Power split multi-rotor aircraft gas turbine hybrid power system, multi-rotor aircraft and control method of multi-rotor aircraft

Abstract

The invention discloses a power split multi-rotor aircraft gas turbine hybrid power system, a multi-rotor aircraft and a control method of the multi-rotor aircraft. The power split multi-rotor aircraft gas turbine hybrid power system comprises a gas turbine engine, a planetary gearbox, a generator, an electric energy distribution unit, a battery, a first motor for driving a first rotor, a second motor for driving a second rotor and a power distribution gearbox. Compared with an existing series hybrid power scheme, the planetary gearbox is introduced for power dividing, the two rotors are mechanically and directly driven, the energy conversion efficiency is higher, the generator only needs to drive the two motors driving the rotors, and the required electric energy is half of that of a traditional hybrid power system under the same condition; therefore, the electric power requirement, the size, the weight and the development difficulty of the generator are greatly reduced, the cruising ability and the energy utilization efficiency are improved, the weight and the design difficulty of a power system are greatly reduced, and low-cost and long-cruising flight is realized.

Description

Technical Field

[0001] The present application relates to the technical field of rotorcraft, and in particular, to a power-split multi-rotor aircraft gas turbine hybrid power system, a multi-rotor aircraft and a control method thereof. Background Art

[0002] Common multi-rotor aircraft include four-axis quadrotor, six-axis six-rotor, eight-axis eight-rotor, etc. Taking the four-axis quadrotor aircraft as an example, the rotor blade angle (pitch) is usually not changed, and the lift of the rotor is changed by independently adjusting the rotation speed of the four rotors. Changing the aircraft attitude and the angle of the rotation plane can achieve aerial movement and change direction. At present, multi-rotor aircraft generally use a combination of batteries and motors as the power system, which has poor power density, energy density, and endurance.

[0003] Gas turbine shaft engines have high power density and endurance, but the output shaft speed of classic gas turbine shaft engines is constant, and it is impossible to achieve flexible speed adjustment when directly used in multi-rotor aircraft. In addition, simple power distribution gearboxes can only achieve an even distribution of engine power and cannot be directly applied to multi-rotor aircraft.

[0004] At present, the gas turbine hybrid power system of multi-rotor aircraft generally uses a series hybrid power system, where the gas turbine shaft engine generates electricity through a generator, and the electricity drives the motor after passing through an inverter. This solution requires energy to be converted twice, and the efficiency is low. The power requirements of the generator and motor are high, making the generator and motor too large, increasing the weight of the power system and the difficulty of design. Summary of the invention

[0005] The present application provides a power-split multi-rotor aircraft gas turbine hybrid power system to solve the technical problems of the prior art such as poor endurance, low efficiency, heavy power system weight and high design difficulty.

[0006] This application is implemented through the following scheme: A power split multi-rotor aircraft gas turbine hybrid power system, comprising: A gas turbine engine, a planetary gearbox, a generator, an electric energy distribution unit, a battery, a first electric motor driving a first rotor, a second electric motor driving a second rotor, and a power distribution gearbox, wherein: the planetary gearbox comprises planetary gears and a planetary carrier, a sun gear output shaft, a planetary gear output shaft, and an outer ring gear; the output shaft of the gas turbine engine is drivingly connected to the planetary gears and the planetary carrier, the sun gear output shaft is drivingly connected to the input end of the generator, the planetary gear output shaft is connected to the outer ring gear, the planetary gear output shaft is drivingly connected to the input end of the power distribution gearbox, the power output end of the generator is respectively connected to the second electric motor driving the second rotor, the first electric motor driving the first rotor, and the battery circuit through the electric energy distribution unit; the output end of the power distribution gearbox is drivingly connected to the third rotor and the fourth rotor, respectively.

[0007] Furthermore, the power distribution unit adopts a power electronic integrated component, including an AC / DC converter, a DC / DC converter, a DC / AC converter, and a DC high-voltage bus. The AC power of the generator is input into the DC high-voltage bus through the AC / DC converter (AC / DC converter), and the battery realizes bidirectional energy exchange of the DC high-voltage bus through the DC / DC converter (DC / DC converter). The power of the DC high-voltage bus is converted into AC power through the DC / AC converter (DC / AC converter) to drive the first motor and the second motor. The first motor drives the first rotor to rotate, and the second motor drives the second rotor to rotate.

