CN109383786B

CN109383786B - Emergency escape power assisting system assembly of aero-engine

Info

Publication number: CN109383786B
Application number: CN201811015569.XA
Authority: CN
Inventors: 周禹博; 高峰; 王景生; 商庆海; 周庆党; 李明赢; 商凯宇
Original assignee: Liaoning Tongxinyuan Technology Co ltd
Current assignee: Shenyang Qiande Electronic Technology Co ltd
Priority date: 2018-08-31
Filing date: 2018-08-31
Publication date: 2022-05-31
Anticipated expiration: 2038-08-31
Also published as: CN109383786A

Abstract

An aeroengine escape assistance system assembly further comprises: the coaxial oil/electricity hybrid driving danger-escaping device specifically comprises: the system comprises a pair of ratchets, a pair of differential sensors and a pair of escape detection units, wherein the pair of ratchets are respectively coaxially connected with the pair of electric/power generation dual-purpose machines and the pair of engines, the pair of differential sensors are connected with the pair of escape detection units, and the pair of escape detection units are contained in the system controller. The danger escaping device ensures that the electric function of any electric/power generation dual-purpose machine can still work when the engine fails, and improves the safety performance of the multi-rotor helicopter.

Description

Emergency escape power assisting system assembly of aero-engine

Technical Field

The application relates to the technical field of helicopters, in particular to an aeroengine power assisting device, an aeroengine danger-escaping power assisting device and an aeroengine danger-escaping power assisting system.

Background

The difference between a multi-rotor helicopter and a single-rotor helicopter is that: the multi-rotor helicopter flies by adopting n pairs of propellers and needs to carry out power control on the attitude balance of each pair of propellers. The main rotor of the single-rotor helicopter is positioned on the center shaft, and naturally droops to obtain balance. The advantages and disadvantages are as follows: when one propeller fails, the rest rotor propellers of the multi-rotor helicopter can continue to finish the work of navigation or safe landing, and the single-rotor helicopter is damaged and killed. Obviously, the practical use of a multi-rotor helicopter to replace a single-rotor helicopter is a technical expectation.

At present, the multi-rotor helicopter driven by a fuel engine has a delayed control reaction after waiting for two strokes of air exhaust and air suction due to the working principle of the engine. After the response comes, the engine continues to suck a small amount of given oil again in an increasing mode at an interval of two strokes by using the rotating speed and the suction force changed by sucking the mixed gas for the previous time, so that the acceleration gradually changes, and then the attitude of the multi-rotor helicopter swaying in the wind is adjusted, which is not time for all, and obviously, the probability of falling is very high.

The problem does not exist in the electric multi-rotor helicopter. However, the specific energy of the storage battery can not reach 300Wh/kg at present, that is to say, in order to increase the load of the electric multi-rotor helicopter and the electricity consumption for the endurance time, the weight of the storage battery is increased according to the restriction of 300Wh/kg, the increased weight of the storage battery becomes the self weight of the multi-rotor helicopter, obviously, the load capacity is reduced in an equivalent way, and the increase and decrease in inverse proportion can reach the takeoff limit that the load weight is zero although the propeller can rotate for a long time. Therefore, the current electric multi-rotor helicopter is limited by a storage battery technology and cannot meet the actual requirement of endurance capacity with large load.

Therefore, an oil/electric hybrid scheme that makes up for the deficiency is a necessary choice. However, one mountain is not compatible with two tigers, and two different power sources are adopted on one multi-rotor helicopter, so that the cruising ability can be increased by configuring the specific energy of the two power sources, and the control of the two power sources can be switched, so that the helicopter is not broken when wind blows, and the technical problem to be solved urgently is formed.

Disclosure of Invention

Therefore, the aircraft engine escape assisting system assembly is needed to solve the problem that the multi-rotor helicopter is poor in cruising ability.

The aero-engine power assist system assembly comprising: an electric power-assisted system with a plurality of rotors driven by an engine, an aircraft engine balance subsystem and a coaxial oil/electricity hybrid driven danger-escaping device.

Therefore, it is necessary to provide an aircraft engine power assisting system for solving the problem of poor cruising ability of the multi-rotor helicopter.

The aeroengine power-assisted system comprises: aeroengine booster unit.

The aeroengine power assisting device comprises: the power generation device comprises an engine, an electric/power generation dual-purpose machine and a propeller, wherein the engine, the electric/power generation dual-purpose machine and the propeller are coaxially connected, and the engine is used for providing total power for the aero-engine power assisting device.

In one embodiment, the engine is further configured to power a propeller.

In one embodiment, the engine is further configured to provide power for charging the battery.

The motor/generator is used for applying 1/2 calibration voltage in a normal state, wherein 1/2 calibration voltage is a reference voltage of a motor power-assisted state, a voltage greater than 1/2 is applied to the motor power-assisted state, a voltage less than 1/2 is applied to the motor power-assisted state, and a voltage of 0 is applied to the generator power-assisted state.

The calibration voltage refers to a voltage value applied to the electric/power generation dual-purpose machine set according to the change of the takeoff load of the multi-rotor helicopter during each takeoff, and the voltage value on the electric/power generation dual-purpose machine is used as an electric power-assisted reference voltage for each different load flight control.

Wherein 1/2 calibration voltage makes lift that each screw produced + lift that each engine produced-many rotor helicopter overcome/n screw 0, promptly: the voltage required by the multi-rotor helicopter for hovering in the air.

In one embodiment, the electric/power generation dual-purpose machine provides instant torque to the propeller through the transmission shaft when an electric boosting state within a calibration voltage range is applied.

In one embodiment, the output power of the electric motor/generator is set to be not more than 5% of the output power of the fuel engine.

In one embodiment, when the dual-purpose motor/generator is in a generating state, the storage battery is supplemented with electric energy.

Aeroengine booster unit still includes: the system comprises a torque sensor, a rotating speed measurer, a flight controller, a system controller and a storage battery, wherein the torque sensor and the rotating speed measurer are used for respectively measuring the torque and the rotating speed of the engine, the electric/power generation dual-purpose machine and the propeller, measurement and control are implemented by an instruction system controller of the flight controller, the storage battery provides electric energy for the aero-engine power assisting device, and the capacity of the storage battery is not more than the power consumption of the electric/power generation dual-purpose machine for 30 seconds.

The system controller includes: the device comprises a power detection unit, a flight control instruction processing unit, an electric/power generation switching unit, an electric quantity detection unit, an engine power allocation unit and a charging and discharging control unit.

The system controller is used for storing an actually measured motor rotating speed/torque/voltage characteristic curve table into the system controller in advance, storing an actually measured fuel engine rotating speed/torque/fuel consumption characteristic curve table into the system controller in advance and storing an actually measured propeller rotating speed/torque/lift force characteristic curve table into the system controller in advance before a measurement and control embodiment.

In one embodiment, the weight is filled to the maximum takeoff weight of the multi-rotor helicopter design, then the motor applies rated voltage, then the power of the engine is adjusted to control the multi-rotor helicopter to hover in the air, and the propeller rotation speed, the torque and the oil consumption parameters of each engine under the full load state of the takeoff weight are saved in a system controller, and the minimum oil motion parameters are marked as A1 group. The motor voltage is then gradually reduced to 0 volts (actual measurement is 0 volts) while the engine power is adjusted to control the multi-rotor helicopter while still hovering in the air, and the rotor speed, torque and power per engine at this full takeoff weight are saved to the system controller, labeled as set a0 maximum oil parameters.

The system controller implements measurement and control, including: executing a set program, measuring dynamic parameters of the multi-rotor helicopter, automatically calibrating voltage and adjusting and controlling the attitude.

In one embodiment, the process of measuring the dynamic parameters of the actual takeoff weight of the multi-rotor helicopter and automatically calibrating the voltage comprises the following steps:

starting the engine, gradually increasing the output power of the engine, and measuring the ground clearance of the multi-rotor helicopter: when the output power of the engine is larger than the maximum oil-dynamic power of the A0 group, the ground clearance is still 0, the engine is judged to be overweight, and the takeoff is stopped; when the output power of the engine is still less than the maximum oil-kinetic power of the A0 group, and the preset ground clearance is reached, the voltage is gradually applied to the motor, the output power of the engine is synchronously reduced, and the multi-rotor helicopter is kept hovering in the air until the output power of the engine is equal to the power of the minimum oil-kinetic parameter of the A1 group, and the voltage applied to the motor at the moment is set as a calibration voltage. And then gradually increasing the output power of the engine, synchronously reducing the voltage on the motor to 1/2 calibration voltage which is the normal working reference voltage of the flight, and ending the process.

