CN220267831U - Hybrid aeroengine - Google Patents
Hybrid aeroengine Download PDFInfo
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- CN220267831U CN220267831U CN202321846218.XU CN202321846218U CN220267831U CN 220267831 U CN220267831 U CN 220267831U CN 202321846218 U CN202321846218 U CN 202321846218U CN 220267831 U CN220267831 U CN 220267831U
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- aircraft engine
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- 230000005540 biological transmission Effects 0.000 claims abstract description 11
- 239000000446 fuel Substances 0.000 description 24
- 238000010586 diagram Methods 0.000 description 8
- 230000008901 benefit Effects 0.000 description 7
- 230000002195 synergetic effect Effects 0.000 description 5
- 239000000284 extract Substances 0.000 description 4
- 239000007921 spray Substances 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000002485 combustion reaction Methods 0.000 description 1
- 230000005611 electricity Effects 0.000 description 1
- 230000007613 environmental effect Effects 0.000 description 1
- 230000003993 interaction Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000001105 regulatory effect Effects 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Abstract
The utility model aims to provide a hybrid aeroengine, which comprises a low-pressure turbine and a core assembly, wherein the core assembly comprises a high-pressure turbine, and further comprises a first fan assembly, a second fan assembly, a motor assembly, a ducted fan and a valve. The first fan assembly is in transmission connection with the low-pressure turbine, the second fan assembly is arranged at the downstream of the first fan assembly along the air inlet direction of the engine and is in transmission connection with the high-pressure turbine, the motor unit is in transmission connection with the low-pressure turbine, and the low-pressure turbine supplies power. The ducted fan is arranged on the outer side of the aero-engine and is in transmission connection with the motor group. The valve is arranged on the air inlet channel casing between the first fan assembly and the second fan assembly and is provided with an opening state and a closing state. Wherein when the valve is opened, the air flow is allowed to flow into the second fan and the low-pressure turbine through the valve. The aero-engine can coordinate a small supersonic speed bypass ratio with a large subsonic speed bypass ratio.
Description
Technical Field
The utility model relates to the field of aeroengines, in particular to a hybrid aeroengine.
Background
The market positioning of the supersonic passenger plane is obviously different from that of the existing high subsonic passenger plane, the flight time of long-distance travel can be obviously shortened, and the travel efficiency and the comfort of passengers are improved. Studies by authorities show that: "developing a commercial supersonic aircraft that is efficient, cost effective, and environmentally friendly may change rules of intercontinental travel, help countries to take a leading role in the field of aviation science and technology, and create economic and social benefits in the world of global affiliation.
However, since the retirement of the aviation services of the synergetic aircraft in 2003, the civil supersonic transportation industry has been in a standstill. Environmental problems (sonic boom, noise, emissions, etc.) of supersonic aircraft are major obstacles to future civil supersonic aircraft. The jet noise intensity of the engine is proportional to the exhaust speed to the eighth power, referring to the n+2 project study of NASA, in order to meet the airport noise demand of the aircraft, in the high-temperature take-off state, the tail nozzle of the engine needs to be in a subcritical state, and the jet speed of the tail nozzle needs to be not more than 335m/s (1100 ft/s), which means that the take-off state engine must have a higher bypass ratio. In addition, another important requirement in the design of the engine of the supersonic civil airliner is to ensure better economy in supersonic speed and subsonic speed flight. Currently, the main flow power device of the subsonic civil aircraft is a large bypass ratio turbofan engine, but when the aircraft flies at supersonic speed, the fuel consumption rate of the large bypass ratio turbofan engine increases dramatically. If a small bypass ratio turbofan or even a turbojet engine is adopted as power, the subsonic propulsion efficiency is obviously reduced due to the high exhaust speed, and the fuel consumption is higher.
In order to solve the foregoing problems, it is desirable to provide an aero-engine capable of achieving a coordinated supersonic low bypass ratio and subsonic high bypass ratio.
Disclosure of Invention
The utility model aims to provide a hybrid aeroengine which can realize coordination of a small supersonic speed bypass ratio and a large subsonic speed bypass ratio.
To achieve the foregoing object, a hybrid aircraft engine comprising a low pressure turbine and a core assembly comprising a high pressure turbine, the hybrid aircraft engine further comprising:
a first fan assembly drivingly connected to the low pressure turbine;
the second fan assembly is arranged at the downstream of the first fan assembly along the air inlet direction of the engine and is in transmission connection with the high-pressure turbine;
a motor unit in driving connection with the low pressure turbine, powered by the low pressure turbine;
the bypass fan is arranged at the outer side of the aero-engine and is in transmission connection with the motor group; and
the valve is arranged on the air inlet channel casing between the first fan assembly and the second fan assembly and is provided with an opening state and a closing state;
wherein when the valve is opened, air flow is allowed to flow into the second fan and the low-pressure turbine through the valve.
