CN105539828B

CN105539828B - Self-generating oil-electricity hybrid power multi-rotor aircraft

Info

Publication number: CN105539828B
Application number: CN201510899766.2A
Authority: CN
Inventors: 赵成俊; 陈蜀乔; 马李斌
Original assignee: Hunan Zhongsheng Mechanical Equipment Co ltd
Current assignee: Hunan Zhongsheng Mechanical Equipment Co ltd
Priority date: 2015-12-08
Filing date: 2015-12-08
Publication date: 2024-05-31
Anticipated expiration: 2035-12-08
Also published as: CN105539828A

Abstract

The invention relates to a self-generating oil-electricity hybrid power multi-rotor aircraft, and belongs to the technical field of unmanned aircrafts. The aircraft mainly comprises a main rotor, an aircraft electric control board, a motor driving auxiliary rotor, a motor, a fuel engine, a rotating speed dual gearbox, a battery, an air door control stepping motor, a generator and the like. The aircraft is provided with flying power jointly by a main rotor and an auxiliary rotor, wherein the main rotor provides main lift-off power, the auxiliary rotor provides small part of lift-off power, the main rotor consists of an upper main rotor and a lower main rotor, the upper main rotor and the lower main rotor have the same size, the rotating speeds are identical, the rotating directions are opposite, and positive and negative torque balance is realized. The fuel engine part power is meshed through the bevel gear and is output to the generator, and the generated power provides power for the whole aircraft. The center of gravity of the main rotor shaft and the oil tank is located at the geometric center of the aircraft. The invention realizes that the multi-rotor aircraft can be accurately controlled, has the characteristics of large load and long endurance, and does not need external power supply for charging.

Description

Self-generating oil-electricity hybrid power multi-rotor aircraft

Technical Field

The invention relates to a self-generating oil-electricity hybrid power multi-rotor aircraft, and belongs to the technical field of unmanned aerial vehicles.

Background

The multi-rotor unmanned aerial vehicle is an aerial vehicle which can take off and land vertically and takes a plurality of rotors as power devices, and does not carry operators. The application is very wide. The main applications are as follows. Such as police service application, fire scene command, rescue and disaster relief, traffic management; news media, aerial photography; photographing wild animals, evaluating environment and performing aerial archaeology; real estate management and pipeline inspection; remotely controlling flying and aerial photography; small articles are dispatched by unmanned aerial vehicles. The multi-axis aircraft comprises a geographic information system and a global positioning system based on the navigation positioning system, can fly according to a preset route through the geographic information system and the global positioning system, and can complete specified actions after reaching a destination, so that the multi-axis aircraft can be used for logistics distribution of small objects.

All applications of multiaxial aircraft are based on long endurance, high load flight of the aircraft, otherwise all applications are of a narrow scope. The existing four-rotor aircraft is driven by a battery, and because the energy stored by a battery with unit mass is far smaller than fossil fuel such as gasoline, the flight time of the existing four-rotor aircraft is very limited, the cruising time is only within 20 minutes under the normal load condition, and the cruising time of the fuel-powered single-rotor helicopter can be as long as a plurality of hours. However, the control of the flight attitude, the stability and the accuracy of the control and the like of the fuel power single-rotor helicopter are still a difficult problem at present. Because accurate control of the rotational speed of the fuel engine by electrical signals is an extremely difficult event, this makes it impossible for a fuel-powered single-rotor helicopter to accomplish the tasks that an electric multi-rotor aircraft can accomplish. Two types of aircrafts, each having advantages and disadvantages, the fuel oil aircrafts have the characteristics of long endurance and large load; the electric multi-rotor aircraft can be accurately controlled, can be preprogrammed to complete various actions along a specified route under the navigation of a GPS, which is not realized by the fuel aircraft, but can not replace a single rotor of fuel power in the application field requiring long cruising time. Therefore, the advantages of the two are combined, so that the aircraft can be accurately controlled, and the aircraft has the characteristics of high load and long endurance. Only in this way can practical use of the multi-rotor aircraft be achieved. In the invention (2015142531. X) of the hybrid electric multi-rotor aircraft, the energy systems of the fuel engine and the motor are two independent systems, and potential risks can exist, for example, the battery power consumption is finished, and the situation that the multi-rotor unmanned aircraft cannot be controlled can occur when the fuel is not consumed at the moment; another situation is that the fuel is exhausted and the battery is not exhausted, so that the multi-rotor unmanned aerial vehicle is likely to fall in the air due to the loss of the main power. Therefore, it is extremely important how to solve the problem of uncoordinated power supply of the multi-rotor unmanned aerial vehicle. Furthermore, consideration must be given to how the flight of the aircraft can be achieved in the absence of charging conditions outdoors at all.

Disclosure of Invention

The invention aims to overcome the common problems of the existing multi-rotor aircraft, namely the defects of short range, small load and limited flight time, can avoid the problem of uncoordinated power supply of the multi-rotor unmanned aircraft, overcomes the inherent defects of an electric multi-rotor aircraft by utilizing self-generating oil-electricity hybrid power, directly drives a main rotor to provide lifting force and simultaneously drives a self-contained generator to generate power, supplies power for a motor of an auxiliary rotor, and is matched with the motor-driven auxiliary rotor to realize that the aircraft can be accurately controlled, has the characteristics of large load and long flight time, and does not need external power supply to charge.

The self-generating oil-electricity hybrid power multi-rotor aircraft mainly comprises an upper main rotor 1, a lower main rotor 2, a multi-rotor aircraft antenna 3, a multi-rotor aircraft electric control board 4, a motor driving auxiliary rotor 5, a motor 6, a fuel engine 7, a rotating speed dual gearbox 8, a carburetor 9, a battery 11, a fuel inlet air door 12, an oil delivery pipe 13, an oil tank 15, a horn 16, an upper main rotor shaft 18, a lower main rotor sleeve shaft 19, an upper bevel gear 20, a lower bevel gear 21, a rotating speed sensor 22, an engine exhaust pipe 23, an air door control stepping motor 24, a generator 25, an electric output line 26, a damping cushion pad 27, a motor driving bevel gear 28 and a motor power input bevel gear 29, wherein the self-generating oil-electricity hybrid power multi-rotor aircraft is provided with flying power jointly by the main rotor and the auxiliary rotor, the main rotor is driven by the fuel generator 7, the auxiliary rotor 5 is driven by the motor 6, the main rotor and the auxiliary rotor 5 jointly provide lift-off power, the main rotor is composed of an upper main rotor 1 and a lower main rotor 2 (see figure 1), an upper main rotor rotating shaft 18 passes through a lower main rotor sleeve shaft 19 from the top end and is welded with a lower bevel gear 21 and is fixedly connected with an engine power output shaft, the lower main rotor 2 is fixedly arranged at the upper end of the lower main rotor sleeve shaft 19, the sleeve shaft and the upper main rotor rotating shaft 18 can mutually independently rotate along the same axis, the lower main rotor sleeve shaft 19 and the upper bevel gear 20 are welded with the upward large end face of the upper bevel gear 20, a left bevel gear and a right bevel gear are arranged between the upper bevel gear 20 and the lower bevel gear 21, four gears are arranged in a speed pair gearbox 8 in total, the left bevel gear rotating shaft and the right bevel gear rotating shaft are fixedly arranged on the side wall of the gearbox, a speed sensor 21 for monitoring the speed of the main rotor rotating shaft 18 is arranged below the lower bevel gear 21, and the gearbox body is fixedly arranged on a fuel engine body, the main rotor shaft 18 passes through the lower bevel gear 21 and then is connected in series (see fig. 2) with a motor driving bevel gear 28, the motor driving bevel gear 28 is fixedly connected with a power output shaft of the fuel engine 7 after passing through the bevel gear, the motor driving bevel gear 28 is meshed with a motor power input bevel gear 29, the rotating shaft of the motor power input bevel gear 29 is the shaft of a rotor of the generator 25, then partial power of the fuel engine is output to the generator 25 through a power output line 26, the rotating shaft of the motor power input bevel gear 29 is the shaft of the rotor of the generator 25, then partial power of the fuel engine is output to the generator 25 through a power output line 26, the main rotor shaft passes through the geometric center of the multi-rotor aircraft, the fuel engine 7 is fixedly mounted on the horn 16, a damping buffer cushion 27 is arranged between the fuel engine 7 and the horn 16, the center of gravity of the fuel tank 15 is positioned at the geometric center of the multi-rotor aircraft, and the landing gear 17 is mounted in an axisymmetric manner by taking the geometric center of the multi-rotor aircraft as the shaft.

