阅读说明：本技术 一种由电机带转增压级实现变涵道比的航空发动机及其控制方法 (Aero-engine with variable bypass ratio realized by motor with booster stage and control method thereof ) 是由 马亚如 顾程 徐鹏 王石柱 于 2021-09-28 设计创作，主要内容包括：本发明属于航空动力技术领域,公开了一种由电机带转增压级实现变涵道比的航空发动机及其控制方法,航空发动机包括内涵道、外涵道、风扇、设置于内涵道内的增压级、高压压气机、燃烧室、高压涡轮以及低压涡轮,高压涡轮通过高压轴驱动高压压气机,低压涡轮通过低压轴驱动风扇,风扇用于向外涵道以及内涵道送风,还包括电机,电机被配置为驱动增压级,增压级用于向内涵道送风。本发明提供的由电机带转增压级实现变涵道比的航空发动机,通过设置电机直接驱动增压级向内涵道送风,省去了传统变涵道发动机在控制涵道比改变过程中复杂的调节机构,有效简化航空发动机整机控制系统繁杂的操作过程,也提升了整机的可靠性与安全性。(The invention belongs to the technical field of aviation power, and discloses an aero-engine with a motor rotating a boosting stage to realize variable bypass ratio and a control method thereof. According to the aero-engine with the variable bypass ratio by the motor to drive the booster stage, the motor is arranged to directly drive the booster stage to supply air to the inner bypass, a complex adjusting mechanism of the traditional variable bypass engine in the process of controlling the change of the bypass ratio is omitted, the complex operation process of the overall control system of the aero-engine is effectively simplified, and the reliability and the safety of the overall engine are improved.)

1. The aero-engine with the variable bypass ratio realized by the motor with the booster stage comprises an inner bypass (100), an outer bypass (200), a fan (300), a booster stage (110) arranged in the inner bypass (100), a high-pressure compressor (120), a combustion chamber (130), a high-pressure turbine (140) and a low-pressure turbine (150), wherein the high-pressure turbine (140) drives the high-pressure compressor (120) through a high-pressure shaft (160), the low-pressure turbine (150) drives the fan (300) through a low-pressure shaft (170), and the fan (300) is used for supplying air to the outer bypass (200) and the inner bypass (100);

characterized in that it further comprises an electric machine (400), said electric machine (400) being configured to drive said booster stage (110), said booster stage (110) being intended to supply air to said endoprosthesis (100).

2. The aircraft engine with variable bypass ratio by means of motor-driven booster stage according to claim 1, characterized in that the output shaft of the motor (400) is connected with the booster stage (110) by means of a coupling.

3. The aircraft engine with variable bypass ratio achieved by turning an electric machine into a boost stage according to claim 1, wherein the output shaft of the electric machine (400) is connected with the boost stage (110) through a speed reduction assembly.

4. The aircraft engine with variable bypass ratio achieved by motor-driven boost stages of claim 3, wherein said speed reduction assembly is a gear assembly.

5. The aircraft engine with variable bypass ratio by means of motor-driven pressure boosting stages as claimed in claim 1, wherein said motor (400) has a through hole passing through in its axial direction, and said low-pressure shaft (170) is disposed through said through hole.

6. The aircraft engine with variable bypass ratio realized by the motor-driven booster stage according to claim 1 is characterized by further comprising a power supply device, wherein the power supply device is used for supplying electric energy required by the motor (400).

7. The aircraft engine with a variable bypass ratio achieved by means of motor-driven boost stages according to claim 6, wherein said power supply device is a generator.

8. A control method of an aircraft engine with a variable bypass ratio realized by a motor-driven booster stage is characterized by being applied to the aircraft engine with the variable bypass ratio realized by the motor-driven booster stage according to any one of claims 1 to 7, and controlling the rotating speed of a motor (400) to be increased when the aircraft engine is in climbing, accelerating and supersonic flight stages so as to increase the air volume of an inner bypass (100) to reduce the bypass ratio; when the aircraft engine is in the takeoff and subsonic flight stage, the rotating speed of the control motor (400) is reduced, so that the air volume of the inner duct (100) is reduced to improve the duct ratio.

Technical Field

The invention relates to the technical field of aviation power, in particular to an aero-engine with a variable bypass ratio realized by a motor with a rotating booster stage and a control method thereof.