[0008] Furthermore, it also includes a third motor connected to the power distribution unit circuit, the output end of the third motor is drivingly connected to the input end of the power distribution gearbox, and a power cut-off device is provided between the planetary gear output shaft and the power distribution gearbox.

[0009] Furthermore, the power cut-off device adopts a clutch, which disconnects or connects the mechanical connection between the planetary gear output shaft and the power distribution gearbox.

[0010] Furthermore, the battery is a power type battery.

[0011] On the other hand, the present application also provides a multi-rotor aircraft, including the power-split multi-rotor aircraft gas turbine hybrid power system.

[0012] On the other hand, the present application also provides a control method for a multi-rotor aircraft, comprising the steps of: When performing pitch motion, the third and fourth rotors are controlled to rotate at the same speed, while the first and second rotors rotate at different speeds. The torques generated by the four rotors are balanced to 0. At this time, the aircraft first tilts to a certain extent, so that the rotor pull produces a horizontal component, causing the aircraft to move forward and backward. When performing yaw motion, the third rotor and the fourth rotor are controlled to have the same rotation speed, and the first rotor and the second rotor have the same rotation speed. At the same time, the rotation speed of the first rotor and the third rotor are different. The torque of the four rotors is unbalanced, causing the aircraft to yaw and turn.

[0013] Furthermore, the method further comprises the steps of: When hovering, the third rotor and the fourth rotor are controlled to have the same rotation speed, the first rotor and the second rotor are controlled to have the same rotation speed, and the first rotor and the third rotor are controlled to have the same rotation speed. The torques of the four rotors are balanced and the thrust is the same as the weight of the aircraft, so that the aircraft hovers in the air. Control the third rotor and the fourth rotor to have the same speed, the first rotor and the second rotor to have the same speed, and the first rotor and the third rotor to have the same speed. The torques of the four rotors are balanced and the thrust is greater than the weight of the aircraft, so that the aircraft takes off vertically. When the torques of the four rotors are balanced and the thrust is less than the weight of the aircraft, the aircraft descends vertically.

[0014] Furthermore, the method further comprises the steps of: During normal operation, the power distribution unit supplies power to the first motor, the second motor and the third motor, the gas turbine engine drives the planetary gearbox, and the planetary gear output shaft of the planetary gearbox and the third motor jointly drive the third rotor and the fourth rotor to rotate.

[0015] Furthermore, the method further comprises the steps of: When the gas turbine engine is facing engine failure or to meet environmental noise requirements, the gas turbine engine, planetary gearbox and generator stop working, and the mechanical connection between the planetary gear output shaft and the power distribution gearbox is disconnected by the clutch. The battery's electrical energy is provided to the first motor, the second motor and the third motor through the power distribution unit to drive all rotors to rotate.

[0016] Compared with the prior art, this application has the following beneficial effects: In this application, a planetary gearbox is introduced to split the power, so that part of the power of the gas turbine engine is allocated to the generator for power generation, and then drives two motors and two rotors; the other part of the power is directly driven by the output shaft of the planetary gear. The battery is only used for peak load regulation of the power grid, and the rotor power is mainly provided by the gas turbine engine. The aircraft is in a vertical take-off state, and the power requirements of the four rotors are the same. Compared with the existing series hybrid scheme, the scheme of this application has higher energy conversion efficiency because the two rotors are mechanically directly driven, while all the rotor power of the series hybrid system needs to be converted twice, and the energy conversion efficiency is relatively low. The generator of this scheme only needs to drive two motors that drive the rotors, and the required electrical energy is half of that of the traditional hybrid system under the same conditions, which greatly reduces the electrical power demand, size, weight and development difficulty of the generator, while the development of high-power and high-speed motors is very difficult, while the development of low-power and high-speed motors is relatively mature; generally, the higher the speed, the smaller the size and weight of the motor, and the higher the power density.

[0017] In addition to the above-described purposes, features and advantages, the present application also has other purposes, features and advantages. The present application will be further described in detail with reference to the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] The drawings constituting a part of the present application are used to provide a further understanding of the present application. The illustrative embodiments and descriptions of the present application are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings: Figure 1 This is a schematic diagram of the composition principle of a power split multi-rotor aircraft gas turbine hybrid power system according to a preferred embodiment of the present application; Figure 2 It is a schematic diagram of the composition principle of a power split multi-rotor aircraft gas turbine hybrid power system according to another preferred embodiment of the present application; Figure 3 It is a flow chart of the control method of the first preferred embodiment of the present application; Figure 4 It is a schematic diagram of the action principle of each rotor in the control method of the first preferred embodiment of the present application; Figure 5 is a flow chart of a control method according to a second preferred embodiment of the present application; Figure 6 It is a flow chart of the control method of the third preferred embodiment of the present application; Figure 7 is a flow chart of a control method according to a fourth preferred embodiment of the present application; Figure 8 It is a flowchart of the control method of the fifth preferred embodiment of the present application.

