CN212615068U - Distributed propulsion turbofan engine - Google Patents
Distributed propulsion turbofan engine Download PDFInfo
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- CN212615068U CN212615068U CN202021549521.XU CN202021549521U CN212615068U CN 212615068 U CN212615068 U CN 212615068U CN 202021549521 U CN202021549521 U CN 202021549521U CN 212615068 U CN212615068 U CN 212615068U
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Abstract
The utility model discloses a distributing type impels turbofan engine, form by main engine and at least one auxiliary fan combination, main engine includes traditional fan rotor, boost level compressor, high-pressure compressor, the combustion chamber, high-pressure turbine, low pressure turbine, the jet, bleed system, the high pressurized air source pipeline, auxiliary fan includes auxiliary fan rotor supporting system, auxiliary fan rotor, air turbine drive arrangement, bleed system draws out the high pressurized air source of a certain proportion from main engine, high pressurized air source passes through in the high pressurized air source pipeline carries air turbine drive arrangement, for directly or through reduction gear drive auxiliary fan rotor. The utility model discloses a nimble adjustment combustion chamber export gas energy distribution form makes its energy supply in a plurality of fan rotors, can effectively reduce the fan rotor diameter of main engine when reaching super high bypass ratio, realizes the optimization regulation to engine equivalent bypass ratio through the regulation to the blade pitch of main engine or auxiliary fan simultaneously.
Description
Technical Field
The utility model belongs to aviation gas turbine engine field, in particular to distributing type impels turbofan engine.
Background
In the field of aviation and gas turbine engines, turbofan engines are widely used due to good fuel economy, wherein turbofan engines with large bypass ratio are adopted in the civil field, and more turbofan engines with small bypass ratio are adopted in the military field, particularly for fighter aircraft.
With the further improvement of the requirement on fuel economy, more civil aviation large ducts are realized by improving the duct ratio than turbofan engines, so that the diameter of a fan is larger and larger, the rotating speed of the fan rotor is also greatly limited due to the limitation of the linear velocity of the blade tip of the fan rotor, and the rotating speed of a low-pressure turbine rotor can only be further reduced, so that the increase of the number of low-pressure turbine stages, the increase of the weight of the engine and the increase of the number of parts and the length of the whole engine are brought; the geared engine of the american boon company (such as PW1100G) alleviates this problem by placing a planetary reducer in series between the fan rotor and the low-pressure turbine rotor, but with higher design bypass ratios, the fan diameter is larger and larger, which results in smaller and smaller distances from the ground after the engine is hoisted, and the mandatory minimum distance requirements of civil aviation regulations cause this approach to end up.
SUMMERY OF THE UTILITY MODEL
In order to solve the problem, the utility model provides a distributing type impels turbofan engine.
The utility model adopts the technical proposal that: a distributed propulsion turbofan engine is composed of a main engine and at least one auxiliary fan, the main engine comprises a traditional fan rotor, a booster-stage compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine, a tail nozzle, an air-entraining system and a high-pressure air source pipeline, the auxiliary fan comprises an auxiliary fan rotor supporting system and an auxiliary fan rotor, the auxiliary fan rotor supporting system is provided with an air turbine driving device, the air turbine driving device is internally provided with an air turbine and a connecting shaft for driving an auxiliary fan rotor, the air entraining system draws a high-pressure air source with a certain proportion from the main engine, the high-pressure air source is conveyed into the air turbine driving device through a high-pressure air source pipeline, the air turbine is used for pushing the air turbine to rotate, and the air turbine drives the auxiliary fan rotor directly through a connecting shaft or through a speed reducer.
Further, the auxiliary fan rotor may be of 1 or more stages, and may rotate in the same direction/in opposite directions, wherein the auxiliary fan rotor is hoisted under the wing together with the main engine, and is combined with the main engine to form a multi-fan propulsion system.
Furthermore, an auxiliary fan casing is arranged on the auxiliary fan rotor to provide accommodation for the auxiliary fan rotor under the limit condition.
Furthermore, the main engine and the auxiliary fan rotor are wholly or partially provided with a variable pitch device.
Drawings
FIG. 1 is a schematic view of a conventional single fan rotor high bypass ratio turbofan engine;
fig. 2 is a schematic structural diagram of the present invention;
shown in the figure: 1-a traditional fan rotor, 2-a booster stage compressor, 3-a high pressure compressor, 4-a combustion chamber, 5-a high pressure turbine, 6-a low pressure turbine, 7-a tail nozzle, 8-a gas-leading system, 9-a high pressure gas source pipeline, 10-an air turbine driving device, 11-an auxiliary fan rotor, 12-an auxiliary fan casing and 13-an auxiliary fan rotor supporting system.
Detailed Description
The following provides a more detailed description of the present invention with reference to the accompanying drawings and examples.
