CN112429249A - Aircraft with a flight control device - Google Patents
Aircraft with a flight control device Download PDFInfo
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- CN112429249A CN112429249A CN202011399424.1A CN202011399424A CN112429249A CN 112429249 A CN112429249 A CN 112429249A CN 202011399424 A CN202011399424 A CN 202011399424A CN 112429249 A CN112429249 A CN 112429249A
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- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 title description 2
- 230000007246 mechanism Effects 0.000 claims abstract description 129
- 230000005540 biological transmission Effects 0.000 claims abstract description 47
- 150000001875 compounds Chemical class 0.000 claims description 30
- 230000001141 propulsive effect Effects 0.000 claims description 4
- 239000000446 fuel Substances 0.000 abstract description 6
- 230000008859 change Effects 0.000 abstract description 4
- 238000000034 method Methods 0.000 abstract description 4
- 230000008569 process Effects 0.000 abstract description 3
- 238000010586 diagram Methods 0.000 description 11
- 230000007704 transition Effects 0.000 description 7
- 125000004122 cyclic group Chemical group 0.000 description 5
- 238000006243 chemical reaction Methods 0.000 description 4
- 238000005265 energy consumption Methods 0.000 description 2
- 239000000126 substance Substances 0.000 description 2
- 208000032953 Device battery issue Diseases 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 238000013329 compounding Methods 0.000 description 1
- 238000011217 control strategy Methods 0.000 description 1
- 230000005611 electricity Effects 0.000 description 1
- 238000004134 energy conservation Methods 0.000 description 1
- 230000009347 mechanical transmission Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
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Classifications
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- 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
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- 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
- B64D35/00—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions
- B64D35/08—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
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- 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
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Mechanical Engineering (AREA)
- Hybrid Electric Vehicles (AREA)
Abstract
The invention relates to a vehicle, and provides an aircraft, wherein the aircraft comprises a fuselage, a main rotor arranged on the fuselage, wings, a propulsion propeller and a driving system, the driving system comprises an engine, a generator, a power distribution mechanism and an electric propulsion mechanism, the electric propulsion mechanism is in transmission connection with the propulsion propeller, the aircraft can operate in a state that the driving system drives the main rotor and the propulsion propeller simultaneously, and the aircraft can operate in a state that the driving system drives only the propulsion propeller. Compared with a single power source in a traditional fuel power system, the aircraft provided by the invention can flexibly adjust the output of the system, optimize the fuel consumption, improve the flight efficiency and reduce the emission and noise pollution in the process of load demand power change.
Description
Technical Field
The invention relates to a vehicle, in particular to an aircraft.
Background
With the continuous development of world economy, ground traffic jam caused by the appearance of super-large-scale cities and satellite city groups is urgently needed to be solved. Besides the development of ground traffic, intelligent stereo traffic is another important way to solve future trips. The urban aircraft can make full use of a low-altitude airspace, a new quick trip mode is provided on the basis of the existing traffic system, and trip efficiency is improved. Based on the application scene of the urban aircraft, the requirements of vertical take-off and landing, efficient cruising, safety, low noise, economic cost, convenience for ground storage and the like are required. Unlike conventional aircraft, urban aircraft place new demands on their power systems.
But combined type rotor craft of VTOL can use as an urban aircraft product. A vertical take-off and landing combined type rotor craft is provided with a main rotor system, wings and a propulsion system arranged along the wings. Wherein, compounding refers to the compound application of main rotor system and wing, provides lift for the aircraft jointly. The main rotor system can operate in two modes, powered to provide lift to the aircraft, and unpowered to provide no lift. When the main rotor is driven by power to generate enough lift force, the vertical take-off and landing of the aircraft can be realized. The propulsion system arranged along the wing is provided with a large-stroke pitch-changing mechanism, can provide forward or backward thrust, and can resist the reaction torque acting on the aircraft body when the main rotor rotates to generate lift force. The propulsion system is also capable of generating propulsion in the forward flight direction of the aircraft, supporting forward flight. When the aircraft climbs and accelerates for a period of time after takeoff, the driving force of the main rotor system can be disconnected, and the aircraft can efficiently cruise in a fixed wing mode. After takeoff, if the main rotor drive system fails, the aircraft can fly to a proper take-off and landing point in a self-rotation rotor mode or a composite self-rotation rotor mode, a fixed wing mode and the like. But combined type rotor craft of VTOL can satisfy the requirement of city flight.
The power system of the traditional aircraft can not meet the urban flight requirements of energy conservation, noise reduction, low emission and low cost, has larger total power for the power system, is generally realized by a turboshaft engine, and has the characteristics of large throttle amount, large noise, uneconomic performance and the like.
Disclosure of Invention
In view of the above, the present invention is directed to an aircraft, so as to solve the problem of large energy consumption of a driving system of the aircraft.
In order to achieve the purpose, the technical scheme of the invention is realized as follows:
an aircraft, wherein the aircraft includes a fuselage, a main rotor disposed on the fuselage, wings, propulsion propellers disposed on the wings, and a drive system, the drive system includes an engine, a generator, a power distribution mechanism, and an electric propulsion mechanism, the engine is drivingly connectable to the generator and the main rotor, the power distribution mechanism is electrically connected to the generator and the electric propulsion mechanism, the electric propulsion mechanism is drivingly connected to the propulsion propellers, the aircraft is capable of operating in a state in which the drive system drives the main rotor and the propulsion propellers simultaneously, and the aircraft is capable of operating in a state in which the drive system drives only the propulsion propellers.
