NL2017971A

NL2017971A - Unmanned aerial vehicle

Info

Publication number: NL2017971A
Application number: NL2017971A
Authority: NL
Inventors: Yu Jinyong; Wang Chao
Original assignee: China Aviation Marine Equipment Yantai Tech Co Ltd
Priority date: 2016-08-09
Filing date: 2016-12-09
Publication date: 2018-02-14
Also published as: CN106114854A; CN106114854B; NL2017971B1; DE102017212716A1

Abstract

The present invention provides an unmanned aerial vehicle, comprising a main body, a fixed wing mounted on both sides of the main body, a plurality of rotors securingly connected to both sides of the fixed wing through respective rotor supporting parts, an airborne sensor system used for gathering flying data of the unmanned aerial vehicle, and a flight control system connected to the airborne sensor system, for adjusting states of the fixed wing and/or the rotors and further flying state of the unmanned aerial vehicle according to the flying data. According to the unmanned aerial vehicle, it is unnecessary for the rotor shaft thereof to rotate relative to the fixed wing, and thus complex mechanical components for controlling rotation of the rotor shaft are no longer necessary. Compared with the existing tilted-rotor aircraft, the aircraft provided herein has a simpler structure and lighter weight. At the same time, the unmanned aerial vehicle adopts one power system to perform vertical taking-off and landing and cruising. Therefore, it can provide much more capacity for meeting the requirements of loading and flying range and time compared with existing aircrafts.

Description

Title: UNMANNED AERIAL VEHICLE

Technical Field

The present invention relates to the field of aircraft, and more particularly, to an unmanned aerial vehicle.

Technical Background

Quadrocopter has been used more and more widely in many fields because of its small volume, light weight, and portability. Since the quadrocopter can easily enter into special spaces where human beings are unable to access, it can be used to perform flying tasks such as aerial shot, real-time monitoring, and geological survey, in addition to being used as plane model.

However, the quadrocopter still has some defects. For example, the speed of the quadrocopter is relatively low, and the time and distance for flying are short, causing the quadrocopter cannot be applied in scenarios where there are strict requirements on high speed and improved range.

Summary of the Invention

To solve the above technical problems, the present invention provides an unmanned aerial vehicle, comprising a main body, a fixed wing mounted on both sides of the main body, a plurality of rotors securingly connected to both sides of the fixed wing through respective rotor supporting parts, an airborne sensor system used for gathering flying data of the unmanned aerial vehicle, and a flight control system connected to the airborne sensor system, for adjusting states of the fixed wing and/or the rotors and further flying state of the unmanned aerial vehicle according to the flying data.

In one example, the flight control system converts the unmanned aerial vehicle from a vertical flying state to a horizontal cruising state based on a vertical-to-horizontal transition flight adjusting model, by which the flight control system is configured to control a head of the unmanned aerial vehicle from a vertical gesture to a horizontal gesture gradually through a differential control on rotating speeds of the rotors, and at the same time increase speed of the unmanned aerial vehicle to a predetermined horizontal cruising speed through increasing the rotating speeds of the rotors.

In one example, the flight control system is configured to convert the unmanned aerial vehicle from the horizontal cruising state to the vertical flying state based on a horizontal-to-vertical transition flight adjusting model, which is converse to the vertical-to-horizontal transition flight adjusting model.

In one example, the flight control system is configured to, when the unmanned aerial vehicle is in a vertical flying or hovering state, adjust the states of the fixed wing and the rotors based on a rotor adjusting model, so as to provide a main lift force for the unmanned aerial vehicle by the rotors. The flight control system is further configured to, when the unmanned aerial vehicle is in a horizontal cruising state, adjust the states of the fixed wing and the rotors based on a fixed wing adjusting model, so as to provide a main lift force for the unmanned aerial vehicle by the fixed wing.

In one example, the flight control system is configured to, in the rotor adjusting model, generate a first motor control instruction and an attitude angle instruction based on a received flight instruction and current location information of the unmanned aerial vehicle detected by the airborne sensor system, generate a second motor control instruction based on the attitude angle instruction and current attitude information of the unmanned aerial vehicle detected by the airborne sensor system, and control the states of the rotors based on the first and second motor control instructions to adjust location and attitude of the unmanned aerial vehicle.

In one example, the flight control system is configured to, in the rotor adjusting model, adjust the location and attitude of the unmanned aerial vehicle through respectively controlling rotating speeds of the rotors.

In one example, the fixed wing adjusted model includes a height control model, in which the flight control system is configured to: calculate a height offset based on an actual height and an expected height of the unmanned aerial vehicle, generate a first control signal using a first predetermined PID adjustor based on the height offset, generate a second control signal based on the first control signal and an actual vertical speed of the unmanned aerial vehicle, and adjust the flight height of the unmanned aerial vehicle through controlling a control surface angle of the unmanned aerial vehicle with the second control signal.

In one example, in the height control model the flight control system is further configured to generate a rotor controlling signal based on the second control signal and thus control the speed of each rotor with the rotor controlling signal.

In one example, the fixed wing adjusting model includes a speed control model, in which the flight control system is configured to: calculate a speed offset based on an actual speed and an expected speed of the unmanned aerial vehicle, generate a third control signal using a second predetermined PID adjustor based on the speed offset, and control the rotating speed of each rotor with the third control signal.

In one example, the rotating speeds of the rotors, when adjusted, have a same changing amount.

Existing vertical taking-off and landing (VTOL) fixed-wing unmanned aerial vehicles are generally categorized in two types. One type of aircraft is provided with a tilt-rotor. However, this kind of aircraft is large and complex in structure, and in the meantime, it is difficult for maintenance and has a high fault rate. Another type of aircraft is provided with two sets of power systems, i.e., one rotor system and one advancing propulsion system. However, this kind of aircraft merely combines two kinds of power systems together, thus reducing the weight of effective load and fuel. Therefore, the flying range and time of the second kind of aircraft cannot be significantly increased than those of the rotor aircraft.