Fig. 1 is a schematic diagram of a conventional single-fan-rotor large-bypass-ratio turbofan engine, in which a conventional fan rotor 1, a booster-stage compressor 2 and a high-pressure compressor 3 continuously boost the airflow entering the engine, and the heat of fossil fuel is absorbed in a combustion chamber 4 to become high-temperature and high-pressure fuel gas, the energy of the fuel gas is divided into three parts, the first part is to impact a high-pressure turbine 5, so that the high-pressure compressor 3 is driven; the second is to strike the low pressure turbine 6, thereby driving the single stage fan rotor 1 and booster stage compressor 2, with the single stage fan rotor 1 providing a portion of the thrust; the third part is rectified by the tail nozzle 7 and then sprayed backwards into the atmosphere, and the other part of thrust is provided. The total thrust of the engine is obtained by synthesizing two thrusts provided by the fan rotor 1 and the tail nozzle 7 for exhausting air backwards; the part of the incoming air flowing into the booster compressor 2 after passing through the single-stage fan rotor 1 is defined as a bypass airflow, the incoming air flows through an outer duct after passing through the single-stage fan rotor 1, and the part of the incoming air not flowing into the booster compressor 2 is defined as a bypass airflow. And defining the mass flow ratio of the outer culvert airflow to the inner culvert airflow as a culvert ratio. As the diameter of the single-stage fan rotor 1 increases and the bypass ratio of the engine increases, the specific gravity of the fuel gas energy with high temperature and high pressure distributed to the second part increases, and the specific gravity of the thrust provided by the fan rotor 1 in the total thrust also increases.
Examples
As shown in fig. 2, the distributed propulsion turbofan engine of the present invention is formed by combining a main engine and at least one auxiliary fan, wherein the main engine includes a conventional fan rotor 1, a booster stage compressor 2, a high pressure compressor 3, a combustion chamber 4, a high pressure turbine 5, a low pressure turbine 6, a nozzle 7, an air-entraining system 8, and a high pressure air source pipeline 9, the auxiliary fan includes an auxiliary fan rotor supporting system 13 and an auxiliary fan rotor 11, an air turbine driving device 10 is disposed on the auxiliary fan rotor supporting system 13, an air turbine and a connecting shaft for driving the auxiliary fan rotor 11 are disposed in the air turbine driving device 10, compressed air generated by the main engine is delivered to the air turbine driving device 10 through the air-entraining system 8 and the high pressure air source pipeline 9, so as to drive the auxiliary fan rotor 11 to rotate, the auxiliary fan rotor 11 is contained by an auxiliary fan casing 12; while the auxiliary fan rotor 11 is supported by an auxiliary fan rotor support system 13. An auxiliary fan rotor 13 is mounted below the wing and together with the main engine forms a multi-fan distributed propulsion system.
In the above-described embodiment, the bleed position of the bleed system 8 may be mounted at the outlet of the fan/booster stage or at the outlet of the high-pressure compressor; when the main engine is selected to be the turbofan engine with a small bypass ratio, the fan can be made into multiple stages, so that the outlet pressure is improved, and the matching with the inlet demand pressure of the air turbine system is facilitated.
In the above embodiment, the bleed air system 8 may add devices such as a control valve and an air flow meter according to the control accuracy requirement; meanwhile, the conventional fan rotor 1 and the auxiliary fan rotor 11 can be wholly or partially provided with a variable pitch device so as to flexibly change the bypass ratio of the engine. The auxiliary fan rotor 11 and the traditional fan rotor 1 are both contained by a fan casing, so that the running noise can be effectively reduced, and the seaworthiness safety is improved.