[0019] In the figure: 1. gas turbine engine; 2. generator; 3. planetary gearbox; 31. sun gear output shaft; 32. outer ring gear; 33. planetary gear and planet carrier; 4. power distribution unit; 5. battery; 6. first rotor; 7. first electric motor; 8. second electric motor; 9. second rotor; 10. third rotor; 11. power distribution gearbox; 12. fourth rotor; 13. second generator. DETAILED DESCRIPTION

[0020] The embodiments of the present application are described in detail below in conjunction with the accompanying drawings, but the present application can be implemented in a variety of different ways defined and covered below.

[0021] like Figure 1 As shown, a preferred embodiment of the present application provides a power-split multi-rotor aircraft gas turbine hybrid power system, comprising a gas turbine engine 1, a planetary gearbox 3, a generator 2, an electric energy distribution unit 4, a battery 5, a first motor 7 for driving a first rotor 6, a second motor 8 for driving a second rotor 9, and a power distribution gearbox 11, wherein: the planetary gearbox 3 comprises planetary gears and a planetary carrier 33, a sun gear output shaft 31, a planetary gear output shaft, and an outer ring gear 32, the output shaft of the gas turbine engine 1 is drivingly connected to the planetary gear and the planetary carrier 33, the sun gear output shaft 31 is drivingly connected to the input end of the generator 2, the planetary gear output shaft is connected to the outer ring gear 32, the planetary gear output shaft is drivingly connected to the input end of the power distribution gearbox 11, the power output end of the generator 2 is respectively connected to the second motor 8 driving the second rotor 9, the first motor 7 driving the first rotor 6, and the battery 5 through the electric energy distribution unit 4; the output end of the power distribution gearbox 11 is drivingly connected to the third rotor 10 and the fourth rotor 12 respectively.

[0022] The power split multi-rotor aircraft gas turbine hybrid power system of this embodiment introduces a planetary gear power split device. The planetary gear is divided into three parts: a sun gear, a planetary gear and a planetary carrier, and an outer gear ring. It has the dual functions of speed regulation and power splitting. The output shaft of the gas turbine engine is connected to the planetary carrier; the generator 2 is connected to the sun gear (center gear); and the outer gear ring is connected to the planetary gear shaft. The planetary gear box 3 as a power split device can adjust the speed of the planetary gear output shaft by adjusting the speed of the generator 2 when the speed of the gas turbine engine output shaft is constant, thereby adjusting the speed of the third rotor 10 and the fourth rotor 12.

[0023] Therefore, it can be understood that in this embodiment, by introducing a planetary gearbox for power diversion, the engine can work in a relatively efficient range for a long time, and at the same time, a part of the power of the gas turbine engine 1 is allocated to the generator 2 for power generation, and then drives two motors and two rotors; the other part of the power is directly driven by the other two rotors through the planetary gear output shaft, wherein the battery is only used for grid peak regulation, the rotor power is mainly provided by the gas turbine engine 1, the aircraft is in a vertical take-off state, and the power requirements of the four rotors are the same. Compared with the existing series hybrid scheme, the scheme of this embodiment has higher energy conversion efficiency because the two rotors are mechanically directly driven, while all the rotor power of the series hybrid system needs to be converted twice, and the energy conversion efficiency is relatively low. The generator of this embodiment only needs to drive two motors that drive the rotors, and the required electrical energy is half of that of the traditional hybrid system under the same conditions, which greatly reduces the electrical power demand, size, weight and development difficulty of the generator, while the development of high-power and high-speed motors is very difficult, while the development of low-power and high-speed motors is relatively mature; generally, the higher the speed, the smaller the size and weight of the motor, and the higher the power density. As long as the thermal management system is reasonable, the power attenuation of the battery 5 is very small under high temperature or plateau conditions. For extreme working conditions such as takeoff / hovering / emergency / plateau, the battery 5 can be combined with the gas turbine engine 1 to provide mechanical power at the same time to meet the power requirements of extreme working conditions and ensure the safety and reliability of flight.