Further, in a state where the driving system drives the main rotor and the propulsion propeller simultaneously, the engine transmits mechanical energy to the main rotor and the generator, respectively, the generator transmits electrical energy to the distribution mechanism, the distribution mechanism transmits electrical energy to the electric propulsion mechanism, and the electric propulsion mechanism transmits mechanical energy to the propulsion propeller;
under the state that the driving system only drives the propulsion propeller, the engine only transmits mechanical energy to the generator, the generator transmits electric energy to the power distribution mechanism, the power distribution mechanism transmits electric energy to the electric propulsion mechanism, and the electric propulsion mechanism transmits mechanical energy to the propulsion propeller.
Further, the aircraft may be operable in a first helicopter mode in which the drive system simultaneously drives the main rotor and the propulsion propeller, wherein the main rotor is capable of providing lift and the propulsion propeller is capable of balancing the torque applied to the fuselage by the main rotor.
Further, the aircraft is operable in a second helicopter mode in which the drive system simultaneously drives the main rotor and the propulsion propeller, wherein the main rotor is capable of providing lift and forward flight power, and the propulsion propeller is capable of balancing the torque applied to the fuselage by the main rotor.
Further, the aircraft is operable in a compound helicopter mode in which the drive system simultaneously drives the main rotor and the propulsion propellers that provide forward flight power and balance the torque applied to the fuselage by the main rotor, the main rotor and the wings providing lift.
Further, the aircraft may be operable in a gyroplane state in which the drive system drives only the propulsion propellers, wherein the main rotor provides lift and the propulsion propellers provide forward flight power.
Further, the aircraft is operable in a compound autogiro state in which the drive system drives only the propulsion propellers, the main rotor and the wings together providing lift, and the propulsion propellers providing forward flight power.
Further, the aircraft may be operated in a fixed wing state with the drive system driving only the propulsion propeller, wherein the wings provide lift and the propulsion propeller provides forward flight power.
Further, the drive system further includes a transmission drivingly connectable to the main rotor, the engine, and the generator.
Further, the power distribution mechanism includes a power distributor and a buffer battery electrically connected to each other, the power distributor being electrically connected to the generator and the electric propulsion mechanism, the electric propulsion mechanism including a motor driver and an electric motor.
Compared with the prior art, the aircraft has the following advantages:
compared with a single power source in a traditional fuel power system, the aircraft provided by the invention can flexibly adjust the output of the system, optimize the fuel consumption, improve the flight efficiency and reduce the emission and noise pollution in the process of load demand power change.
Additional features and advantages of the invention will be set forth in the detailed description which follows.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate an embodiment of the invention and, together with the description, serve to explain the invention and not to limit the invention. In the drawings:
FIG. 1 is a schematic structural diagram of an aircraft in accordance with an embodiment of the present invention;
FIG. 2 is a schematic diagram of a drive system according to an embodiment of the present invention;
FIG. 3 is an energy flow diagram of the drive system according to an embodiment of the present invention, wherein the main rotor and the propulsor are driven simultaneously, the aircraft operating in a helicopter mode or a compound helicopter mode;
FIG. 4 is a power flow diagram of the drive system according to an embodiment of the present invention in which only the propulsion propellers are driven and the aircraft is operating in either the autogiro state or the compound autogiro state or the fixed-wing state;
FIG. 5 is a power flow diagram of the drive system according to an embodiment of the present invention, wherein one of the motors has failed;
FIG. 6 is a power flow diagram of the drive system according to an embodiment of the present invention, wherein one of the motors has failed;
FIG. 7 is a power flow diagram of the drive system according to an embodiment of the present invention, wherein one of the generators fails;
FIG. 8 is an energy flow diagram of the drive system according to an embodiment of the present invention, wherein one of the generators fails;
FIG. 9 is an energy flow diagram of the drive system according to an embodiment of the present invention, wherein one of the generators fails;
FIG. 10 is a power flow diagram of the drive system according to an embodiment of the present invention, wherein one electric propulsion mechanism is malfunctioning;
FIG. 11 is a power flow diagram of the drive system according to the embodiment of the present invention, wherein one electric propulsion mechanism is malfunctioning;
description of reference numerals:
10-main rotor, 11-transmission, 12-first engine, 13-second engine, 14-first generator, 15-second generator, 16-distribution mechanism, 17-electric propulsion mechanism, 18-propulsion propeller, 20-fuselage, 30-wing, 141-first generator body, 142-first motor controller, 151-second generator body, 152-second motor controller, 161-distributor, 162-buffer battery, 171-first motor driver, 172-second motor driver, 173-third motor driver, 174-fourth motor driver, 175-first motor, 176-second motor, 177-third motor, 178-fourth motor.
Detailed Description
In addition, the embodiments of the present invention and the features of the embodiments may be combined with each other without conflict.
The present invention will be described in detail below with reference to the accompanying drawings in conjunction with embodiments.
The invention provides a driving system of an aircraft, wherein the aircraft comprises a fuselage 20, a main rotor 10 arranged on the fuselage 20, wings 30, propeller propellers 18 arranged on the wings 30 and a driving system, the driving system comprises an engine, a generator, a transmission 11, a power distribution mechanism 16 and an electric propulsion mechanism 17, the engine is in transmission connection with the generator and the main rotor 10, the power distribution mechanism 16 is in electrical connection with the generator and the electric propulsion mechanism 17, and the electric propulsion mechanism 17 is in transmission connection with the propeller propellers 18.
The aircraft comprises a fuselage 20, wings 30 are connected to two sides of the fuselage 20, a rotatable main rotor wing 10 is arranged on the top of the fuselage 20, a propulsion propeller 18 is arranged on the wings 30, and the central axis of the propulsion propeller 18 extends approximately in the front-rear direction. Wherein the main rotor 10 can provide lift and forward flight power, the wings 30 can also provide lift, and the propulsor propeller 18 can provide forward flight power.