According to the unmanned aerial vehicle provided herein, it is unnecessary for the rotor shaft thereof to rotate relative to the fixed wing, and thus complex mechanical components for controlling rotation of the rotor shaft are no longer necessary. Compared with the existing tilted-rotor aircraft, the aircraft provided herein has a simpler structure and lighter weight. At the same time, the unmanned aerial vehicle provided herein adopts one power system (i.e., rotor and its electrical motor) to perform vertical taking-off and landing and cruising. Therefore, it can provide much more capacity for meeting the requirements of loading and flying range and time compared with existing aircrafts.

The unmanned aerial vehicle provided herein combines the VTOL technology with the normal control technology for the fixed-wing aircraft together, thus it can not only perform VTOL and hovering as a helicopter, but also has advantages of fast speed and wide range as a fixed-wing unmanned aerial vehicle. However, both the existing unmanned aerial vehicle and the existing manned aircraft can possess only one of the advantages mentioned above. For example, a helicopter can vertically takeoff and land, but its speed is low and the flying time and range are short, while a fixed-wing aircraft with high speed and satisfactory flying time and range needs airport runways or complex launch and recovery equipments.

Since the unmanned aerial vehicle provided herein is able to vertically take off and land, it can be applied in most ships or occasions without airport runways (such as islands), and thus can meet both military and civil needs.

Moreover, during high-speed flight, the unmanned aerial vehicle can fly through using the fixed wing, thus it can possess advantages of long flying-distance and flying-time like the existing fixed-wing aircraft. In this case, the unmanned aerial vehicle can reach the destination in a short time. In addition, when the aircraft reach the destination, it can perform hovering or cruising through using the rotor. Accordingly, the aircraft provided herein is especially suitable for reconnaissance, survey, patrol, and so on.

Other features and advantages of the present invention will be further explained in the following description, and will partly become self-evident therefrom, or be understood through the implementation of the present invention. The objectives and advantages of the present invention will be achieved through the structures specifically pointed out in the description, claims, and the accompanying drawings.

Brief Description of the Drawings

The accompanying drawings, together with the embodiments, are provided for a further understanding of the present invention, and constitute a part of the description, and are not intended to limit, the present invention.

Figs. 1 to 3 show a front view, a side view and a top view of an unmanned aerial vehicle according to an embodiment of the present invention, respectively;

Fig. 4 schematically shows a rotation state of rotors of the unmanned aerial vehicle;

Figs. 5 to 7 schematically show an electrical system, an airborne sensor system, and a power and actuation system of the unmanned aerial vehicle, respectively;

Fig. 8 schematically shows a flight procedure of the unmanned aerial vehicle;

Fig. 9 schematically shows a rotor adjusting model according to the embodiment;

Fig. 10 schematically shows attitude control of the unmanned aerial vehicle;

Fig. 11 and 12 schematically shows a flying-height control circuit and a flying-speed control circuit in a fixed-wing adjusting model of the unmanned aerial vehicle, respectively; and

Figs. 13 and 14 show flow charts of flight control of the unmanned aerial vehicle, respectively.

Detailed Description of the Embodiments

The present invention will be explained in detail below with reference to the accompanying drawings, so that the objective, technical solutions and advantages thereof can be understood more clearly. It should be noted that the embodiments and features disclosed here can be combined with each other in any manner as long as there is no conflict, and the technical solutions obtained thereby all fall within the scope of the present invention.

In the meantime, many details are illustrated for the sake of explanation in the following, in order to provide a thorough comprehension to the embodiments of the present invention. However, it is obvious to persons skilled in the art that the present invention can be implemented without all of the details or specific modes disclosed herein.

In addition, steps as shown in the flow charts of the drawings can be performed by a computer system loaded a set of computer executable instructions. Moreover, while the flow chart shows a logic order of the steps being performed, in some instances these steps can be performed in an order different from that as shown.

Existing VTOL fixed-wing unmanned aerial vehicles are generally categorized in two types. One type of aircraft is provided with a tilt-rotor. However, this kind of aircraft is large and complex in structure, and in the meantime, it is difficult for maintenance and has a high fault rate. Another type of aircraft is provided with two sets of power systems, i.e., one rotor system and one advancing propulsion system. However, this kind of aircraft merely combines two kinds of power systems together, thus reducing the weight of effective load and fuel. Therefore, the flying range and time of the second kind of aircraft cannot be significantly increased than those of the rotor aircraft.

To solve the above defects existing in the prior art, the present invention provides a kind of novel VTOL fixed-wing unmanned aerial vehicle. The unmanned aerial vehicle is not only able to vertically take off and land and hover like a helicopter, but also possesses advantages of fast speed and long range like a fix-wing plane.

Fig. 1, Fig. 2, and Fig. 3 respectively show a front view, a side view and a top view of an unmanned aerial vehicle according to one embodiment of the present invention.

With reference to Figs. 1 to 3, the unmanned aerial vehicle provided herein preferably includes a main body 101, a fixed wing 102, a plurality of main undercarriages 103, and a plurality of rotors, wherein the fixed wing 102 is fixed on both sides of the main body 101. In the embodiment, the fixed wing 102 is preferably provided with an elevon 107 and two ailerons, i.e., a first aileron 106a and a second aileron 106b.

As shown in Fig. 1, the elevon 107, the first aileron 106a, and the second aileron 106b are all arranged at a longitudinal end (along y direction in Fig. 1) of the fixed wing 102, i.e., the end of the fixed wing 102 far away from the head of the aircraft, and the first and second ailerons are disposed at both sides of the elevon 107 respectively. The elevon 107 is used to control tilt angle of the main body through swing of control surface thereof during flight, and the ailerons 106a, 106b are used to control rolling of the aircraft through swing of control surfaces thereof during flight.