In one or more embodiments, a first variable area bypass ejector is disposed between the second fan assembly and the high pressure turbine.
In one or more embodiments, the hybrid aircraft engine has a first outer duct component and a second outer duct component within an inlet duct of the hybrid aircraft engine, the first variable area duct injector being disposed between the first outer duct component and the second outer duct component.
In one or more embodiments, the aircraft engine further includes an intermediate casing load frame and a turbine inter-stage casing load frame, the low pressure turbine rotor being supported on the intermediate casing load frame and the turbine inter-stage casing load frame by bearings.
In one or more embodiments, the core assembly is supported on the intermediate casing load frame and turbine interstage casing load frame by bearings.
In one or more embodiments, the valve is disposed on the inlet casing between the first fan assembly and the intermediate casing load frame.
In one or more embodiments, the first fan assembly is drivingly connected to a rotor of the low pressure turbine by a low pressure turbine shaft.
In one or more embodiments, the second fan assembly is drivingly connected to the rotor of the high pressure turbine by a high pressure turbine shaft.
In one or more embodiments, the aircraft engine is provided with a tail pipe at its end in the intake direction.
In one or more embodiments, a second variable area bypass eductor is disposed at the high pressure turbine.
The utility model has the beneficial effects that:
when the engine is in the working condition of taking off or subsonic flight, the mode selection valve is opened, the main thrust is generated by the ducted fan, the engine has extremely high equivalent ducted ratio, and the thrust is large and the fuel consumption is low. And because the low-pressure turbine extracts most of gas energy, the temperature and pressure of the gas at the outlet of the low-pressure turbine are reduced, the exhaust speed of the spray pipe is also reduced, and the jet noise of the engine can be obviously reduced. When the engine is in a supersonic flight working condition, the mode selection valve is closed, and the low-pressure turbine outlet air flow and the outer duct air flow are mixed and then discharged through the tail jet pipe to generate all thrust. In this mode, the low pressure turbine only extracts the gas energy required for operation of the first fan assembly, all thrust forces are generated by expansion of the tail pipe, and the bypass ratio is small and the fuel consumption is low.
The foregoing description is only an overview of the technical solutions of the present application, and may be implemented according to the content of the specification in order to make the technical means of the present application more clearly understood, and in order to make the above-mentioned and other objects, features and advantages of the present application more clearly understood, the following detailed description of the present application will be given.
Drawings
Various other advantages and benefits will become apparent to those of ordinary skill in the art upon reading the following detailed description of the preferred embodiments. The drawings are only for purposes of illustrating the preferred embodiments and are not to be construed as limiting the application. Also, like reference numerals are used to designate like parts throughout the accompanying drawings. In the drawings:
FIG. 1 illustrates a schematic diagram of a hybrid aircraft engine according to some embodiments of the present application;
FIG. 2 illustrates a schematic diagram of a takeoff state engine jet speed versus an existing engine jet speed in accordance with some embodiments of the present hybrid aircraft engine;
FIG. 3 illustrates a schematic diagram of a takeoff state engine bypass ratio versus an existing engine bypass ratio in accordance with some embodiments of the present hybrid aircraft engine;
FIG. 4 illustrates a schematic diagram of engine fuel consumption versus existing engine fuel consumption at take-off according to some embodiments of the present hybrid aircraft engine;
FIG. 5 illustrates a graphical representation of supersonic cruise state engine fuel consumption versus existing engine supersonic cruise state engine fuel consumption in accordance with some embodiments of the present hybrid aircraft engine.
Detailed Description
Embodiments of the technical solutions of the present application will be described in detail below with reference to the accompanying drawings. The following examples are only for more clearly illustrating the technical solutions of the present application, and thus are only examples, and are not intended to limit the scope of protection of the present application.
Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this application belongs; the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the application; the terms "comprising" and "having" and any variations thereof in the description and claims of the present application and in the description of the figures above are intended to cover non-exclusive inclusions.
To achieve a coordinated supersonic low bypass ratio and subsonic high bypass ratio, according to some embodiments of the present application, a hybrid aero-engine is provided, as shown in fig. 1, which is a schematic diagram of a hybrid aero-engine according to some embodiments of the present application.
