CN109383782B - Danger-escaping energy-saving power-assisting system of aero-engine - Google Patents
Danger-escaping energy-saving power-assisting system of aero-engine Download PDFInfo
- Publication number
- CN109383782B CN109383782B CN201811012821.1A CN201811012821A CN109383782B CN 109383782 B CN109383782 B CN 109383782B CN 201811012821 A CN201811012821 A CN 201811012821A CN 109383782 B CN109383782 B CN 109383782B
- Authority
- CN
- China
- Prior art keywords
- power
- engine
- speed
- electric
- pair
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 230000002829 reductive effect Effects 0.000 claims abstract description 35
- 238000003860 storage Methods 0.000 claims abstract description 30
- 238000010248 power generation Methods 0.000 claims description 77
- 238000001514 detection method Methods 0.000 claims description 50
- 230000001276 controlling effect Effects 0.000 claims description 20
- 230000005611 electricity Effects 0.000 claims description 19
- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 claims description 14
- 230000001133 acceleration Effects 0.000 claims description 10
- 230000003247 decreasing effect Effects 0.000 claims description 9
- 238000012545 processing Methods 0.000 claims description 9
- 230000001105 regulatory effect Effects 0.000 claims description 7
- 230000005540 biological transmission Effects 0.000 claims description 6
- 230000000295 complement effect Effects 0.000 claims description 6
- 238000007599 discharging Methods 0.000 claims description 3
- 230000000630 rising effect Effects 0.000 claims 1
- 230000000694 effects Effects 0.000 abstract description 10
- 238000006243 chemical reaction Methods 0.000 abstract description 6
- 238000000034 method Methods 0.000 description 42
- 230000008569 process Effects 0.000 description 22
- 238000005259 measurement Methods 0.000 description 17
- 239000000446 fuel Substances 0.000 description 16
- 230000008859 change Effects 0.000 description 11
- 230000033228 biological regulation Effects 0.000 description 9
- 238000010586 diagram Methods 0.000 description 9
- 239000004020 conductor Substances 0.000 description 5
- 230000008878 coupling Effects 0.000 description 5
- 238000010168 coupling process Methods 0.000 description 5
- 238000005859 coupling reaction Methods 0.000 description 5
- 230000005484 gravity Effects 0.000 description 5
- 230000007423 decrease Effects 0.000 description 4
- 238000005265 energy consumption Methods 0.000 description 4
- 230000002441 reversible effect Effects 0.000 description 4
- 238000002360 preparation method Methods 0.000 description 3
- 230000001360 synchronised effect Effects 0.000 description 3
- 238000004364 calculation method Methods 0.000 description 2
- 238000012937 correction Methods 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 230000009977 dual effect Effects 0.000 description 2
- 238000004134 energy conservation Methods 0.000 description 2
- 230000007246 mechanism Effects 0.000 description 2
- 230000005693 optoelectronics Effects 0.000 description 2
- 238000012546 transfer Methods 0.000 description 2
- WHXSMMKQMYFTQS-UHFFFAOYSA-N Lithium Chemical compound [Li] WHXSMMKQMYFTQS-UHFFFAOYSA-N 0.000 description 1
- 241000282376 Panthera tigris Species 0.000 description 1
- 230000009471 action Effects 0.000 description 1
- 230000002411 adverse Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000013016 damping Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 230000003111 delayed effect Effects 0.000 description 1
- 230000008030 elimination Effects 0.000 description 1
- 238000003379 elimination reaction Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 230000000670 limiting effect Effects 0.000 description 1
- 229910052744 lithium Inorganic materials 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000003287 optical effect Effects 0.000 description 1
- 230000004044 response Effects 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
- 239000000725 suspension Substances 0.000 description 1
- 238000012549 training Methods 0.000 description 1
- 230000001052 transient effect Effects 0.000 description 1
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/04—Helicopters
- B64C27/12—Rotor drives
- B64C27/14—Direct drive between power plant and rotor hub
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/02—Aircraft characterised by the type or position of power plants
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/02—Aircraft characterised by the type or position of power plants
- B64D27/026—Aircraft characterised by the type or position of power plants comprising different types of power plants, e.g. combination of a piston engine and a gas-turbine
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/02—Aircraft characterised by the type or position of power plants
- B64D27/24—Aircraft characterised by the type or position of power plants using steam or spring force
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
Landscapes
- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Mechanical Engineering (AREA)
- Control Of Eletrric Generators (AREA)
- Hybrid Electric Vehicles (AREA)
Abstract
The invention relates to an escape energy-saving power-assisted system of an aero-engine. The aero-engine danger-escaping energy-saving power assisting system provides instantaneous power assisting torque required by straightening through adjusting the working voltage of the motor, then the instantaneous power assisting torque provided by the motor is gradually replaced by the engine torque coming after delay, the instantaneous power assisting effect is realized, on one hand, the torque coming after oil movement delay is compensated by means of the electric sensitive reaction characteristic, on the other hand, the oil movement follows the instantaneous power assisting torque provided by the motor in time, the electric power consumption is small, a large-capacity storage battery does not need to be configured, the dead weight of the multi-rotor helicopter with oil movement does not need to be increased, the advantages are gained and the disadvantages are compensated, namely, the falling probability is reduced, and meanwhile, the cruising ability is improved.
Description
Technical Field
The invention relates to the technical field of helicopters, in particular to an aero-engine power assisting device, an aero-engine danger-escaping power assisting device, an aero-engine energy-saving power assisting device and an aero-engine danger-escaping energy-saving 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, because the working principle of an engine of a multi-rotor helicopter driven by a fuel engine determines that the engine has delayed control reaction after waiting for two strokes of air exhaust and air suction. After the response comes, the engine continues to suck a small amount of given oil again in an increasing mode at an acceleration gradually changed by changing the rotating speed and the suction force of the previous suction of the mixed gas at every two strokes, and then adjusts the attitude of the multi-rotor helicopter swaying in the wind, which is too late and obviously has high probability of breaking.