In one embodiment, the attitude adjustment control process of the multi-rotor helicopter under the actual takeoff weight comprises the following steps:

reading the instant torque and the instant rotating speed of a torque sensor and a rotating speed measurer according to the inclination angle alpha and the angular momentum j of the inclination speed of the multi-rotor helicopter notified by a flight controller, searching the actual rotating speed/torque/voltage characteristic curve table of the actual measurement motor, the actual rotating speed/torque/oil consumption characteristic curve table of the actual measurement fuel engine, the related parameters of the actual measurement propeller rotating speed/torque/lift characteristic curve table, the minimum oil motion parameters marked as A1 group and the maximum oil motion parameters marked as A0 group, and calculating the oil quantity and the voltage required for overcoming the angular momentum and the torque required for straightening according to the relative parameters, the voltage increased and decreased by constantly and quantitatively adjusting the motor, and the synchronous quantitative addition, And reducing the oil supply of the engine. When the multi-rotor helicopter inclination angle alpha notified by the flight controller is equal to 0, the adjustment is ended.

The aero-engine power assisting device obtains the inclination parameters of the multi-rotor helicopter and calculates the power output by the synchronous adjusting motor and the engine according to the relevant parameters stored in advance. It is clear that the instantaneous boost torque required for the swing is provided by adjusting the operating voltage of the electric motor, and then the instantaneous boost torque provided by the electric motor is gradually replaced by the engine torque coming later. Therefore, the instant power assisting effect is realized through the mode that the electric power and the oil power are successively relayed, on one hand, the torque that the oil power lags behind is compensated by means of the electric sensitive reaction characteristic, on the other hand, the oil power timely replaces the instant power assisting torque provided by the motor immediately after the oil power, so that the electric power consumption is very small, a large-capacity storage battery is not needed to be configured, the dead weight of the oil power multi-rotor helicopter is not needed to be increased, the advantages and the disadvantages are made up, the probability of breaking is reduced, and the endurance capacity is improved by hundreds of times.

The aero-engine power assisting system further comprises a power balance controller, a pair of forward and reverse rotating engines, a pair of forward and reverse rotating motors, a pair of forward and reverse rotating propellers and a pair of forward and reverse rotating system controllers, wherein the controlled end of the power balance controller is connected with the fly controller, the left side of the power balance controller is connected with the left electric/power generation switching unit to control the left propeller to rotate leftwards, and the right side of the power balance controller is connected with the right electric/power generation switching unit to control the right propeller to rotate rightwards. When the command of the flight controller rises, the power balance controller controls the pair of electric/power generation dual-purpose machines to be switched into the motor mode simultaneously through the pair of system controllers. When the flight controller commands the correction of the inclination angle, the power balance controller controls the pair of electric/power generation dual-purpose machines to respectively switch into the voltage increasing and decreasing complementary modes through the pair of system controllers, so that the left propeller increases the speed, the right propeller reduces the speed, and the speed increasing value is equal to the speed reducing value.

Above-mentioned aeroengine helping hand system through setting up the power balance control ware for when aeroengine helping hand system takes place the slope, realize controlling aeroengine helping hand system and ajust sensitively, avoid appearing falling down the phenomenon, improved aeroengine helping hand system's flight stability.

The aeroengine danger-escaping power assisting device comprises: a backstop, a differential sensor and a danger escaping detection unit. The method is characterized in that: the backstop is arranged between the electric/power generation dual-purpose machine and the fuel engine and is coaxially connected, the differential sensor is connected with the danger escape detection unit, and the danger escape detection unit is contained in the system controller.

The danger escaping detection unit is characterized in that: the rotation speeds of the shafts at the two ends of the backstop are monitored by the differential sensor at any time, the shafts at the two ends of the backstop are rigidly connected when the fuel engine rotates normally, and when the fuel engine stops due to faults or the rotation speed of the electric/power generation dual-purpose machine is greater than that of the fuel engine, the shafts at the two ends of the backstop are disconnected, the rotation speed difference is monitored by the danger escaping detection unit, and the system controller carries out emergency treatment.

The danger-escaping power assisting device for the aircraft engine ensures that the electric function of the electric/power generating dual-purpose machine can still work when the fuel engine breaks down, and improves the safety performance of the multi-rotor helicopter.

The aero-engine danger-escaping power-assisted system further comprises a power balance controller, a pair of forward and reverse rotating engines, a pair of forward and reverse rotating motors, a pair of forward and reverse rotating propellers and a pair of forward and reverse rotating system controllers, wherein the controlled end of the power balance controller is connected with the fly controller, the left side of the power balance controller is connected with the left electric/power generation switching unit to control the left propeller on the left side to rotate leftwards, and the right side of the power balance controller is connected with the right electric/power generation switching unit to control the right propeller on the right side to rotate rightwards. When the command of the flight controller rises, the power balance controller controls the pair of direct current dual-purpose machines to be switched into the motor mode simultaneously through the pair of system controllers. When the flight controller commands to correct the inclination angle, the power balance controller controls the pair of direct current dual-purpose machines to respectively switch into a voltage increasing and decreasing complementary mode through the pair of system controllers, so that the left propeller increases the speed, the right propeller reduces the speed, and the speed increasing value is equal to the speed reducing value.

According to the aero-engine danger-escaping power-assisted system, the power balance controller is arranged, so that when the aero-engine power-assisted system assembly inclines, the aero-engine power-assisted system can be sensitively controlled to be aligned, the phenomenon of falling is avoided, and the flight stability of the aero-engine power-assisted system assembly is improved.

Drawings

FIG. 1 is a schematic diagram of an embodiment of an integrated system for a paired oil/electric hybrid rotor;

FIG. 2 is a schematic structural diagram of an assembly of a paired oil/electric hybrid rotor according to another embodiment;

FIG. 3 is a schematic flow chart of measuring dynamic parameters of actual takeoff weight of a multi-rotor helicopter and automatically calibrating voltage in one embodiment;

FIG. 4 is a schematic view of a process flow of attitude adjustment control under actual takeoff weight of the multi-rotor helicopter in one embodiment;

FIG. 5 is a flow chart illustrating a method of boosting an engine according to an embodiment;

FIG. 6 is a schematic structural diagram of an exemplary aero-engine power assist system assembly;

FIG. 7 is a schematic structural diagram of an aircraft engine power assist system assembly in another embodiment;

FIG. 8 is a schematic structural diagram of an aero-engine emergency power assist system assembly according to an embodiment;

FIG. 9 is a schematic structural diagram of an aero-engine danger-escaping power-assisted system assembly in another embodiment.

Detailed Description

For better understanding of the objects, technical solutions and technical effects of the present invention, the present invention will be further explained with reference to the accompanying drawings and examples. Meanwhile, the following described examples are only for explaining the present invention, and are not intended to limit the present invention.

Referring to fig. 1 and 2, the assembly system of the oil/electric hybrid rotor includes: the combined type hybrid power rotor comprises an engine 101, a dual electric/power generation machine 102 and a propeller 100, wherein the engine 101, the dual electric/power generation machine 102 and the propeller 100 are coaxially connected, and the engine 101 is used for providing total power for an assembly system of the paired oil/electric hybrid power rotors.

In one embodiment, the engine 101 is also used to power the propeller 100.

In one embodiment, the engine 101 is also used to power the battery 105 for charging.

The electric/power generation dual-purpose machine 102 is used for applying 1/2 calibration voltage in a normal state, wherein the 1/2 calibration voltage is a reference voltage in an electric power assisting state, the calibration voltage greater than 1/2 is applied in an electric power-on state, the calibration voltage less than 1/2 is applied in an electric power-off state, and the calibration voltage 0 is applied in a power generation state.

The calibration voltage refers to a voltage value applied to the electric/power generation dual-purpose machine 102 set according to the change of the takeoff load of the multi-rotor helicopter during each takeoff, and the voltage value on the electric/power generation dual-purpose machine 102 is used as an electric power-assisted reference voltage for each different load flight control.

In one embodiment, the electric/electric generator 102 provides a transient torque to the propeller 100 through the transmission shaft when applying the electric power assist condition within the calibrated voltage range.

In one embodiment, the output power of the electric/electric power generator 102 is set to be not more than 5% of the output power of the fuel engine 101.

In one embodiment, when the motor/generator 102 is in a generating state, the battery 105 is replenished with electric energy.