The hybrid aircraft engine 100 comprises a low pressure turbine 1, a core assembly 2, a first fan assembly 3, a second fan assembly 4, a motor assembly 5, a ducted fan 6 and a shutter 7.
The core assembly 2 comprises a high-pressure compressor 21, a combustion chamber 22 and a high-pressure turbine 23, the first fan assembly 3 is in transmission connection with the low-pressure turbine 1, and the second fan assembly 4 is arranged at the downstream of the first fan assembly 3 along the air inlet direction a of the engine and in transmission connection with the high-pressure turbine 23.
The motor group 5 is in driving connection with the low-pressure turbine 1 and provides power via the low-pressure turbine 1. In some specific embodiments, the motor group 5 includes a generator and a driving motor, and the generator is in transmission connection with the low-pressure turbine shaft of the low-pressure turbine 1, so that mechanical energy generated by rotation of the low-pressure turbine 1 can be converted into electric energy by the motor group 5 and output electric power. In a specific embodiment, the electric motor set 5 further includes an electric storage unit, such as a battery, electrically connected to the generator. The electric power provided by the generator can be stored by the electricity storage unit and provide the drive motor to output power from the power source. In one embodiment, the electrical connection referred to herein is a wired electrical connection through wires.
The ducted fan 6 is arranged outside the aeroengine and is in driving connection with the motor unit 5, and in a specific embodiment, the ducted fan 6 is in driving connection with a driving motor in the motor unit 5, so that electric power is extracted from the motor unit 5 to generate rotation. In a specific embodiment, the ducted fans 6 may be provided on the aircraft wing.
The shutter 7 is provided on the casing outside the intake duct 101 between the first fan assembly 3 and the second fan assembly 4, and has an open state and a closed state. It will be appreciated that, similar to existing valve arrangements, the valve 7 allows airflow through in the open condition and prevents airflow through the valve 7 in the closed condition.
When the valve 7 is opened, air flow is allowed to flow into the second fan 4 and the low-pressure turbine 1 through the valve 7, and the valve 7 can be opened or closed according to the flight working condition, so that the air flow entering the core engine assembly 2 is adjusted, the engine flow is matched with the capture flow of the air inlet channel 101, and the flight resistance is reduced.
Specifically, when the engine is in a takeoff or subsonic flight condition, the mode selection valve 7 is opened, the flow of the second fan assembly 4 and the low pressure turbine 1 is increased, the flow of air into the core assembly 2 is increased, and the output power of the low pressure turbine 1 is increased. The output power of the low-pressure turbine 1 is supplied to the motor set 5 for the most part except for the first fan assembly 3, and the generated electric power is sequentially transmitted to the motor set 5 and the ducted fan 6, and finally, the ducted fan 6 generates most of the thrust. And meanwhile, the outlet airflow of the low-pressure turbine 1 is mixed with the outer duct airflow and then discharged through the tail nozzle, so that a small part of thrust is generated. In this mode, the main thrust is generated by the ducted fan 6, the engine has extremely high equivalent ducted ratio, and the thrust is large and the fuel consumption is low. And because the low-pressure turbine 1 extracts most of gas energy, the temperature and pressure of the gas at the outlet of the low-pressure turbine 1 are reduced, the exhaust speed of the spray pipe is also reduced, and the jet noise of the engine can be obviously reduced.
When the engine is in supersonic flight, the mode selection valve 7 is closed, the flow of the second fan assembly 4 and the low-pressure turbine 1 is reduced, the output power of the low-pressure turbine 1 is fully supplied to the first fan assembly 3 to work, and the motor group 5 and the ducted fan 6 are closed. The outlet air flow of the low-pressure turbine 1 is mixed with the outer duct air flow and then discharged through the tail nozzle, so that all thrust is generated. In this mode, the low pressure turbine 1 extracts only the gas energy required for the operation of the first fan assembly 3, all thrust forces are produced by the expansion of the tail pipe, the bypass ratio is small and the fuel consumption is low.
FIG. 2 illustrates a schematic diagram of a takeoff state engine jet speed versus an existing engine jet speed in accordance with some embodiments of the present hybrid aircraft engine. In the figure, a column A1 shows the jet speed of the 'synergistic' engines, a column A2 shows the jet speed of the existing horizontal supersonic turbofan engines, a column A3 shows the jet speed of the hybrid aero-engine, and a dotted line shows the aircraft take-off noise limit value.