It should be noted that in other embodiments, the elevon and the ailerons can be arranged at other suitable positions, and the number thereof can also be other suitable values respectively. The present invention is not limited in this aspect.

In the embodiment, the main undercarriages 103 are each connected to a rotor supporting part 109, and symmetrically placed on both sides of the main body 101. The main undercarriages 103 extend along a head-tail direction of the main body 101. When the aircraft is in a state of parking on the ground, the main undercarriages 103 can support the aircraft, so that the main body 101 and the fixed wing 102 of the aircraft can vertically stand on the horizontal ground.

Additionally, in the embodiment, the main undercarriages 103 are further arranged symmetrically with respect to the fixed wing 102 so as to better support the main body of the aircraft. In other words, the two main undercarriages 103 on the same side of the main body 101 are respectively arranged on two sides of the fixed wing 102 which is at the same side of the main body 101 as said two main undercarriages 103.

In the embodiment, each main undercarriage 103 and its rotor supporting part 109 are preferably formed into one piece. Of course, in other embodiments the main undercarriage 103 can be connected to the rotor supporting part 109 by other means, and the present invention is not limited in this aspect.

It should be also noted that the main undercarriages 103 may be provided with corresponding buffering members, such as hydraulic absorbers, so as to reduce impact on the aircraft when landing. Similarly, the present invention is not limited in this aspect.

From the foregoing it can be seen that the unmanned aerial vehicle provided by the embodiment adopts a non-tail arrangement. That is to say, there is no vertical tail at the tail section of the main body like the existing aircrafts. Instead, the main undercarriages 103 function as the vertical tail during flight, so as to determine the flying direction for the aircraft. In this way, the structure of the aircraft can be simplified, and the weight thereof can be reduced significantly.

With reference to Fig. 3, the unmanned aerial vehicle is further provided with auxiliary undercarriages 109a each located at an end of the fixed wing 102 along a lateral direction (x direction in Fig. 1). The auxiliary undercarriages 109a can play a role for auxiliary supporting the aircraft when it is in a parking state (i.e., stays on the ground). Specifically, if the main body of the aircraft inclines to the right or left, the auxiliary undercarriages 109a can support the main body when contacting the ground, so as to avoid the aircraft falls down on the ground due to excessive inclination.

In the embodiment, when the aircraft is in a flying state, the auxiliary undercarriages 109a can also function to reduce a downwash airstream and increase lift force. It should be noted that in different embodiments the auxiliary undercarriages 109a can have different shapes and sizes according to actual needs, and the present invention is not limited in this aspect.

With reference to Figs. 1 to 3, the unmanned aerial vehicle preferably includes four rotors with identical structure, each comprising a driving motor 104 and a propeller 105. The driving motor 104 is fixedly connected to a corresponding rotor supporting part 109, and can drive the propeller 105 mounted thereon to rotate, so as to power the aircraft.

In the embodiment, the four rotors are symmetrically arranged on both sides of the fixed wing. In particular, as shown in Fig. 3, a first propeller 105a and a third propeller 105c respectively corresponding to a first rotor and a third rotor are symmetrically arranged on both sides of a first fixed wing (i.e., the fixed wing located at the left side of the main body as shown in Fig. 3), and a second propeller 105b and a fourth propeller 105d respectively corresponding to a second rotor and a fourth rotor are symmetrically arranged on both sides of a second fixed wing (i.e., the fixed wing at the right side of the main body as shown in Fig. 3).

Of course, in other embodiments the number of the rotors in the aircraft can be other values, and the rotors can be driven by other suitable devices (for example, an engine using fossil fuel).

In the embodiment, when the rotors run normally, the propellers of two neighboring rotors will rotate in opposite directions. For example, as shown in Fig. 4, the first propeller 105a and the fourth propeller 105d rotate clockwise, while the second propeller 105b and the third propeller 105c rotate counterclockwise. As such, torques of the rotors will be counteracted with each other, so that the aircraft can fly steadily.

Still referring to Figs. 1 to 3, the unmanned aerial vehicle further includes a cabin 108, where some devices such as image monitoring device are placed therein. When the aircraft is in a horizontal flying state, the cabin 108 will be at an underside of the main body 101, thus the devices contained therein can face towards the ground for conveniently monitoring targets on the ground.

The aircraft according to the embodiment inventively adopts a main body structure in which a quad-rotor boost structure is combined with a conventional fixed-wing arrangement. The taking-off and landing of the aircraft can be performed through a lift force provided by the four rotors, and other flying states can be achieved mainly through a lift force provided by a combination control of the four rotors and the fixed wing. Moreover, the aircraft has no extra power means except the rotors, and no tilting mechanism, either. Therefore, compared with the existing aircrafts, the aircraft provided herein has a simpler structure.

Fig. 5 schematically shows an electric system of the unmanned aerial vehicle provided herein.

As shown in Fig. 5, the electric system preferably includes an airborne sensor system 501, a data communication system 502, a flight control system 503, and a power and actuation system 504. The airborne sensor system 501 is used for gathering flying data of the aircraft, sending the gathered flying data to the flight control system 503 which is electrically connected thereto, so as to allow the flight control system 503 to adjust the flying state of the aircraft according to the received flying data.

In specific, as shown in Fig. 6, the airborne sensor system 501 preferably includes an inertial navigation measuring unit 501a, a wireless altimeter 501b, a pressure altimeter 501c, an airspeed meter 501d, and a GPS receiver 501e, etc. The inertial navigation measuring unit 501a preferably includes a triaxial accelerometer, a triaxial gyro, and a triaxial magnetometer, etc. After the flight control system 503 processes the flying data gathered by the airborne sensor system 501, the information such as flying attitude, attitude angular velocity, flying speed, latitude, longitude, height or the like can be obtained, and then used for adjustment of the flying state of the aircraft.