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, a complementary oil/electric hybrid scheme 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 aero-engine risk-escaping energy-saving power-assisted system is needed to be provided for solving the problem that the multi-rotor helicopter is poor in cruising ability.
The energy-conserving helping hand system of aeroengine escape from danger includes: an aero-engine power assisting device, an aero-engine danger-escaping power assisting device and an aero-engine energy-saving power assisting device.
The aircraft engine power assisting device comprises: 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 an electric power-assisted system with the engine driving a plurality of rotors.
In one embodiment, the engine is further configured to power a propeller.
In one embodiment, the engine is further used to power the battery charge.
The electric/power generation dual-purpose machine 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 to be in an electric power-on state, the calibration voltage less than 1/2 is applied to be in an electric power-off state, and the voltage of 0 is applied to be in a power generation 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 power assisting 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 power generation state, the electric energy is supplemented to the storage battery.
The 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 is used for providing electric energy for an electric power assisting system of the engine driving multiple rotors, and the capacity of the storage battery is not more than the power consumption of the electric/power generation dual-purpose machine for working 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, the motor applies rated voltage, 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, which is marked as A1 set of minimum oil motion parameters. 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 to remain 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 of 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, and ending the process, wherein the nominal voltage is the normal working reference voltage of the flight.
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 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-assisted effect is realized through the sequential relay mode of electric power and oil power, on one hand, the torque of oil power lagging is compensated by means of the electric sensitive reaction characteristic, on the other hand, the oil power immediately replaces the instant power-assisted torque provided by the motor in time, so that the electric power consumption is low, 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 are gained and the shortages are compensated, namely, the probability of falling the helicopter is reduced, and the endurance capacity is improved by hundreds of times.
The aero-engine danger-escaping energy-saving power assisting device further comprises: 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, 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 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.
Above-mentioned aeroengine energy-conserving booster unit of escaping danger through setting up the power balance control ware for when aeroengine helping hand system assembly takes place the slope, realize controlling aeroengine helping hand system to put in a sensitive way, avoid appearing breaking down the phenomenon, improved aeroengine helping hand system assembly's flight stability.
The aero-engine danger-escaping assistance system 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 pair of danger escaping detection units is used for: the rotation 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 rotation speeds of the pair of electric/power generation dual-purpose machines are greater than the rotation speeds of the pair of engines, the shafts at the two ends of the pair of ratchets are disconnected, so that rotation speed difference occurs; when the rotational speed difference is detected by the pair of escape detection units, respectively, an emergency process is performed by the system controller.
The arrangement of the emergency-escaping energy-saving power assisting device for the aero-engine ensures that the electric function of the corresponding electric/power-generating dual-purpose machine can still work when any engine fails, and improves the safety performance of the multi-rotor helicopter.
The emergency escape energy-saving power assisting device of the aero-engine 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 flying 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 voltage reducing 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 energy-saving power assisting device for escape of the aircraft engine, the power balance controller is arranged, so that when the power assisting system of the aircraft engine inclines, the power assisting system of the aircraft engine is sensitively controlled to be aligned, the phenomenon of falling is avoided, and the flight stability of the energy-saving power assisting system of the aircraft engine is improved.
The emergency escape energy-saving power assisting device of the aero-engine further comprises a permanent magnet speed regulating device. The permanent magnet speed regulation device comprises a permanent magnet speed regulator, an automatic speed locker, an auxiliary speed regulator, a rotating speed sensor and an output power measurer, wherein the engine is coaxially and mechanically connected with the permanent magnet speed regulator, the automatic speed locker is respectively in signal connection with the engine, the rotating speed sensor and the output power measurer, and the auxiliary speed regulator is in signal connection with the permanent magnet speed regulator and the automatic speed locker.
The automatic speed locker comprises: a method for locking the output of an engine at an optimal constant speed.
In one embodiment, the optimum speed of the engine is stored in the automatic speed locker.
The optimal constant rotating speed is the rotating speed with the best energy conversion efficiency which is always output by the engine.
In one embodiment, a map of engine characteristic parameters is stored in the automatic transmission.
In one embodiment, the automatic governor continuously acquires the actually required rotating speed in the auxiliary governor, the instant rotating speed of the engine in the rotating speed sensor and the instant power in the output power measurer, calculates the difference between the actual required rotating speed and the optimal rotating speed, calculates the kinetic energy difference between the actual required rotating speed and the optimal rotating speed according to the characteristic curve parameters of the engine, performs linear compensation, and adjusts the accelerator/electric quantity according to the formula of torque 9550 × power/rotating speed so that the engine outputs at the optimal rotating speed and the actually required torque is ensured to be output.
The auxiliary governor includes: a method for controlling a permanent magnet speed regulator to output an actually required rotating speed.
In one embodiment, a characteristic curve parameter chart of the permanent magnet speed regulator is stored in the auxiliary speed regulator.
In one embodiment, after the auxiliary speed regulator receives the speed regulation information, the auxiliary speed regulator performs linear compensation according to parameters corresponding to a characteristic curve of the permanent magnet speed regulator, and then adjusts an air gap of the permanent magnet speed regulator so that the permanent magnet speed regulator outputs a specified rotating speed.
The smaller the air gap between the permanent magnet speed regulator and a conductor in the permanent magnet speed regulator is, the larger the torque transmitted by coupling is. When the air gap is determined, the rotating speed and the torque output by the permanent magnet speed regulator are obtained by the automatic speed regulator through calculating and adjusting the throttle/electric quantity of the engine.
The revolution speed sensor includes: photoelectric speed measuring structure and mechanical speed measuring structure.
In one embodiment, the mechanical speed measuring structure includes a centrifugal wheel speed measuring device, the centrifugal wheel speed measuring device is connected with the automatic speed locker, and the automatic speed locker is used for performing linear compensation on the error of the rotating speed measured by the centrifugal wheel speed measuring device according to the characteristic parameters of the engine
The automatic speed locker and the auxiliary speed regulator are matched with each other to carry out real-time measurement and control on the engine and the permanent magnet speed regulator, so that the rotating speed and the torque which are actually required by the permanent magnet speed regulator are ensured to be output, and the engine always runs at the optimal rotating speed. In addition, the electric quantity/accelerator of the prime motor is adjusted, the electric quantity/accelerator is calculated according to the operation characteristic and the torque of the prime motor which are 9550 multiplied by the power/rotating speed formula, and the electric quantity/accelerator is converted into the torque output, so the energy-saving effect is achieved. Therefore, the cruising ability of the multi-rotor helicopter is improved.