The oil delivery pipe 13 of the oil tank 15 is connected with the carburetor 9, the carburetor is connected with the fuel air inlet air door 12, the wind shield rotating shaft of the air inlet air door is connected with the air door control stepping motor 24 for controlling the opening and closing angle of the wind shield, the generator 25 is arranged on the generator damping bracket 14, the damping bracket is provided with a damping cushion pad 27, and the generator 25 is fixed on the horn 16.

This many rotor crafts of many rotor crafts antenna 3 and many rotor crafts automatically controlled board 4 are installed on horn 16, battery 11 installs in the frame below, adjust the battery mount and make antenna, automatically controlled board and battery constitute integrative focus at rotor craft's geometric center, motor cabinet 10 is installed to the end of horn 16, motor cabinet fixed motor 6, motor power output shaft installs motor drive pair rotor 5, the projection interval of main rotor and pair rotor on the plane that the cross horn constitutes is greater than 2cm for the air current between them does not take place mutual disturbance.

The power management system comprises a voltage stabilizing and rectifying module, a current limiting module, a charging module, a relay system, a load circuit, a battery pack, an undervoltage protection module, a control module and a fuel oil stock display module; the fuel engine 7 drives the generator 25 to generate alternating current through bevel gear transmission, excitation current of the generator is regulated through the voltage stabilizing rectifying module, so that the generator outputs relatively stable direct current, output current is limited through the current limiting circuit, overlarge load of the generator is avoided, the current enters the charging module, current is distributed through the relay system, the relay system is controlled by the controller to intelligently control charge and discharge of the load and the storage battery, charging and power supply currents are intelligently regulated, the battery pack is provided with an undervoltage protection circuit, and the controller is also responsible for transmitting fuel stock and charge and discharge data of the load and the storage battery to the ground control and monitoring platform through the wireless radio frequency module (see fig. 4).

The running program of the self-generating oil-electricity hybrid power multi-rotor aircraft is that a system is initialized firstly, then a channel is tested, after the channel is normal, a fuel engine is started, then the rotating speed of the engine is thoroughly carried out, if the rotating speed of the engine is abnormal, the rotating speed of the engine is regulated to reach a specified range, after the rotating speed of the engine is normal, the battery voltage, the fuel quantity and the voltage parameters of the engine of the whole aircraft are tested to reach the specified range, then the taking-off gesture is detected, theoretical rotating speed values of different motors and fuel generators are calculated according to the gesture, the rotating speeds of the motors and the fuel generators are regulated to take off, and remote control operation is carried out (see fig. 5 and 6).

Working principle: the hybrid electric vehicle is divided into a main rotor and an auxiliary rotor, wherein the main rotor provides main lift power, generally bears more than 50% of lift force, the auxiliary rotor provides a small part of lift power, the provided lift force is less than 50%, and the main lift power and the auxiliary lift power have the following relation: the higher the lift force duty ratio provided by the main rotor wing, the longer the dead time, the larger the load capacity, and the disadvantage is that the handling capacity is weakened; the higher the lift duty cycle provided by the secondary rotor, the shorter the dead time, the poorer the load carrying capacity, but the better the handling performance. Therefore, the two are matched with each other, and the proportional relation of the generating force between the two is adjusted according to the actual situation. For example, when the aircraft is not required to fly flexibly and the long-endurance high-load flight is emphasized, the duty ratio of the main rotor force must be increased, and if the performance of the aircraft is emphasized, the duty ratio of the main rotor force must be reduced so as to realize flexible control.

The main rotor is composed of an upper main rotor 1 and a lower main rotor 2 (see figure 1), an upper main rotor rotating shaft 18 passes through a lower main rotor sleeve shaft 19 of the lower main rotor 2 from the top end and then is fixedly connected with a power output shaft of an engine after being welded with the lower bevel gear 21, the lower main rotor sleeve shaft 19 is welded with the upward large end face of an upper bevel gear 20, a left bevel gear and a right bevel gear are arranged between the upper bevel gear 20 and the lower bevel gear 21, four gears in total are arranged and are arranged in a rotating speed dual gear box 8, and the left bevel gear rotating shaft and the right bevel gear rotating shaft are fixed with the side wall of the gear box. The structure makes the rotational speed of the upper main rotor wing and the rotational speed of the lower main rotor wing identical, the rotational speed direction is opposite, positive and negative torque balance is realized, and the configuration makes the multi-rotor aircraft not rotate in the air.

The main shaft of the lower main rotor 2 is a lower main rotor sleeve shaft 19, a cavity is arranged in the sleeve shaft, only the upper, middle and lower 3 end surfaces are in contact with the inner upper main rotor shaft 18 to form constraint, so that the lower main rotor sleeve shaft 19 and the upper main rotor shaft 18 can rotate along the same axis independently, and the sleeve structure aims to reduce mechanical friction loss caused by high-speed movement as much as possible. During use, a suitable lubricant must be added to reduce friction losses.

A rotation speed sensor 21 for monitoring the rotation speed of the main rotor shaft 18 is arranged below the lower bevel gear 21, the gear box body is fixed on the fuel engine body, and the fuel engine 7 is fixed on the horn 16 through an engine fixing bracket 14 (see fig. 2). The engine with the power output shaft positioned at the center of gravity of the engine is selected as far as possible, the main rotor shaft passes through the geometric center of the multi-rotor aircraft, and the arrangement ensures that the lift force of the main rotor and the auxiliary rotor are in axisymmetric distribution, so that the control is convenient; the center of gravity of the oil tank 15 is located at the geometric center of the multi-rotor aircraft, and oil in the oil tank is continuously reduced in the working process of the engine, so that the center of gravity of the mailbox is still located at the geometric center of the multi-rotor aircraft, and the aircraft still keeps horizontal flight in the flying process.

The main rotor shaft 18 passes through the lower bevel gear 21 and is connected in series (see fig. 2) with a motor-driven bevel gear 28, the motor-driven bevel gear 28 is fixedly connected with a power output shaft of the fuel engine 7 after passing through the bevel gear, the motor-driven bevel gear 28 is meshed with a motor power input bevel gear 29, the rotating shaft of the motor power input bevel gear 29 is the shaft of a rotor of the generator 25, so that partial power of the fuel engine is output to the generator 25, kinetic energy is converted into electric energy, and generated electric power is output through an electric power output line 26 to provide electric energy for the whole aircraft.

The oil delivery pipe 13 of the oil tank 15 is connected with the carburetor 9, the carburetor is connected with the fuel air inlet air door 12, the wind shield rotating shaft of the air inlet air door is connected with the air door control stepping motor 24 for controlling the opening and closing angle of the wind shield, PID closed loop control is adopted for the control, and the rotating speed signal is from the engine rotating speed sensor.

Be equipped with shock attenuation blotter 27 between fuel engine 7 and the horn 16, generator 25 installs on generator shock attenuation bracket 14, and this shock attenuation bracket is equipped with shock attenuation blotter 27, fixes generator 25 on installing on the horn 16, and shock attenuation cushioned setting can effectively alleviate the vibrations that generator and engine brought to the flight gesture of aircraft is not influenced.