The utility model has the advantages that:
1. compared with the gear transmission engine in the prior art, the utility model can significantly reduce the size of the fan of the main engine, thereby reducing the rigidity requirement on the height of the undercarriage;
2. as the size of the fan increases, the manufacturing technology of the large-sized fan becomes more and more complex and difficult, and the cost becomes higher. The utility model can greatly reduce the size of the fan, reduce the manufacturing cost, and simultaneously avoid the risk of falling of the whole engine project caused by not mastering the high-difficulty fan blade preparation technology;
3. compared with an open rotor engine (or a paddle fan engine), the utility model discloses when realizing super high bypass ratio, reduced the size of main engine for main engine and auxiliary fan still can hoist and mount in the wing below, need not make big change to the structure of current civil aviation passenger plane, reduced the research and development cost of new model;
4. compared with the mechanical gear transmission adopted in the patent CN 201610821255-a multi-fan propulsion device, the utility model adopts the technical approach of energy transmission based on compressed air to reduce the complexity of mechanical transmission;
5. compared with the engine-generator-motor energy transmission path widely used in the extended range electric automobile for the ground, the utility model can abandon the requirement on the motor and the generator with huge weight and reduce the weight of the whole power device because of the limited power density of the motor and the huge transmission power of the airplane engine;
6. compared with a fuel gas shunting device adopted in a patent CN 95112599-removal type super fan engine, the utility model can effectively reduce the use of high-value high-temperature alloy materials and the complexity of a heat management system of the engine, thereby effectively reducing the cost of the engine;
7. taking the traditional layout of hanging left and right hairs under wings at present as an example, matching an auxiliary fan at the left and right sides of two main engines respectively as an embodiment, the layout can fully utilize the injection effect of the main engines, improve the total inlet pressure of the auxiliary fan, and further improve the thrust of the auxiliary fan, which is similar to the assistance of head geese in a goose array on small geese; meanwhile, the auxiliary fan can also improve the total pressure of the inlet of the main engine by working, and the purpose of '1 +1> 2' cooperative propulsion is achieved when the auxiliary fan works together;
8. in practical application, a turbofan engine with a small bypass ratio can be selected as a main engine, and when an aircraft system needs power devices with different bypass ratios, the equivalent bypass ratio of the whole power device can be adjusted only by adjusting the proportion of air entraining and the size, the number of stages and the number of auxiliary fans, so that the power devices can be flexibly matched during model selection, the structural change of the main engine is reduced, and the research and development cost is greatly reduced;
9. furthermore, in practical application, the flexible adjustment of the bypass ratio can be realized through the control of the blade pitch of the fan blades of the main engine or the auxiliary fan, the performance of the engine under different flying heights is improved, and the optimization of the matching of the flying is realized.
It should be noted that many modifications and adaptations of the present invention may occur to one of ordinary skill in the art in light of the present disclosure, and are intended to be within the scope of the present invention as defined by the following claims.
Claims (4)
1. A distributed propulsion turbofan engine is formed by combining a main engine and at least one auxiliary fan, wherein the main engine comprises a traditional fan rotor (1), a booster-stage compressor (2), a high-pressure compressor (3), a combustion chamber (4), a high-pressure turbine (5), a low-pressure turbine (6) and a tail nozzle (7), the auxiliary fan comprises an auxiliary fan rotor supporting system (13) and an auxiliary fan rotor (11), and the distributed propulsion turbofan engine is characterized in that the main engine further comprises an air entraining system (8) and a high-pressure air source pipeline (9), an air turbine driving device (10) is arranged on the auxiliary fan rotor supporting system (13), an air turbine and a connecting shaft used for driving the auxiliary fan rotor (11) are arranged in the air turbine driving device (10), and the air entraining system (8) guides out a high-pressure air source with a certain proportion from the main engine, the high-pressure air source is conveyed into an air turbine driving device (10) through a high-pressure air source pipeline (9) to push an air turbine to rotate, and the air turbine drives an auxiliary fan rotor (11) directly through a connecting shaft or through a speed reducer.
2. A distributed propulsion turbofan engine according to claim 1 wherein the auxiliary fan rotor (11) can be of stage 1 or multi-stage, co/counter rotating, wherein the auxiliary fan rotor (11) is hoisted under the wing together with the main engine, and is combined with the main engine as a multi-fan propulsion system.
3. A distributed propulsion turbofan engine according to claim 1 or 2 wherein the auxiliary fan rotor (11) is provided with an auxiliary fan case (12).
4. A distributed propulsion turbofan engine according to claim 1 or 2 wherein the main and auxiliary fan rotors (11) are wholly or partially variable pitch devices.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202021549521.XU CN212615068U (en) | 2020-07-30 | 2020-07-30 | Distributed propulsion turbofan engine |
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| CN202021549521.XU CN212615068U (en) | 2020-07-30 | 2020-07-30 | Distributed propulsion turbofan engine |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113847165A (en) * | 2021-10-12 | 2021-12-28 | 中国电子科技集团公司第三十八研究所 | Series-connection supercharged electric turbine engine double-output-shaft motor turbofan system |
| CN117871789A (en) * | 2024-03-08 | 2024-04-12 | 天津市金晶气体压缩机制造有限公司 | Detection device applied to large-size air compressor |
-
2020
- 2020-07-30 CN CN202021549521.XU patent/CN212615068U/en active Active
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113847165A (en) * | 2021-10-12 | 2021-12-28 | 中国电子科技集团公司第三十八研究所 | Series-connection supercharged electric turbine engine double-output-shaft motor turbofan system |
| CN117871789A (en) * | 2024-03-08 | 2024-04-12 | 天津市金晶气体压缩机制造有限公司 | Detection device applied to large-size air compressor |
| CN117871789B (en) * | 2024-03-08 | 2024-05-31 | 天津市金晶气体压缩机制造有限公司 | Detection device applied to large-size air compressor |
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