Background

With the increasing requirements on the performance of the airplane and the continuous progress of the related technologies of the engines, barriers among engines of different types are broken gradually, the variable-cycle/variable-bypass-ratio engine is generated by taking the advantages of the turbojet engine and the turbofan engine into consideration, adapts to different flight states by changing the bypass ratio of the engine, and has considerable advantages in indexes such as oil consumption, propulsion efficiency and operation range compared with the fixed-bypass-ratio turbojet engine/turbofan engine.

In the prior art, in order to realize variable bypass ratio, an engine needs to have a plurality of adjustable external bypasses, and is combined with a plurality of adjusting mechanisms such as a variable-area bypass ejector, a variable-pitch blade, a middle bypass valve, a bypass valve and the like, the addition of the adjusting mechanisms not only greatly increases the design difficulty of a mechanical structure, but also seriously reduces the safety and reliability of the whole engine, and due to the addition of various adjusting variables, the control on corresponding parts is more complicated, and the difficulty is greatly improved no matter the control rule is determined or the control system is designed, so that the bypass ratio changing mode of the engine which is realized by adopting a multi-actuating mechanism adjusting mode at present has great defects in performance, efficiency, stability and safety.

Disclosure of Invention

The invention aims to provide an aircraft engine which changes the flow of an inner duct by adopting a motor to adjust a booster stage so as to realize a variable duct ratio and a control method thereof, so as to solve the problems of complicated duct ratio adjusting mechanism and high design difficulty of a control system of the aircraft engine in the prior art.

In order to achieve the purpose, the invention adopts the following technical scheme:

the aero-engine with the motor rotating a supercharging stage to realize variable bypass ratio comprises an inner bypass, an outer bypass, a fan, a supercharging stage arranged in the inner bypass, a high-pressure compressor, a combustion chamber, a high-pressure turbine and a low-pressure turbine, wherein the high-pressure turbine drives the high-pressure compressor through a high-pressure shaft, the low-pressure turbine drives the fan through a low-pressure shaft, and the fan is used for supplying air to the outer bypass and the inner bypass;

further included is an electric machine configured to drive the boost stage for supplying air to the inner duct.

Optionally, the output shaft of the motor is connected to the boost stage by a coupling.

Optionally, the output shaft of the electric machine is connected to the boost stage via a speed reduction assembly.

Optionally, the speed reduction assembly is a gear assembly.

Optionally, the motor has a through hole penetrating along an axis direction thereof, and the low-pressure shaft is disposed through the through hole.

Optionally, the electric vehicle further comprises a power supply device for supplying the electric energy required by the motor.

Optionally, the power supply device is a generator.

When the aircraft engine is in the stages of climbing, accelerating and supersonic flying, the rotating speed of the motor is controlled to increase so as to increase the air volume of the inner duct to reduce the duct ratio; when the aircraft engine is in the takeoff and subsonic flight stages, the rotating speed of the motor is controlled to be reduced so as to reduce the air volume of the inner duct and improve the bypass ratio.

Has the advantages that:

according to the aero-engine with the motor for realizing the variable bypass ratio by rotating the booster stage, the motor is arranged to directly drive the booster stage to supply air to the inner bypass, the aero-engine can change the bypass ratio by actively adjusting the rotating speed of the motor according to different flight working conditions, so that the aero-engine can realize good use performance under different working conditions, the structure is simple and practical, a complicated adjusting mechanism of the traditional variable bypass engine in the process of changing the control bypass ratio is omitted, the complicated operation process of a whole machine control system of the aero-engine is effectively simplified, and the reliability and the safety of the whole machine are also improved.

Drawings

FIG. 1 is a schematic structural diagram of an aircraft engine with a variable bypass ratio realized by a motor-driven rotating and pressurizing stage.

In the figure:

100. an inner duct; 110. a pressurization stage; 120. a high pressure compressor; 130. a combustion chamber; 140. a high pressure turbine; 150. a low pressure turbine; 160. a high pressure shaft; 170. a low pressure shaft;

200. an outer duct;

300. a fan;

400. an electric motor.

Detailed Description

The present invention will be described in further detail with reference to the accompanying drawings and examples. It is to be understood that the specific embodiments described herein are merely illustrative of the invention and are not limiting of the invention. It should be further noted that, for the convenience of description, only some of the structures related to the present invention are shown in the drawings, not all of the structures.

In the description of the present invention, unless expressly stated or limited otherwise, the terms "connected," "connected," and "fixed" are to be construed broadly, e.g., as meaning permanently connected, removably connected, or integral to one another; can be mechanically or electrically connected; either directly or indirectly through intervening media, either internally or in any other relationship. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.