The oil delivery pipe 13 of the oil tank 15 is connected with the carburetor 9, the carburetor is connected with the fuel inlet air door 12, and the wind shield rotating shaft of the improved air door is connected with the air door control stepping motor 24 for controlling the opening and closing angles of the wind shield. In the flying process, the accelerator is increased by the remote controller, and two control modes can be selected

1. Constant rotation speed control mode of main rotor wing

The control mode is characterized in that after the aircraft is lifted off, the fuel engine works at a certain constant rotating speed to provide a constant lifting force, the weight of the aircraft is reduced to more than half of the original weight, all other control is realized by the auxiliary rotor wing driven by the motor, and the control mode has the advantages that a control circuit is extremely simple, but the control performance is not very good;

2. main rotor non-constant rotation speed control mode

The control mode is that after the aircraft is lifted, the rotation speed of the fuel engine is non-constant, a strict proportional relation exists between the rotation speed of the main rotor wing and the rotation speed of the auxiliary rotor wing driven by the motor, in other words, when the throttle is increased, the rotation speed of the main rotor wing and the rotation speed of the auxiliary rotor wing are simultaneously lifted, and the control mode is more superior than the above mode, and is relatively more complicated to control.

The projection interval of the main rotor wing and the auxiliary rotor wing on the plane formed by the cross horn is at least more than 2cm, so that the air flows of the main rotor wing and the auxiliary rotor wing do not generate mutual disturbance. The fuel engine exhaust pipe 23 is downward, and air flow disturbance can be reduced. In practical application, the interval distance is as far as possible, so that air flow disturbance is avoided, and the flight attitude of the aircraft is damaged. Determination of lift force of a fuel engine:

let the lift of the rotor wing be Y _１

Y_１＝∫¹ ₀(△P_{Lower part(s)}cosθ_{Lower part(s)}－△P_{Upper part}cosθ_{Upper part})cosa dx

=1/2 ρV²∫¹ ₀(p_{Lower part(s)}cosθ_{Lower part(s)}－P_{Upper part}cosθ_{Upper part})cosa dx

=1/2 ρV²b∫¹ ₀(P_{Lower part(s)}cosθ_{Lower part(s) －}P_{Upper part}cosθ_{Upper part})cosa dx/b

Ream (P _{Lower part(s)}cosθ_{Lower part(s)}－P_{Upper part}cosθ_{Upper part})cosa dx/b=Cy₁ (1)

Cy ₁ is the lift coefficient of the rotor profile, then the unit extender blade She Shengli can be written as

Y₁＝Cy₁1/2.ρv²b1

Where b1 is the blade area per unit span length, so helicopter rotor lift can be written in a manner analogous to this

Y=Cy1/2.p v²S （2）

The formula is a rotor wing lift force formula of the helicopter. Cy is the lift coefficient and S is the total area of the blade.

Total lift of hybrid power

Y_{Total (S)}= Y₁+Y₂+Y₃+…+Y_n+Y_{Fuel oil} (3)

And n is the number of the rotor wings of the multi-rotor aircraft. According to the formula (2), the lift force of the fuel engine and the motor rotor wing can be calculated, and the proper lift force ratio Y _{Fuel oil}/ Y_{Total (S)} is determined.

For example, the main rotor provides 60% of the lift and the auxiliary rotor provides 40% of the lift. The fuel engine uses 55% of the power to provide lift and converts 45% of the power to electrical energy for driving the secondary rotor motor and battery charging.

Unmanned aerial vehicle remote sensing control platform circuit

In the unmanned aerial vehicle system, ① is formed by expanding multiple serial ports and USB interfaces so as to realize the communication between the system and peripheral equipment; ② The engine air door controls the stepping motor driving module; ③ The generator voltage-stabilizing power supply module; ④ A power management monitoring system; ⑤ A DC brushless motor circuit module; ⑥ A charging module; ⑦ A power supply undervoltage protection circuit module; ⑧ An adjustable current-limiting DC regulated power supply module (see FIG. 3).

① USB interface expansion circuit: USB port expansion is implemented by CH375 chip. CH375 is a universal interface chip for a USB bus. For a common USB storage device, the built-in firmware of CH375 can automatically process the special communication protocol of Mass-Storag Mass storage device, and in general, an external singlechip does not need to write a firmware program. There are 2 ways of communicating CH375 with a single chip: parallel and serial modes. The schematic diagram of the USB expansion circuit is shown in FIG. 7, and the CH375 chip is set to be in a built-in firmware mode, and a 12 MHz crystal is used. The P0 port of the singlechip is connected with D0-D7 of CH375 to be used as a data bus, the decoder outputs the chip connected with CH375, the singlechip A0 is connected with A0 of CH375, and the address of CH375 or data input and output can be selected. The address is transmitted when A0 is high and D0 to D7 are transmitted, and the data is transmitted when A0 is low. P3.6 and P3.7 control the read and write operations of CH375, respectively. CH375 is connected with the input end of the singlechip, and generates an interrupt signal when data is input through the USB port to inform the singlechip to process the data. After the CH375 chip is initialized and successfully communicated with the host, the indicator light is turned on.

② Step motor driving circuit: in the invention, the size of the throttle of the engine is controlled by adopting a stepping motor in a closed loop manner, so that the control of the engine speed is realized. The stepping motor is driven by the THB6128 chip, and the single chip microcomputer can control the stepping motor only by outputting the running direction and pulse signals of the stepping motor. THB6128 is a special chip for driving a high-subdivision two-phase hybrid stepping motor, and a control signal is output through a singlechip, so that a high-performance multi-subdivision driving circuit can be designed. The dual full-bridge MOSFET is characterized in that the dual full-bridge MOSFET is driven, the low on-resistance Ron=0.55Ω, the highest withstand voltage is 36V, the large current is 2.2A (peak value), various subdivisions are selectable, the highest subdivision can be 128 subdivisions, the dual full-bridge MOSFET has an automatic half-current locking function, 3 attenuation modes of fast attenuation, slow attenuation and hybrid attenuation are selectable, and the dual full-bridge MOSFET is internally provided with temperature protection and overcurrent protection. Fig. 8 shows a motor throttle stepper motor drive circuit, as are course angle stepper motor drive circuit, pitch angle, roll angle stepper motor drive. In the figure, CP1 and U/D are respectively drive pulse and motor running direction control signals given by a singlechip. M1, M2, M3 are motor drive subdivision number selection signal input, and are manually controlled by a dial switch. FDT1 and VREG1 are attenuation mode selection voltage and current control voltage inputs, respectively. A slow decay mode when 3.5V; when in a mixed attenuation mode; a fast decay mode is when FDT1< 0.8V. The driving current value of the stepping motor can be set by adjusting the voltage of the VREG1 terminal.

③ Voltage-stabilizing power supply module of generator

The invention employs a small-sized excitation alternator. The operating principle is that when the voltage of the generator rises to a specified value, the voltage applied to the point "a" of the voltage divider, namely the reverse voltage born by the voltage stabilizing tube VS2, exceeds the reverse breakdown voltage and is reversely broken down and conducted, and the transistor VT1 is also conducted. VTl is conducted to reduce the potential of 'b', the diode VD2 is stopped after being subjected to reverse voltage, VT2 and VT3 are also stopped, an excitation circuit of the generator is cut off, excitation current is interrupted, a magnetic field of the generator is disappeared, and the voltage of the generator is reduced. When the voltage drops below the voltage regulation value, the voltage stabilizing tube VS2 is turned off again, then VTl is also turned off, VT2 and VT3 are turned on again, and the voltage of the generator rises again. The repeated circulation is performed in this way, and the on-off of the exciting circuit is controlled, so that the voltage of the generator can be kept constant when the rotating speed of the generator changes, as shown in fig. 9.