The assembly system of the pair of oil/electricity hybrid rotors further comprises: the system comprises a torque sensor 106, a rotating speed measurer 107, a flight controller 104, a system controller 108 and a storage battery 105, wherein the torque sensor 106 and the rotating speed measurer 107 are used for respectively measuring the torque and the rotating speed of the engine 101, the electric/power generation dual-purpose machine 102 and the propeller 100, measurement and control are carried out by the instruction system controller 108 of the flight controller 104, the storage battery 105 provides electric energy for the assembly system of the pair of oil/electricity hybrid power rotors, and the capacity of the storage battery 105 is not more than the electric quantity of the electric/power generation dual-purpose machine 102 for 30 seconds.

The system controller 108 includes: power detection unit 200, flight control instruction processing unit 301, electric/power generation switching unit 300, electric quantity detection unit 400, engine power allocation unit 401, and charge/discharge control unit 500.

The system controller 108 is configured to, before a measurement and control embodiment, store the actually measured motor speed/torque/voltage characteristic curve table in the system controller 108 in advance, store the actually measured fuel engine speed/torque/fuel consumption characteristic curve table in the system controller 108 in advance, and store the actually measured propeller speed/torque/lift characteristic curve table in the system controller 108 in advance.

In one embodiment, the weight is filled to the maximum takeoff weight of the multi-rotor helicopter design, the motor is then applied with a rated voltage, the power of the engine 101 is adjusted to control the multi-rotor helicopter to hover in the air, and the rotating speed and torque of the propeller 100 and the fuel consumption parameters of each engine 101 in the full load state of the takeoff weight are saved in the system controller 108, labeled as group a1 minimum oil-kinetic parameters. The motor voltage is then gradually reduced to 0 volts (actual measurement is 0 volts) while the engines 101 are adjusted to control the multi-rotor helicopter while still hovering in the air, and the speed, torque and power of each engine 101 of the propellers 100 at full takeoff weight are saved to the system controller 108, labeled as set a0 of maximum oil-dynamic parameters.

The system controller 108 performs measurement and control, including: executing a set program, measuring dynamic parameters of the multi-rotor helicopter, automatically calibrating voltage and adjusting and controlling the attitude.

In one embodiment, referring to fig. 3, a process of measuring a dynamic parameter of an actual takeoff weight of a multi-rotor helicopter and automatically calibrating a voltage includes:

starting engine 101, gradually increasing the output power of engine 101, and measuring the ground clearance of the multi-rotor helicopter: when the output power of the engine 101 is larger than the maximum oil-dynamic power of the A0 group, the ground clearance is still 0, the engine is judged to be overweight, and the takeoff is stopped; when the output power of the engine 101 is less than the maximum oil-dynamic power of the A0 group, and the preset ground clearance is reached, gradually applying voltage to the electric motor, synchronously reducing the output power of the engine 101, and keeping the multi-rotor helicopter hovering in the air until the output power of the engine 101 is equal to the power of the minimum oil-dynamic parameter of the A1 group, and setting the voltage applied to the electric motor at the moment as a calibration voltage. Then, the output power of the engine 101 is gradually increased, the voltage on the motor is synchronously reduced to 1/2, and the process is ended, wherein the nominal voltage is the normal working reference voltage of the flight.

In one embodiment, referring to fig. 4, the attitude adjustment control process under the actual takeoff weight of the multi-rotor helicopter includes:

reading the instant torque and the instant rotation speed of the torque sensor 106 and the rotation speed measurer 107 according to the inclination angle alpha and the angular momentum j of the inclination speed of the multi-rotor helicopter notified by the flight controller 104, searching the measured motor rotation speed/torque/voltage characteristic curve table stored in the system controller 108 in advance before one measurement and control embodiment, the measured fuel engine 101 rotation speed/torque/fuel consumption characteristic curve table stored in the system controller 108 in advance, the measured propeller 100 rotation speed/torque/lift characteristic curve table stored in the system controller 108 in advance, the related parameters marked as A1 group minimum oil motion parameters and marked as A0 group maximum oil motion parameters, calculating the oil quantity and voltage required for overcoming the angular momentum and the torque required for straightening, and constantly and quantitatively adjusting the increasing and decreasing voltages of the motor, and synchronously and quantitatively adding and subtracting oil supply of the engine 101. When the multi-rotor helicopter pitch angle α notified by the flight controller 104 is equal to 0, the adjustment is ended.

The assembly system of the pair of oil/electricity hybrid rotors obtains the inclination parameters of the multi-rotor helicopter and calculates the power output by the synchronous adjusting motor and the engine according to the relevant parameters stored in advance. It is clear that the instantaneous boost torque required for the swing is provided by adjusting the operating voltage of the electric motor, and then the instantaneous boost torque provided by the electric motor is gradually replaced by the engine torque coming later. Therefore, the instant power assisting effect is realized through the mode that the electric power and the oil power are successively relayed, on one hand, the torque that the oil power lags behind is compensated by means of the electric sensitive reaction characteristic, on the other hand, the oil power timely replaces the instant power assisting torque provided by the motor immediately after the oil power, so that the electric power consumption is very small, a large-capacity storage battery is not needed to be configured, the dead weight of the oil power multi-rotor helicopter is not needed to be increased, the advantages and the disadvantages are made up, the probability of breaking is reduced, and the endurance capacity is improved by hundreds of times.

The aircraft engine balancing subsystem comprising: the power balance controller 109, the controlled end of the power balance controller 109 is connected to the flight controller 104, the left side is connected to the left motor/generator switching unit 300, the left propeller 100 is controlled to rotate left, the right propeller 100 is controlled to rotate right, and the right propeller 100 is controlled to rotate right. When the flight controller 104 instructs to go up, the power balance controller 109 controls the pair of motors/generators 102 to simultaneously switch to the motor mode through the pair of system controllers 108. When the flight controller 104 instructs to correct the pitch angle, the power balance controller 109 controls the pair of motor/generators 102 to switch to the power generation reduction increase/decrease voltage complementation mode, respectively, through the pair of system controllers 108, so that the left-side propeller 100 is increased in speed while the right-side propeller 100 is decreased in speed, and the increase value is equal to the decrease value.

In one embodiment, the present invention provides a multi-rotor helicopter power system comprising: comprises a propeller 100, an engine 101 and an electric/power generation dual-purpose machine 102;

the motor/generator 102 is coaxially connected to the engine 101 and the propeller 100 at 20 c;

the electric/power generation dual-purpose machine 102 is used for receiving working voltage and switching to an electric working state or a power generation working state according to the working voltage.

Among them, the propeller 100 is rigidly connected to the shaft 20c as a system-driven end member of the multi-rotor helicopter power system. When the propeller 100 is in the working state, the power thereof is equal to the output power of the engine + the output power of the electric/power generation dual-purpose machine in the electric working state; or, the propeller power is the output power of the engine, which is the power consumed by the electric/electric generator in the power generation operation state.

The motor 101 is in contact with the shaft 20c, and drives the shaft 20c to rotate when operating. The engine 101 is a fuel-powered engine.

The motor/generator 102 includes two states, namely a motor and a generator, and the two operating states are respectively completed by the motor and the generator. Specifically, when the motor/generator 102 is in an electric operating state, the motor operates according to external power supply to drive the shaft 20 c; when the motor/generator 102 is in the generating operation state, the generator is driven by the shaft 20c to convert the mechanical energy into the electric energy for generating electricity. Due to the operating characteristics of the motor, the motor may generate a transient torque for the coaxial shaft 20c after starting.

After the electric/power generation dual-purpose machine 102 is connected to the working voltage, the electric/power generation dual-purpose machine is switched to the electric working state or the power generation working state according to the working voltage, that is, the electric/power generation dual-purpose machine 102 is switched to the motor or the power generator state according to the working voltage. When the operating voltage is greater than 0, the electric/power generation dual-purpose machine 102 is an electric motor. Specifically, when the working voltage is equal to one-half of the rated voltage, the output of the motor correspondingly enables the multi-rotor helicopter to keep a normal working state; when the working voltage is greater than one-half of the calibration voltage, the output of the motor correspondingly enables the multi-rotor helicopter to be in an acceleration state; when the working voltage is less than one-half of the rated voltage, the output of the motor correspondingly enables the multi-rotor helicopter to be in a deceleration state. When the operating voltage is equal to 0, the motor/generator 102 is a generator.

In the present embodiment, the motor is a momentary torque generated coaxially with 20c, and is used only for attitude correction for compensating for an energizing lag of the engine 101.