FIG. 3 illustrates a schematic diagram of a takeoff state engine bypass ratio versus an existing engine bypass ratio in accordance with some embodiments of the present hybrid aircraft engine. In the figure, a column A4 shows the duct ratio of the existing horizontal supersonic turbofan engine, a column A5 shows the duct ratio of the hybrid aeroengine, and a 'synergistic' aircraft uses a turbojet engine, wherein the duct ratio is 0. As can be seen from the figures, the hybrid aeroengine of the present configuration can significantly increase the bypass ratio, with a great advantage in terms of bypass ratio over existing engine configurations.
FIG. 4 illustrates a schematic diagram of a comparison of takeoff state engine fuel consumption versus existing engine fuel consumption for some embodiments of the present hybrid aircraft engine. In the figure, a column A6 shows the fuel consumption rate of the 'synergetic' engine, a column A7 shows the fuel consumption rate of the existing horizontal supersonic turbofan engine, a column A8 shows the fuel consumption rate of the hybrid aero-engine, and the hybrid aero-engine with the configuration can obviously reduce the fuel consumption rate of the engine and has a larger advantage in the fuel consumption rate compared with the configuration of the existing engine.
FIG. 5 illustrates a graphical representation of supersonic cruise state engine fuel consumption versus existing engine supersonic cruise state engine fuel consumption in accordance with some embodiments of the present hybrid aircraft engine. In the figure, a column A9 shows the supersonic cruising state fuel consumption rate of the 'synergetic' engine, a column A10 shows the supersonic cruising state fuel consumption rate of the prior horizontal supersonic turbofan engine, a column A11 shows the supersonic cruising state fuel consumption rate of the hybrid aeroengine, and the hybrid aeroengine of the configuration can obviously reduce the supersonic cruising state fuel consumption rate and has a larger advantage in the supersonic cruising state fuel consumption rate compared with the prior engine configuration.
In the description of the embodiments of the present application, the technical terms "first," "second," etc. are used merely to distinguish between different objects and are not to be construed as indicating or implying a relative importance or implicitly indicating the number of technical features indicated, a particular order or a primary or secondary relationship. In the description of the embodiments of the present application, the meaning of "plurality" is two or more unless explicitly defined otherwise.
Reference herein to "an embodiment" means that a particular feature, structure, or characteristic described in connection with the embodiment may be included in at least one embodiment of the present application. The appearances of such phrases in various places in the specification are not necessarily all referring to the same embodiment, nor are separate or alternative embodiments mutually exclusive of other embodiments. Those of skill in the art will explicitly and implicitly appreciate that the embodiments described herein may be combined with other embodiments.
In some embodiments of the present hybrid aeroengine, a first variable area bypass injector 8 is disposed between the second fan assembly 4 and the high-pressure turbine 21, and the first variable area bypass injector 8 can be opened or closed or adjusted to open or close according to the flight condition, so as to adjust the air flow entering the core engine assembly 2, realize that the engine flow matches the capture flow of the air inlet channel 101, and reduce the flight resistance.
In some embodiments of the present hybrid aeroengine, the first outer duct member 91 and the second outer duct member 92 are provided in the air intake duct 101 of the hybrid aeroengine, the first variable area duct 8 ejector is provided between the first outer duct member 91 and the second outer duct member 92, wherein the first outer duct member 91 and the second outer duct member 92 are formed on the core casing with an air intake passage formed therebetween, and the first variable area duct 8 is provided at the air intake passage to achieve adjustment of the air flow rate into the core assembly 2.
In some embodiments of the present hybrid aircraft engine, the aircraft engine further comprises an intermediate casing load frame 93 and a turbine inter-stage casing load frame 94, on which the rotor of the low pressure turbine 1 is supported by bearings, in particular by front end ball bearings and rear end roller bearings.
In some embodiments of the present hybrid aircraft engine, core engine assembly 2 is supported by bearings on intermediate casing load frame 93 and turbine inter-stage casing load frame 94, in particular, by front end ball bearings and rear end roller bearings.
In some embodiments of the present hybrid aircraft engine, the shutter 7 is provided on the inlet casing between the first fan assembly 3 and the intermediate casing load frame 93.
In some embodiments of the present hybrid aircraft engine, the first fan assembly 3 is drivingly connected to the rotor of the low pressure turbine 1 by a low pressure turbine shaft.
In some embodiments of the present hybrid aircraft engine, the second fan assembly 4 is drivingly connected to the rotor of the high pressure turbine 23 by a high pressure turbine shaft.
In some embodiments of the present hybrid aircraft engine, the aircraft engine is provided with a tail nozzle 95 at the end in the intake direction, through which the air flow is ejected through the tail nozzle 95 to provide thrust. In one particular embodiment, the tail nozzle 95 is an adjustable exit area tail nozzle.