④ Power management monitoring system

And (3) power management: in order to improve the safety of the aircraft, a set of equipment monitoring system needs to be designed for monitoring the attitude information of the aircraft, the conditions of the on-board equipment, the conditions of a power supply and the like in real time. The power supply used by the platform is a battery pack formed by connecting two lithium batteries in series, and a charge and discharge management system taking Atmega16l as a core is adopted by utilizing the charge and discharge characteristics of the lithium ion batteries. The unmanned aerial vehicle power management system framework is shown in fig. 4. In the system, the self-contained small generator 25 converts kinetic energy into alternating current, and outputs 11.6V direct current voltage after rectification and voltage stabilization, so that two lithium batteries can be charged by the output voltage. The controller of the power management system is an Atmega161 single-chip microcomputer, and the controller controls the relay switch to carry out charge and discharge management on the batteries by detecting the voltage of the two lithium batteries.

After the controller collects the information in the power supply system, the data are transmitted to the ground in real time through the wireless transmission equipment. The ground monitoring platform can also send some instructions to the mega16l, and the battery is controlled to charge and discharge by controlling the relay switch, so that the purpose of monitoring and controlling the aircraft is achieved. The on-board power supply module consists of two lithium batteries produced by Interman battery limited company, the voltage is 8.4V when the electric quantity of the battery pack is sufficient, the electric charge quantity of the battery is closely related to the reliability of the whole power supply system, and the more the residual electric quantity of the battery is, the higher the reliability of the system is, so the residual electric quantity of the battery can be obtained in real time when the aircraft flies, and the reliability of the aircraft is greatly improved.

And (3) power supply monitoring: the helicopter can smoothly complete the flight task, and sufficient power supply is indispensable. As is apparent from the characteristics of lithium batteries, in the case of overdischarge, the electrolyte is decomposed to deteriorate the battery characteristics and reduce the number of times of charging. Therefore, in order to protect the safety of the battery, the power supply system needs to pass through the under-voltage protection module and the voltage stabilizing module before supplying power to the control system. In order to predict the residual electric quantity in the power supply system, a method for detecting the voltage of the power supply system is adopted, and after the power supply voltage of the system is detected, a database established by a discharge curve is searched, so that the residual electric quantity in the power supply system can be estimated. The power supply voltage required by the singlechip is 2.7-5.5V, so that an external reference voltage can be designed for meg a16l to be 2.5V, and the reference voltage stabilizing circuit is shown in figure 16. Therefore, the system needs to divide the voltage of the battery by using a resistor and the maximum divided voltage value cannot exceed 2.5V, and the real-time voltage in the power supply system can be obtained after the voltage value measured by the controller is multiplied by the reduced multiple of the voltage division. The power consumption condition of the lithium battery is monitored at any time, the phenomenon of excessive use of the battery is prevented, and the purposes of effectively using the battery capacity and prolonging the service life can be achieved.

⑤ DC brushless motor circuit: the brushless DC motor consists of motor main body and driver and is one electromechanical integrated product. The DC brushless motor has the same working principle and application characteristics as a general DC motor, and the composition is different, besides the motor, the motor is provided with a plurality of reversing circuits, and the motor of the DC brushless motor is an electromechanical energy conversion part and is provided with a sensor besides a motor armature and a permanent magnet excitation part. A portion of the AC-DC circuit of the motor is shown in fig. 17.

⑥ A charging circuit: the charging characteristics of the lithium ion battery are different from those of the nickel-cadmium and nickel-hydrogen batteries, when the lithium ion battery is charged, the battery voltage slowly rises, the charging current gradually decreases, and when the voltage reaches about 4.2V, the voltage is basically unchanged, and the charging current continuously decreases. Therefore, the modified charger can be charged by a constant-current-first-constant-voltage charging mode, and a specific charging circuit is shown in fig. 10. The circuit adopts LM2575 ADJ to form a chopper-type switch voltage stabilizer, and the maximum charging current is 1A.

The working principle of the circuit is as follows: when the battery is connected to the charger, the circuit outputs a constant current to charge the battery. The constant current control part of the charger consists of half of a double operational amplifier LM358, gain setting resistors R3 and R4, a current sampling resistor R5 and a 1.23V feedback reference voltage source. Immediately after the battery is connected, the operational amplifier LM358 outputs a low level, the output voltage of the switching regulator LM2575-ADJ is high, and the battery starts to charge. When the charging current rises to 1A, the voltage drop at two ends of the sampling resistor R5 (50 mEurope) reaches 50mV, the voltage is amplified by an operational amplifier with the gain of 25, 1.23V voltage is output, and the voltage is applied to the feedback end of the LM2575 to stabilize the feedback circuit. When the battery voltage reaches 8.4V, LM3420 starts to control the feedback leg of LM2575 ADJ. LM3420 makes the charger turn into constant voltage charging process, and the battery both ends voltage is stabilized at 8.4V.R6, R7 and C3 constitution compensation network, guarantees that the charger is under constant current/steady voltage state steady operation. If the input supply voltage is interrupted, the diode D2 and the PNP input stage in the op amp LM358 are reverse biased, thereby isolating the battery from the charging circuit, ensuring that the battery will not discharge through the charging circuit. When the charge is shifted to the constant voltage charge state, the diode D3 is reverse biased, so no sink current is generated in the op amp.

⑦ And the under-voltage protection circuit of the power supply, namely the lithium battery, needs to consider the safety during charging and discharging so as to prevent the characteristic degradation. Therefore, in order to protect the safety of the lithium battery during the system operation, a set of under-voltage protection circuit is required to prevent the battery characteristics and durability characteristics of the power management system from being deteriorated due to the over-use. The under-voltage protection of the power supply is known from the battery discharge characteristic of the lithium battery, when the battery is at 3.5V, the battery power is about to be used up, the battery is charged in time, and otherwise, the battery voltage is rapidly reduced until the battery is damaged. The undervoltage protection circuit is shown in fig. 11, and the comparison result is compared with the reference voltage designed by TL431 by utilizing the voltage division of the resistor, and the comparison result is sent to the LM324 amplifying circuit to trigger a switching system formed by a triode, so that the on-resistance of the load loop is controlled. Experiments prove that when the system voltage reaches the critical dangerous voltage of 7V, the output current of the system is only 4mA, so that the overdischarge phenomenon of the lithium battery of the system is prevented. Since lithium ion batteries have high energy density, it is difficult to ensure the safety of the batteries. In the overcharged state, the energy is excessive after the battery temperature rises, and the electrolyte is decomposed to generate gas, so that there is a risk of spontaneous combustion or rupture due to the rise of the internal pressure; conversely, in the overdischarged state, the electrolyte is decomposed to deteriorate battery characteristics and durability, thereby reducing the number of times of charging. The charging circuit and the management system can effectively prevent and treat overcharge and overuse of the lithium battery, thereby ensuring the safety of the battery and prolonging the service life of the lithium battery. The system has the functions of automatically controlling charge and discharge management, monitoring the voltage of the battery in real time and the like. The feasibility of the system is verified through debugging and experiments, but in order to ensure the safety of the aircraft, more experiments are performed to ensure the safety and stability of the autonomous flight of the unmanned aerial vehicle.

⑧ Adjustable current-limiting DC stabilized power supply: the adjustable stabilized voltage supply has the current limiting function, the voltage and current values are digitally displayed, and the current limiting function is very useful in the process of experiments, so that the loss caused by errors is avoided. When the output current exceeds the preset current, the output voltage drops, and the current limiting indicator lamp is turned on, which indicates that the current exceeds the set value, and the rechargeable battery can be charged by using the constant current function (see fig. 12).