In one embodiment, in the electric operating state, the output power of the electric/electric generator 102 is smaller than the preset power; wherein the preset power is smaller than the output power of the engine 101.

The output power of the motor/generator 102, that is, the output power of the motor is limited so that the instantaneous torque generated by the motor can satisfy the compensation requirement for the stress application delay of the engine 101, thereby preventing the output power of the motor from being excessive and reducing the energy consumption of the motor. Based on the low-energy-consumption motor, the capacity of the storage battery 105 for supplying power to the motor can be reduced, and the overall weight of the multi-rotor helicopter can be effectively controlled.

Further, the motor/generator 102 with output power smaller than the preset power has smaller volume and weight, and is also beneficial to controlling the overall weight of the multi-rotor helicopter.

Generally, the preset power is 3% to 7% of the output power of the engine 101 in order to balance the power consumption and the performance of the motor. Preferably, the preset power is 5% of the output power of the engine 101.

The multi-rotor helicopter power system adopts two power sources of oil power and electric power, and provides instant power for the propeller when the electric/power generation dual-purpose machine is in an electric working state so as to overcome the defect of stress application lag generated by long power stroke of the engine and facilitate the sensitive state control of the multi-rotor helicopter. Meanwhile, when the power of the engine is sufficiently prepared, the electric motor/generator in the power generation working state can be used for generating power, and the electric quantity is reserved for the subsequent electric motor/generator in the power generation working state. Based on the method, the sensitive state control of the multi-rotor helicopter is met, meanwhile, a large-capacity storage battery does not need to be configured for the electric/power generation dual-purpose machine, the whole machine weight of the multi-rotor helicopter is effectively controlled, and the cruising ability of the multi-rotor helicopter is favorably ensured.

The embodiment of the present invention further provides a paired oil/electricity hybrid rotor assembly system based on the multi-rotor helicopter power system, please refer to fig. 1 and fig. 2, the paired oil/electricity hybrid rotor assembly system includes a flight controller 104, a storage battery 105, a plurality of torque sensors 106, a plurality of rotation speed measuring devices 107, a plurality of system controllers 108, and a plurality of multi-rotor helicopter power systems according to any of the above embodiments; the system controller 108, the torque sensor 106 and the rotating speed measurer 107 are in one-to-one correspondence with the multi-rotor helicopter power systems;

the system controller 108 is connected to the flight controller 104, the battery 105, the corresponding torque sensor 106, the corresponding rotational speed measuring device 107, the corresponding electric motor/generator 102, and the corresponding engine 101, respectively;

the torque sensor 106 is used for measuring the torque of the propeller 100, and the rotating speed measurer 107 is used for measuring the rotating speed of the propeller 100;

the battery 105 is also connected to the flight controller 104 and the electric motor/generator 102, respectively.

Typically, a multi-rotor helicopter includes an overall control system, i.e., flight controller 104, for coordinating and managing the various units of the overall helicopter. In the flight controller 104 provided in the present embodiment, the flight controller 104 is configured to obtain various parameters of the multi-rotor helicopter in flight, including an inclination angle, an angular momentum of an inclination speed, an actual takeoff weight, a calibration voltage of a motor during takeoff, a ground clearance, and the like.

The system controller 108 is at the intermediate control position of the multi-rotor helicopter, and is configured to pre-store the pre-measured rotational speed/torque/voltage characteristic curve data of the motor and the rotational speed/torque/voltage characteristic curve data of the propeller 100, and to receive various relevant parameters sent by the flight controller 104, the storage battery 105, the corresponding torque sensor 106, and the corresponding rotational speed measuring device 107 connected thereto, analyze and process the various relevant parameters, and control the power system of the corresponding multi-rotor helicopter, that is, the electric/electric generator 102 and the engine 101, according to the relevant parameters.

The torque sensor 106 is used to measure the instantaneous torque of the propeller 100. Wherein the torque sensor 106 includes a torque sensor 106. The torque sensor 106 is fixed to the shaft 20c and detects the instantaneous torque of the shaft 20 c.

The rotational speed measurer 107 is used to measure the instantaneous rotational speed of the propeller 100. The rotation speed measuring device 107 includes a rotation speed measuring device 107, such as an optical code disc or a hall sensor.

The battery 105 is a chargeable and dischargeable power source, and includes the battery 105 or a lithium battery or the like. The battery 105 is connected to the flight controller 104 and supplies power to the flight controller 104, including chip-level power supply or element-level power supply. In a preferred embodiment, the battery 105 can be a small-capacity battery 105 on the multi-rotor helicopter for supplying power to the flight control system, i.e., the battery 105 of the multi-rotor helicopter is provided, so that the multi-rotor helicopter does not add extra weight.

The battery 105 is also connected to the motor/generator 102, and it should be noted that the link formed by connecting the battery 105 to the motor/generator 102 includes a charging link and a discharging link. When the motor/generator 102 is an electric motor, the battery 105 supplies power to the electric motor through a discharge link; when the motor/generator 102 is a generator, the generator charges the battery 105 through a charging link.

In one embodiment, the size and weight of the battery 105 are controlled to control the overall weight of the multi-rotor helicopter. The storage battery 105 is selected from the storage batteries 105 with the capacity smaller than the preset capacity; the preset capacity is the electricity consumption of the electric/power generation dual-purpose machine 102 in the electric working state for a preset time period. When the electric/power generation dual-purpose machine 102 is in an electric working state, the electric/power generation dual-purpose machine works for a preset time period, so that the requirement for compensating the stress application lag of the engine 101 can be met, and sufficient torque is provided for controlling the flight attitude of the multi-rotor helicopter. In a preferred embodiment, the predetermined period of time is 30 seconds to ensure a balance between the performance, volume and weight of the battery 105. It should be noted that the capacity of battery 105 may be adapted to the type of helicopter.

Referring to fig. 6 and 7, in one embodiment, an aircraft engine power assist system assembly is provided, including: a pair of engines 101, a pair of electric motor/generator 102 and a pair of propellers 100, the engines 101, the electric motor/generator 102 and the propellers 100 are in one-to-one correspondence, and the corresponding engines 101, the electric motor/generator 102 and the propellers 100 are coaxially connected, the engines 101 provide total power for the aeroengine 101 power-assisted system assembly, characterized in that: the electric/power generation dual-purpose machine 102 sets a calibration voltage according to a takeoff calibration voltage program, wherein 1/2 calibration voltage is a reference voltage in an electric power assisting state, if the applied calibration voltage is more than 1/2, the electric power assisting state is achieved, if the applied calibration voltage is less than 1/2, the electric power reducing state is achieved, and if the applied voltage is 0, the output power of the electric/power generation dual-purpose machine 102 is not more than 5% of the output power of the engine 101;

aeroengine 101 helping hand system assembly still includes: a pair of torque sensors 106, a pair of rotation speed measuring devices 107, a battery 105, and a pair of system controllers 108, wherein the torque sensors 106 and the rotation speed measuring devices 107 respectively detect the torque and the rotation speed of the engine 101, the electric motor/generator 102, and the propeller 100, respectively, and the capacity of the battery 105 is not greater than the electric power consumption of the electric motor/generator 102 for 30 seconds;

the system controller 108 is configured to store the corresponding motor speed/torque/voltage characteristic curve table, the engine speed/torque/oil consumption characteristic curve table, the propeller speed/torque/lift characteristic curve table, the a1 set minimum oil motion parameter table, and the a0 set maximum oil motion parameter table;

the system controller 108 is further configured to perform: measuring dynamic parameters of the multi-rotor helicopter, setting a take-off calibration voltage of the multi-rotor helicopter and controlling the multi-rotor helicopter to adjust the attitude;

aeroengine 101 helping hand system assembly still includes: and the dynamic balance controllers 109 are respectively connected with the pair of system controllers 108, and are used for respectively controlling the rotating states of the corresponding propellers 100 through the system controllers 108.

Further, in one embodiment, the controlled end of the dynamic balance controller 109 is connected to the flight controller 104, a first side of the dynamic balance controller 109 is connected to the corresponding system controller 108 to control the left-hand rotation of the first side propeller 100, and another side of the dynamic balance controller 109 is connected to the corresponding system controller 108 to control the right-hand rotation of the second side propeller 100.

