1 and 2, when the engine works, the high-pressure rotor system rotates clockwise according to the clockwise direction, and the tangential speed of the second-stage fan 3 is not too high due to the fact that the rotating speed of the air compressor 6 is high and the second-stage fan is limited by the size; a second speed reducer 16 is designed between the compressor 6 and the second stage fan 3, and is used for reducing the output rotation speed of the compressor 6 to a proper degree of the second stage fan 3, so that the advantages of the respective components of the compressor 6 and the second stage fan 3 can be fully utilized, and the performance of the engine is improved.
Embodiment two: referring to fig. 1 and 3, when the engine works, the low-pressure rotor system rotates in a clockwise direction, and the tangential speed of the first-stage fan 2 is not excessively high due to the fact that the rotating speed of the low-pressure turbine 10 is relatively high and the first-stage fan is limited by the size; a first speed reducer 13 is designed between the low-pressure turbine 10 and the first-stage fan 2, and is used for reducing the output rotation speed of the low-pressure turbine 10 to a proper degree of the first-stage fan 2, so that the advantages of the components of the low-pressure turbine 10 and the first-stage fan 2 can be fully utilized, and the performance of the engine is improved.
According to the novel aviation turbofan engine, due to the adoption of the design concept of contra-rotating high-pressure and low-pressure rotor systems, stator guide blades are omitted between the first-stage fan 2 and the second-stage fan 3 and between the high-pressure turbine 9 and the low-pressure turbine 10, so that the gyroscopic moment of the engine is reduced, the stability of an airplane is improved, the axial length of the engine is shortened, the weight of a structure is reduced, the development cost is reduced, and the thrust-weight ratio is improved.
Claims (2)
1. The novel aviation turbofan engine is characterized by comprising a first-stage fan, a second-stage fan, a gas compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine, a first speed reducer and a second speed reducer, wherein the first-stage fan, the first speed reducer and the low-pressure turbine form a low-pressure rotor system, the low-pressure rotor system is supported on a first bearing, a third bearing, a fourth bearing and a seventh bearing, one end of the first-stage fan is mounted on the first bearing, the other end of the first-stage fan is connected with the first speed reducer, the first bearing is mounted on a front casing, the other end of the first speed reducer is connected with the low-pressure turbine through a low-pressure turbine shaft, the low-pressure turbine is supported on the seventh bearing, and the seventh bearing is mounted on a rear casing; the third bearing and the fourth bearing are intermediate bearings, the third bearing is arranged between the first speed reducer and the second-stage fan, and the fourth bearing is arranged between the front journal of the compressor and the low-pressure turbine shaft; the second-stage fan, the second speed reducer, the air compressor and the high-pressure turbine form a high-pressure rotor system, the high-pressure rotor system is supported on a second bearing, a fifth bearing and a sixth bearing, the second-stage fan is connected with one end of the second speed reducer, the other end of the second speed reducer is connected with a front journal of the air compressor, the high-pressure turbine is supported on the sixth bearing, and the second bearing and the fifth bearing are mounted on an intermediate casing; the sixth bearing is an intermediate bearing and is mounted on the rear journal of the high-pressure turbine and the low-pressure turbine shaft.
2. The novel aviation turbofan engine of claim 1 wherein the front casing, intermediate casing, compressor casing, combustion chamber, outer culvert casing, low pressure turbine casing, rear casing and nozzle comprise the stator system of the engine, wherein the front casing, intermediate casing and rear casing are the primary load bearing casings.
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| Application Number | Priority Date | Filing Date | Title |
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| CN201910274675.8A CN109854377B (en) | 2019-04-08 | 2019-04-08 | Novel aviation turbofan engine |
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| Application Number | Priority Date | Filing Date | Title |
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| CN201910274675.8A CN109854377B (en) | 2019-04-08 | 2019-04-08 | Novel aviation turbofan engine |
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| CN109854377A CN109854377A (en) | 2019-06-07 |
| CN109854377B true CN109854377B (en) | 2024-05-03 |
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Families Citing this family (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113123881B (en) * | 2019-12-31 | 2022-05-31 | 中国航发商用航空发动机有限责任公司 | Support structure of engine |
| CN111305952A (en) * | 2020-02-26 | 2020-06-19 | 北京航空航天大学 | Mixed exhaust turbofan engine propulsion system based on heating of outer duct |
| CN112983651B (en) * | 2021-04-26 | 2023-07-28 | 黄锴 | Small aviation double-rotor unmanned aerial vehicle engine |
| CN113864239A (en) * | 2021-10-27 | 2021-12-31 | 中国航发沈阳发动机研究所 | Double-duct aero-engine high-low pressure gas engine and intermediate casing part thereof |
| CN115405421B (en) * | 2022-11-01 | 2023-02-03 | 北京航空航天大学 | Three-rotor variable-cycle engine overall structure with interstage combustion chamber |
| CN116085142B (en) * | 2023-04-11 | 2023-05-30 | 北京航空航天大学 | Novel overall structure of interstage combustion variable cycle engine |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| US7395657B2 (en) * | 2003-10-20 | 2008-07-08 | General Electric Company | Flade gas turbine engine with fixed geometry inlet |
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|---|---|---|---|---|
| CN1453466A (en) * | 2002-03-01 | 2003-11-05 | 通用电气公司 | Reverse rotary aero-gas turbine of air-compressor with high general pressure ratio |
| CN1900508A (en) * | 2005-06-06 | 2007-01-24 | 通用电气公司 | Integrated counterrotating turbofan |
| CN101149026A (en) * | 2007-11-08 | 2008-03-26 | 北京航空航天大学 | Turbine fan engine and design method thereof |
| CN105765166A (en) * | 2013-11-21 | 2016-07-13 | 斯奈克玛 | Modular engine, such as a jet engine, with a speed reduction gear |
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| CN109854377A (en) | 2019-06-07 |
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