Technical parameters: input voltage: 24VAC input current: 3A (maximum) output voltage: 0-24V adjustable output current: 2mA-3A adjustable output voltage ripple: 0.01% (maximum)

The working principle is shown in fig. 1, the primary rectified current is connected to a wiring terminal J4, rectified by a rectifier bridge BRl, smoothed and filtered by EC1 and R1, and output to the collector of a voltage regulating tube Q2. This circuit has unique characteristics that are different from other regulated power supplies. The reference voltage of the power supply is provided by an operational amplifier U2 with fixed gain, and a voltage stabilizing tube with a voltage stabilizing value of 5.6V is selected as DZ 1. After the power is turned on, the output voltage of the op-amp U2 increases to turn on DZ1, and is stabilized around 5.6V by R7, and since r8=r1o, the output voltage of U2 is 11.2V. The magnification of U3 is about 3 times, according to the formula a= (r1 3+r18)/R13. The reference voltage of 11.2V can be amplified to approximately over 30V, and potentiometer VR3 and resistor R14 constitute an output voltage zero regulator that can output the voltage of OV.

Another very important feature of the circuit is that the maximum output current can be preset and can be effectively converted from a constant voltage source to a constant current source. The circuit detects the voltage drop across resistor R19 connected in series across the load through U1, the inverting input of U1 is connected to the reference OV through R9, while the non-inverting input voltage can be regulated by VRl, which is adjusted to make the non-inverting input of U1V, assuming an output voltage of a few volts. The voltage amplifying section of the circuit keeps the output voltage constant while the effect of R19 in the string at the output loop is negligible because the resistance of R19 is small and outside the circuit voltage control feedback loop. When the load and the output voltage are unchanged, the circuit is in a voltage stabilizing state, when the voltage drop on the R19 is larger than 1V due to the increase of the load current, the U1 output is in a low level, and as the output end of the U1 is connected to the non-inverting input end of the U3 through the D2, the U1 forcedly pulls down the potential of the non-inverting input end of the U3, so the output voltage is reduced until the voltage drop on the two ends of the current sampling resistor R19 is reduced to 1V, and the circuit is switched into a constant current mode. Limiting the output current by monitoring the voltage drop across R19 to reduce the output voltage is an effective way to keep the output current constant and very accurate, allowing the current to be controlled to 2mA. The function of C1 here is to increase the stability of the circuit, Q1 is used to indicate whether the current limiting circuit is active, and Q1 will drive the LED to emit light whenever it enters the current limiting state. In order for U3 to control the output voltage to OV, a negative supply voltage is required, which is made up of a simple voltage pump circuit, EC3, EC4 and associated components. Regulated by R21 and DZ2, the negative voltage provides power to both U1 and U3, and U2 is powered by a single power supply. In order to avoid the out-of-control of the circuit when the power supply is turned off, a protection circuit is formed by the Q4 and related elements, and when the alternating voltage disappears, the negative voltage also disappears immediately, so that the Q4 is conducted, the output voltage becomes OV, and the circuit and a load connected with the circuit are effectively protected. During normal operation, the base electrode of the Q4 is connected to negative voltage through the R22 to cut off, an output short-circuit protection circuit is arranged in the U3, the Q4 is conducted and cannot damage the IC, so that charges stored in the filter capacitor can be quickly discharged, and the function is very beneficial to experiments, because most of stabilized power supplies tend to have output voltage rising instantaneously when a power switch is turned off, and the output voltage is heavy.

To prevent VR2 from getting bad contact, R25 is connected to the non-inverting terminal of U3 to raise the output voltage to the maximum value, and when VR2 is open, the non-inverting terminal voltage of U3 is pulled to 0V to make the output to 0V.

Multi-rotor unmanned aerial vehicle hardware

For example four-axis aircraft

1. List of device list

An F330 frame; 4 Langyan angel series brushless motors; 1 aeromodelling helicopter fuel engine; 1 miniature excitation generator; the power generator throttle controls 1 stepping motor, and 1 power management system (comprising a voltage stabilizing rectifying module, a current limiting module, a charging module, a relay system, a load circuit, a battery pack, an undervoltage protection module and a fuel oil stock display module) is provided; 6 bevel gears with different specifications; in 30A, tewei AL electric tuning (4), positive and negative paddles, one lithium battery 2200mah, one STM32 chip, MPU6050 and NRF24l01 transceiver module. The connection of the various parts is shown in figure 5.

The implementation process comprises the following steps: referring to fig. 3, the upper computer is turned on to establish a connection with the flight controller. The MPU6050 feeds back the attitude angle and acceleration of the current aircraft to the upper computer, and the attitude angle and the acceleration are communicated by using NRF24l01, wherein the communication frequency is 2.4G, so that the requirement of real-time communication can be met. When you need to change the flight state of the aircraft, the rotation speed of each motor of the aircraft is changed through mouse and keyboard control, and the process is realized by adjusting the duty ratio of the motors through STM32 chips (PWM adjustment).

2. Attitude measurement system

The state of a four-axis aircraft at a certain moment is described by 6 physical quantities, including 3 position quantities in three-dimensional coordinates and attitude quantities along 3 axes (i.e., referred to as six degrees of freedom). The sensor is used as a detecting device, can sense the measured information, and can convert the information sensed by detection into an electric signal or other information output in a required form according to a certain rule so as to meet the requirements of information transmission, processing, storage, display, recording, control and the like.

(1) Acceleration sensor: the acceleration sensor is used for measuring the inclination angle of the machine body relative to the horizontal plane, utilizing the universal gravitation of the earth, projecting the gravitational acceleration to the X, Y and Z axes and measuring the posture of an object;

(2) A gyroscope: the influence of external force on the object is measured by utilizing invariance of the direction pointed by the rotating shaft of the rotating object when the rotating shaft is not influenced by the external force. Since the rotational axis direction of a rotating object is uncertain unlike the earth's gravitational force and the earth's magnetic force of north and south, the angular velocity sensor can only measure a change in position, and cannot measure an absolute angle and posture of the object like an acceleration sensor and a geomagnetic sensor.

3. Motor and fuel engine driving module

And driving each motor to reach a specified rotating speed according to the command of the central control module, feeding back the speeds of the motors and the fuel engines to the controller module through the speed measurement feedback device, and controlling the rotating speeds of the motors and the fuel engines to be expected values by utilizing closed-loop control. Thereby realizing different flight states of the four-axis aircraft.

4. Main control module

The central control module, namely a core processor of the flight control system, is used as a core control part of the whole system and is mainly responsible for acquiring and resolving the attitude angular rate (pitch angle rate and roll angle rate) detected by the sensor, the linear acceleration and heading information of the three axes in real time; according to the detected flight information, the output control quantity is calculated by combining with a given control scheme, the output control quantity is converted into corresponding PWM signals, the corresponding PWM signals drive the four motors to work after passing through a driving circuit, the four-axis aircraft is kept to fly stably, and data transmission is carried out between the four-axis aircraft and a ground station through a wireless communication module, so that the flight state is changed by receiving control commands, and the flight state data is downloaded. The PWM pulse control mode is to control the on-off of the switching device of the inverter circuit, so that a series of pulses with equal amplitude are obtained at the output end, and the pulses are used for replacing sine waves or needed waveforms. That is, a plurality of pulses are generated in a half period of the output waveform, so that the equivalent voltage of each pulse is a sine waveform, and the obtained output is smooth and has few low harmonics. The width of each pulse is modulated according to a certain rule, so that the output voltage of the inverter circuit can be changed, and the output frequency can be changed.

5. Motor driving module

The control of the actual equivalent voltage applied to the two ends can be realized by controlling the PWM, so that the control speed is realized, the higher the PWM duty ratio is, the higher the equivalent voltage is, and the lower the duty ratio is, the lower the equivalent voltage is.

6. A wireless communication module: a communication link is established between the ground control station and the aircraft via the wireless network. The ground control station transmits flight and task control instructions to the main control module, and the central control module transmits flight states, tasks, fuel oil storage and other conditions.