[0024] Specifically, the power distribution unit 4 adopts a power electronic integrated component, including an AC / DC converter, a DC / DC converter, a DC / AC converter, and a DC high-voltage bus. The AC power of the generator 2 is input into the DC high-voltage bus through the AC / DC converter (AC / DC converter), and the battery 5 realizes bidirectional energy exchange of the DC high-voltage bus through the DC / DC converter (DC / DC converter). The electric energy of the DC high-voltage bus is converted into AC power through the DC / AC converter (DC / AC converter) to drive the first motor 7 and the second motor 8. The first motor 7 drives the first rotor 6 to rotate, and the second motor 8 drives the second rotor 9 to rotate.

[0025] In this embodiment, the generator 2 and the first motor 7, the second motor 8, and the battery 5 are connected through the power distribution unit 4. The power distribution unit 4 is a power electronic integrated component, and the energy flow process inside it is summarized as follows: the alternating current of the generator 2 is input to the DC high-voltage bus through the AC / DC converter (AC / DC converter), and the battery can realize the bidirectional energy exchange of the DC high-voltage bus through the DC / DC converter (DC / DC converter). The electric energy of the DC high-voltage bus is converted into alternating current through the DC / AC converter (DC / AC converter) to drive the first motor 7 and the second motor 8. The first motor 7 drives the first rotor 6, and the second motor 8 drives the second rotor 9. The power of the planetary gear output shaft is evenly distributed to the third rotor 10 and the fourth rotor 12 through the power distribution gearbox. The power of the battery 5 is only used to adjust the working state of the engine. In high-power demand states such as take-off / emergency / landing / hovering, the battery 5 can provide high power for a short time. In low-power demand states such as level flight, the electric energy of the generator 2 can charge the battery 5. Since the motor only needs to drive two rotors, the size and weight of the battery 5 and the motor can be greatly reduced in this case compared to the traditional series hybrid mode.

[0026] like Figure 2 As shown, in another preferred embodiment of the present application, the power-split multi-rotor aircraft gas turbine hybrid power system also includes a third motor 13 connected to the circuit of the power distribution unit 4, the output end of the third motor 13 is drivingly connected to the input end of the power distribution gearbox 11, and a power cut-off device is arranged between the planetary gear output shaft and the power distribution gearbox 11, and the power cut-off device adopts a clutch, which disconnects or connects the mechanical connection between the planetary gear output shaft and the power distribution gearbox 11 through the clutch.

[0027] Different from the above-mentioned embodiment, the power split multi-rotor aircraft gas turbine hybrid power system in this embodiment further includes a third motor 13 connected to the electric energy distribution unit 4 circuit, and a power cut-off device arranged between the planetary gear output shaft and the power distribution gearbox 11. By adding a third motor 13, when the gas turbine engine 1 is not suitable for working in special occasions, such as when the gas turbine engine 1 fails or to meet environmental noise requirements, all rotor power can be temporarily provided directly by the power battery, and the gas turbine engine 1, the planetary gearbox 3 and the generator 2 are not working. At this time, the mechanical connection between the planetary gear output shaft and the power distribution gearbox can be disconnected by the clutch, and the rotor of the gas turbine engine 1 can also be locked, the generator 2 is not excited, and the generator rotor and the planetary gearbox 3 are idling. In normal operation, the second generator 13 and the planetary gear output shaft jointly drive the third rotor 10 and the fourth rotor 12.

[0028] Preferably, the battery 5 is a power type battery, which can provide high power output to meet the large demand for energy in a short period of time. It is very suitable for applications that require fast response and large current discharge, such as electric vehicles, hybrid power systems, acceleration and braking energy recovery systems, frequency regulation and peak compliance management of power grids, and fast response components in renewable energy systems. Similarly, the power type battery used in this embodiment can provide high power output to meet the large demand for energy in a short period of time during flight, and ensure the energy needs and safety requirements of the aircraft.

[0029] Another preferred embodiment of the present application also provides a multi-rotor aircraft, comprising the power-split multi-rotor aircraft gas turbine hybrid power system.