Further, in one embodiment, the dynamic balance controller 109 is configured to control the pair of motors/generators 102 to be simultaneously switched to the motor mode by the pair of system controllers 108 when the flight controller 104 instructs to ascend;

the dynamic balance controller 109 is further configured to control the pair of electric/power generation dual-purpose machines 102 to switch to a complementary mode of generating electricity and reducing the rotation speed and increasing the rotation speed electrically, respectively, through the pair of system controllers 108 when the flight controller 104 instructs to correct the tilt angle, so that the first side propeller 100 is increased in speed while the second side propeller 100 is reduced in speed, and the increased speed value is equal to the reduced speed value.

As shown in fig. 2, the system controller 108 includes a power detection unit 200. Power detection section 200 is connected to corresponding torque sensor 106 and corresponding rotational speed measuring device 107. The power detection unit 200 acquires the instantaneous torque detected by the torque sensor 106 through the connection with the torque sensor 106; meanwhile, the instantaneous rotation speed detected by the rotation speed measuring device 107 is obtained through connection with the rotation speed measuring device 107. Further, the power detection unit 200 may calculate the power of the coaxial line 20c according to the instant torque and the instant rotation speed. The instantaneous torque and the instantaneous rotational speed are converted into one of the types of control commands of the system controller 108 by the power detection unit 200.

The system controller 108 also includes a motor/power generation switching unit 300. The electric power/electric power generation switching unit 300 is connected to the power detection unit 200, the flight controller 104, and the electric power/electric power generation dual-purpose machine 102, respectively. The motor/generator switching unit 300 is configured to switch the operating state of the motor/generator 102 according to a switching control instruction. The switching control instruction includes a control instruction provided by the power detection unit 200 and a control instruction provided by the flight controller 104.

The system controller 108 also includes a flight control instruction processing unit 301. The electric power generation/conversion unit 300 is connected to the flight controller 104 via the flight control instruction processing unit 301. Flight control instruction processing unit 301 is configured to process an instruction set provided by flight controller 104. The flight controller 104 transmits to the instruction set of the system controller 108, and the flight control instruction processing unit 301 transmits, to the electric/power generation switching unit 300, an instruction for controlling the electric/power generation switching unit 300, which is one of the control instructions of the electric/power generation dual-purpose machine 102.

The system controller 108 also includes a power detection unit 400. The power amount detection unit 400 is connected to the electric/power generation switching unit 300 and the storage battery 105, respectively. The power detecting unit 400 is used for detecting the power of the battery 105 connected to the battery and providing another control instruction to the power/generation switching unit 300 according to the detected power. The control command provided by the electric quantity detection unit 400 for the electric/power generation switching unit 300 includes: when the electric quantity of the storage battery 105 is lower than the preset threshold value, the control instruction provided by the electric quantity detection unit 400 causes the electric/power generation switching unit 300 to switch the electric/power generation dual-purpose machine 102 to the power generation operation state.

Wherein, since the electric energy of the electric/power generation switching unit 300 in the power generation operation state is derived from the mechanical energy of the engine 101 driving the shaft 20c, in one embodiment, the system controller 108 further includes an engine power allocating unit 401; the engine power allocation unit 401 is connected to the electric quantity detection unit 400, the power detection unit 200, and the engine 101, respectively.

The instantaneous torque and the instantaneous rotation speed acquired by the power detection unit 200 mainly depend on the output power of the engine 101. The power detecting unit 200 is connected to the engine power allocating unit 401, and can send a control command to the engine power allocating unit 401 to adjust the output power of the engine 101 according to the measured instantaneous torque and instantaneous speed of the coaxial 20 c.

Meanwhile, the output power of the propeller 100 is the output power of the engine 101 — the power consumed by the electric/electric generator 102 in the electric power generating operation state. The engine power allocating unit 401 is connected to the electric quantity detecting unit 400, that is, the engine power allocating unit 401 can adjust the output power of the engine 101 according to the electric quantity of the battery 105.

The following explains the operation logic among the power detection unit 200, the motor/generator switching unit 300, the electric quantity detection unit 400, and the engine power allocation unit 401 by a specific application example, and it should be noted that the present application example is only for explanation and is not intended to limit the operation logic among the units.

The engine 101 is connected to the motor/generator 102 through a shaft 20c, and is connected to the propeller 100 through a shaft 20 c. The torque sensor 106 and the rotation speed measuring device 107 transmit the instantaneous torque and the instantaneous rotation speed to the power detection unit 200, respectively. The flight controller 104 sends out a control command to control acceleration, and when the rotating speed of the coaxial 20c is less than the hovering rotating speed n of the multi-rotor helicopter in the air, the engine power allocating unit 401 controls the rotating speed of the engine 101 to be increased or decreased, and the motor idles along with the engine 101. When the rotation speed of the shaft 20c is greater than n, the motor/generator switching unit 300 controls the motor to reach the incremental rotation speed nx without changing the oil supply amount of the engine 101.

When the voltage of the battery 105 is V0, the engine power allocating unit 401 controls the rotation speed of the engine 101 to be the sum of n and nx, and the motor/power generation switching unit 300 controls the motor to gradually switch to the generator to charge the battery 105. When the voltage of the battery 105 is V, the generator stops generating electricity. When the nx is equal to 0 in the circulation mode, the motor and the generator return to the normal state.

The system controller 108 further includes a charge and discharge control unit 500. The charge/discharge control unit 500 is connected to the electric quantity detection unit 400, the power detection unit 200, and the battery 105, respectively. The charge/discharge control unit 500 controls a charge/discharge link between the battery 105 and the motor/generator 102. The charge/discharge control unit 500 controls the charge/discharge link between the battery 105 and the motor/generator 102, including controlling the magnitude of the charge/discharge current, the magnitude of the charge/discharge voltage, the charge/discharge duration, and the like, based on the control command transmitted from the electric power detection unit 400 and the control command transmitted from the power detection unit 200.

In one embodiment, an aircraft engine escape assistance system assembly is also provided, and referring to fig. 8 and 9 together, the multi-rotor helicopter further includes a pair of dynamic balance controllers 109, and the dynamic balance controllers 109 are respectively connected to the flight controller 104 and each system controller 108.

In one embodiment, the aero-engine risk clearance assistance system assembly further comprises: the coaxial oil/electricity hybrid driving danger-escaping device specifically comprises: a pair of ratchets 111, a pair of differential sensors 112, and a pair of escape detection units 400, wherein the pair of ratchets 111 are coaxially connected between the pair of electric/power generators 102 and the pair of engines 101, the pair of differential sensors 112 are connected to the pair of escape detection units 400, and the pair of escape detection units 400 are included in the pair of system controllers 108;

the pair of escape detection units 400 is configured to: the rotation speeds of the shafts at the two ends of the pair of ratchets 111 are monitored in real time through the pair of differential sensors 112, the shafts at the two ends of the pair of ratchets 111 are rigidly connected when the pair of engines 101 rotate normally, and the shafts at the two ends of the pair of ratchets 111 are disconnected when the pair of engines 101 stop rotating due to faults or the pair of electric/power generation dual-purpose machines 102 rotate at a speed higher than that of the pair of engines 101, so that a rotation speed difference occurs; when the difference in the rotational speeds is detected by the pair of danger avoidance detecting units 400, emergency processing is performed by the pair of system controllers 108, respectively.

The danger escaping device ensures that the electric function of any electric/power generation dual-purpose machine can still work when the engine fails, and improves the safety performance of the multi-rotor helicopter.

In one embodiment, the multi-rotor helicopter further comprises a dynamic balance controller, which is connected to the flight controller 104 and each system controller 108, respectively. In one embodiment, the multi-rotor helicopter further comprises a gyroscope, and the gyroscope is connected with the dynamic balance controller.

According to the multi-rotor helicopter, the system controller is in the middle control position of the multi-rotor helicopter, various parameters required by the multi-rotor helicopter when different flight states are controlled are obtained through the connection of the flight controller, the storage battery, the torque sensor and the rotating speed measurer, and the power system of the multi-rotor helicopter is controlled through the parameters, namely the electric/power generation dual-purpose machine and the engine are comprehensively controlled, so that the multi-rotor helicopter can realize sensitive state control. The motor/generator is in an electric working state and provides instant power for the engine so as to solve the problem of power application lag of the engine. Because the engine stress application lag duration is short, the power generation power can be provided for the electric/power generation dual-purpose machine in the power generation working state after the engine power preparation is finished, and the flight controller only needs chip-level power supply, so that the multi-rotor helicopter does not need to select a large-capacity storage battery, the whole machine weight of the multi-rotor helicopter is effectively controlled, and the endurance capacity of the multi-rotor helicopter is favorably ensured.