It should be noted that in other embodiments, the airborne sensor system 501 may include one or more of the above-mentioned devices only, or include other devices not mentioned here, or include a combination of one or more of the above-mentioned devices and other devices not mentioned here. The present invention is not limited in this aspect.

Fig. 7 schematically shows the power and actuation system 504 in the unmanned aerial vehicle of the present embodiment.

As shown in Fig. 7, the power and actuation system 504 preferably includes an electric power module 701, a signal adjusting circuit 702, a driving motor 104, ailerons 106, and an elevon 107, wherein the electric power module 701 is used for providing electric power for each electric device in the aircraft.

Specifically, in present embodiment the electric power module 107 includes an engine 701a and a generator 701b, wherein the engine 701a, which is connected to the generator 701b through shaft and located in the main body of the aircraft, drives the generator 701b in operation as a power source, so as to enable the generator can generate electrical power. The signal adjusting circuit 702b, being electrically connected with the generator 701b, can adjust the electric signal from the generator 701b to generate corresponding electrical signals and transmit them to the driving motor 104, the elevon 107 and the ailerons 106, so as to control speed of the driving motor 104, and the tilt angles of the elevon 107 and the ailerons 106.

As shown in Fig. 7, the electric power module 701 further includes a battery 701c, which is electrically connected to the signal adjusting circuit 701c for transmitting the energy stored therein to the signal adjusting circuit 702, and finally providing electric power for the operation of each electric device in the aircraft. In the present embodiment, preferably, the electric energy generated by the generator 701b can be transmitted to the battery 701c, if necessary, to charge the battery 701c.

From the above it can be seen that the aircraft of the present embodiment adopts two power supply devices (one is composed by the engine 701a and the generator 701b, and the other is the battery 701c) to supply power for the electric devices in the aircraft. However, in other embodiments it is conceivable to use only one power supply device to provide electric power for the electric devices. In this way, the number of the devices in the aircraft can be effectively reduced, the structure of the aircraft can be simplified, and in the meantime, the whole weight of the aircraft can be decreased.

Still referring to Fig. 5, the data communication system 502 of the aircraft according to the present embodiment is connected to the flight control system 503 to receive control instructions from outside, and transmit them to the flight control system 503 for adjustment of the flying state of the aircraft according to the instructions.

It still should be noted that in the present embodiment, the electric system of the unmanned aerial vehicle may further include other suitable modules, and the present invention is not limited in this aspect. For example, in one embodiment of the present invention, the electric system of the aircraft further includes a landing-assisted subsystem, etc. Moreover, a ground station subsystem which is placed on the ground or ships can be used with the aircraft. The landing-assisted subsystem is used for guiding the aircraft to land at a specified location on the ground or ship, and the ground station subsystem can control the flying state of the aircraft through respective instructions transmitted by the data communication system 502 to the aircraft.

The present unmanned aerial vehicle adopts a compound control strategy, in which a thrust vector multiple-rotor mode and a fixed-wing mode are combined together, to control the flying state of the aircraft. For taking off and landing, the rotor shaft always faces forwardly with tail sitter. In other words, the aircraft is supported by the main undercarriages before taking-off and after landing. In this case, the head of the aircraft faces up vertically to the sky, while the tail thereof faces down vertically to the ground, which forms the state of tail sitter. During taking off, cruising and landing, the rotor shaft always faces forwardly, i.e., toward the head of the aircraft, and therefore the rotor shaft will not rotate relative to the fixed wing.

The aircraft of the embodiment is a tension reversing VOTL aircraft. The aircraft takes off in a multiple-rotor tail sitter mode, thus the lift force generated by the rotors is in good linear relationship with the control signal of the flight control system (i.e., PWM wave received by the electric motor). Therefore, the power system model of the rotors can be considered as a linear model, which facilitates the design of the control program significantly.

The power of the rotors originates simply from tensions and torques caused by rotation of four rotors. If neglecting the problem in which those rotors are not vertically mounted, forces on the rotors in a body axes coordinate system can be intuitively expressed as follows:

(1) wherein Fb indicates the lift force of the aircraft body, and E indicates the lift force produced by rotor i.

The force on the four rotors can be converted into a different coordinate system by virtue of the rotation matrix between different coordinate systems, for control the flying state of the unmanned aerial vehicle.

In specific, from the flying procedure of the unmanned aerial vehicle as shown in Fig. 8, it can be seen that the aircraft firstly takes off from the ground vertically (i.e., the head of the aircraft is along the direction of y axis), and in this case, the rotors mainly provide the lift force for the aircraft. When the aircraft arrives at a certain height, it enters into a multiple-rotor mode to accelerate horizontally. Four rotors produce a downward moment by differential control at the same time, and control the elevon to produce another downward moment also. The amount of forces generated by the rotors and the elevon may be calculated according to a compound control law about the flying speed and the flying attitude, thereby distributing the control amount with respect to the surfaces of the rotors and elevon based on the calculation.

With the downward moment, the head of the aircraft may be pulled lower and lower till the body of the aircraft tends to be horizontal. In this procedure, the flight control system may adjust the rotating speed of the rotors to increase the airspeed of the aircraft up to a horizontal cruising speed finally. When arriving at the horizontal cruising speed, the aircraft may enter into the fixed-wing mode and fly horizontally (i.e., the head of the aircraft is along the direction of x axis). Since the speed of the aircraft in the horizontal flying mode is considerably high, the fixed wing can be used to provide the main lift force.