[0030] like Figure 3 and Figure 4 As shown, the first preferred embodiment of the present application also provides a control method for a multi-rotor aircraft, comprising the steps of: S1. When performing pitching motion (forward and backward motion), the third rotor 10 and the fourth rotor 12 are controlled to rotate at the same speed, and the first rotor 6 and the second rotor 9 rotate at different speeds. The torques generated by the four rotors are mutually balanced to be 0. At this time, the aircraft first tilts to a certain extent (pitch motion), so that the rotor pull produces a horizontal component (forward and backward motion), causing the aircraft to move forward and backward; S2. When performing yaw motion, the third rotor 10 and the fourth rotor 12 are controlled to have the same rotation speed, the first rotor 6 and the second rotor 9 have the same rotation speed, and at the same time the rotation speed of the first rotor 6 is different from the rotation speed of the third rotor 10. The torque of the four rotors is unbalanced, causing the aircraft to yaw and turn.

[0031] The control method of this embodiment realizes forward and backward movement and yaw turning by controlling the rotation speeds of the third rotor 10, the fourth rotor 12, the first rotor 6 and the second rotor 9 to be different, wherein, when performing the forward and backward movement, it is only necessary to control the rotation speeds of the third rotor 10 and the fourth rotor 12 to be the same, and the rotation speeds of the first rotor 6 and the second rotor 9 to be different, and the aircraft first tilts to a certain degree (pitch movement), so that the rotor pull produces a horizontal component (forward and backward movement), causing the aircraft to produce forward and backward movement; and when performing the yaw movement, it is only necessary to control the rotation speeds of the third rotor 10 and the fourth rotor 12 to be the same, and the rotation speeds of the first rotor 6 and the second rotor 9 to be the same, and at the same time the rotation speeds of the first rotor 6 The speed of the first rotor 10 is different from that of the third rotor 10, which can make the torque of the four rotors unbalanced, causing the aircraft to yaw and turn. It can be seen that this embodiment can ensure that the aircraft can complete all flight actions such as forward, backward, yaw and turn by adopting an appropriate control strategy through a hybrid power system combined with the working principle of a multi-axis aircraft. The control is simple and reliable. Since the third rotor 10 and the fourth rotor 12 are powered by a gas turbine engine 1 and a planetary gearbox 3, and the first rotor 6 and the second rotor 9 are powered by a battery 5 and a first motor 7 and a second motor 8, the two are independent and can be easily adjusted. Therefore, it provides convenience and feasibility for the flexible control of the flight actions of the multi-rotor aircraft. Since the motor only needs to drive two rotors, compared with the traditional series hybrid mode, the size and weight of the battery and the motor can be greatly reduced in this case.

[0032] Preferably, if Figure 5 As shown, in the second preferred embodiment of the present application, the control method of the multi-rotor aircraft further includes the steps of: S3. When hovering, control the third rotor 10 and the fourth rotor 12 to have the same rotation speed, the first rotor 6 and the second rotor 9 to have the same rotation speed, and the rotation speed of the first rotor 6 and the third rotor 10 to have the same rotation speed. The torques of the four rotors are balanced and the thrust is the same as the weight of the aircraft, so that the aircraft hovers in the air.

[0033] In this embodiment, by controlling the speed of the third rotor 10 and the fourth rotor 12 to be the same, the speed of the first rotor 6 and the second rotor 9 to be the same, and the speed of the first rotor 6 and the speed of the third rotor 10 to be the same, the aircraft can remain stable without pitching or yaw motion. At this time, when the torques of the four rotors are balanced and the thrust is the same as the weight of the aircraft, the aircraft can be kept stationary in the air, meeting the needs of hovering operations.

[0034] Preferably, if Figure 6 As shown, in the third preferred embodiment of the present application, the control method of the multi-rotor aircraft further includes the steps of: S4. When taking off, control the third rotor 10 and the fourth rotor 12 to have the same speed, the first rotor 6 and the second rotor 9 to have the same speed, and the first rotor 6 and the third rotor 10 to have the same speed. When the torques of the four rotors are balanced and the thrust is greater than the weight of the aircraft, the aircraft takes off vertically, and when the torques of the four rotors are balanced and the thrust is less than the weight of the aircraft, the aircraft descends vertically.

[0035] In this embodiment, by controlling the speeds of the third rotor 10 and the fourth rotor 12 to be the same, the speeds of the first rotor 6 and the second rotor 9 to be the same, and the speeds of the first rotor 6 and the third rotor 10 to be the same, the aircraft can remain stable without pitching or yaw motion. At this time, when the torques of the four rotors are balanced and the thrust is greater than the weight of the aircraft, the aircraft can take off vertically, thereby meeting the take-off requirements of the aircraft; conversely, when the torques of the four rotors are balanced and the thrust is less than the weight of the aircraft, the aircraft descends vertically, thereby meeting the descent requirements of the aircraft.