Referring to fig. 5, in one embodiment, a detailed procedure for shimming a multi-rotor helicopter in which a pitch condition is to occur is described. In this embodiment, a power assisting method for an engine is provided, where the power assisting method specifically includes the following steps:

s202, acquiring the inclination parameters of the multi-rotor helicopter, wherein the inclination parameters comprise the inclination angle and the inclination speed of the multi-rotor helicopter, the calibration voltage of a motor during takeoff and the rotating speed of a propeller.

In particular, multi-rotor helicopters are highly likely to experience a pitch condition during flight, especially in windy environments. The collectable inclination parameters mainly comprise an inclination angle of the multi-rotor helicopter, an angular momentum of an inclination speed of the multi-rotor helicopter, a calibration voltage of a motor corresponding to a takeoff weight of the multi-rotor helicopter, a rotating speed of a propeller and the like. It should be noted that the calibration voltage of the motor is set as a reference point for each takeoff control according to the change of the takeoff weight of the multi-rotor helicopter when the multi-rotor helicopter takes off each time.

As an alternative embodiment, a flight controller in the multi-rotor helicopter can be used for acquiring the inclination angle of the multi-rotor helicopter in an inclined state and the angular momentum of the inclination speed of the multi-rotor helicopter; the calibration voltage of the motor corresponding to the takeoff weight of the multi-rotor helicopter can be preset in a database in a system controller, and the calibration voltage of the motor corresponding to the actual takeoff weight of the multi-rotor helicopter can be obtained by matching the actual takeoff weight measured during takeoff of the multi-rotor helicopter with a preset database in the system controller; the rotational speed of the propellers can be obtained using a rotational speed measurer in a multi-rotor helicopter.

And S204, adjusting the working voltage of the motor in the multi-rotor helicopter according to the tilt parameter, and adjusting the output power of the engine, wherein when the multi-rotor helicopter is not subjected to tilt adjustment, the working voltage of the motor is one-half of the calibration voltage.

Specifically, it is to be noted that, when the multi-rotor helicopter is flying at a constant speed (in a normal state), that is, when the multi-rotor helicopter is not performing tilt adjustment or acceleration/deceleration adjustment, the operating voltage of the motor is one-half of the calibration voltage. When the multi-rotor helicopter is judged to incline, the system controller immediately adjusts the working voltage of the motor in the multi-rotor helicopter according to the inclination parameters, and simultaneously, the output power of the engine is synchronously adjusted, so that the motor and the engine jointly complete the alignment of the multi-rotor helicopter.

And S206, adjusting the rotating speed of the propeller according to the adjusted working voltage and the output power until the multi-rotor helicopter is righted.

Specifically, the working voltage of the motor is adjusted, so that the corresponding torque is generated by changing the output power of the motor, and the torque is transmitted to the propeller through the transmission shaft, so that the propeller can rapidly change the rotating speed to straighten the multi-rotor helicopter. And meanwhile, the output power of the engine is adjusted to gradually replace the output power of the motor, so that the multi-rotor helicopter is straightened. The adjusting mode realizes that the torque required by the multi-rotor helicopter for straightening is obtained quickly by using the motor, then the output power of the engine is gradually adjusted until the output power of the engine can replace the output power of the motor, and the straightening of the multi-rotor helicopter is completed by coordinating the mode of combining the motor and the output power of the engine.

And S208, after the multi-rotor helicopter is straightened, adjusting the output power of the engine again, and enabling the working voltage of the motor to return to one-half of the calibrated voltage.

Specifically, after the multi-rotor helicopter is straightened, the output power of the engine is adjusted again, and the working voltage of the motor is returned to one-half of the rated voltage, namely, the multi-rotor helicopter is returned to a flying state of flying at a constant speed.

According to the power assisting method of the engine, the inclination parameters of the multi-rotor helicopter are obtained, the working states of the motor and the engine are adjusted according to the inclination parameters, the requirement is clear, the aligning torque required by aligning can be instantly provided for the propeller by adjusting the working voltage of the motor so as to quickly align the multi-rotor helicopter, meanwhile, the engine is gradually used for replacing the motor to provide the aligning torque required by aligning for the propeller, and therefore, the effect of aligning the multi-rotor helicopter is achieved through the combined action of the motor and the engine. The power assisting method of the engine and the multi-rotor helicopter adopt two power sources of oil motion and electric motion, on one hand, instantaneous power assisting is provided by means of an electric control mode, state control on the multi-rotor helicopter sensitively is achieved, and the aligning torque required by aligning the propeller can be quickly generated.

The power assisting method of the engine does not need a power supply with large capacity of electricity, so that the dead weight of the multi-rotor helicopter is reduced, the cruising ability of the multi-rotor helicopter is further improved, and the actual requirements of users are met. Further, in a preferred embodiment, the method relates to specific application parameters of the motor, the engine and the power supply, and specifically, the boosting method further comprises the following steps:

and controlling a power supply in the multi-rotor helicopter to apply voltage to the motor, wherein the output power of the motor is not more than 5% of the output power of the engine, and the capacity of the power supply is not more than the power consumption of the motor for 30 seconds.

As an alternative embodiment, a generator which is provided with an engine and provides power for a flight control system in the multi-rotor helicopter is changed into a dual-purpose motor/generator which is controlled and switched by a system controller. The improved electric/power generation dual-purpose machine is coaxially connected with the engine and the propeller. And still adopt the small capacity battery of original flight control system power consumption. This equates to no increase in weight of the electric system, maintaining the range of the engine drive. The battery removes the power consumption for a flight control system, the battery capacity is smaller than the power consumption of the motor for 30 seconds, and the cruising ability of the multi-rotor helicopter is ensured by adopting the lightest weight. And the motor is only used for compensating the posture correction of the stress application lag of the engine, the power of the motor is less than 5% of the power of the engine, the energy consumption of the cart drawn by the large horse is avoided, and the effect of improving the endurance capacity is obtained after the lightest weight is adopted to ensure that the oil/electric power is mixed.

In one embodiment, the method relates to a specific process for adjusting the working states of an engine and a motor. Wherein, S204 specifically comprises the following steps:

s2042, calculating a straightening torque required by straightening the multi-rotor helicopter according to the inclination angle and the inclination speed of the multi-rotor helicopter, the calibration voltage of a motor during takeoff and the rotating speed of a propeller;

and S2044, adjusting the working voltage of the motor in the multi-rotor helicopter according to the straightening torque, and adjusting the output power of the engine.

Specifically, when the multi-rotor helicopter inclines, the straightening torque required by straightening the multi-rotor helicopter is calculated according to the inclination angle and the inclination speed of the multi-rotor helicopter, the calibration voltage of the motor during takeoff and the rotating speed of the propeller, and the rotating speed/torque/lift characteristic relation of the propeller is obtained from the system controller. The system controller further quantitatively adjusts the operating voltage of the motor to be increased or decreased according to the swing torque, and quantitatively increases or decreases the fuel supply amount of the engine (i.e., increases or decreases the output of the engine) in synchronization.

Further, when the propeller reaches the required increased or decreased rotating speed by receiving the aligning torque, the motor correspondingly increases or decreases the working voltage along with the continuously increased output power of the engine until the voltage returns to one half of the rated voltage. The accuracy of the multi-rotor helicopter in the aligning process is improved by taking a plurality of inclination parameter indexes as the basis.

In one embodiment, the data preparation process for obtaining the calibration voltage of the motor specifically comprises the following steps:

s302, acquiring the maximum load parameter of the multi-rotor helicopter, and enabling the multi-rotor helicopter to be in a full-load state according to the maximum load parameter.

Specifically, the maximum load parameters may include a maximum load capacity, an accuracy of a center of gravity of the weight, and a gravity corresponding to the maximum load capacity. The multi-rotor helicopter is in a full-load state, namely a maximum load state. It should be noted that the accuracy of the center of gravity is a concern when placing weights.

S304, rated voltage is applied to the motor in the multi-rotor helicopter in the full-load state, and the output power of the engine is adjusted according to the maximum load parameter, so that the multi-rotor helicopter can hover in the air.

Specifically, rated voltage is applied to a motor in the multi-rotor helicopter, the output power of an engine is continuously adjusted, and the multi-rotor helicopter in a full-load state is controlled to hover in the air.

S306, acquiring the rotating speed of a propeller, the torque of the propeller and the output power of the engine when the fully loaded multi-rotor helicopter hovers in the air, wherein the rotating speed, the torque of the propeller and the output power of the engine are used as first hovering parameters of the multi-rotor helicopter.