Drawings
FIG. 1 is a schematic structural diagram of a first aero-engine power assisting device provided in one embodiment;
FIG. 2 is a schematic structural diagram of an aircraft engine escape assisting device 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 according to an embodiment;
fig. 4 is a schematic view of a flow chart of attitude adjustment control under actual takeoff weight of the multi-rotor helicopter according to the embodiment;
FIG. 5 is a schematic flow chart diagram of a method of assisting power in an engine according to an embodiment;
FIG. 6 is a schematic structural diagram of a first emergency escape energy-saving power assisting device of an aircraft engine according to an embodiment;
FIG. 7 is a schematic structural diagram of a second emergency escape energy-saving power assisting device of an aircraft engine according to an embodiment;
FIG. 8 is a schematic structural diagram of a third aero-engine risk elimination energy-saving power-assisted system according to an embodiment;
FIG. 9 is a schematic structural diagram of a fourth aero-engine risk-escaping energy-saving power-assisted system according to an embodiment
FIG. 10 is a schematic diagram of a permanent magnet governor according to an embodiment;
FIG. 11 is a characteristic diagram of an engine provided in an embodiment.
Detailed Description
For better understanding of the objects, technical solutions and effects of the present invention, the present invention will be further explained with reference to the accompanying drawings and examples. It is to be noted that the following examples are provided only for illustrating the present invention and are not intended to limit the present invention.
Referring to fig. 1 and 2 together, the assembly system of the paired oil/electric hybrid rotor includes: the combined power system 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 the combined system of the paired oil/electric hybrid rotor.
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 charging of the battery 105.
The electric/power generation dual-purpose machine 102 is used for applying 1/2 calibration voltage in a normal state, wherein 1/2 calibration voltage is a reference voltage in an electric power assisting state, a voltage greater than 1/2 is applied to be in an electric power-on state, a voltage less than 1/2 is applied to be in an electric power-off state, and a voltage of 0 is applied to be 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.
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/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/electric dual-purpose machine 102 and the propeller 100, measurement and control are carried out by the command system controller 108 of the flight controller 104, the storage battery 105 provides electric energy for the assembly system of the pair of the oil/electric hybrid power rotors, and the capacity of the storage battery 105 is not more than the electric quantity of the electric/electric dual-purpose machine 102 for 30 seconds.
The system controller 108 includes: power detection means 200, flight control command processing means 301, electric/power generation switching means 300, electric power amount detection means 400, engine power allocation means 401, and charge/discharge control means 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 rotation 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, which is labeled as a1 set of minimum fuel economy 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 inclination angle α notified by the flight controller 104 is equal to 0, the adjustment is ended.
The assembly system of the pair of the oil/electricity hybrid power rotor wing obtains the inclination parameters of the multi-rotor wing 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 yaw-up 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 after the lag. 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 motors/generators 102 to switch to a complementary mode of generating electricity and decreasing the rotation speed and electrically increasing the rotation speed, respectively, through the pair of system controllers 108, so that the left-hand propeller 100 is decelerated while the right-hand propeller 100 is decelerated, and the increased speed value is equal to the decelerated value.
In one embodiment, the present invention provides a multi-rotor helicopter power system: comprises a propeller 100, an engine 101 and a motor/generator 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.
The propeller 100 is rigidly connected to the shaft 20c as a system-driven end fitting of the multi-rotor helicopter power system. When the propeller 100 is in a working state, the power of the propeller 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 of a motor and a generator, and the two operating states are respectively performed 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 a generating operation state, the generator is driven by the shaft 20c to convert the mechanical energy into electric energy for generating electricity. Due to the operating characteristics of the motor, the motor can produce a momentary 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 motor/generator 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 power system of the multi-rotor helicopter 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 solve the defect of force application lag caused by long power stroke of an 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.
An embodiment of the present invention further provides a paired oil/electricity hybrid rotor assembly system based on the above 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 data of the speed/torque/voltage characteristic curve of the motor and the data of the speed/torque/voltage characteristic curve 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 speed measurer 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 the torque sensor 106. The torque sensor 106 is fixed to the coaxial shaft 20c and picks up the instantaneous torque of the coaxial 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 storage 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 for powering the flight control system on a multi-rotor helicopter, i.e., the battery 105 is provided on the multi-rotor helicopter, 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, wherein 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 aero-engine 101 power-assisted system assembly, and the aero-engine 101 power-assisted system assembly is 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;
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, a1 set of minimum oil-motion parameter table, and a0 set of 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;
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 switched to the motor mode simultaneously by the pair of system controllers 108 when the flight controller 104 commands an increase;
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 the 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 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 detection unit 400 is configured to detect the power of the battery 105 connected thereto, and provide 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 amount of electricity in the storage battery 105 is lower than a preset threshold value, the control instruction provided by the electricity amount 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 according to the measured instantaneous torque and instantaneous speed of the coaxial 20c, so that the engine power allocating unit 401 can adjust the output power of the engine 101.
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 allocation 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 coaxial 20c is greater than n, the motor/generation 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 power 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 an embodiment, an energy-saving assistance system for aircraft engine escape is further provided, please refer to fig. 7, 8 and 9, 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 the system controllers 108.
In one embodiment, the system for assisting in escape and energy conservation of an aircraft engine 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 mechanism 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, during flight, especially in windy environments, a tilt situation is likely to occur in a multi-rotor helicopter. 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 to obtain 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 the 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 the preset database in the system controller; the rotational speed of the propellers can be obtained using a tachometer 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 output power of the engine by adjusting the working voltage of the motor in the multi-rotor helicopter according to the inclination parameters, so that the motor and the engine finish the alignment of the multi-rotor helicopter together.