7. Power supply module

The power module provides power for an onboard control system, a motor and the like. The battery includes nickel-hydrogen battery and lithium battery. The current discharge C coefficient of lithium batteries is generally larger than that of nickel-hydrogen batteries and is relatively constant, so that lithium batteries are selected. The choice of battery is mainly seen in two properties: first, capacity and second, multiplying power. The greater the capacity, the longer the cruising ability of the aircraft. But the larger the capacity, the heavier the battery weight, so the capacity selection is required to satisfy a certain cruising ability and is lighter. The sensors used are the most common accelerometers and gyroscopes.

Function of basic module

1. Upper computer

The upper computer is mainly used for displaying the flight attitude of the aircraft and sending out control information for the aircraft. By using the upper computer, the PID control on the aircraft can be conveniently realized, and the parameters of the aircraft can be changed in real time. The upper computer also has the basic functions of waveform display, calibration and the like;

2. NRF24l01 transceiver module

NRF24.L01 is a novel monolithic radio frequency transceiver (see figure 13) which works in the 2.4 GHz-2.5 GHz ISM frequency band. And a frequency synthesizer, a power amplifier, a crystal oscillator, a modulator and other functional modules are built in the frequency synthesizer. The NRF24L01 has low power consumption, the working current is only 9mA when the power is transmitted at-6 dBm, and the working current is only 12.3mA when the power is received, so that the energy-saving design is more convenient due to various low-power working modes (power-down mode and idle mode). In the process of communication debugging of the wireless module, a plurality of problems are encountered, namely, firstly, programs at a receiving and transmitting end must be matched for use, and the wireless module can receive and transmit information only by setting the same baud rate.

4. MPU6050 module

MPU-6050 integrates a 3-axis gyroscope, 3-axis accelerator, and includes external sensors such as accelerators, magnetic sensors, etc. that can be connected to other brands via a second I2C port (see FIG. 14).

The sensing range is that the MPU-6050 has the angular velocity full-grid sensing ranges of + -250, + -500, + -1000 and + -2000 DEG/sec (dps), the fast and slow actions can be accurately tracked, and the user-programmable accelerator Quan Ge has the sensing ranges of + -2 g, + -4 g + -8 g and + -16 g. MPU-6050 can work under different voltages, and VDD power supply voltage is 2.5V+ -5%, 3.0V+ -5% or 3.3V+ -5%, and logic interface VVDIO power supply is 1.8V+ -5%. MPU-6050 has a package size of 4x4x0.9mm (QFN) and is a revolutionary size in the industry. Other features include built-in temperature sensors, oscillators that have only + -1% variation in the operating environment. MPU6050 is a guarantee that the entire aircraft can fly, just as if it were a sensory of an aircraft, reflecting various conditions of the aircraft at all times. Then through the processing of stm32 chip, output signal goes control motor's rotational speed to guarantee that the aircraft can fly steadily, and adjust oneself according to the flight requirement.

5. Flight control module

The system comprises a brain, a core component and an information exchange and processing center of the aircraft. For example, STM32 is used, the specific model being STM32F103VET6.STM32 series is based on ARM Cortex-M3 cores specifically designed for embedded applications requiring high performance, low cost, and low power consumption. The properties are divided into two different families: STM32F103 "enhanced" series and STM32F101 "basic" series. The frequency of the enhanced serial clock reaches 72MHz; the basic clock frequency is 36MHz, and the performance greatly improved by the price of 16-bit products compared with 16-bit products is the best choice for 16-bit product users. Both families have 32K to 128K flash memories built in, except for the combination of SRAM maximum capacity and peripheral interfaces. At a clock frequency of 72MHz, the code is executed from the flash memory, STM32 consumes 36mA, which is the lowest-power-consumption product in the 32-bit market, and is equivalent to 0.5mA/MHz. STM32F103xx enhancement series uses a high performance ARM Cortex-M3 32 bit RISC core with a working frequency of 72MHz, built-in high speed memory (flash memory up to 128K bytes and SRAM 20K bytes), rich enhancement I/O ports and peripherals coupled to two APB buses. All models of devices contain 2 12-bit ADCs, 3 general 16-bit timers and one PWM timer, and also standard and advanced communication interfaces: up to 2I 2C and S, 3 USART, one USB and one CAN (see fig. 15).

6. Electric regulator

Electric speed regulator is called electronic speed regulator, english electronic speed controller, ESC for short. The motor can be divided into a brush electric motor and a brushless electric motor according to different motors. Which adjusts the rotational speed of the motor in response to the control signal. For their connection, this is generally the case:

(1) The input line of the electric regulator is connected with the battery; (2) The output lines (two brushes and three brushless lines) of the electric regulator are connected with the motor; (3) electrically modulated signal lines are connected to the receiver. Meanwhile, the electric regulator can output a voltage-stabilizing direct current power supply of about 5V, can directly supply power for the main control chip, and can omit a voltage-stabilizing circuit. The signal wire supplies power for the receiver, and the receiver supplies power for control equipment such as steering engine.

In this example, a brushless power conditioner is used, with three output lines, a 5V power supply, a ground line, and a PWM signal to control the output lines. The 5V voltage output by the electric regulator is stable enough and can be directly used for supplying power to the flight control.

7. Brushless motor

The brushless DC motor is composed of a motor main body and a driver, and is a typical electromechanical integrated product. The Brushless motor and the brush motor have opposite structures, a rotor of the Brushless motor is permanent magnet steel, the permanent magnet steel is connected with an output shaft together with a shell, a stator is a winding coil, and a reversing brush used for alternating an electromagnetic field of the brush motor is removed, so that the Brushless motor is called a Brushless motor (brush motor). The simple operation principle of the brushless motor is that simply, by changing the alternating frequency and waveform of current waves input to the stator coil of the brushless motor, a magnetic field rotating around the geometric axis of the motor is formed around the winding coil, the magnetic field drives permanent magnet steel on the rotor to rotate, the motor rotates, the performance of the motor is related to factors such as the number of the magnetic steel, the magnetic flux intensity of the magnetic steel, the input voltage of the motor and the like, and the control performance of the brushless motor is greatly related to the input of direct current, the current needs to be changed into 3-phase alternating current by an electronic speed regulator, and a control signal needs to be received from a remote controller receiver to control the rotating speed of the motor so as to meet the use requirement of a model. The brushless motor has the advantages that: the brushless motor KV value is defined as a rotating speed/V, and means a rotating speed value that an input voltage is increased by 1 volt and an idle rotating speed of the brushless motor is increased. From this definition, it is known that the input of the brushless motor voltage follows a strict linear proportional relationship with the motor idle speed. In the design, the KV value of the brushless motor is 980, and the rated working voltage is 11.1v. When the accelerator is regulated to the maximum, the rotating speed of the motor can reach more than two tens of thousands of revolutions per minute, and the east-west aircraft with a load of about 1kg can still fly normally. Due to the high rotational speed of the motor, care should be taken to be scraped by the paddle during the debugging process.

8. Fuel engine

The model is 26CC air-mode petrol engine, 26CC air-mode petrol engine domestic configuration (RUIXING carburetor) 26CC air-mode petrol engine inlet configuration (original WALBRO carburetor). Cooling mode, namely air cooling. The number of strokes of the internal combustion engine is two strokes. Cylinder number is single cylinder.

Control part

The design has only three sensors, can only measure two degrees of freedom of the PITCH and ROLL of the four-axis aircraft and the rotation speed of the fuel engine, and cannot measure the YAW. The idea of the program control is shown in fig. 6.