The driving system of the aircraft comprises an engine, a generator, a transmission 11, a power distribution mechanism 16 and an electric propulsion mechanism 17, wherein the engine can convert chemical energy of chemical fuel into mechanical energy, the engine can transmit the mechanical energy to a main rotor wing 10 (driving the main rotor wing to rotate) and the generator through the transmission 11, the generator can convert the mechanical energy into electric energy and transmit the electric energy to the power distribution mechanism 16, the power distribution mechanism 16 can directly transmit the electric energy to the electric propulsion mechanism 17, and the electric propulsion mechanism 17 converts the electric energy into the mechanical energy to drive a propulsion screw to rotate 18.
In addition, the drive system comprises a transmission 11 which can be drivingly connected to the main rotor 10, the engine and the generator. The transmission 11 is provided with a clutch that can selectively connect and disconnect the engine to and from the main rotor 10 and the engine to and from the generator to change the power transmission path. And the transmission 11 can change the transmission ratio between the engine and the main rotor 10 and between the engine and the generator.
Specifically, the power distribution mechanism 16 includes a power distributor 161 and a buffer battery 162 electrically connected to each other, and the power distributor 161 is electrically connected to the generator and the electric propulsion mechanism 17. Distributor 161 may distribute the electrical energy generated by the generator, wherein a portion of the electrical energy may be directly transferred to electric propulsion means 17, another portion of the electrical energy may be stored in buffer battery 162, and in some cases distributor 161 may transfer the electrical energy of buffer battery 162 to electric propulsion means 17 and the generator.
Wherein the electric propulsion means 17 comprise a motor drive and an electric motor. The motor drive may convert the direct current to alternating current and provide the alternating current to the electric motor, and the motor drive may control the rotational speed of the electric motor, and thus the rotational speed of the propulsion propeller 18.
Specifically, the driving system comprises a first engine 12, a second engine 13, a first generator 14 and a second generator 15 which can be in transmission connection with the main rotor 10, wherein the first engine 12 can be in transmission connection with the first generator 14, and the second engine 13 can be in transmission connection with the second generator 15. The drive system comprises two engines, namely a first engine 12 and a second engine 13, and two corresponding generators, namely a first generator 14 and a second generator 15, i.e. the drive system comprises two sets of hybrid power units, and when one of the sets fails, the power output can be maintained through the other set, ensuring the operation of the main rotor 10 and the propeller 18.
Wherein the aircraft comprises two of said wings 30, each of said wings 30 being provided with two of said propeller propellers 18. The two wings 30 are symmetrically arranged, and the propeller rotors 18 on the two wings 30 are also symmetrically arranged.
Wherein the driving system comprises four electric propulsion mechanisms 17 corresponding to the propulsion propellers 18 one by one, and the four electric propulsion mechanisms 17 operate independently of each other. The number of the electric propulsion mechanisms 17 may be four, and the electric propulsion mechanisms include motor drivers and electric motors, that is, a first motor driver 171 and a first electric motor 175, a second motor driver 172 and a second electric motor 176, a third motor driver 173 and a third electric motor 177, and a fourth motor driver 174 and a fourth electric motor 178, and may drive the four propulsion propellers 18 in a one-to-one correspondence.
Wherein the generator comprises a generator body and a motor controller, the motor controller being capable of converting alternating current and direct current. The first generator 14 includes a first generator body 141 and a first motor controller 142, and the second generator 15 includes a second generator body 151 and a second motor controller 152. The generator body can convert mechanical energy into alternating current electric energy, the motor controller can convert alternating current into direct current to be transmitted to the power distribution mechanism 16, and direct current from the power distribution mechanism 16 can be converted into alternating current to be transmitted to the generator body so as to drive the generator body to operate, namely, the electric energy is converted into mechanical energy.
Wherein the generator is configured to receive electrical energy from the power distribution mechanism 16 and provide starting power to the engine. During the start-up phase of the drive system, the buffer battery 162 may provide electric energy, which is transmitted to the motor controller through the distributor 161 to convert the dc power into ac power to operate the generator body, and the engine body may be drivingly connected to the engine to drive the engine to start operation, and then the engine drives the generator to operate.
Wherein the electric propulsion mechanism 17 is arranged on the wing 30, and the engine, the generator and the power distribution mechanism 16 are arranged on the fuselage 20. In the scheme, the engine is a power component, and after mechanical energy output by the engine is converted into electric energy through the generator, the electric energy can be transmitted to the electric propulsion mechanism 17 on the wing 30 through the power distribution mechanism 16 and the lead, so that the mechanical transmission component is prevented from being arranged on the wing 30, and the weight and the complexity of the wing 30 can be reduced.
During operation, the aircraft can be operated with the drive system driving both the main rotor 10 and the propulsion propeller 18, and with the drive system driving only the propulsion propeller 18.
Energy flow
1. Simultaneous drive of the main rotor and propulsion propeller
Wherein, in a state that the driving system drives the main rotor 10 and the propeller 18 simultaneously, the engine transmits mechanical energy to the main rotor 10 and the generator, respectively, the generator transmits electrical energy to the distribution mechanism 16, the distribution mechanism 16 transmits electrical energy to the electric propulsion mechanism 17, and the electric propulsion mechanism 17 transmits mechanical energy to the propeller 18.