[0036] Preferably, if Figure 7 As shown, in the fourth preferred embodiment of the present application, the control method of the multi-rotor aircraft further includes the steps of: S5. During normal operation, the power distribution unit 4 supplies power to the first motor 7, the second motor 8 and the third motor 13, the gas turbine engine 1 drives the planetary gearbox 3, and the planetary gear output shaft of the planetary gearbox 3 and the third motor 13 jointly drive the third rotor 10 and the fourth rotor 12 to rotate.

[0037] The control method of this embodiment is mainly aimed at Figure 2 The power-split multi-rotor aircraft gas turbine hybrid power system shown in the figure is provided with a third electric motor 13 which is provided with electric energy by a battery 5 and an electric energy distribution unit 4. When operating normally, the gas turbine engine 1 drives the planetary gear output shaft of the planetary gearbox 3 and the third electric motor 13 to jointly drive the third rotor 10 and the fourth rotor 12 to rotate, thereby improving the power performance and providing strong power support during high-speed flight or rapid acceleration. At the same time, the combination of the two can reduce fuel consumption, improve fuel economy, and achieve low-cost and long-endurance flight.

[0038] Preferably, if Figure 8 As shown, in the fifth preferred embodiment of the present application, the control method of the multi-rotor aircraft further includes the steps of: S6. When the gas turbine engine is facing engine failure or to meet environmental noise requirements, the gas turbine engine 1, the planetary gearbox 3 and the generator 2 stop working, and at the same time, the mechanical connection between the planetary gear output shaft and the power distribution gearbox 11 is disconnected by the clutch, and the electric energy of the battery 5 is provided to the first motor 7, the second motor 8 and the third motor 13 through the power distribution unit 4 to drive all the rotors to rotate.

[0039] In this embodiment, when the gas turbine engine 1 faces engine failure or meets environmental noise requirements, in order to enable the aircraft to continue to perform flight operations safely and reliably, this embodiment, on the basis of stopping the operation of the gas turbine engine 1, the planetary gearbox 3 and the generator 2, simultaneously disconnects the mechanical connection between the planetary gear output shaft and the power distribution gearbox 11 through the clutch, and the electric energy of the battery 5 is provided to the first motor 7, the second motor 8 and the third motor 13 through the power distribution unit 4 to drive all the rotors to rotate. At this time, even if the gas turbine engine faces engine failure or has to be shut down to meet environmental noise requirements, this embodiment can still provide electric energy to the first motor 7, the second motor 8 and the third motor 13 through the battery 5 and the electric energy through the power distribution unit 4 to drive all the rotors to rotate, thereby ensuring that all the rotors can be temporarily rotated by directly providing electric energy to the battery 5, ensuring that the aircraft can still fly normally without being affected in special occasions where the gas turbine engine 1 is not suitable for operation (such as engine failure or to meet environmental noise requirements), thereby improving the adaptability, robustness and reliability of aircraft control.

[0040] Although the preferred embodiments of the present application have been described, those skilled in the art may make other changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications falling within the scope of the present application.

[0041] Obviously, those skilled in the art can make various changes and modifications to the present application without departing from the spirit and scope of the present application. Thus, if these modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is also intended to include these modifications and variations.

Claims

1. A power split multi-rotor aircraft gas turbine hybrid power system, characterized in that: The invention comprises a gas turbine engine (1), a planetary gearbox (3), a generator (2), an electric energy distribution unit (4), a battery (5), a first electric motor (7) for driving a first rotor (6), a second electric motor (8) for driving a second rotor (9), and a power distribution gearbox (11), wherein: the planetary gearbox (3) comprises planetary gears and a planetary carrier (33), a sun gear output shaft (31), a planetary gear output shaft, and an outer gear ring (32); the output shaft of the gas turbine engine (1) is drivingly connected to the planetary gears and the planetary carrier (33); the sun gear output shaft (31) is drivingly connected to the input end of the generator (2); and the planetary gear output shaft (32) is connected to the input end of the generator (2). The output shaft is connected to the outer gear ring (32); the planetary gear output shaft is drivingly connected to the input end of the power distribution gearbox (11); the power output end of the generator (2) is respectively connected to the second motor (8) driving the second rotor (9), the first motor (7) driving the first rotor (6), and the battery (5) circuit through the power distribution unit (4); the output end of the power distribution gearbox (11) is drivingly connected to the third rotor (10) and the fourth rotor (12); the power distribution unit (4) adopts a power electronic integrated component, including an AC / DC converter, a DC / DC converter, a DC / AC converter, and a DC high-voltage bus. The alternating current of the generator (2) is input into the direct current high voltage bus through the AC / DC converter, and the battery (5) realizes bidirectional energy exchange of the direct current high voltage bus through the DC / DC converter. The electric energy of the direct current high voltage bus is converted into alternating current through the DC / AC converter and then drives the first motor (7) and the second motor (8). The first motor (7) drives the first rotor (6) to rotate, and the second motor (8) drives the second rotor (9) to rotate.