Similarly, during landing, the aircraft firstly climbs to enable the head of the aircraft facing toward the sky. In this procedure, the aircraft is converted from the fixed-wing mode to the multiple-rotor mode by means of a compound control of the rotors and the elevon. In the multiple-rotor mode, the aircraft may adjust the lift force generated by the rotors through adjusting the speed of the rotors, so as to slowly lower the height till landing on the ground.

It should be noted that in other embodiments the unmanned aerial vehicle may fly with other suitable schemes, and the present invention is not limited in this aspect. For example, in one embodiment of the present invention the unmanned aerial vehicle may adopt a vertical-to-horizontal transition stall mode when it is switched from the vertical flying mode to the horizontal cruising mode.

As shown in Fig. 8, in the horizontal stall mode (as indicated by the dash lines), when the aircraft vertically flies up to a certain height, the flight control system will control the body to turn 90 degrees by adjusting the rotating speed of the rotors. Since the airspeed is relatively low at this time, the unmanned aerial vehicle will fall down due to stall, and thus perform diving acceleration by virtue of gravity. During the diving acceleration, with the airspeed of the aircraft increasing, the lift force provided by the fixed wing will be increased, eventually causing the aircraft to stay in the horizontal cruising state.

In the embodiment, the flight control system generally resolves the mathematical model of the aircraft into a vertical subsystem and a lateral subsystem when controlling the flying state of the aircraft. During the transition flight, the state values such as rolling angle, yaw angle, and sideslip angle etc. of the lateral subsystem of the unmanned aerial vehicle maintain unchanged, but pitching angle, airspeed, and angle of attack of the lateral subsystem are changed. Therefore, for conveniently controlling the aircraft, the mathematical model of the aircraft in the transition flight mode is simplified in the present embodiment by ignoring the lateral subsystem of the aircraft and taking it as a disturbance. That means only the vertical subsystem of the aircraft is considered, so that the six degree of freedom model of the aircraft can be simplified as a two degree of freedom model.

When the aircraft stays a VTOL state or a hovering state, the flight control system may adjust the states of the rotors and the fixed wing according to the rotor adjusting mode, so that the main lift force for the whole aircraft can be provided by the rotors.

When the aircraft stays a VTOL state or a hovering state, the head of the aircraft faces upwardly, with a pitching angle of nearly 90 degrees. In this flight mode, the speed of the aircraft is relatively low, and the influence of the aileron surfaces is weak. Therefore, adjusting and maintaining all flying attitudes of the aircraft is mainly performed by the rotors, and the weight management of the aircraft is mainly balanced by the pull force generated by the rotors. In the embodiment, the vertical flight mode of the aircraft is mainly used for the vertical taking-off and landing, hovering, and horizontal low-speed maneuver of the aircraft.

In the embodiment, a coordinate system X-Y-Z is defined to be static relative to the ground when the aircraft is in the vertical flight mode, wherein X axis and Y axis are in the horizontal plane (when the aircraft is vertical), Z axis points to the ground, and X axis points to the direction vertical to the surface of the fixed wing in the horizontal plane. Rotation of four rotors around X axis is called as rolling, rotation around Y axis is called as pitching, and rotation around Z axis is called as yawing. Attitude angles are defined by Euler angles, in the order of rolling-pitching-yawing.

In the embodiment, the flight control system changes movement of the aircraft along the vertical direction through changing the total of lift forces provided by the four rotors, and changes vertical attitude thereof through changing the difference between lift forces of the propellers, so as to change the speed and position thereof along the vertical direction.

Fig. 9 schematically shows a rotor adjusting model according to the present embodiment.

As shown in Fig. 9, according to the present embodiment, the flight control system preferably controls the flying state of the aircraft through an inner-external loop control mode. The flight control system mainly adjusts the attitude of the aircraft in the inner loop and the location of the aircraft in the spatial coordinate system in the external loop.

In specific, when the flight control system controls the attitude of the aircraft according to the rotor adjusting model, it will generate a first motor control instruction and an attitude angle instruction based on a received flight instruction and current location information of the aircraft. According to the attitude angle instruction, the flight control system will adjust the attitude of the aircraft in the inner loop, and generate a second motor control instruction based on the attitude angle instruction and current attitude information of the aircraft. Then, the electric motor will adjust its running parameters, such as rotary speed, according to the first and second motor control instructions. And in the four-rotor model, changes of the running parameters of the electric motor will render changes of the running parameters of the rotors, and thus changes of the location and attitude of the aircraft.

Since the unmanned aerial vehicle adopts a symmetric arrangement of four rotors, it is possible to decouple the relationships among the height, pitching, rolling, and yawing of the aircraft, so as to control the flying state of the aircraft.

In specific, as shown in Fig. 10, when the aircraft rises up, the flight control system simultaneously increases rotating speed of each rotor, so that each rotor provides the same lift force. In this way, there is an upward acceleration for the aircraft, but the attitude of the aircraft remains unchanged.

If the aircraft needs to pitch, the flight control system will decrease the rotating speeds of the third and fourth rotors at the time of increasing those of the first and second rotors. In this way, the lift forces generated by the first and second rotors will be increased, while those generated by the third and fourth rotors will be decreased, so that the aircraft will have a positive angular acceleration in the direction of pitching.

When the aircraft needs to roll, the flight control system will decrease the rotating speeds of the second and third rotors at the time of increasing those of the first and fourth rotors. As such, the lift forces generated by the first and fourth rotors will be increased, while those generated by the second and third rotors will be decreased, so that the aircraft will have an angular acceleration in the direction of rolling.

When the aircraft needs to yaw, the flight control system will decrease the rotating speeds of the second and fourth rotors at the time of increasing those of the first and third rotors. As such, the lift forces generated by the first and third rotors will be increased, while those generated by the second and fourth rotors will be decreased, so that the aircraft will have an angular acceleration in the direction of yawing.