Specifically, the buffer battery 162 transmits electric energy to the first motor controller 142 and the second motor controller 152 through the distributor 161, converts direct current into alternating current, and drives the first generator body 141 and the second generator body 151 to operate, so as to transmit mechanical energy to the first engine 12 and the second engine 13, thereby realizing starting of the first engine 12 and the second engine 13; after the first engine 12 and the second engine 13 are started, a part of mechanical energy is transmitted to the main rotor 10 through the transmission 11 to drive the main rotor 10 to rotate, and the other part of mechanical energy is transmitted to the first generator 14 and the second generator 15 to generate alternating current, the alternating current is converted into direct current and then transmitted to the distributor 161, electric energy is transmitted to the electric propulsion mechanism 17 (part of electric energy can be transmitted to the buffer battery 162), and the motor is controlled by the motor driver to operate, so that the propulsion propeller 18 is driven to rotate.
Wherein the aircraft is operable in a first helicopter state wherein the main rotor 10 is capable of providing lift and the propulsor propellers 18 are capable of balancing the torque applied to the fuselage 20 by the main rotor 10. In this state, the horizontal speed of the aircraft is substantially 0, and mainly the lift provided by the main rotor 10 is increased, while the rotating speed of the propeller 18 or the pitch of the propeller 18 on the wings 30 on both sides is different, and the acting force of the wings 30 on the fuselage 20 is different, so that the torque applied to the fuselage 20 by the rotation of the main rotor 10 can be balanced.
The aircraft is operable in a second helicopter mode wherein the main rotor 10 is capable of providing lift and forward flight power and the propulsor propellers 18 are capable of balancing the torque applied to the fuselage 20 by the main rotor 10. By varying the disk angle of attack of the main rotor 10, the main rotor 10 can generate forward flight power while providing lift, and the propulsion propellers 18 still provide a balancing torque.
The aircraft is operable in a compound helicopter mode in which the main rotor 10 is only capable of providing lift, the propulsor rotor 18 is only capable of providing forward flight power and is capable of balancing the torque applied to the fuselage 20 by the main rotor 10, and the main rotor 10 and the wings 11 are capable of providing lift. At this time, the horizontal speed of the aircraft is increased, the wings 30 can provide lift force, the propulsion propeller 18 can provide forward flying power, and the required driving energy is increased; for the main rotor 10, the input energy can be reduced and the drive system enters a more economical energy consumption stage.
2. Driving only the propulsion propeller
In a state where the drive system drives only the propulsion propeller 18, the engine transmits only mechanical energy to the generator, the generator transmits electric energy to the distribution mechanism 16, the distribution mechanism 16 transmits electric energy to the electric propulsion mechanism 17, and the electric propulsion mechanism 17 transmits mechanical energy to the propulsion propeller 18. When the engine is no longer providing mechanical energy to main rotor 10, a state is entered in which only propeller 18 is driven. The transmission connection between the first engine 12 and the main rotor 10 and the transmission connection between the second engine 13 and the main rotor 10 are disconnected through the transmission 11, the mechanical energy is completely transmitted to the first generator 14 and the second generator 15 through the first engine 12 and the second engine 13 respectively, the mechanical energy is converted into alternating current electric energy through the generator body, the alternating current electric energy is converted into direct current electric energy through the motor controller and is transmitted to the distributor 161, the distributor 161 transmits the electric energy to the electric propulsion mechanism 17 (part of the electric energy can be transmitted to the buffer battery 162), and the operation of the electric motor is controlled through the motor driver, so that the propulsion propeller 18 is driven to rotate.
Wherein the aircraft is operable in a compound autogyro state. Wherein the main rotor 10 and the wings 30 together provide lift and the propulsion propellers 18 provide forward flight power. The main rotor (10) is air driven to spin providing lift, the wings (30) providing more lift due to the faster flight speed, and the propulsion propellers (18) providing forward flight power. The state of the compound autogyro is a transition state after the power transmission to the main rotor 10 is stopped.
Wherein the aircraft is operable in a rotorcraft state. Wherein the main rotor (10) is driven by air to idle and provide lift, the wings (30) do not provide lift, and the propulsion propellers (18) provide forward flight power.
In addition, the aircraft may operate in a fixed wing state, wherein the wings 30 provide lift and the propulsor propellers 18 provide forward flight power. The main rotor 10 is driven to idle by air, and basically provides no lift and no forward flying power. Substantially all of the energy is delivered to the propulsion propeller 18 and the aircraft flies at a higher horizontal speed.
Fault handling
1. Simultaneous drive of the main rotor and propulsion propeller
The aircraft can be operated in a state where the drive system simultaneously drives the main rotor 10 and the propulsion propeller 18, and the aircraft can be operated in the event of a failure of one of the engine, the generator, the electric propulsion mechanism 17, and the buffer battery 162. In this case, the aircraft operates in a first helicopter state, a second helicopter state or a compound helicopter state, or transitions to an autogyro state, a compound autogyro state, a fixed wing state.
As shown in fig. 5, when the first engine 12 fails, the aircraft can operate in a first helicopter state, a second helicopter state or a compound helicopter state, wherein the transmission connection between the first engine 12 and the main rotor 10 and the first generator 14 is disconnected, the power sources of the drive system are the second engine 13 and the buffer battery 162, and the main rotor 10 and the propulsion propeller 18 are driven simultaneously. The clutch of the transmission 11 connected to the first engine 12 is disconnected, the first engine 14 in driving connection with the first engine 12 is stopped, the power source of the aircraft is changed into the buffer battery 162 and the second engine 13, the high-power load end is mainly the main rotor 10 in the stage, the propulsion propeller 18 only needs small power to balance the reaction torque of the main rotor 10 on the fuselage 20, and the power of the second engine 13 can maintain the aircraft to be operated to a proper lifting point and landed in the state of the first helicopter or the second helicopter.