In the present invention, unless otherwise expressly stated or limited, "above" or "below" a first feature means that the first and second features are in direct contact, or that the first and second features are not in direct contact but are in contact with each other via another feature therebetween. Also, the first feature being "on," "above" and "over" the second feature includes the first feature being directly on and obliquely above the second feature, or merely indicating that the first feature is at a higher level than the second feature. A first feature being "under," "below," and "beneath" a second feature includes the first feature being directly under and obliquely below the second feature, or simply meaning that the first feature is at a lesser elevation than the second feature.

In the description of the present embodiment, the terms "upper", "lower", "right", etc. are used in an orientation or positional relationship based on that shown in the drawings only for convenience of description and simplicity of operation, and do not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be construed as limiting the present invention. Furthermore, the terms "first" and "second" are used only for descriptive purposes and are not intended to have a special meaning.

The embodiment provides an aircraft engine with a variable bypass ratio realized by a motor with a boosting stage, as shown in fig. 1, the aircraft engine comprises an inner bypass 100, an outer bypass 200, a fan 300, a boosting stage 110 arranged in the inner bypass 100, a high-pressure compressor 120, a combustion chamber 130, a high-pressure turbine 140 and a low-pressure turbine 150, wherein the high-pressure turbine 140 drives the high-pressure compressor 120 through a high-pressure shaft 160, the low-pressure turbine 150 drives the fan 300 through a low-pressure shaft 170, the fan 300 is used for supplying air to the outer bypass 200 and the inner bypass 100, the boosting stage 110 is used for supplying air to the inner bypass 100, the aircraft engine further comprises a motor 400, and the motor 400 is configured to drive the boosting stage 110 to directly supply air to the inner bypass 100. In this embodiment, the motor 400 is arranged to directly drive the booster stage 110 to supply air to the inner duct 100, and the duct ratio of the aircraft engine can be changed by actively adjusting the rotating speed of the motor 400 according to different working conditions during operation, so that the requirement that the aircraft engine can ensure good use performance under different operating conditions is met, and a control mode of directly driving the booster stage 110 by using the motor 400 replaces a complex adjusting mechanism of a traditional variable-duct-ratio engine in the control duct ratio changing process, so that the complex operation process of the overall aircraft engine control system is effectively simplified, the structure is simple, and the operation is convenient. In addition, the structure of the present embodiment in which the motor directly drives the supercharging stage to rotate may be used in a dual-rotor turbofan engine, or may be used in a multi-rotor turbofan engine, which is not limited herein.

Further, the motor 400 directly drives the booster stage 110 to rotate, so that the aircraft engine can stably run under unstable working conditions such as starting, acceleration and deceleration and the like, and surge of the high-pressure compressor 120 in the working process is avoided, thereby avoiding damage to parts or over-temperature of the parts caused by severe vibration of the aircraft engine during working, and effectively ensuring the reliability and safety of the whole aircraft engine.

Further, an output shaft of the motor 400 is coupled to the boost stage 110 by a coupling. The coupler is used as a reliable and efficient torque transmission device, has the characteristics of high sensitivity, high transmission efficiency, long service life, good damping performance, good corrosion resistance and the like, and is widely applied to connection of shafts.

Further, the output shaft of the motor 400 may also be coupled to the boost stage 110 through a speed reduction assembly. In this embodiment, when the rotation speed output by the output shaft of the motor 400 is not matched with the rotation speed required for driving the boost stage 110, a speed reduction assembly needs to be arranged to realize the connection between the two, and by arranging the speed reduction assembly on the output shaft of the motor 400, the speed reduction assembly can provide a suitable transmission ratio, so that the boost stage 110 can be driven to rotate at the actually required rotation speed. Particularly, the speed reducing assembly is a gear assembly which has light weight and compact structure, has good vibration resistance and shock resistance in operation and has high transmission efficiency. In this embodiment, the gear assembly includes at least two gears that mesh with each other, wherein the gear connected to the output shaft of motor 400 is a driving gear, the gear connected to boost stage 110 is a driven gear, and the gear assembly is configured such that the driving gear and the driven gear are appropriately sized according to a desired gear ratio to enable boost stage 110 to rotate at a desired target rotational speed. Set up the speed reduction subassembly in addition and be the gear assembly, drive the more driven gear of number of teeth by the less driving gear meshing of number of teeth, driven gear compares the driving gear rotational speed and reduces, and the torque increase makes the operation steady when pressure boost level 110 starts, and dynamic nature is stronger.