Specifically, the propeller rotation speed, the propeller torque, and the output power of each engine of the multi-rotor helicopter in the above-described state (i.e., the full-load state of the multi-rotor helicopter hovering in the air) are taken as the first hovering parameters of the multi-rotor helicopter. Further, the first hover parameter is saved in the system controller for quick recall when subsequently needed. Through obtaining each item flight parameter under the full load state of many rotor helicopters, guaranteed that many rotor helicopters can not damage its inside part because of the overload, avoided the risk that its working life reduces by a wide margin.

In another embodiment, the data preparation process is involved to further obtain a nominal voltage of the motor. Wherein, S306 comprises the following steps:

and S308, reducing the working voltage of the motor to zero volt, and adjusting the output power of the engine to keep the multi-rotor helicopter hovering in the air.

Specifically, in combination with the previous embodiment, after the first hovering parameter of the multi-rotor helicopter is obtained, the operating voltage of the motor is gradually decreased to zero volts (the actual measurement value is zero volts), and the output power of the engine is adjusted to ensure that the torque required for the propeller to be straightened is transmitted, so as to control the multi-rotor helicopter to hover in the air.

And S310, acquiring the rotating speed of a propeller, the torque of the propeller and the output power of the engine when the multi-rotor helicopter in the full-load state is suspended in the air when the multi-rotor helicopter does not have the motor and taking the rotating speed of the propeller, the torque of the propeller and the output power of the engine as second suspension parameters of the multi-rotor helicopter.

Specifically, the rotor speed, the rotor torque, and the output power of each engine of the multi-rotor helicopter in the above state (i.e., when the multi-rotor helicopter in a full load state is suspended without an electric motor) are used as the second hovering parameters of the multi-rotor helicopter. The second hover parameter is saved in the system controller for quick invocation at a later time if needed. By further acquiring various flight parameters of the multi-rotor helicopter in a full-load state when the multi-rotor helicopter works without a motor, the maximum load capacity of the multi-rotor helicopter during takeoff can be accurately obtained so as to provide reference for placing heavy objects.

In one embodiment, the method relates to a specific process for setting the calibration voltage of the motor in the multi-rotor helicopter. Wherein, this helping hand method still includes:

and S402, acquiring the takeoff weight of the multi-rotor helicopter.

Specifically, the takeoff weight of the multi-rotor helicopter includes the weight and load weight of the multi-rotor helicopter. It should be noted that when placing a load (heavy object), the accuracy of the center of gravity needs to be considered, and the higher the accuracy, the higher the measurement accuracy of the calibration voltage. It is also noted that the multi-rotor helicopter needs to take off in a windless environment, so that the multi-rotor helicopter cannot tilt, and the measurement accuracy of the calibration voltage is ensured.

S404, adjusting the output power of the engine according to the takeoff weight, enabling the multi-rotor helicopter to fly away from the ground, and detecting the ground clearance of the multi-rotor helicopter.

Specifically, the engine is started first, and the output power of the engine is gradually adjusted according to the takeoff weight, so that the multi-rotor helicopter can successfully fly off the ground. Meanwhile, the ground clearance of the multi-rotor helicopter can be detected in real time by using a distance measuring sensor, and the distance measuring sensor optionally comprises but is not limited to an ultrasonic sensor, an infrared sensor and a laser sensor.

And S406, when the ground clearance reaches a target distance, applying an operating voltage to a motor in the multi-rotor helicopter and reducing the output power of the engine so as to keep the multi-rotor helicopter hovering in the air.

Specifically, when the ground clearance of the multi-rotor helicopter reaches a set target distance, the output power of the engine is adjusted first, so that the multi-rotor helicopter can hover in the air. Then, the working voltage is applied to the motor in the multi-rotor helicopter, and the output power of the engine is reduced, so that the multi-rotor helicopter always hovers in the air.

S408, when the output power of the engine is reduced to meet the corresponding output power in the preset first hovering parameter, taking the working voltage of the motor at the moment as the calibration voltage of the motor.

Specifically, when the output power of the engine is reduced to meet the corresponding output power in the first hovering parameter, that is, the output power of the engine reaches the output power of the engine when the multi-rotor helicopter in a full-load state hovers in the air, the working voltage of the motor at the time is used as the calibration voltage of the motor. Further, the nominal voltage of the motor is saved to the system controller. It is clear that, due to the different takeoff weights of the multi-rotor helicopters, different nominal voltages of the motors are obtained, and in general, each takeoff weight corresponds to a nominal voltage of one motor. Of course, it is also possible to correspond to the nominal voltages of a plurality of motors for each takeoff weight to meet different practical requirements, and the embodiment is not limited herein.

In one embodiment, S408 is followed by the step of:

and S409, reducing the working voltage of the motor and increasing the output power of the engine so as to keep the multi-rotor helicopter hovering in the air.

Specifically, in combination with the above embodiment, after obtaining the calibration voltage of the motor in the multi-rotor helicopter, it is necessary to further obtain the operating voltage of the motor in the normal state of the multi-rotor helicopter, that is, in the constant speed state. Firstly, the working voltage of the motor is reduced from the nominal voltage, and meanwhile, the output power of the engine is improved, so that the multi-rotor helicopter always hovers in the air.

And S410, when the working voltage of the motor is reduced to one half of the rated voltage, stopping adjusting the motor and the generator.

Specifically, when the operating voltage of the motor is reduced to one-half of the rated voltage, the operating voltage of the motor in the normal state of the multi-rotor helicopter, that is, in the constant speed state, is obtained, and at this time, the motor and the generator are not adjusted any more. The multi-rotor helicopter takes the working voltage of the motor and the output power of the generator as initial conditions for acceleration or deceleration.

With reference to the above embodiments, the present application fully describes the setting process of the calibration voltage through an embodiment. In the embodiment, an actually measured motor rotating speed/torque/voltage characteristic curve table is stored in a system controller in advance; the actually measured propeller rotating speed/torque/lift force characteristic curve table is also stored in the system controller in advance. The full load of N rotor wing complete machines (the accuracy of the center of gravity needs to be considered when heavy objects are placed) is measured in advance, rated voltage is applied to a motor, the power of an engine is adjusted to control the complete machine to hover in the air, and the rotating speed and the torque of a propeller and the power of each engine in the full load state of the takeoff weight at the moment are stored in a system controller and marked as first hovering parameters. And then gradually reducing the voltage of the motor to 0 volt (the actual measurement value is 0 volt), adjusting the power of the engine to control the whole machine to be suspended in the air, and storing the rotating speed and the torque of the propeller and the power of each engine in the full-load state of the takeoff weight at the moment into a system controller, wherein the parameters are marked as second suspension parameters.

Furthermore, the multi-rotor helicopter takes off in the windless environment during actual flight, and the take-off weight is T. And starting the engine, gradually adjusting the engine power p, and simultaneously measuring the ground clearance s of the multi-rotor helicopter. When the ground clearance reaches the preset height, the power of the engine is adjusted to keep the multi-rotor helicopter hovering in the air.

And then powering up the motor, adjusting the power of the engine to keep the multi-rotor helicopter hovering in the air at the same time, determining the voltage value v on the motor at the moment as a calibration voltage when the power of the engine is reduced to a value corresponding to the first hovering parameter, and storing the voltage value v into a system controller and marking the voltage value v as the calibration voltage. Then, the voltage of the motor is gradually reduced to 1/2 nominal voltage, and the power of the engine is adjusted to keep the multi-rotor helicopter hovering in the air. Whereby the calibration voltage setting is ended.

In one particular embodiment, a particular process is involved of how the propellers of a multi-rotor helicopter accelerate. In this embodiment, the method includes the following steps:

applying a working voltage to the motor to be more than one half of a calibration voltage so as to accelerate the propeller corresponding to the motor;

in the process of accelerating the propeller corresponding to the motor, increasing the output power of the engine and reducing the working voltage of the motor;

and stopping adjusting the motor and the generator until the working voltage of the motor is reduced to one-half of the rated voltage.

Specifically, when the multi-rotor helicopter inclines, one side of the multi-rotor helicopter is higher and the other side of the multi-rotor helicopter is lower, so that the propeller on the lower side needs to be accelerated, and the torque required by the acceleration of the propeller can be quickly obtained by applying a voltage to the motor corresponding to the propeller, wherein the voltage is greater than one half of a calibration voltage, so that the acceleration of the propeller corresponding to the motor is realized. Meanwhile, the output power of the engine is gradually increased, and the applied voltage of the motor is slowly reduced, so that the engine gradually replaces the motor to provide the torque required by acceleration for the propeller until the working voltage of the motor is reduced to one-half of the rated voltage, and the acceleration process of the multi-rotor helicopter is completed.