As shown in fig. 6, when the first engine 12 fails, the aircraft may be switched to an autogiro state, a compound autogiro state, and a fixed-wing state, wherein the transmission connection between the first engine 12 and the main rotor 10 and the first generator 14 is disconnected, the transmission connection between the second engine 13 and the second generator 15 and the main rotor 10 is disconnected, the power sources of the drive system are the second engine 13 and the buffer battery 162, and only the propulsion propeller 18 is driven. The aircraft may be transitioned from the current operating state to the autogyro state operation, depending on the operational needs. When the horizontal speed of the aircraft is relatively high, such as in a compound helicopter state, it may be converted to a compound autogyroplane state and further to a fixed wing state, i.e., the power to the main rotor 10 is stopped to provide forward flight power through the propulsion propellers 18 and lift through the wings 30.
When the buffer battery 162 fails, if the flying height of the aircraft is 0, the aircraft is restarted after the failure is eliminated; if the aircraft is in flight, the electrical connection between the power distributor 161 and the buffer battery 162 is broken and no power is supplied to the buffer battery 162.
As shown in fig. 7, when the first generator 14 fails, the aircraft may operate in a first helicopter state, a second helicopter state or a compound helicopter state, wherein the transmission connection between the first generator 12 and the first generator 14 is disconnected, the electrical connection between the first generator 14 and the power distribution mechanism 16 is disconnected, the first generator 12 and the second generator 13 drive the main rotor 10, and the second generator 15 and the buffer battery 162 drive the electric propulsion mechanism 17. At this time, the aircraft maintains the first helicopter state, the torque signal output by the flight controller to the first motor controller 142 is adjusted to minimize the output power of the first generator 14, the aircraft controller outputs a control command to disconnect the electrical connection between the distributor 161 and the first generator 14, disconnect the clutch connecting the first generator 14 and the transmission 11 (so that the engine no longer provides input power to the first generator 14), the propulsion power is provided by the second generator 15 and the buffer battery 162, and the main rotor system 10 is driven by the engine.
As shown in fig. 8, when the first generator 14 fails, the aircraft may operate in a first helicopter state, a second helicopter state or a hybrid helicopter state, wherein the transmission connection between the first engine 12 and the first generator 14 is disconnected, the electrical connection between the first generator 14 and the power distribution mechanism 16 is disconnected, the transmission connection between the second engine 13 and the main rotor 10 is disconnected, the first engine 12 drives the main rotor 10, and the second generator 15 and the buffer battery 162 drive the electric propulsion mechanism 17. At this point, the aircraft may transition to a second helicopter state, wherein the control of the aircraft controller is adjusted to: the torque signal output by the flight controller to the first motor controller 142 is adjusted to minimize the output power of the first generator 14, the aircraft controller outputs a control command to disconnect the electrical connection between the distributor 161 and the first generator 14, disconnect the clutch connecting the first generator 14 to the transmission 11 (so that the first generator 12 no longer provides input power to the first generator 14), the first generator 12 provides the full driving force of the main rotor 10, disconnect the clutch connecting the second engine 13 and the main rotor 10 included in the transmission 11, and the second engine 13 only drives the second generator 15 on the local side to generate power for driving the propeller 18 to rotate, so as to provide the reactive torque and the forward flight power for balancing the main rotor 10 to the fuselage 20.
As shown in fig. 9, when the first generator 14 fails, the aircraft may be switched to an autogiro state, a compound autogiro state, and a fixed wing state, wherein the transmission connection between the first generator 12 and the main rotor 10 and the first generator 14 is disconnected, the electrical connection between the first generator 14 and the power distribution mechanism 16 is disconnected, the transmission connection between the second generator 13 and the main rotor 10 is disconnected, and the second generator 15 and the buffer battery 162 drive the electric propulsion mechanism 17. At this point, the aircraft may transition to a compound autogyro state, where the control strategy of the aircraft controller is adjusted to: the aircraft controller issues a control command to adjust a torque signal output by the flight controller to the first motor controller 142 so as to minimize the output power of the first generator 14, and the aircraft controller outputs a control command to disconnect the electrical connection between the distributor 161 and the first generator 14, disconnect the clutch connecting the first generator 14 and the transmission 11 (so that the first generator 12 does not provide input power for the first generator 14 any more), and stop the output of the first generator 12. The clutch contained in the transmission 11 and connecting the second engine 13 with the main rotor 10 is disconnected, and the second engine 13 only drives the second generator 15 on the side to generate electricity for driving the propulsion propeller 18 to rotate so as to provide forward flight power for the aircraft.
Wherein the wings 30 are provided with two sets of symmetrically arranged propulsion propellers 18, the drive system comprising a first and a fourth electric propulsion mechanism for driving a first set of symmetrically arranged propulsion propellers 18 and a second and a third electric propulsion mechanism for driving a second set of symmetrically arranged propulsion propellers 18. The two-sided wing 30 is a substantially symmetrical structure including an electric propulsion mechanism and a propulsive propeller 18 disposed thereon.
Wherein, as shown in fig. 10, when the first electric propulsion mechanism fails, the aircraft may maintain a first helicopter state, a second helicopter state, or a compound helicopter state operation wherein the electrical distribution mechanism 16 is electrically disconnected from the first electric propulsion mechanism and the fourth electric propulsion mechanism. At this time, the aircraft may be switched to a first helicopter state or a second helicopter state, in which the rotation speed command issued by the flight controller to the first motor driver 171 becomes 0, the rotation speed of the fourth motor driver 174 symmetrically arranged with respect to the first motor driver 171 is reduced to 0, the electrical connection between the power distribution mechanism 16 and the first motor driver 171 and the electrical connection between the power distribution mechanism 174 are disconnected, and the reaction torque of the dynamic main rotor 10 acting on the fuselage 20 is balanced by the second electrical propulsion mechanism and the third electrical propulsion mechanism.