2. The power split multi-rotor aircraft gas turbine hybrid power system according to claim 1, characterized in that: It also includes a third motor (13) connected to the circuit of the power distribution unit (4), the output end of the third motor (13) is drivingly connected to the input end of the power distribution gearbox (11), and a power cut-off device is provided between the planetary gear output shaft and the power distribution gearbox (11).

3. The power split multi-rotor aircraft gas turbine hybrid power system according to claim 2, characterized in that: The power disconnection device adopts a clutch, and the mechanical connection between the planetary gear output shaft and the power distribution gear box (11) is disconnected or connected by the clutch.

4. The power split multi-rotor aircraft gas turbine hybrid power system according to claim 1, characterized in that: The battery (5) is a power type battery.

5. A multi-rotor aircraft, characterized in that: The invention comprises a power split multi-rotor aircraft gas turbine hybrid power system as claimed in any one of claims 1 to 4.

6. A control method for a multi-rotor aircraft as claimed in claim 5, characterized in that: Includes steps: When performing pitching motion, the third rotor (10) and the fourth rotor (12) are controlled to rotate at the same speed, the first rotor (6) and the second rotor (9) rotate at different speeds, the torques generated by the four rotors are mutually balanced to be 0, and the aircraft first tilts to a certain degree, so that the rotor pull produces a horizontal component, causing the aircraft to move forward and backward; when performing yaw motion, the third rotor (10) and the fourth rotor (12) are controlled to rotate at the same speed, the first rotor (6) and the second rotor (9) rotate at the same speed, and at the same time the first rotor (6) and the third rotor (10) rotate at different speeds, the torques of the four rotors are unbalanced, causing the aircraft to yaw and turn.

7. The control method according to claim 6, characterized in that: Also includes the steps: When hovering, the third rotor (10) and the fourth rotor (12) are controlled to rotate at the same speed, the first rotor (6) and the second rotor (9) are controlled to rotate at the same speed, and the rotation speed of the first rotor (6) and the rotation speed of the third rotor (10) are controlled to be the same, the torques of the four rotors are balanced and the thrust is the same as the weight of the aircraft, so that the aircraft is hovering in the air; The third rotor (10) and the fourth rotor (12) are controlled to have the same rotational speed, the first rotor (6) and the second rotor (9) are controlled to have the same rotational speed, and at the same time, the rotational speed of the first rotor (6) and the rotational speed of the third rotor (10) are controlled to have the same rotational speed. The torques of the four rotors are balanced and the thrust is greater than the weight of the aircraft, so that the aircraft takes off vertically. When the torques of the four rotors are balanced and the thrust is less than the weight of the aircraft, the aircraft descends vertically.

8. The control method according to claim 6, characterized in that: The invention also includes the following steps: during normal operation, the power distribution unit (4) supplies power to the first motor (7), the second motor (8) and the third motor (13); the gas turbine engine (1) drives the planetary gearbox (3); the planetary gear output shaft of the planetary gearbox (3) and the third motor (13) jointly drive the third rotor (10) and the fourth rotor (12) to rotate.

9. The control method according to claim 8, characterized in that: The method also includes the following steps: When the gas turbine engine is facing engine failure or in order to meet environmental noise requirements, the gas turbine engine (1), the planetary gearbox (3) and the generator (2) stop working, and at the same time, the mechanical connection between the planetary gear output shaft and the power distribution gearbox (11) is disconnected by the clutch, and the electric energy of the battery (5) is provided to the first motor (7), the second motor (8) and the third motor (13) through the power distribution unit (4) to drive all the rotors to rotate.

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