S206, adjusting the rotating speed of the propeller according to the adjusted working voltage and output power until the multi-rotor helicopter is straightened.
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 put in place, the output power of the engine is adjusted again, and the working voltage of the motor returns to one-half of the rated voltage, that is, the multi-rotor helicopter returns 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, and meanwhile, the engine is gradually used for replacing the motor to provide the aligning torque required by aligning for the propeller, so that 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 power and electric power, on one hand, the instant power assisting is provided by means of an electric control mode, the state control of the multi-rotor helicopter is sensitively realized, the aligning torque required by the aligning of the propeller can be quickly generated, on the other hand, the engine gradually replaces a motor to provide the aligning torque required by the aligning for the propeller, the power consumption of the motor is very small, a power supply with very large capacity is not needed, and therefore the cruising ability of the multi-rotor helicopter is ensured.
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, firstly, a direct current generator provided by an engine of the multi-rotor helicopter and used for providing power for a flight control system is changed into a motor/power generation direct current dual-purpose machine controlled and switched by a system controller. The improved electric/power generation direct current 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 capacity of the battery 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 attitude correction of the stress application lag of the engine, the power of the motor is less than 5 percent of the power of the engine, the energy consumption of the cart is avoided being pulled by a big horse, and the effect of improving the cruising ability 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 correcting torque required by the multi-rotor helicopter to correct 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 correcting torque, and adjusting the output power of the engine.
Specifically, when the multi-rotor helicopter inclines, the yaw torque required by the yaw of 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 relationship of rotating speed/torque/lift force characteristic 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:
and 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. And (3) putting the multi-rotor helicopter 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 rotor rotation speed, the rotor torque, and the output power of each engine of the multi-rotor helicopter in the above-described state (i.e., when the multi-rotor helicopter is hovering in the air in a fully loaded state) are used as the first hovering parameter of the multi-rotor helicopter. Further, the first hover parameter is saved in the system controller for quick recall in the future as 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, a data preparation process is involved to further obtain a calibration voltage for 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 does not have an electric motor and hovers in the air as second hovering 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 is, the higher the measurement accuracy of the calibration voltage is. 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 working voltage to the 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 hovers 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, there is one nominal voltage of each motor for each takeoff weight. Of course, it is also possible that each takeoff weight corresponds to the calibration voltage of a plurality of motors 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 previous 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 simultaneously, 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 this embodiment, the actually measured motor speed/torque/voltage characteristic curve table is stored in the 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 complete rotary wing machines (the accuracy of the gravity center 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 under 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 power p of the engine, 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 comprises 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 to the motor which is less than one half of a calibration voltage so as to decelerate a propeller corresponding to the motor;
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 specific embodiment, a specific process is involved how a multi-rotor helicopter generates electricity. In this embodiment, the method comprises 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 born by the work of the motor, at the moment, the working voltage of the motor is zero volt, the electric/power generation direct current dual-purpose machine is converted into a power generator state from the motor state, and the power supply in the multi-rotor helicopter is charged. It should be clear that in the process of generating electricity, the propeller input power is equal to the engine output power, i.e., the power consumed by the generator.
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 aircraft engine risk-escaping energy-saving boosting system is provided, which includes the steps described in S202-S208 above, and further includes:
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.
Referring to fig. 10, in one embodiment, the permanent magnet speed adjusting device further comprises a pair of permanent magnet speed adjusting devices, one of the permanent magnet speed adjusting devices 113 comprises a permanent magnet speed adjuster 1004, an automatic speed locker 1006 and an auxiliary speed adjuster 1008;
one of the motors 1002 is connected to one of the permanent magnet governors 1004 and one of the automatic governors 1006, respectively;
one of the permanent magnet governors 1004 is coupled to one of the automatic governors 1006 and to the system controller 108 via the auxiliary governor 1008.
The automatic speed locker 1006 is configured to obtain a characteristic parameter table of the engine in advance, store the table in the automatic speed locker 1006, obtain a rotation speed of the rotation speed sensor 1100 when the automatic speed locker 1006 operates, perform linear compensation by comparing the characteristic parameter stored in the automatic speed locker 1006, and control an accelerator to adjust a working state of the engine so that the engine operates at a preset optimal rotation speed;
the auxiliary speed regulator 1008 is used for acquiring a characteristic parameter chart of the permanent magnet speed regulator in advance and storing the characteristic parameter chart in the auxiliary speed regulator, acquiring a control instruction signal when the auxiliary speed regulator works, and controlling electric quantity to adjust the rotating speed required by the output of the permanent magnet speed regulator 1004 after linear compensation is carried out on the characteristic parameter stored in the auxiliary speed regulator 1008.
In one embodiment, further comprising: and one of the speed sensors 1100 is connected with one of the permanent magnet speed regulators 1004 and one of the automatic speed lockers 1006 respectively, and is used for acquiring the speed information of the engine 1002 through the permanent magnet speed regulator 1004 and feeding the speed information back to the automatic speed lockers.
In one embodiment, further comprising: a pair of output power measuring devices 1009, one of the output power measuring devices 1009 is connected to one of the permanent magnet speed controllers 1004 and one of the automatic speed lockers 1006, respectively, and is configured to obtain output power information of the engine through the permanent magnet speed controllers 1004 and send the output power information to the automatic speed lockers 1006, so that the automatic speed lockers can regulate and control the engine to output an optimal rotating speed and output a required torque.
In one embodiment, the permanent magnet governor 1004 is coaxially connected to the engine 1002, and the permanent magnet governor 1004 includes an output drive shaft for outputting the rotational power regulated by the permanent magnet governor 1004 to a load device.
The automatic speed locker 1006 is configured to obtain a characteristic parameter table of the engine in advance, store the table in the automatic speed locker 1006, obtain a rotation speed of a rotation speed sensor when the automatic speed locker 1006 operates, and control an accelerator to adjust a working state of the engine after linear compensation is performed by comparing the characteristic parameter table stored in the automatic speed locker 1006, so that the engine operates at a preset optimal rotation speed.