Alternatively, as shown in fig. 11, when the first electric propulsion mechanism 17 fails, the aircraft may be switched to autogiro state, compound autogiro state, fixed-wing state operation, wherein the electrical distribution mechanism 16 is disconnected from the first electric propulsion mechanism 17 and the fourth electric propulsion mechanism 17, and the transmission connection of the first engine 12 and the second engine 13 to the main rotor 10 is disconnected. The rotation speed command issued by the flight controller to the first motor driver 171 is minimized, the rotation speed of the fourth motor driver 174 symmetrically arranged with the first motor driver 171 is minimized, the electrical connection between the first motor driver 171 and the fourth motor driver 174 of the power distribution mechanism 16 is disconnected, and only two second electric propulsion mechanisms and two third electric propulsion mechanisms symmetrically distributed on the wing 30 are reserved. According to flight requirements, the driving system also supports the aircraft to be switched to a self-rotation gyroplane/fixed wing state to work, and an aircraft controller needs to issue the following commands: the propulsion of the electric propulsion mechanism is increased, the cyclic pitch system of the main rotor 10 is controlled, the driving force required by the main rotor 10 is gradually reduced, the clutch in which the first engine 12 and the second engine 13 are connected with the main rotor 10 is disconnected through the transmission 11, and the propulsion power of the aircraft is mainly provided by the second motor 176 and the third motor 177 which are remained in the electric propulsion mechanism.
2. Driving only the propulsion propeller
The aircraft can be operated in a state where the drive system drives only the propulsion propeller 18, and the aircraft can be operated when one of the engine, the generator, the electric propulsion mechanism 17, and the buffer battery 162 fails. The aircraft is capable of operating in a rotorcraft, compound rotorcraft state, fixed wing state, or transitioning to a first helicopter state, a second helicopter state, or a compound helicopter state.
Wherein the aircraft can transition to the compound helicopter/helicopter state when said first engine 12 fails, as shown in figure 5. The power of the second engine 13 is preferably as follows: failure of the first engine 12 can support transition of the aircraft from the compound autogiro or fixed wing state to flight and landing in the first or second helicopter state. Disconnecting the transmission connection between the first engine 12 and the first generator 14, disconnecting the electrical connection between the power distribution mechanism 16 and the first generator 14, and powering the second engine 13 and the buffer battery 162, wherein the propulsion propeller 18 and the main rotor 10 are both driven. Wherein, aircraft controller control does: the given torque value sent to the first generator 14 is reduced to the minimum, the electric connection between the distribution mechanism 16 and the first generator 14 is disconnected, the clutch in the transmission 11, which connects the first generator 12 with the first generator 14, is disconnected, the first generator 12 stops working, the clutch in the transmission 11, which connects the first generator 12 with the main rotor 10, is switched to the closed state, and the power source of the aircraft becomes the buffer battery 162 and the second engine 13.
As shown in fig. 6, when the first engine 12 fails, the aircraft may operate in a combined autogiro/fixed wing state in which the drive connection between the first engine 12 and the first generator 14 is disconnected, the electrical connection between the power distribution mechanism 16 and the first generator 14 is disconnected, the drive connection between the second engine 13 and the generator 15 is maintained, the drive system has power sources of the second engine 13 and the buffer battery 162, and the propulsion propeller 18 is driven. Specifically, the aircraft controller controls as follows: the given value of the torque sent to the first generator 14 is reduced to the minimum, the electric connection between the power distribution mechanism 16 and the first generator 14 is disconnected, the power source of the aircraft is changed into the buffer battery 162 and the second engine 13, and the throttle amount of the first engine 12 and the second engine 13 of the aircraft, the output torque of the second generator 15, the rotating speed of the motor of the electric propulsion mechanism 17, the pitch system of the propulsion propeller 18 and the like need to be controlled simultaneously in the conversion process.
As shown in fig. 7, when the first generator 14 fails, the aircraft may transition to a compound helicopter/helicopter state in which the first engine 12 is disconnected from the first generator 14, and the first generator 14 is disconnected from the power distribution mechanism 16, so that the first engine 12 and the second engine 13 are drivingly connected to the main rotor 10, respectively, and the second generator 15 and the buffer battery 162 drive the electric propulsion mechanism 17. The aircraft controller controls: adjusting the torque signal output by the flight controller to the first motor controller 142 so that the first generator 14 does not output electric energy, the flight controller outputting a control command to disconnect the electrical connection between the power distribution mechanism 161 and the first generator 14; disconnecting the clutch of the transmission 11, which connects the first engine 12 with the first generator 14, adjusting the difference between the rotational speed of the first engine 12 and the main rotor 10 to an allowable range, closing the clutch of the transmission 11, which connects the first engine 12 with the main rotor 10, gradually adjusting the power output from the first engine 12) to the main rotor 10, adjusting the difference between the rotational speed of the second engine 13 and the main rotor 10 to an allowable range, closing the clutch of the transmission 11, which connects the second engine 13 with the main rotor 10, gradually adjusting the power output from the second engine 13 to the main rotor 10, simultaneously providing the main rotor 10 with driving force by the first engine 12 and the second engine 13, supporting the aircraft to switch to the first helicopter state, while adjusting the cyclic control system of the main rotor 10, the propeller 18 and the power output by the second engine 13 to the second generator 15 according to the thrust requirements of the aircraft.