It should be noted that in the procedure of changing one of the attitude parameters during control, the increment and/or decrement of the rotating speed of each rotor is preferably kept the same at the time of adjusting the speed of each rotor, so as to not affect other parameters.

In the horizontal flight mode, a large fixed wing of the aircraft can generate enough lift force to balance its gravity due to the high airspeed, thereby presenting advantages of high speed and high efficiency. Since the aircraft mainly relies on aerodynamic force to balance the gravity during this period, the flight control system controls the elevon and the ailerons mainly based on the fixed-wing adjusting model, thus controlling the flight state of the aircraft.

Fig. 11 schematically shows a flying-height control circuit in the fixed-wing adjusting model of the unmanned aerial vehicle.

As shown in Fig. 11, when the flight control system controls the height of the aircraft based on the fixed-wing adjusting model, it will firstly obtain an expected height Hg of the aircraft, and then limit the amplitude of the expected height Hg to calculate a difference between an actual height H and the limited height Hg, so as to obtain a height offset AH. The height offset AH is a height that the aircraft needs to be changed.

After obtaining the height offset AH, the flight control system will use a first PID regulator to generate a first control signal Ci based on the height offset AH. After obtaining the first control signal Ci, the flight control system will generate a second control signal Θ based on the first control signal Ci and an actual vertical speed H of the aircraft. Lastly, the second control signal 6g is input to a pitching control circuit, so that a surface signal Slon for controlling a deflection state of the elevon surface can be obtained. After the elevon receives the surface signal Slon, it will adjust its surface to have a corresponding deflection angle. Therefore, the aircraft can be controlled to reach the expected height Hg.

In the actual control, if the elevon and the ailerons of the aircraft, move, both the yawing angle and the airspeed thereof will be significantly changed. When the rotating speed of the rotors is constant, the control to the aircraft actually turns to a conversion between the kinetic energy and the potential energy of the aircraft. In this situation, the control range will be limited, and thus at this time it is necessary to control the rotating speed of the rotors to change the total energy of the aircraft, so as to enable the height and speed of the aircraft to reach respective expected values.

Therefore, as shown in Fig. 11, the flight control system will adjust the rotating speed of the rotors when adjusting the surface of the elevon of the aircraft. In specific, after receiving the second control signal Θ , the flight control system will generate a rotor control signal δ , which is used to control the rotating speed of the electric motor corresponding to each rotor of the aircraft. Therefore, the pull force generated by each rotor can be adjusted, so that the speed of the aircraft can be adjusted.

Fig. 12 schematically shows a flying-speed control circuit in the fixed-wing adjusting model of the unmanned aerial vehicle.

As shown in Fig. 12, when controlling the aircraft’s speed based on the fixed-wing adjusting model, the flight control system will firstly obtain an expected speed V of the aircraft, and then limit the amplitude thereof to calculate a difference value between the expected speed after amplitude limitation and the actual speed of the aircraft, so as to obtain a speed offset AV.

After obtaining the speed offset AV, the flight control system will use a second PID regulator to generate a third control signal G?based on the speed offset AV. After obtaining the third control signal C:i, the flight control system will generates a rotor control signal Sp of the aircraft, which can be used to control the rotating speed of the electric motor corresponding to each rotor of the aircraft. Therefore, the pull force generated by each rotor can be adjusted, so that the speed of the aircraft can be adjusted.

As described above, when the state of the aileron of the aircraft is unchanged, increasing or decreasing the rotor speed will cause changes of the pitching angle of the aircraft, and thus changes of the height thereof. In this case, the flight control system will regulate the state of the elevon at the time when adjusting the rotating speed of the rotors. In specific, as shown in Fig. 12, after obtaining the third control signal (¾ the flight control system will generate a surface signal Slon to control a surface deflection state of the elevon. The elevon will adjust the surface thereof to have a corresponding deflection angle after receiving the surface signal Slon.

Figs. 13 and 14 each show a flight control flow chart of the unmanned aerial vehicle.

As shown in Fig. 13, the flight control system firstly performs initialization after the aircraft is activated, and then collect a remote control signal after a predetermined time period. In the embodiment, the remote control signal is transmitted from the ground station system to control the flying state of the aircraft.

After receiving the remote control signal, the flight control system will determine whether there is a taking-off instruction. If there is the taking-off instruction, it is further determined whether the current height of the aircraft reaches a predetermined height. In the embodiment, the predetermined height is characterized as a height of the aircraft when the taking-off ends, which is preferably 25 cm. It should be noted that in different embodiments the above predetermined height can be any other suitable value according to actual requirements on the flight and parameters of the aircraft, and the present invention is not limited in this regard.

If the aircraft reaches the predetermined height, it indicates that the taking-off of the aircraft has been fulfilled. If not, it indicates that the taking-off has not been fulfilled, so that the flight control system will still collect the remote control signal, and then updates the attitude of the aircraft.

When updating the attitude of the aircraft, the flight control system determines whether there is a taking-off instruction. If there is the taking-off instruction, the flight control system will command the aircraft to take off with a fixed lift force through controlling the rotating speed of the rotors. In this procedure, the flight control system will continuously control the flying state of the aircraft and determine whether the aircraft reaches the predetermined height. If there is no taking-off instruction, the flight control system will control the attitude of the aircraft, and determine again if there is a taking-off instruction received.

As shown in Fig. 14, the flight control system will continuously collect the remote control signal and determine whether it is necessary for the aircraft to be changed into a manual mode after the taking-off procedure ends. If it is necessary for the aircraft to be changed into the manual mode, the flight control system will obtain a flying state instruction of the aircraft according to the remote control signal received, and then control the flying state of the aircraft, such as the flying height and the flying attitude, according to the flying state instruction. During the procedure of controlling the flying state of the aircraft, the flight control system will further determine whether it is necessary to land according to the received remote control signal. If there is a need to land, the flight control system will control the aircraft to be changed into a rotor flight mode, and then stop the rotors when landing requirements are met. As such, the whole controlling procedure ends.