As another embodiment, the permanent magnet speed governor 1004 receives the engine speed in real time and stores it so that the automatic speed governor can directly read the data stored in the permanent magnet speed governor to obtain the engine speed.
The auxiliary speed regulator 1008 is used for acquiring a characteristic parameter chart of the permanent magnet speed regulator in advance and storing the characteristic parameter chart in the auxiliary speed regulator 1008, and the auxiliary speed regulator 1008 acquires a control instruction signal when working, and controls electric quantity to adjust the rotating speed required by the output of the permanent magnet speed regulator after linear compensation is carried out on the characteristic parameter stored in the auxiliary speed regulator 1008.
The characteristic parameters of the engine 1002 include all the operation characteristic parameters related to the engine operation state, such as the rotation speed and the like. The characteristic parameters of the permanent magnet speed regulator comprise all the working characteristic parameters related to the working state of the permanent magnet speed regulator, such as rotating speed and the like.
The prime mover refers to all machines which utilize energy to generate prime power, and is the main source of power needed in modern production and living fields. The prime mover may include an engine and an electric motor, with the engine including a dc engine and a dc electric motor that are capable of adjusting the output speed. The permanent magnet speed regulator mainly comprises three components: permanent magnet rotor, conductor rotor, speed control mechanism, its theory of operation does: the adjuster adjusts the relative position of the cylindrical permanent magnet rotor and the cylindrical conductor rotor in the axial direction to change the coupling effective part of the permanent magnet rotor and the conductor rotor, namely the torque transmitted between the permanent magnet rotor and the conductor rotor can be changed, repeatable, adjustable and controllable output torque and rotating speed can be realized, and the purposes of speed regulation and energy conservation are realized.
Specifically, the output end of the engine 1002 is connected to the input end of a permanent magnet speed regulator 1004, and the output end of the permanent magnet speed regulator 1004 is connected to the input end of a load. The permanent magnet speed controller 1004 adopts an air gap to transfer torque, and utilizes permanent magnet coupling force to transfer torque in a non-contact manner, so that stepless speed regulation is realized. There is no rigid connection between the engine and the load device and there is damping by slip during mechanical shock, thus greatly reducing vibration and noise.
The automatic speed locker 1006 obtains an actual characteristic parameter of the engine in real time, and adjusts a working state of the engine 1002, such as an instant rotation speed, according to the actual characteristic parameter, so that the engine 1002 can work at a preset optimal rotation speed. It is understood that the characteristic parameters of engine 1002 include, but are not limited to, engine power, speed, torque, specific energy consumption, etc. The optimum rotational speed of the engine 1002 can be finally obtained by largely performing rotational speed training on the engine 1002 to obtain a rotational speed at which the efficiency thereof is optimum, and the optimum rotational speed is previously stored in the system.
The auxiliary speed regulator 1008 obtains a speed regulation control signal from the speed regulation control end, and adjusts the coupling magnetic gap based on the permanent magnet coupling characteristic of the permanent magnet speed regulator 1004 according to the speed regulation control signal to adjust the output rotating speed of the permanent magnet speed regulator 1004 and ensure that the permanent magnet speed regulator 1004 outputs the rotating speed meeting the requirement of the speed regulation control end.
The engine speed regulating system obtains characteristic parameters of the engine in real time, and then the automatic speed locker 1006 controls electric quantity or an accelerator, so that the engine 1002 always runs at a preset optimal rotating speed, the output rotating speed is ensured to be a constant rotating speed, and the engine is transformed into a prime mover outputting the constant rotating speed. The speed of the permanent magnet speed regulator 1004 is regulated by the auxiliary speed regulator 1008, and the constant rotating speed output by the engine is changed to the working rotating speed required by the work and is output. At this time, the power output by the permanent magnet speed regulator 1004 is equal to the power output by the direct current motor or the engine, the change of the load is equal to the change of the torque output by the direct current motor or the engine, the rotating speed of the engine which always works at the optimal efficiency is not changed, the engine is ensured to always work at the optimal efficiency state, and therefore, the energy-saving effect is achieved.
In one embodiment, the prime mover speed regulation system further comprises a pair of speed sensors 1100, wherein the speed sensors 1100 are respectively connected to the permanent magnet speed regulator 1004 and the automatic speed regulator 1006, and are used for acquiring the speed information of the engine 1002 through the permanent magnet speed regulator 1004 and feeding the speed information back to the automatic speed regulator 1006.
Specifically, the rotation speed sensor 1100 obtains the rotation speed information of the engine in real time and sends the rotation speed information to the automatic speed locker 1006, the automatic speed locker receives the rotation speed information, compares the instant rotation speed in the rotation speed information with the preset optimal rotation speed based on the characteristic curve of the engine, calculates the rotation speed difference between the instant rotation speed and the optimal rotation speed of the engine, and then quantitatively compensates the electric quantity or the oil quantity of the engine in real time according to the rotation speed difference to ensure that the engine always works at the optimal rotation speed. By adopting the structural mode of information feedback, the engine can be accurately controlled to work at the optimal rotating speed.
Further, as an embodiment, the speed sensor 1100 includes an optoelectronic speed measuring structure (not shown) and a mechanical speed measuring structure (not shown), and the optoelectronic speed measuring structure and the mechanical speed measuring structure are respectively connected to the automatic speed locker, and the automatic speed locker is configured to calculate a difference between an instant speed and an optimal speed of the engine.
Furthermore, in an embodiment, the mechanical speed measuring structure includes a centrifugal wheel speed measuring device (not shown), and the centrifugal wheel speed measuring device is connected to the automatic speed locker, and the automatic speed locker is configured to linearly compensate for an error of a rotating speed measured by the centrifugal wheel speed measuring device according to a characteristic parameter of the engine.