Wherein, as shown in fig. 8, when the first generator 14 fails, the aircraft can be switched to a compound helicopter/helicopter state in which the drive connection of the first engine 12 to the first generator 14 is disconnected, and the electrical connection of the first generator 14 to the power distribution mechanism 16 is disconnected, so that the first engine 12 is drive-connected to the main rotor 10, and the output power of the second engine 13 is adjusted. Wherein the aircraft controller controls: feeding back a fault signal of the first generator 14 to the aircraft controller, adjusting a torque signal output by the flight controller to the first motor controller 142 so that the first generator 14 does not output electric energy, and outputting a control command by the flight controller to disconnect the electric connection of the distributor 161 and the first generator 14; the method comprises the steps of disconnecting a clutch in the transmission 11, connecting a first engine 12 with a first generator 14, adjusting the difference between the rotation speed of the first engine 12 and the rotation speed of the main rotor 10 to an allowable range, closing the connection between the first engine 12 and the main rotor 10, gradually adjusting the power output from the first engine 12 to the main rotor 10 to the helicopter state of the aircraft, simultaneously adjusting a cyclic control system of the main rotor 10, a propeller 18, and adjusting the power output from a second engine 13 to a second generator 15 according to the propulsive force demand of the aircraft.
As shown in fig. 9, when the first generator 14 fails, the aircraft may operate in a combined autogiro/fixed wing state, in which the transmission connection between the first generator 12 and the first generator 14 is disconnected, the electrical connection between the first generator 14 and the power distribution mechanism 16 is disconnected, so that the first generator 12 stops operating, and the second generator 15 and the buffer battery 162 drive the electric propulsion mechanism 17. Wherein, the torque signal output by the flight controller to the first motor controller 142 is adjusted to make the first generator 14 not output electric energy, and the flight controller outputs a control command to disconnect the electric connection between the power distribution mechanism 161 and the first generator 14; the clutch connecting the first engine 12 and the first generator 14 in the transmission 11 is disconnected, the first engine 12 stops working, the main high-power load of the aircraft is the propulsion propeller 18, the propulsion power is provided by the second generator 15 and the buffer battery 162, and the main rotor 10 is in a self-rotating state during landing.
Wherein, when the buffer battery 162 is out of order, if the aircraft is in a flight state, the electrical connection between the power distributor 161 and the buffer battery 162 is broken. Wherein, the flight controller controls: the electric connection between the buffer battery 162 and the distributor 161 is broken, and the generator directly driven by the engine generates electric energy which is distributed to the electric propulsion mechanism 17 through the power distribution mechanism 16 to drive the propulsion propeller 18 to rotate so as to generate the forward flying power. Fuel economy may be degraded due to buffer battery failure.
Wherein, as shown in fig. 10, when said first electric propulsion mechanism fails, the aircraft can switch to a compound helicopter/helicopter state, wherein said distributor 161 is electrically disconnected from said first electric propulsion mechanism and said fourth electric propulsion mechanism. The rotation speed command issued by the flight controller to the first motor driver 171 is minimized, and simultaneously, the rotation speed of the fourth motor driver 174 arranged symmetrically to the first motor driver 171 is minimized, and then the electrical connection between the power distribution mechanism 16 and the first motor driver 171 and the fourth motor driver 174 is disconnected, only two second electric propulsion mechanisms and two third electric propulsion mechanisms which are symmetrically distributed on the wing 30 are reserved in the electric propulsion mechanism 17 to work, and the propulsion power of the aircraft is mainly provided by the remaining second electric motors 176 and the third electric motors 177 in the electric propulsion mechanism 17. Adjusting the difference in the rotational speed of the first engine 12 and the main rotor 10 to an allowable range, closing the connection of the first engine 12 and the main rotor 10, adjusting the difference in the rotational speed of the first engine 13 and the main rotor 10 to an allowable range, closing the connection of the first engine 13 and the main rotor 10, gradually adjusting the power output by the first engines 12 and 13 to the main rotor 10 to a helicopter state of the aircraft, and simultaneously adjusting the cyclic control system of the main rotor 10, the pitch system of the propeller 18.
Wherein, as shown in fig. 11, when the first electric propulsion mechanism fails, the electrical connection of the power distributor to the first electric propulsion mechanism and the fourth electric propulsion mechanism is disconnected, and the aircraft can operate in a compound autogiro/fixed wing state. The rotation speed command issued by the flight controller to the first motor driver 171 is minimized, the rotation speed of the fourth motor driver 174 symmetrically arranged with the first motor driver 171 is minimized, the electrical connection between the power distribution mechanism 16 and the first motor driver 171 and the electrical connection between the fourth motor driver 174 are disconnected, and the cyclic pitch system of the main rotor 10 is adjusted to enable the main rotor 10 to operate in the air-driven autorotation state, so as to adjust the propulsive force of the electric propulsion mechanism 17. At this time, the aircraft may be operating in a compound autorotor state or an autorotor state or a fixed wing state.
The above description is only for the purpose of illustrating the preferred embodiments of the present invention and is not to be construed as limiting the invention, and any modifications, equivalents, improvements and the like that fall within the spirit and principle of the present invention are intended to be included therein.