Attitude control system: the four-axis aircraft is an underactuated system with 6 degrees of freedom and 4 inputs, has the characteristics of instability, strong coupling and the like, and is easily interfered by the outside besides being influenced by the mechanical structure and the rotor aerodynamics of the aircraft. The gesture of the unmanned aerial vehicle is finally regulated by regulating the rotating speeds of the 4 motors, the flight control system obtains gesture information of the unmanned aerial vehicle through each sensor, calculates the rotating speeds of the 4 motors through a certain control algorithm, sends the rotating speeds to a motor speed regulator (called electric regulation for short) through an I2C interface, and regulates the rotating speeds of the 4 motors so as to realize the gesture control of the unmanned aerial vehicle; the control of the fuel engine can be controlled at a constant speed, so that the fuel engine is always at a constant rotating speed, the control procedure is simplified most, and the weight of the whole multi-rotor unmanned aircraft is reduced by 1/3 of that of the original multi-rotor unmanned aircraft. The attitude control is the basis of the whole flight control, and according to the mathematical model of the attitude control subsystem, the states required to be detected by the attitude control system are the angular speeds, the angles and the heights of the unmanned aerial vehicle relative to the ground in the 3 axial directions under the machine body coordinate system. The flight control system is responsible for various tasks such as sensor information acquisition, control algorithm calculation and communication, and is a core of the whole unmanned aerial vehicle, and has the main functions of:

The system comprises a main controller, a sensor, a PC and a system, wherein the main controller can quickly acquire data of each sensor and process the data, the sensor detects the state of the unmanned aerial vehicle in real time, the state comprises information such as posture, position and speed, the main controller can exchange data with the PC, and the system can conduct wireless data transmission.

Advantageous effects

The invention has the advantages of two types of aircrafts, and has the characteristics of long endurance and large load of the fuel oil aircrafts; meanwhile, the electric multi-rotor aircraft has the performance that the electric multi-rotor aircraft can be accurately controlled, can complete the pre-programming, can complete various actions along a specified route under the navigation of a GPS, and has the characteristics of large load and long endurance. The self-generating oil-electricity hybrid power driving mode is adopted, charging is not needed, the multi-rotor aircraft is suitable for field work, the multi-rotor aircraft can fly in a larger range, and the multi-rotor aircraft has a good application prospect.

Drawings

FIG. 1 is a schematic view of an engine main rotor;

FIG. 2 is a schematic diagram of the overall structure;

FIG. 3 is an electrical schematic;

FIG. 4 power management system framework;

FIG. 5 is a block diagram of an electrical apparatus of a self-generating electric hybrid multi-rotor aircraft;

FIG. 6 is a diagram of a self-generating electric hybrid multi-rotor aircraft operating sequence;

FIG. 7 USB is a schematic diagram of an expansion circuit;

FIG. 8 engine damper stepper motor drive circuit;

FIG. 9 generator voltage regulator;

Fig. 10 high-efficiency switching type constant current/constant voltage charger circuit;

FIG. 11 undervoltage protection circuit;

FIG. 12 is a regulated current-limiting DC regulated power supply;

FIG. 13 is a schematic circuit diagram of NRF24l 01;

FIG. 14 is a schematic circuit diagram of the MPU6050 module;

FIG. 15 STM32 chip circuit;

FIG. 16 is a reference voltage circuit diagram;

Fig. 17 is a brushless motor AC-DC circuit.

In fig. 1-17: 1. upper main rotor, 2, lower main rotor, 3, multi-rotor aircraft antenna, 4, multi-rotor aircraft electric control board, 5, motor driven auxiliary rotor, 6, motor, 7, fuel engine, 8, rotation speed dual gearbox, 9, carburetor, 10, motor mount, 11, battery, 12, fuel intake damper, 13, oil delivery pipe, 14, generator damping bracket, 15, oil tank, 16, horn, 17, landing gear, 18, upper main rotor shaft, 19, lower main rotor sleeve shaft, 20, upper bevel gear, 21, lower bevel gear, 22, rotation speed sensor, 23, engine exhaust pipe, 24, throttle control stepper motor, 25, generator, 26, power output line, 27, damping cushion, 28, motor driven bevel gear, 29, motor power input bevel gear.

The specific embodiment is as follows: the invention is further described below with reference to the drawings and examples.

Example one: the self-generating oil-electricity hybrid power multi-rotor aircraft is characterized in that flying power is jointly provided by a main rotor and an auxiliary rotor, the main rotor is driven by a fuel generator 7, the auxiliary rotor 5 is driven by a motor 6, the main rotor and the auxiliary rotor 5 jointly provide lift-off power, the main rotor is composed of an upper main rotor 1 and a lower main rotor 2 (see the figure 1), an upper main rotor rotating shaft 18 passes through a lower main rotor sleeve shaft 19 from the top end and is fixedly connected with an engine power output shaft after being welded with a lower bevel gear 21, the lower main rotor 2 is fixedly arranged at the upper end of the lower main rotor sleeve shaft 19, the sleeve shaft and the upper main rotor rotating shaft 18 can rotate along the same axial lead independently, the lower main rotor sleeve shaft 19 is welded with the upper large end face of an upper bevel gear 20, a left bevel gear and a right bevel gear are arranged between the upper bevel gear 20 and the lower bevel gear 21, and the total four gears are arranged in a speed dual gearbox 8, the left and right bevel gear shafts are fixed with the side wall of the gear box, a rotation speed sensor 21 for monitoring the rotation speed of the main rotor shaft 18 is arranged below the lower bevel gear 21, the gear box body is fixed on the engine body of the fuel engine, the main rotor shaft 18 passes through the lower bevel gear 21 and is connected in series (see figure 2) with a motor driving bevel gear 28, the motor driving bevel gear 28 is fixedly connected with a power output shaft of the fuel engine 7 after passing through the bevel gear, the motor driving bevel gear 28 is meshed with a motor power input bevel gear 29, the rotation shaft of the motor power input bevel gear 29 is the shaft of a rotor of the generator 25, so that partial power of the fuel engine is output to the generator 25, generated power is output through a power output line 26, the main rotor shaft passes through the geometric center of the multi-rotor aircraft, the gravity center of the oil tank 15 is positioned at the geometric center of the multi-rotor aircraft, the landing gear 17 takes the geometric center of the multi-rotor aircraft as the shaft, and (5) axisymmetrically installing.

Example two: the self-generating oil-electricity hybrid power multi-rotor aircraft is characterized in that flying power is jointly provided by a main rotor and an auxiliary rotor, the main rotor is driven by a fuel generator 7, the auxiliary rotor 5 is driven by a motor 6, the main rotor and the auxiliary rotor 5 jointly provide lift-off power, the main rotor is composed of an upper main rotor 1 and a lower main rotor 2 (see the figure 1), an upper main rotor rotating shaft 18 passes through a lower main rotor sleeve shaft 19 from the top end and is fixedly connected with an engine power output shaft after being welded with a lower bevel gear 21, the lower main rotor 2 is fixedly arranged at the upper end of the lower main rotor sleeve shaft 19, the sleeve shaft and the upper main rotor rotating shaft 18 can rotate along the same axial lead independently, the lower main rotor sleeve shaft 19 is welded with the upper large end face of an upper bevel gear 20, a left bevel gear and a right bevel gear are arranged between the upper bevel gear 20 and the lower bevel gear 21, and the total four gears are arranged in a speed dual gearbox 8, the left and right bevel gear shafts are fixed with the side wall of the gear box, a rotation speed sensor 21 for monitoring the rotation speed of the main rotor shaft 18 is arranged below the lower bevel gear 21, the gear box body is fixed on the engine body of the fuel engine, the main rotor shaft 18 passes through the lower bevel gear 21 and is connected in series (see figure 2) with a motor driving bevel gear 28, the motor driving bevel gear 28 is fixedly connected with a power output shaft of the fuel engine 7 after passing through the bevel gear, the motor driving bevel gear 28 is meshed with a motor power input bevel gear 29, the rotation shaft of the motor power input bevel gear 29 is the shaft of a rotor of the generator 25, so that partial power of the fuel engine is output to the generator 25, generated power is output through a power output line 26, the main rotor shaft passes through the geometric center of the multi-rotor aircraft, the gravity center of the oil tank 15 is positioned at the geometric center of the multi-rotor aircraft, the landing gear 17 takes the geometric center of the multi-rotor aircraft as the shaft, and (5) axisymmetrically installing.