Specifically, one or more centrifugal wheel speed measuring devices are arranged in the rotating speed sensor, the rotating speed of the centrifugal wheel speed measuring devices is the same as that of the permanent magnet speed regulator, and when the rotating speed of the permanent magnet speed regulator is increased, the centrifugal wheel speed measuring devices can drive a valve of the engine to be opened due to the fact that centrifugal force is increased, and therefore the rotating speed of the engine is increased. When the rotating speed of the permanent magnet speed regulator is reduced, the centrifugal wheel speed measuring device can drive a valve of the engine to be slowly closed due to the fact that centrifugal force is reduced, and the rotating speed of the engine is reduced.
As an alternative embodiment, the centrifugal wheel speed measuring device comprises an even number of centrifugal wheels (not shown), and the even number of centrifugal wheels are connected with a valve of the engine and used for controlling the opening and closing of the valve.
Specifically, there are an even number of centrifugal wheels in the rotational speed sensor, such as 2, 4, etc. The rotating speed of the centrifugal wheel is the same as that of the permanent magnet speed regulator, and when the rotating speed of the permanent magnet speed regulator is increased, the centrifugal wheel drives a valve of the engine to be opened due to the fact that the centrifugal force is increased, so that the rotating speed of the engine is increased. When the rotating speed of the permanent magnet speed regulator is reduced, the centrifugal wheel drives a valve of the engine to be slowly closed due to the fact that centrifugal force is reduced, and the rotating speed of the engine is reduced. An even number of centrifugal wheels are arranged, so that the adverse effect caused by vibration of the centrifugal wheels can be eliminated.
As another alternative, the centrifugal wheel speed measuring device includes an even number of weighted spherical conical pendulums, and the working principle of the centrifugal wheel speed measuring device is the same as that of the centrifugal wheel, and the description thereof is omitted.
It should be clear that, the above-mentioned control method for controlling the valve of the engine by the speed measuring device of the centrifugal wheel may also include that when the rotation speed of the permanent magnet speed regulator is increased, the speed measuring device of the centrifugal wheel drives the valve of the engine to close slowly due to the increase of the centrifugal force, so that the rotation speed of the engine is reduced. When the rotating speed of the permanent magnet speed regulator is reduced, the centrifugal wheel speed measuring device can drive a valve of the engine to be slowly opened due to the fact that centrifugal force is reduced, and the rotating speed of the engine is increased.
Further, in an embodiment, the permanent magnet speed adjusting device further includes an output power measurer 1009, and the output power measurer 1009 is respectively connected to the permanent magnet speed regulator 1004 and the automatic speed locker 1006, and is configured to obtain output power information of the engine through the permanent magnet speed regulator and send the output power information to the automatic speed locker 1006, so that the automatic speed locker 1006 can regulate and control the engine 1002 to output an optimal rotation speed while outputting a required torque.
Specifically, in combination with the previous embodiment, the output power measurer 1009 obtains the output power information of the engine in real time and sends the output power information to the automatic speed locker 1006, the automatic speed locker 1006 receives the output power information, compares the actual output power in the output power information with the preset optimal output power based on the operating characteristic curve of the engine 1002, calculates the power difference between the actual output power and the optimal output power of the engine, and then quantitatively compensates the electric quantity or the oil quantity of the engine in real time according to the power difference to ensure that the engine works at the optimal rotation speed. By further acquiring the actual output power of the engine, the engine can be further accurately controlled to be maintained at the optimum rotation speed.
In one embodiment, the permanent magnet speed regulator 1004 is coaxially connected to the engine 1002, and the permanent magnet speed regulator 1004 includes an output transmission shaft 1042 for outputting the rotational power (such as the rotational speed and the torque) regulated by the permanent magnet speed regulator 1004 to a load device, it is clear that the power output by the permanent magnet speed regulator 1004 is equal to the power output by the dc motor or the engine, and the change of the load is equal to the change of the torque output by the dc motor or the engine.
In one embodiment, the automatic speed locker 1006 includes a comparing unit, configured to obtain a characteristic parameter of the engine, compare an instant speed in the characteristic parameter with a preset optimal speed of the engine, obtain a speed difference between the instant speed and the optimal speed, and operate the engine at the optimal speed according to the speed difference. The characteristic parameters include instant rotation speed and output power.
According to the danger-escaping energy-saving power-assisting system for the aircraft engine, firstly, the rotating speed of the best efficiency of a prime mover capable of regulating speed, such as a direct current motor or an engine, is preset, then the automatic speed locker 1006 controls electric quantity or an accelerator, so that the direct current motor or the engine always operates at the preset optimal rotating speed, the output is ensured to be constant rotating speed, and the direct current motor or the engine is transformed into the prime mover capable of outputting the constant rotating speed. And then the permanent magnet speed regulator 106 is adopted to regulate the speed, and the constant rotating speed output by the direct current motor or the engine is changed to the rotating speed required by the work. At this time, the power output by the permanent magnet speed regulator 106 is equal to the power output by the direct current motor or the engine, the change of the load is equal to the change of the torque output by the direct current motor or the engine, and the rotating speed of the direct current motor or the engine is not changed when the direct current motor or the engine always works at the optimal efficiency, so that the direct current motor or the engine is ensured to always work at the optimal efficiency state, and the energy-saving effect is achieved.
Referring to fig. 11, according to the figure, it can be obtained:
4500 rpm, 21.8Nm torque, oil consumption 517g/Kwh
6500 rpm, 21.5Nm torque, 370g/Kwh fuel consumption
The torque therefore has little influence on the fuel consumption, which is mainly determined by the rotational speed.
Calculating the oil consumption difference at the same output power and two rotating speeds:
power: 4500 rpm × 21.8Nm/9550 ═ 10.3KW,
oil consumption at 4500 rpm: 10.3KW × 517g/KWh 5325g
Fuel consumption at 6500 rpm: 10.3KW × 370g/KWh 3811g
6500 r/min oil saving per hour: 5325 g-3811 g 1514g
Power: 6500 rpm x 21.5Nm/9550 is 14.6KW,
oil consumption at 4500 rpm: 14.6KW × 517g/Kwh 7548g
6500 oil consumption at rpm: 14.6KW × 370g/KWh 5417g
6500 r/min oil saving per hour: 7548 g-5417 g ═ 2131g
Please refer to table 1, which is a comparison table of the above calculation results.