In this embodiment, the motor 400 has a through hole penetrating along its axis, and the low voltage shaft 170 is disposed through the through hole. Specifically, since the output shaft of the motor 400 needs to drive the pressure increasing stage 110 to rotate, that is, the output shaft of the motor 400 and the pressure increasing stage 110 are coaxially arranged, the motor 400 needs to be inserted into the low-pressure shaft 170 during design, the motor 400 in this embodiment is an insertion motor, and the aperture of the insertion hole is larger than the shaft diameter of the low-pressure shaft 170, so as to ensure that the rotation of the pressure increasing stage 110 driven by the motor 400 and the rotation of the fan 300 driven by the low-pressure turbine 150 do not interfere with each other. Meanwhile, the rotation of the booster stage 110 in the embodiment is no longer controlled by the low-pressure turbine 150, so that the load of the low-pressure turbine 150 is effectively reduced, the number of design stages of the low-pressure turbine 150 can be reduced, the design size of the aircraft engine is reduced, and the overall weight is reduced.

Further, the aero-engine with a variable bypass ratio realized by the motor-driven boost stage of the embodiment further includes a power supply device, and the power supply device is used for providing electric energy required by the motor 400 to ensure continuous operation of the motor 400. Specifically, the power supply device may be a generator, and in this embodiment, a rotating shaft of the generator is preferably connected to an end of the low-voltage shaft 170, and when the aircraft engine is running, the low-voltage shaft 170 rotates to drive the generator to rotate at a high speed to generate electric power, so as to provide continuous electric power for the motor 400. The invention is not limited to this, and the generator may also rotate by other driving forms to generate electric energy, which is not further limited herein.

The embodiment further provides a control method of an aircraft engine with a motor rotating and pressurizing stage to realize a bypass ratio, which is applied to the aircraft engine with the motor rotating and pressurizing stage to realize the bypass ratio, and specifically, when the aircraft engine is in climbing, accelerating and ultrasonic flight stages, the rotation speed of the motor 400 is controlled to be increased so as to increase 100 air volume of the inner bypass to reduce the bypass ratio; when the aircraft engine is in the takeoff and subsonic flight stages, the rotating speed of the motor 400 is controlled to be reduced, so that the air quantity of the inner duct 100 is reduced to improve the bypass ratio.

Specifically, when the aircraft engine is in the climbing, accelerating and ultrasonic flight stages, the rotating speed of the motor 400 is controlled to be increased, so that the air volume of the inner duct 100 is obviously increased, the duct ratio is reduced, the thrust of the aircraft engine is effectively enhanced, and the performance of the aircraft engine can be closer to that of a turbojet engine; when the aircraft engine is in the takeoff and subsonic flight stages, the rotating speed of the control motor 400 is reduced, the air quantity of the inner duct 100 is reduced, the duct ratio is increased, and the aircraft engine can be used for less oil consumption, so that more flight mileage is increased.

Taking an aircraft engine with a thrust level of 3000kgf as an example, the main performance parameters of the aircraft engine of the present invention were compared with those of the engines with a small bypass ratio and a large bypass ratio, and the results are shown in table 1.

TABLE 13000 kgf thrust Engine Main Performance parameters

As can be seen from the table 1, the aero-engine with the variable bypass ratio realized by the motor with the rotating booster stage works in a small bypass ratio mode of 0.8 during takeoff, and the thrust of 3000kgf is kept during takeoff; the bypass ratio is increased to 5.5 at cruise conditions. Compared with an engine with a small bypass ratio, the fuel consumption rate is remarkably reduced (-42%) under the condition that the thrust is slightly reduced in the whole process; compared with an engine with a large bypass ratio, the fuel consumption is slightly higher, the thrust in a cruising state is obviously increased, and meanwhile, the inlet flow is only half of that of the engine with the large bypass ratio, so that the size and the weight of the whole body of the aircraft engine are effectively reduced; and meanwhile, the temperature before the turbine is effectively reduced.

It should be understood that the above-described embodiments of the present invention are merely examples for clearly illustrating the present invention, and are not intended to limit the embodiments of the present invention. Numerous obvious variations, adaptations and substitutions will occur to those skilled in the art without departing from the scope of the invention. And are neither required nor exhaustive of all embodiments. Any modification, equivalent replacement, and improvement made within the spirit and principle of the present invention should be included in the protection scope of the claims of the present invention.