The oil delivery pipe 13 of the oil tank 15 is connected with the carburetor 9, the carburetor is connected with the fuel air inlet air door 12, and the wind shield rotating shaft of the air inlet air door is connected with the air door control stepping motor 24 for controlling the opening and closing angles of the wind shield.

The generator 25 is mounted on a generator shock mount 14 provided with shock absorbing cushions 27 to secure the generator 25 to the horn 16.

Example three: the self-generating oil-electricity hybrid power multi-rotor aircraft is characterized in that flying power is jointly provided by a main rotor and an auxiliary rotor, the main rotor is driven by a fuel generator 7, the auxiliary rotor 5 is driven by a motor 6, the main rotor and the auxiliary rotor 5 jointly provide lift-off power, the main rotor is composed of an upper main rotor 1 and a lower main rotor 2 (see the figure 1), an upper main rotor rotating shaft 18 passes through a lower main rotor sleeve shaft 19 from the top end and is fixedly connected with an engine power output shaft after being welded with a lower bevel gear 21, the lower main rotor 2 is fixedly arranged at the upper end of the lower main rotor sleeve shaft 19, the sleeve shaft and the upper main rotor rotating shaft 18 can rotate along the same axial lead independently, the lower main rotor sleeve shaft 19 is welded with the upper large end face of an upper bevel gear 20, a left bevel gear and a right bevel gear are arranged between the upper bevel gear 20 and the lower bevel gear 21, and the total four gears are arranged in a speed dual gearbox 8, the left and right bevel gear shafts are fixed with the side wall of the gear box, a rotation speed sensor 21 for monitoring the rotation speed of the main rotor shaft 18 is arranged below the lower bevel gear 21, the gear box body is fixed on the engine body of the fuel engine, the main rotor shaft 18 passes through the lower bevel gear 21 and is connected in series (see figure 2) with a motor driving bevel gear 28, the motor driving bevel gear 28 is fixedly connected with a power output shaft of the fuel engine 7 after passing through the bevel gear, the motor driving bevel gear 28 is meshed with a motor power input bevel gear 29, the rotation shaft of the motor power input bevel gear 29 is the shaft of a rotor of the generator 25, so that partial power of the fuel engine is output to the generator 25, generated power is output through a power output line 26, the main rotor shaft passes through the geometric center of the multi-rotor aircraft, the gravity center of the oil tank 15 is positioned at the geometric center of the multi-rotor aircraft, the landing gear 17 takes the geometric center of the multi-rotor aircraft as the shaft, and (5) axisymmetrically installing.

Claims

1. A self-generating oil-electricity hybrid power multi-rotor aircraft is characterized in that: the motor-driven multi-rotor aircraft comprises an upper main rotor (1), a lower main rotor (2), a multi-rotor aircraft electric control board (4), a motor driving auxiliary rotor (5), a motor (6), a fuel engine (7), a rotating speed dual gearbox (8), a carburetor (9), a battery (11), a fuel inlet air door (12), an oil delivery pipe (13), an oil tank (15), a horn (16), an upper main rotor shaft (18), a lower main rotor sleeve shaft (19), an upper bevel gear (20), a lower bevel gear (21), a rotating speed sensor (22), an air door control stepping motor (24), a generator (25) and a damping cushion pad (27); the self-generating oil-electricity hybrid power multi-rotor aircraft is characterized in that a main rotor and a motor drive auxiliary rotor (5) jointly provide flying power, the main rotor is driven by a fuel engine (7), and the motor drive auxiliary rotor (5) is driven by a motor (6); the main rotor consists of an upper main rotor (1) and a lower main rotor (2), an upper main rotor shaft (18) passes through a lower main rotor sleeve shaft (19) from the top end and then is welded with a lower bevel gear (21) and then is fixedly connected with a power output shaft of a fuel engine (7), the lower main rotor (2) is fixedly arranged at the upper end of the lower main rotor sleeve shaft (19), the lower main rotor sleeve shaft (19) and the upper main rotor shaft (18) can rotate independently along the same axial lead, the lower main rotor sleeve shaft (19) is welded with the upward large end face of an upper bevel gear (20), and a left bevel gear and a right bevel gear are arranged between the upper bevel gear (20) and the lower bevel gear (21); the motor-driven multi-rotor aircraft is arranged in a rotating speed dual gearbox (8), a rotating shaft of a left bevel gear and a rotating shaft of a right bevel gear are fixed on the side wall of the rotating speed dual gearbox (8), a rotating speed sensor (22) for monitoring the rotating speed of an upper main rotor shaft (18) is arranged below a lower bevel gear (21), the rotating speed dual gearbox (8) is fixed on an engine body of the fuel engine (7), the upper main rotor shaft (18) penetrates through the lower bevel gear (21) and is connected with a motor-driven bevel gear (28) in series, the upper main rotor shaft (18) penetrates through the motor-driven bevel gear (28) and is fixedly connected with a power output shaft of the fuel engine (7), the motor-driven bevel gear (28) is meshed with a motor power input bevel gear (29), the rotating shaft of the motor power input bevel gear (29) is the shaft of a rotor of the generator (25), part of power of the fuel engine (7) is output to the generator (25), power generated by the generator (25) is output through a power output line (26), the upper main rotor shaft (18) penetrates through the geometrical center of the self-powered multi-rotor aircraft, the fuel engine (7) is fixedly arranged on an upper horn (16), and the geometrical center of the self-powered multi-rotor aircraft is arranged between the horn (16) and the self-powered multi-rotor aircraft, and the vibration-damping center of the motor-driven multi-rotor aircraft is arranged between the horn (16 and the motor-driven vibration absorber; the landing gear (17) is arranged in an axisymmetric way by taking the geometric center of the self-generating oil-electricity hybrid power multi-rotor aircraft as an axis;

The oil delivery pipe (13) of the oil tank (15) is connected with the carburetor (9), the carburetor is connected with the fuel air inlet air door (12), the wind shield rotating shaft of the air inlet air door is connected with the air door control stepping motor (24) for controlling the opening and closing angle of the wind shield, the generator (25) is arranged on the generator damping bracket (14), the generator damping bracket (14) is provided with a damping buffer cushion (27), and the damping buffer cushion (27) is used for fixing the generator (25) on the horn (16) arranged on the generator damping bracket (14);

The multi-rotor aircraft antenna (3) and the multi-rotor aircraft electric control board (4) are mounted on the horn (16), the battery (11) is mounted below the frame, the battery fixing frame is adjusted to enable the center of gravity of the multi-rotor aircraft antenna (3), the multi-rotor aircraft electric control board (4) and the battery to be integrated into a whole, a motor seat (10) is mounted at the end of the horn (16) and is used for fixing a motor (6), a motor power output shaft is used for driving a secondary rotor (5), and the projection interval of the primary rotor and the secondary rotor (5) on a plane formed by the crossed horn (16) is larger than 2cm, so that air flows of the two are not mutually disturbed;

The power management system comprises a voltage-stabilizing rectification module, a current-limiting module, a charging module, a relay system, a load circuit, a battery pack, an undervoltage protection module, a control module and a fuel oil stock display module; the fuel engine (7) drives the generator (25) to generate alternating current through bevel gear transmission, exciting current of the generator is regulated through the voltage stabilizing rectifying module, so that the generator outputs relatively stable direct current, output current is limited through the current limiting circuit, overlarge load of the generator is avoided, the current enters the charging module, the current is distributed through the relay system, the relay system is controlled by the controller to intelligently control charge and discharge of the load and the storage battery, charging and power supply current are intelligently regulated, the battery pack is provided with an undervoltage protection circuit, and the controller is also responsible for transmitting fuel stock and charge and discharge data of the load and the storage battery to the ground control and monitoring platform through the wireless radio frequency module.