TABLE 1
From the above calculation results, in one embodiment, under the condition that the engine works at the optimal speed of 6500 rpm, the energy is saved more than that of the engine at other speeds under the condition of the same output power.
All possible combinations of the technical features in the above embodiments may not be described for the sake of brevity, but should be considered as being within the scope of the present disclosure 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, for a person skilled in the art, several variations and modifications can be made without departing from the inventive concept, which falls within the scope of the present invention. Therefore, the protection scope of the present patent shall be subject to the appended claims.
Claims (9)
1. An aeroengine danger-escaping energy-saving power-assisted system comprises: the 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 are in one-to-one correspondence, the corresponding engines, the electric/power generating dual-purpose machines and the propellers are coaxially connected, the engines provide total power for the risk-escaping energy-saving power assisting system of the aircraft engine, and the 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 aeroengine escaping energy-saving power-assisted system 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 aeroengine escaping energy-saving power-assisted system 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 danger-escaping energy-saving power-assisted system of the aero-engine further comprises a coaxial oil/electricity hybrid-driven danger escaping device, wherein the coaxial oil/electricity hybrid-driven danger escaping device 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 pair of escape detection units is used for: the rotating speeds of the shafts at the two ends of the pair of backstops 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 backstops are rigidly connected, and when the pair of engines stop rotating due to faults or the rotating speed of the pair of electric/power generation dual-purpose machines is greater than that of the pair of engines, the shafts at the two ends of the pair of backstops are disconnected, so that a rotating speed difference occurs; when the rotating speed difference is detected by the pair of danger escaping detection units respectively, emergency treatment is carried out by a system controller;
the danger-escaping energy-saving power-assisted system of the aero-engine further comprises:
a pair of permanent magnet speed adjusting devices, each of which comprises a permanent magnet speed adjuster, an automatic speed locker and an auxiliary speed adjuster;
one said motor is connected with one said permanent magnet speed regulator and one said automatic speed-locking device respectively;
one said permanent magnet governor being connected to one said automatic governor and to one said system controller via one said auxiliary governor;
the automatic speed locker is used for pre-storing a characteristic parameter chart of the engine, acquiring the rotating speed of the rotating speed measurer when the automatic speed locker works, and controlling an accelerator to adjust the working state of the engine after linear compensation is carried out by comparing the characteristic parameter stored in the automatic speed locker so as to enable the engine to work at the preset optimal rotating speed;
the auxiliary speed regulator is used for pre-storing a characteristic parameter chart of the permanent magnet speed regulator, acquiring a control instruction signal when the auxiliary speed regulator works, and controlling electric quantity to adjust the rotating speed required by the output of the permanent magnet speed regulator after linear compensation is carried out on the characteristic parameter stored in the auxiliary speed regulator.
2. The aero-engine risk-escaping energy-saving boosting system according to claim 1, further comprising: and the flight controller is connected with the system controller.
3. The aircraft engine danger-escaping energy-saving power-assisted system according to claim 2, characterized in that the controlled end of the dynamic balance controller is connected with the flight controller, the first side of the dynamic balance controller is connected with the corresponding system controller to control the left rotation of the first side propeller, and the other side of the dynamic balance controller is connected with the corresponding system controller to control the right rotation of the second side propeller.
4. The aircraft engine danger-eliminating energy-saving power-assisted system according to claim 3, wherein the dynamic 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 flight controller commands rising;
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 voltage decreasing complementary mode through the pair of system controllers when the flight controller orders 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 system of claim 2, wherein the system controller is connected to the flight controller, the battery, the torque sensor, the rotational speed measurer, the electric/power generator, and the engine, respectively;
the storage battery is respectively connected with the flight controller and the electric/power generation dual-purpose machine.
6. The aircraft engine danger-escaping energy-saving power-assisted system according to claim 2, wherein the system controller 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 power detection unit is respectively connected with the torque sensor 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;
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;
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.
7. The aero-engine risk-escaping energy-saving boosting system according to claim 1, further comprising: and one of the rotating speed sensors is respectively connected with one of the permanent magnet speed regulators and one of the automatic speed lockers, and is used for acquiring rotating speed information of the engine through the permanent magnet speed regulators and feeding the rotating speed information back to the automatic speed lockers.
8. The aero-engine risk-escaping energy-saving boosting system according to claim 1, further comprising: and one output power measurer is respectively connected with one permanent magnet speed regulator and one automatic speed locker and is used for acquiring the output power information of the engine through the permanent magnet speed regulators and sending the output power information to the automatic speed locker, so that the automatic speed locker can regulate and control the optimal output rotating speed of the engine and simultaneously output the required torque.