However, if there is unnecessary for the aircraft to be changed into the manual mode, the flight control system will read a predetermined attitude instruction from the memory of the aircraft through a serial port, and convert the rotor flight mode into the fixed-wing flight mode based on the attitude instruction. And in the fixed-wing flying mode, the flight control system will control the flying attitude of the aircraft and the flying height according to the specific control instruction.

During the procedure of controlling the flying state of the aircraft, the flight control system will further determine whether there is a need to land according to control instructions stored in the aircraft. If it is necessary to land, the flight control system will control the aircraft to be changed into the rotor flight mode, and when the lading conditions are met (such as the height of the aircraft is zero), the rotors are controlled to stop. As such, the whole controlling procedure ends.

The unmanned aerial vehicle provided herein combines the VTOL technology with the normal control technology for the fixed-wing aircraft together, thus it can not only perform VTOL and hovering as a helicopter, but also has advantages of fast speed and wide range as a fixed-wing unmanned aerial vehicle. However, both the existing unmanned aerial vehicle and the existing manned aircraft can possess only one of the advantages mentioned above. For example, a helicopter can vertically take off and land, but its speed is low and the flying time and range are short, while a fixed-wing aircraft with high speed and satisfactory flying time and range needs airport runways or complex launch and recovery equipments.

It should be understood that the embodiment provided here is not limited to the specific structures or process steps disclosed herein, but should be extended equivalents of the technical features which persons skilled in the art can appreciate. It should be still understood, terms used herein are merely for describing specific embodiments, and not intended to be restrictive. “One embodiment” or “embodiments” mentioned in the description indicate that specific features, structures, or characteristics are involved in at least one embodiment of the present invention. Therefore, the phrases “one embodiment” or “embodiments” in each place throughout the description do not always indicate the same embodiment.

Although the above examples are intended for explaining a principle of the present invention in one or multiple applications, it is obvious for the person skilled in the art to make various modifications to formations, usages, or details of implementation without departing away from the concept and idea of the present invention on the condition that there is no need for inventive labors.

Claims

1. Een onbemand luchtvaartuig, omvattende: een hoofdlichaam; een vaste-vleugel gemonteerd aan beide kanten van het hoofdlichaam; een meervoudig aantal rotoren stevig bevestigd aan beide kanten van de vaste-vleugel door respectieve rotor dragende onderdelen; een airborne-sensorsysteem dat wordt gebruikt voor het verzamelen van vluchtdata van het onbemande luchtvaartuig; en een vluchtbesturingssysteem verbonden met het airborne -sensorsysteem, voor het aanpassen van standen van de vaste-vleugel en/of de rotoren en verdere vliegstand van het onbemande luchtvaartuig aan de hand van de vluchtgegevens.An unmanned aircraft, comprising: a main body; a fixed wing mounted on both sides of the main body; a plurality of rotors securely attached to both sides of the fixed wing by respective rotor bearing members; an airborne sensor system used to collect flight data from the unmanned aircraft; and a flight control system connected to the airborne sensor system, for adjusting positions of the fixed wing and / or the rotors and further flight position of the unmanned aircraft on the basis of the flight data.

2. Het onbemande luchtvaartuig volgens conclusie 1, waarbij het vluchtbesturingssysteem het onbemande luchtvaartuig van een verticale vliegstand naar een horizontale kruisstand brengt aan de hand van een verticale-naar-horizontale transitie vlucht aanpas model, waarbij het vluchtbesturingssysteem is ingericht voor het besturen van een kop van het onbemande luchtvaartuig geleidelijk aan van een verticale houding naar een horizontale houding door middel van een differentieel-regeling van rotatiesnelheden van de rotoren, en voor het tegelijkertijd de snelheid van het onbemande luchtvaartuig verhogen tot een vooraf bepaalde horizontale kruissnelheid door het verhogen van de rotatiesnelheid van de rotoren.The unmanned aircraft according to claim 1, wherein the flight control system brings the unmanned aircraft from a vertical flight position to a horizontal cross position on the basis of a vertical-to-horizontal transition flight adaptation model, wherein the flight control system is adapted to control a head of the unmanned aircraft gradually from a vertical position to a horizontal position by means of a differential control of rotational speeds of the rotors, and simultaneously increasing the speed of the unmanned aircraft to a predetermined horizontal cruising speed by increasing the rotational speed of the rotors.

3. Het onbemande luchtvaartuig volgens conclusie 2, waarbij het vluchtbesturingssysteem is ingericht om het onbemande luchtvaartuig van de horizontale kruisstand naar de verticale vliegstand te brengen op basis van een horizontale-naar-verticale transitie vlucht aanpas model, welke tegenovergesteld is aan het verticale-naar-horizontale transitie vlucht aanpas model.The unmanned aircraft according to claim 2, wherein the flight control system is arranged to take the unmanned aircraft from the horizontal cruising position to the vertical flight position based on a horizontal-to-vertical transition flight adaptation model, which is opposite to the vertical-to - horizontal transition flight adaptation model.