In another particular embodiment, a particular process is involved in how the propellers of a multi-rotor helicopter slow down. In this embodiment, the method includes the following steps:

applying a working voltage of less than one-half of a calibration voltage to the motor to enable the propeller corresponding to the motor to decelerate;

in the process of decelerating the propeller corresponding to the motor, reducing the output power of the engine and increasing the working voltage of the motor;

and stopping adjusting the motor and the generator until the working voltage of the motor is increased to one-half of the rated voltage.

Specifically, in combination with the above embodiment, when the propeller of the multi-rotor helicopter needs to be decelerated, the propeller of the multi-rotor helicopter can be rapidly decelerated by applying a voltage to the motor that is less than one-half of the rated voltage. During the deceleration process, the output power of the engine is gradually reduced, and the applied voltage of the motor is slowly increased, so that the engine gradually replaces the motor to provide the torque required by the deceleration for the propeller, until the working voltage of the motor is increased to one-half of the rated voltage, and the deceleration process of the multi-rotor helicopter is completed.

In another particular embodiment, a particular process is involved in how a multi-rotor helicopter generates electricity. In this embodiment, the method includes the following steps:

when the voltage applied to the motor is zero volts, the motor is converted into a generator to charge a power supply in the multi-rotor helicopter.

Specifically, when the electric quantity detecting unit detects that the electric quantity of the power supply is lower than the preset electric quantity, the system controller increases the output power of the engine, namely, the accelerator is increased, so that the torque generated by the engine is increased, and simultaneously, the working voltage of the motor is reduced along with the increased torque of the engine until the torque corresponding to the output power of the engine can replace the torque borne by the motor during working, at the moment, the working voltage of the motor is zero volt, the dual-purpose motor/generator is switched from the motor state to the generator state, and the power supply in the multi-rotor helicopter is charged. Clearly, in the process of generating power, the propeller input power is equal to the engine output power, and the generator consumes power.

In one embodiment, the power assist method further comprises: and if the output power of the engine is greater than the preset output power in the second hovering parameter, judging that the multi-rotor helicopter is overweight and cannot take off.

In one embodiment, an aero-engine risk-escaping assistance system assembly is provided, including the steps described in S202 to S208 above, and further including:

when any engine 101 stops rotating due to a fault, the coaxial 20c is separated from the coaxial 20b, the corresponding danger escape detection unit 400 automatically starts a landing program, and the corresponding electric/power generation dual-purpose 102 machine independently drives the landing.

The technical features of the above embodiments can be arbitrarily combined, and for the sake of brevity, all possible combinations of the technical features in the above embodiments are not described, but should be considered as the scope of the present specification as long as there is no contradiction between the combinations of the technical features.

The above examples only show some embodiments of the present invention, and the description thereof is more specific and detailed, but not construed as limiting the scope of the invention. It should be noted that various changes and modifications can be made by those skilled in the art without departing from the spirit of the invention, and these changes and modifications are all within the scope of the invention. Therefore, the protection scope of the present patent should be subject to the appended claims.

Claims

1. An aero-engine escape assistance system assembly comprising: the aircraft engine emergency power assisting system comprises a pair of engines, a pair of electric/power generating dual-purpose machines and a pair of propellers, wherein the engines, the electric/power generating dual-purpose machines and the propellers correspond to one another one to one, the corresponding engines, the electric/power generating dual-purpose machines and the propellers are coaxially connected, the engines provide total power for the aircraft engine emergency power assisting system assembly, and the aircraft engine emergency power assisting system is characterized in that: the electric/power generation dual-purpose machine sets a calibration voltage according to a take-off calibration voltage program, wherein 1/2 calibration voltage is a reference voltage in an electric power-assisted state, if the applied voltage is more than 1/2, the electric power-assisted state is realized, if the applied voltage is less than 1/2, the electric power-reduced state is realized, and if the applied voltage is 0, the output power of the electric/power generation dual-purpose machine is not more than 5% of the output power of the engine;

the aero-engine danger-escaping power-assisted system assembly further comprises: the system comprises a pair of torque sensors, a pair of rotating speed measuring devices, a storage battery and a pair of system controllers, wherein the torque sensors and the rotating speed measuring devices respectively detect the torque and the rotating speed of the corresponding engine, the electric motor/generator and the propeller, and the capacity of the storage battery is not more than the power consumption of the electric motor/generator for 30 seconds;

the system controller is used for storing the corresponding motor rotating speed/torque/voltage characteristic curve table, the engine rotating speed/torque/oil consumption characteristic curve table, the propeller rotating speed/torque/lift characteristic curve table, an A1 group minimum oil motion parameter table and an A0 group maximum oil motion parameter table;

the system controller is further configured to perform: measuring dynamic parameters of the multi-rotor helicopter, setting a take-off calibration voltage of the multi-rotor helicopter and controlling the multi-rotor helicopter to adjust the attitude;

the aero-engine danger-escaping power-assisted system assembly further comprises: the power balance controllers are respectively connected with the pair of system controllers and are used for respectively controlling the rotating states of the corresponding propellers through the system controllers; the power balance controller is used for controlling the pair of electric/power generation dual-purpose machines to be switched into a motor mode simultaneously through the pair of system controllers when the instruction of the flight controller rises;

the aero-engine danger-escaping power-assisted system assembly further comprises: the coaxial oil/electricity hybrid driving danger-escaping device specifically comprises: the system comprises a pair of backstops, a pair of differential sensors and a pair of danger escaping detection units, wherein the pair of backstops are respectively coaxially connected with the pair of corresponding electric/power generation dual-purpose machines and the pair of corresponding engines;

the pair of danger escaping detection units is used for: the rotating speeds of the shafts at the two ends of the pair of ratchets are respectively monitored in real time through the pair of differential sensors, when the pair of engines rotate normally, the shafts at the two ends of the pair of ratchets are rigidly connected, and when the pair of engines stop rotating due to faults or the rotating speeds of the pair of electric/power generation dual-purpose machines are greater than the rotating speeds of the pair of engines, the shafts at the two ends of the pair of ratchets are disconnected, so that a rotating speed difference is generated; when the rotational speed difference is detected by the pair of danger escaping detection units, respectively, emergency processing is performed by the system controller.

2. The aircraft engine risk-escape assistance system assembly according to claim 1 further comprising: and the flight controller is connected with the system controller.

3. The aircraft engine escape assistance system assembly according to claim 2 wherein the controlled end of the dynamic balance controller is connected to the flight controller, the first side of the dynamic balance controller is connected to a corresponding system controller to control the left-hand rotation of the first side propeller, and the other side of the dynamic balance controller is connected to a corresponding system controller to control the right-hand rotation of the second side propeller.

4. An aircraft engine escape assistance system assembly according to claim 3,

the power balance controller is also used for controlling the pair of electric/power generation dual-purpose machines to be respectively switched into a voltage increasing and reducing complementary mode through the pair of system controllers when the flight controller instructs to correct the inclination angle, so that the first side propeller is accelerated, the second side propeller is decelerated, and the acceleration value is equal to the deceleration value.

5. The aircraft engine risk-escaping power-assisted system assembly according to claim 2, wherein the system controller is connected to the flight controller, the battery, a torque measurer, the rotational speed measurer, the electric/power-generating dual-purpose machine, and the engine, respectively.

6. The aircraft engine escape assistance system assembly of claim 5,

the storage battery is respectively connected with the flight controller and the electric/power generation dual-purpose machine.

7. The aircraft engine escape assistance system assembly according to claim 2, wherein the system controller includes a power detection unit, a flight control instruction processing unit, a motoring/generating switching unit, a charge amount detection unit, an engine power allocation unit, and a charge/discharge control unit.

8. The aircraft engine escape assistance system assembly of claim 7,

the power detection unit is respectively connected with a torque measurer and the rotating speed measurer;

the electric/power generation switching unit is respectively connected with the power detection unit, the flight controller and the electric/power generation dual-purpose machine.

9. The aircraft engine escape assistance system assembly of claim 7,

the flight control instruction processing unit is connected with the flight controller;

the electric quantity detection unit is respectively connected with the electric/power generation switching unit and the storage battery.

10. The aircraft engine escape assistance system assembly of claim 7,

the engine power allocation unit is respectively connected with the electric quantity detection unit, the power detection unit and the engine;

the charge and discharge control unit is respectively connected with the electric quantity detection unit, the power detection unit and the storage battery.