Claims (10)
1. An aircraft, characterized in that it comprises a fuselage (20), a main rotor (10) arranged on the fuselage (20), a wing (11), a propeller (18) arranged on the wing (11), and a drive system, the driving system comprises an engine, a generator, a power distribution mechanism (16) and an electric propulsion mechanism (17), the engine being drivingly connected to the generator and to the main rotor (10), the electrical distribution mechanism (16) being electrically connected to the generator and to the electrical propulsion mechanism (17), the electric propulsion mechanism (17) is connected with the propulsion propeller (18) in a transmission way, the aircraft being capable of operating with the drive system simultaneously driving the main rotor (10) and the propulsive propellers (18), and the aircraft is capable of operating in a state in which the drive system drives only the propulsion propeller (18).
2. The aircraft of claim 1, characterized in that, in a condition in which said drive system drives said main rotor (10) and said propulsion propellers (18) simultaneously, said engine transmits mechanical energy to said main rotor (10) and said generator, respectively, said generator transmits electrical energy to said distribution mechanism (16), said distribution mechanism (16) transmits electrical energy to said electric propulsion mechanism (17), said electric propulsion mechanism (17) transmits mechanical energy to said propulsion propellers (18);
in a state in which the drive system drives only the propulsion propeller (18), the engine transmits mechanical energy only to the generator, which transmits electrical energy to the distribution mechanism (16), the distribution mechanism (16) transmits electrical energy to the electric propulsion mechanism (17), and the electric propulsion mechanism (17) transmits mechanical energy to the propulsion propeller (18).
3. The aircraft of claim 2, wherein the aircraft is operable in a first helicopter state in which the drive system drives the main rotor (10) and the propulsion propellers (18) simultaneously, wherein the main rotor (10) is capable of providing lift and the propulsion propellers (18) are capable of balancing the torque acting on the fuselage (20) by the main rotor (10).
4. The aircraft of claim 2, wherein the aircraft is operable in a second helicopter state in which the drive system simultaneously drives the main rotor (10) and the propulsion propeller (18), wherein the main rotor (10) is capable of providing lift and forward flight power, and the propulsion propeller (18) is capable of balancing the torque acting on the fuselage (20) by the main rotor (10).
5. The aircraft of claim 2, characterized in that it is operable in a compound helicopter state, in a state in which said drive system drives said main rotor (10) and said propulsion propellers (18) simultaneously, wherein said propulsion propellers (18) are capable of providing forward flight power and of balancing the torque acting on said fuselage (20) by said main rotor (10), said main rotor (10) and said wings (11) being capable of providing lift.
6. The aircraft of claim 2, wherein the aircraft is operable in a gyroplane state in which the drive system drives only the propulsor propellers (18), wherein the main rotor (10) provides lift and the propulsor propellers (18) provide forward flight power.
7. The aircraft of claim 2, wherein the aircraft is operable in a compound autogiro state in which the drive system drives only the propulsor propellers (18), wherein the main rotor (10) and the wings (11) together provide lift and the propulsor propellers (18) provide forward flight power.
8. The aircraft of claim 2, wherein the aircraft is operable in a fixed wing state with the drive system driving only the propulsion propeller (18), wherein the wing (11) provides lift and the propulsion propeller (18) provides forward flight power.
9. The aircraft of claim 1, wherein the drive system further comprises a transmission (11) drivingly connectable to the main rotor (10), the engine and the generator.
10. The aircraft of claim 1, characterized in that the power distribution mechanism (16) comprises a power distributor (161) and a buffer battery (162) electrically connected to each other, the power distributor (161) being electrically connected to the generator and to the electric propulsion mechanism (17), the electric propulsion mechanism (17) comprising a motor drive and an electric motor.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011399424.1A CN112429249A (en) | 2020-12-01 | 2020-12-01 | Aircraft with a flight control device |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011399424.1A CN112429249A (en) | 2020-12-01 | 2020-12-01 | Aircraft with a flight control device |
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| CN112429249A true CN112429249A (en) | 2021-03-02 |
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20190337612A1 (en) * | 2018-05-03 | 2019-11-07 | Carter Aviation Technologies, Llc | Compound Rotorcraft with Propeller |
| WO2020096254A1 (en) * | 2018-11-07 | 2020-05-14 | 문창모 | Vertical takeoff and landing aircraft using hybrid electric propulsion system and control method therefor |
| CN111498104A (en) * | 2020-04-20 | 2020-08-07 | 飞的科技有限公司 | Aircraft with a flight control device |
| CN111498101A (en) * | 2020-04-20 | 2020-08-07 | 飞的科技有限公司 | Aircraft with a flight control device |
| CN111498103A (en) * | 2020-04-20 | 2020-08-07 | 飞的科技有限公司 | Aircraft with a flight control device |
| US20200307779A1 (en) * | 2019-03-26 | 2020-10-01 | Bell Textron Inc. | Compound Helicopters having Auxiliary Propulsive Systems |
-
2020
- 2020-12-01 CN CN202011399424.1A patent/CN112429249A/en not_active Withdrawn
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20190337612A1 (en) * | 2018-05-03 | 2019-11-07 | Carter Aviation Technologies, Llc | Compound Rotorcraft with Propeller |
| WO2020096254A1 (en) * | 2018-11-07 | 2020-05-14 | 문창모 | Vertical takeoff and landing aircraft using hybrid electric propulsion system and control method therefor |
| US20200307779A1 (en) * | 2019-03-26 | 2020-10-01 | Bell Textron Inc. | Compound Helicopters having Auxiliary Propulsive Systems |
| CN111498104A (en) * | 2020-04-20 | 2020-08-07 | 飞的科技有限公司 | Aircraft with a flight control device |
| CN111498101A (en) * | 2020-04-20 | 2020-08-07 | 飞的科技有限公司 | Aircraft with a flight control device |
| CN111498103A (en) * | 2020-04-20 | 2020-08-07 | 飞的科技有限公司 | Aircraft with a flight control device |
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