4. Het onbemande luchtvaartuig volgens één der conclusies 1-3, waarbij het vluchtbesturingssysteem ingericht is om, wanneer het onbemande luchtvaartuig zich in een verticale vliegstand of zweefstand bevindt, de standen van de vaste-vleugel en de rotoren aan te passen aan de hand van een rotor aanpas model, om door middel van de rotoren een hoofd liftkracht voor het onbemande luchtvaartuig te bewerkstelligen; en waarbij het vluchtbesturingssysteem ingericht is om, wanneer het onbemande luchtvaartuig zich in een horizontale kruisstand bevindt, de standen van de vaste-vleugel en de rotoren aan te passen aan op basis van een vaste-vleugel aanpas model, om door middel van de vaste-vleugel een hoofd liftkracht voor het onbemande luchtvaartuig te bewerkstelligen.The unmanned aircraft according to any one of claims 1-3, wherein the flight control system is adapted to adjust the positions of the fixed wing and the rotors on the basis of the flight control system when the unmanned aircraft is in a vertical flight position or gliding position a rotor adjustment model, to achieve a main lift force for the unmanned aircraft by means of the rotors; and wherein the flight control system is adapted, when the unmanned aircraft is in a horizontal cruising position, to adjust the positions of the fixed wing and the rotors to a fixed-wing adjustment model, by means of the fixed wing wing a head lifting force for the unmanned aircraft.

5. Het onbemande luchtvaartuig volgens conclusie 4, waarbij het vluchtbesturingssysteem is ingericht voor, in het rotor aanpas model, het genereren van een eerste motor besturingsinstructie en een houding hoek instructie aan de hand van een ontvangen vluchtinstructie en actuele locatie-informatie van het onbemande luchtvaartuig die door het airborne-sensorsysteem zijn waargenomen, het genereren van een tweede motor besturingsinstructie aan de hand van de houding instructie en actuele houding-informatie van het onbemande luchtvaartuig die door het airborne-sensorsysteem zijn waargenomen, en het besturen van de standen van de rotoren aan de hand van de eerste en tweede motor besturingsinstructies om de locatie en houding van het onbemande luchtvaartuig aan te passen.The unmanned aircraft according to claim 4, wherein the flight control system is adapted to, in the rotor fitting model, generate a first engine control instruction and an attitude angle instruction based on a received flight instruction and current location information of the unmanned aircraft sensed by the airborne sensor system, generating a second engine control instruction based on the attitude instruction and current attitude information of the unmanned aircraft observed by the airborne sensor system, and controlling the positions of the rotors on the basis of the first and second engine control instructions to adjust the location and attitude of the unmanned aircraft.

6. Het onbemande luchtvaartuig volgens conclusie 5, waarbij het vluchtbesturingssysteem ingericht is voor, in het rotor aanpas model, het aanpassen van de locatie en houding van het onbemande luchtvaartuig door respectieve rotatiesnelheden van de rotoren te regelen.The unmanned aerial vehicle according to claim 5, wherein the flight control system is adapted for, in the rotor fitting model, adjusting the location and attitude of the unmanned aerial vehicle by controlling respective rotational speeds of the rotors.

7. Het onbemande luchtvaartuig volgens één der conclusies 4-6, waarbij het vaste-vleugel aanpas model een hoogte besturingsmodel omvat, waarin het vluchtbesturingssysteem is ingericht voor: het berekenen van een hoogte-afwijking aan de hand van een werkelijke hoogte en een verwachtte hoogte van het onbemande luchtvaartuig, het genereren van een eerste besturingssignaal met behulp van een eerste vooraf bepaalde PID-aanpasser aan de hand van de hoogte-afwijking, het genereren van een tweede besturingssignaal aan de hand van het eerste besturingssignaal en een werkelijke verticale snelheid van het onbemande luchtvaartuig, en het aanpassen van de vlieghoogte van het onbemande luchtvaartuig door het met het tweede besturingssignaal besturen van een stuurvlak hoek van het onbemande luchtvaartuig.The unmanned aircraft according to any of claims 4-6, wherein the fixed wing adaptation model comprises a height control model, wherein the flight control system is adapted to: calculate a height deviation on the basis of an actual height and an expected height of the unmanned aircraft, the generation of a first control signal with the aid of a first predetermined PID adjuster on the basis of the altitude deviation, the generation of a second control signal on the basis of the first control signal and an actual vertical speed of the unmanned aircraft, and adjusting the flight height of the unmanned aircraft by controlling a pilot plane angle of the unmanned aircraft with the second control signal.

8. Het onbemande luchtvaartuig volgens conclusie 7, waarbij in het vlieghoogte besturingsmodel het vluchtbesturingssysteem verder is ingericht om een rotor besturingssignaal te genereren aan de hand van het tweede besturingssignaal en zodanig de snelheid van elke rotor te besturen met het rotor besturingssignaal.The unmanned aerial vehicle according to claim 7, wherein in the flight height control model the flight control system is further adapted to generate a rotor control signal on the basis of the second control signal and thus control the speed of each rotor with the rotor control signal.

9. Het onbemande luchtvaartuig volgens conclusie 7 of 8, waarbij het vaste-vleugel aanpas model een snelheid besturingsmodel omvat, waarin het vluchtbesturingssysteem ingericht is voor: het berekenen van een snelheidsafwijking aan de hand van een werkelijke snelheid en een verwachtte snelheid van het onbemande luchtvaartuig, het genereren van een derde besturingssignaal met behulp van een tweede vooraf bepaalde PID-aanpasser aan de hand van de snelheidsafwijking, en het besturen van de rotatiesnelheid van elke rotor met het derde besturingssignaal.The unmanned aircraft according to claim 7 or 8, wherein the fixed wing adaptation model comprises a speed control model, wherein the flight control system is adapted to: calculate a speed deviation based on an actual speed and an expected speed of the unmanned aircraft generating a third control signal with the aid of a second predetermined PID adjuster from the speed deviation and controlling the rotational speed of each rotor with the third control signal.

10. Het onbemande luchtvaartuig volgens één der conclusies 1-9, waarbij de rotatiesnelheden van de rotoren, wanneer aangepast, eenzelfde verander grootte hebben.The unmanned aircraft according to any of claims 1-9, wherein the rotational speeds of the rotors, when adjusted, have the same change size.