In some embodiments of the present hybrid aircraft engine, a second variable area bypass injector is provided at the high pressure turbine 23 through which the air flow into the high pressure turbine 23 can be further regulated.
In a specific embodiment, the motor unit 5 is also provided on the intermediate casing load frame 93.
In the description of the embodiments of the present application, unless explicitly specified and limited otherwise, the terms "mounted," "connected," "secured" and the like are to be construed broadly and may be, for example, fixedly connected, detachably connected, or integrally formed; or may be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communicated with the inside of two elements or the interaction relationship of the two elements. The specific meaning of the above terms in the embodiments of the present application will be understood by those of ordinary skill in the art according to the specific circumstances.
It is to be understood that the two directions "perpendicular", "coincident", "parallel", etc. mentioned herein do not need to meet mathematically strict angular requirements, but rather allow a certain tolerance range, e.g. within 20 ° compared to mathematically required angles, whereas "along" a certain direction means that there is at least a component in that direction, preferably an angle to that direction within 45 °, more preferably an angle within 20 °.
Finally, it should be noted that: the above embodiments are only for illustrating the technical solution of the present application, and not for limiting the same; although the present application has been described in detail with reference to the foregoing embodiments, it should be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some or all of the technical features thereof can be replaced by equivalents; such modifications and substitutions do not depart from the spirit of the embodiments, and are intended to be included within the scope of the claims and description. In particular, the technical features mentioned in the respective embodiments may be combined in any manner as long as there is no structural conflict. The present application is not limited to the specific embodiments disclosed herein, but encompasses all technical solutions falling within the scope of the claims.
Claims (10)
1. A hybrid aircraft engine comprising a low pressure turbine and a core assembly comprising a high pressure turbine, the hybrid aircraft engine further comprising:
a first fan assembly drivingly connected to the low pressure turbine;
the second fan assembly is arranged at the downstream of the first fan assembly along the air inlet direction of the engine and is in transmission connection with the high-pressure turbine;
a motor unit in driving connection with the low pressure turbine, powered by the low pressure turbine;
the bypass fan is arranged at the outer side of the aero-engine and is in transmission connection with the motor group; and
the valve is arranged on the air inlet channel casing between the first fan assembly and the second fan assembly and is provided with an opening state and a closing state;
wherein when the valve is opened, air flow is allowed to flow into the second fan and the low-pressure turbine through the valve.
2. The hybrid aircraft engine of claim 1, wherein a first variable area bypass ejector is disposed between the second fan assembly and the high pressure turbine.
3. The hybrid aircraft engine of claim 2, wherein the hybrid aircraft engine has a first outer duct member and a second outer duct member within an inlet duct of the hybrid aircraft engine, the first variable area duct injector being disposed between the first outer duct member and the second outer duct member.
4. The hybrid aircraft engine of claim 1, further comprising an intermediate casing load frame and a turbine inter-stage casing load frame, the low pressure turbine rotor being supported on the intermediate casing load frame and the turbine inter-stage casing load frame by bearings.
5. The hybrid aircraft engine of claim 4, wherein the core assembly is supported on the intermediate casing load frame and turbine inter-stage casing load frame by bearings.
6. The hybrid aircraft engine of claim 4, wherein the valve is disposed on an inlet casing between the first fan assembly and the intermediate casing load frame.
7. The hybrid aircraft engine of claim 1, wherein the first fan assembly is drivingly connected to a rotor of the low pressure turbine by a low pressure turbine shaft.
8. The hybrid aircraft engine of claim 1, wherein the second fan assembly is drivingly connected to a rotor of the high-pressure turbine by a high-pressure turbine shaft.
9. The hybrid aircraft engine according to claim 1, characterized in that the aircraft engine is provided with a tail pipe at its end in the intake direction.
10. The hybrid aircraft engine of claim 1, wherein a second variable area bypass ejector is provided at the high pressure turbine.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202321846218.XU CN220267831U (en) | 2023-07-13 | 2023-07-13 | Hybrid aeroengine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202321846218.XU CN220267831U (en) | 2023-07-13 | 2023-07-13 | Hybrid aeroengine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN220267831U true CN220267831U (en) | 2023-12-29 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202321846218.XU Active CN220267831U (en) | 2023-07-13 | 2023-07-13 | Hybrid aeroengine |
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|---|---|
| CN (1) | CN220267831U (en) |
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2023
- 2023-07-13 CN CN202321846218.XU patent/CN220267831U/en active Active
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