9. The aero-engine danger-escaping energy-saving power-assisted system according to claim 1, wherein the permanent magnet speed governor is coaxially connected with the engine, and the permanent magnet speed governor comprises an output transmission shaft for outputting the rotating power after the speed of the permanent magnet speed governor is regulated to a load device.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811012821.1A CN109383782B (en) | 2018-08-31 | 2018-08-31 | Danger-escaping energy-saving power-assisting system of aero-engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811012821.1A CN109383782B (en) | 2018-08-31 | 2018-08-31 | Danger-escaping energy-saving power-assisting system of aero-engine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN109383782A CN109383782A (en) | 2019-02-26 |
| CN109383782B true CN109383782B (en) | 2022-07-26 |
Family
ID=65418535
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201811012821.1A Active CN109383782B (en) | 2018-08-31 | 2018-08-31 | Danger-escaping energy-saving power-assisting system of aero-engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN109383782B (en) |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101056777A (en) * | 2004-11-16 | 2007-10-17 | 大众汽车股份公司 | Method for controlling the operation of a hybrid motor vehicle and hybrid vehicle |
| CN102232268A (en) * | 2011-06-27 | 2011-11-02 | 华为技术有限公司 | Method, device for realizing energy saving control of generating system, and generating system |
| US9212625B2 (en) * | 2010-11-19 | 2015-12-15 | Rudolph Allen SHELLEY | Hybrid gas turbine propulsion system |
| CN105329448A (en) * | 2015-11-24 | 2016-02-17 | 中国航空工业集团公司沈阳飞机设计研究所 | Oil-electricity mixed power system of vertical taking-off and landing unmanned aerial vehicle |
| CN105752344A (en) * | 2016-03-15 | 2016-07-13 | 电子科技大学 | Plug-in hybrid power driving device for tilt-rotor aircraft |
| CN107878762A (en) * | 2017-11-28 | 2018-04-06 | 北京正兴弘业科技有限公司 | A kind of long endurance unmanned aircraft oil electric mixed dynamic system and control method |
Family Cites Families (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2997382B1 (en) * | 2012-10-29 | 2014-11-21 | Eurocopter France | METHOD FOR MANAGING AN ENGINE FAILURE ON A MULTI-ENGINE AIRCRAFT PROVIDED WITH A HYBRID POWER PLANT |
| US9193451B2 (en) * | 2013-04-22 | 2015-11-24 | Ival O. Salyer | Aircraft using turbo-electric hybrid propulsion system for multi-mode operation |
| WO2016049027A1 (en) * | 2014-09-23 | 2016-03-31 | Sikorsky Aircraft Corporation | Hybrid electric power drive system for a rotorcraft |
| CN104670504B (en) * | 2015-02-13 | 2016-06-29 | 吉林大学 | Oil/optical electrical multi power source Fixed Wing AirVehicle |
| CN104760696B (en) * | 2015-04-22 | 2016-07-27 | 深圳市艾特航空科技股份有限公司 | A kind of multi-rotor aerocraft |
| CN207631503U (en) * | 2017-11-28 | 2018-07-20 | 北京正兴弘业科技有限公司 | A kind of long endurance unmanned aircraft oil electric mixed dynamic system |
| CN108263618A (en) * | 2017-12-22 | 2018-07-10 | 成都才智圣有科技有限责任公司 | A kind of hybrid power multiaxis rotor wing unmanned aerial vehicle |
| CN108082500A (en) * | 2018-01-29 | 2018-05-29 | 吉林大学 | A kind of fixed-wing formula hybrid power aircraft driving device and driving method |
-
2018
- 2018-08-31 CN CN201811012821.1A patent/CN109383782B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101056777A (en) * | 2004-11-16 | 2007-10-17 | 大众汽车股份公司 | Method for controlling the operation of a hybrid motor vehicle and hybrid vehicle |
| US9212625B2 (en) * | 2010-11-19 | 2015-12-15 | Rudolph Allen SHELLEY | Hybrid gas turbine propulsion system |
| CN102232268A (en) * | 2011-06-27 | 2011-11-02 | 华为技术有限公司 | Method, device for realizing energy saving control of generating system, and generating system |
| CN105329448A (en) * | 2015-11-24 | 2016-02-17 | 中国航空工业集团公司沈阳飞机设计研究所 | Oil-electricity mixed power system of vertical taking-off and landing unmanned aerial vehicle |
| CN105752344A (en) * | 2016-03-15 | 2016-07-13 | 电子科技大学 | Plug-in hybrid power driving device for tilt-rotor aircraft |
| CN107878762A (en) * | 2017-11-28 | 2018-04-06 | 北京正兴弘业科技有限公司 | A kind of long endurance unmanned aircraft oil electric mixed dynamic system and control method |
Also Published As
| Publication number | Publication date |
|---|---|
| CN109383782A (en) | 2019-02-26 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US20220194584A1 (en) | Method and apparatus for lifting a payload | |
| US10046853B2 (en) | Hybrid gyrodyne aircraft employing a managed autorotation flight control system | |
| CN107878762B (en) | Long-endurance unmanned aerial vehicle oil-electricity hybrid power system and control method | |
| US20190263519A1 (en) | Hybrid aircraft | |
| US8282044B2 (en) | Actuator device for varying the attitude of a spacecraft | |
| JP2019501057A (en) | Micro hybrid generator system drone | |
| CN109383787B (en) | Power assisting system of aero-engine | |
| US10577095B2 (en) | Active rotor damping system | |
| CN109383783B (en) | Energy-saving power-assisting system of aero-engine | |
| CN109383785B (en) | Aerial platform with energy-saving power assisting system | |
| CN109335003B (en) | Power assisting device for aircraft engine | |
| CN109383782B (en) | Danger-escaping energy-saving power-assisting system of aero-engine | |
| CN109383784B (en) | Aerial platform with power assisting system | |
| CN109573061B (en) | Aerial platform with escape energy-saving power-assisted system | |
| CN109278996B (en) | Danger-escaping energy-saving power assisting device for aero-engine | |
| CN109204797B (en) | Energy-saving power assisting device of aero-engine | |
| JP7094232B2 (en) | Multicopter | |
| CN109398690B (en) | Emergency escape power assisting system of aero-engine | |
| CN109455294B (en) | Aerial platform with escape assisting system | |
| CN109383786B (en) | Emergency escape power assisting system assembly of aero-engine | |
| US20230249839A1 (en) | Control device for flight vehicle | |
| JP7439274B2 (en) | Hybrid drive system for helicopters | |
| JP2022167542A (en) | multicopter | |
| CN114172421A (en) | Unmanned aerial vehicle deicing and power supply control method, device and system and storage medium | |
| JP2023092169A (en) | Flight vehicle controller |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant | ||
| TR01 | Transfer of patent right | ||
| TR01 | Transfer of patent right |
Effective date of registration: 20230911 Address after: 801, No.19-1 Huahai Road, Shenyang Economic and Technological Development Zone, Shenyang, Liaoning 110000 Patentee after: Shenyang Qiande Electronic Technology Co.,Ltd. Address before: No. 19-13, Canghai Road, economic and Technological Development Zone, Tiexi District, Shenyang, Liaoning 110021 Patentee before: LIAONING TONGXINYUAN TECHNOLOGY Co.,Ltd. |