A vertical take-off and landing ground effect vehicle and control method thereof

Abstract

The invention relates to a vertical take-off and landing ground effect vehicle and a control method thereof. The vehicle comprises a fuselage, a ducted ground effect wing, an outer wing, a vertical tail, a buoy, an embedded ducted rotor system and a propulsion rotor power system; the control method of the vehicle comprises a longitudinal attitude control method, a lateral attitude control method, a motion control method on the water surface and an emergency power jump control method. The vertical take-off and landing ground effect vehicle of the invention adopts a vertical take-off / landing mode. Compared with a traditional ground effect vehicle, it has no taxiing process, greatly improves the riding comfort of the passengers in the take-off and landing stages, and can seamlessly fly from flat ground to shallow water to water surface. The vehicle adopts automatic control auxiliary driving, which reduces the risk of the wing tip colliding with the water surface / ground during lateral maneuvers such as turning, thereby improving safety. At the same time, it has a power jump capability, which greatly improves flight safety; it can directly take off and land on buildings and islands without docks, greatly improving the application field.

Description

Technical Field

[0001] The present invention relates to the field of aviation technology, and in particular to a vertical take-off and landing ground effect vehicle and a control method thereof. Background Art

[0002] The current ground effect flight technology is relatively mature. During the take-off and landing phases of traditional ground effect vehicles, the waves have a great impact on the aircraft's attitude. The passengers can feel obvious turbulence and the ride comfort is very poor. In serious cases, it will also affect the attitude control of the aircraft during take-off and landing, and even cause flight safety accidents.

[0003] During the ground effect cruise phase, when the pilot controls the aircraft to roll and turn, it is easy for the pilot to over-control the aircraft, causing the wing tip of one side of the aircraft to be too close to the water surface, affecting flight safety, or even causing the wing tip to enter the water, resulting in an accident in which the aircraft is destroyed and people are killed.

[0004] At the same time, existing ground effect vehicles need a dock to load and unload passengers and cargo, and cannot reach some cliff islands where there are no conditions for building a dock.

[0005] When a traditional vertical take-off and landing aircraft flies forward, it lowers its head to create forward thrust for the rotor blades. The faster it flies, the greater the angle of the head lowering. When flying at high speeds, the passengers' field of vision and comfort become very poor. Summary of the invention

[0006] The technical problem solved by the invention is to provide a vertical take-off and landing ground effect vehicle and a control method thereof.

[0007] The technical solution adopted by the present invention to solve its technical problem is:

[0008] A vertical take-off and landing ground effect vehicle, comprising: a fuselage, a ducted ground effect wing, an outer wing, a vertical tail, a float, an embedded ducted rotor system and a propulsion rotor power system;

[0009] The interior of the fuselage is used to arrange airborne electronic equipment, and other equipment or loads such as passengers or cargo;

[0010] The ducted ground effect wing is symmetrically arranged on the left and right and fixedly connected to the upper side of the middle and rear section of the fuselage, and a plurality of ducts are regularly and symmetrically arranged on the ducted ground effect wing;

[0011] The buoys are arranged symmetrically on the left and right sides of the fuselage and are installed at the left and right wing tips of the ducted ground effect wing to form left and right buoys;

[0012] The outer wings are arranged symmetrically on the left and right sides of the fuselage and tilted outwards, and are installed on the outer sides of the left and right floats to form left and right outer wings; or, the outer wings are rotatably installed on the outer sides of the left and right floats through hinge components to form foldable left and right outer wings;

[0013] The vertical tail is arranged symmetrically on the left and right sides of the fuselage to form left and right vertical tails, and the left and right vertical tails are arranged and installed on the upper side of the rear part of the ducted ground effect wing through tail fin buoys, and the tail fin buoys are arranged symmetrically on the left and right sides of the fuselage and installed on the rear part of the ducted ground effect wing; the propulsion rotor power system is arranged tilted forward through a strut and installed between the left and right vertical tails;

[0014] Alternatively, the vertical tail is directly arranged and installed on the upper side of the tail of the fuselage, and the propulsion rotor power system is arranged forwardly tilted through the strut and installed on the left and right sides of the vertical tail to form a left and right propulsion rotor power system;

[0015] The embedded ducted rotor power system includes a plurality of rotor power devices, which are embedded in a plurality of ducts of the ducted ground effect wing through a rotor power mounting bracket to form a plurality of embedded ducted rotor power devices; the rotor power mounting bracket is fixedly connected to the lower part or the upper part of the ducted ground effect wing, and the plurality of rotor power devices are mounted on the rotor power mounting bracket and embedded in the duct;

[0016] The plurality of embedded ducted rotor power units are arranged to be tilted outwards, and the angle at which the embedded ducted rotor power units are tilted outwards is 5 to 10 degrees.

[0017] Furthermore, the ducted ground effect wing adopts an arched wing structure, the left and right pontoons are symmetrically and horizontally arranged through the wingtips of the ducted ground effect wing, and the tail pontoon is also symmetrically and horizontally arranged and installed at the rear of the ducted ground effect wing.

[0018] Furthermore, the propulsion rotor power system is arranged to be tilted forward through a strut and fixedly installed between the left and right vertical tails or on the left and right sides of the vertical tail, and the forward tilt angle of the propulsion rotor power system is 25°~35°; or it is arranged between the left and right vertical tails or on the left and right sides of the vertical tail through a strut and a tilt mechanism, and the propulsion rotor power system is tilted through the tilt mechanism, and its tilt angle range is -5°~100°.

[0019] Furthermore, the propeller disc area of ​​the embedded ducted rotor power device at the front section of the ducted ground effect wing is greater than or equal to the propeller disc area of ​​the embedded ducted rotor power device at the rear section thereof.

[0020] Furthermore, the embedded ducted rotor power system includes four, six or eight embedded ducted rotor power devices in an even order, and are installed in the duct of the ducted ground effect wing through a left-right symmetrical arrangement of rotor power mounting brackets.

[0021] Furthermore, the center of gravity of the vertical take-off and landing ground effect vehicle is set at 0 to 0.05 times the average aerodynamic chord length in front of the aerodynamic focus, and the rotor center is set at 0.02 to 0.08 times the average aerodynamic chord length behind the center of gravity of the vertical take-off and landing ground effect vehicle.

[0022] A control method for a vertical take-off and landing ground effect vehicle, comprising a longitudinal attitude control method, a lateral attitude control method, a water surface movement control method and an emergency power jump control method;

[0023] The longitudinal attitude control method includes pitch attitude control in a hovering mode and a ground effect cruise mode, and the pitch attitude control is accomplished by a combined rotor system composed of the embedded ducted rotor power system and the propulsion rotor power system;

[0024] In the ground effect cruise mode, the resultant nose-down moment of the combined rotor system located behind the center of gravity balances most of the aerodynamic nose-up moment, and each rotor works near the optimal state; in the flight from the hovering mode to the ground effect cruise mode, by increasing the fixed tilt angle of the propulsion rotor power system or increasing the power of the tiltable propulsion rotor power system, the thrust for forward flight is provided while suppressing the nose-up aerodynamic moment of the aircraft, so that the pitch angle of the aircraft is controlled within 5°, so as to realize the attitude conversion control of the aircraft from the hovering mode to the ground effect cruise mode;

[0025] In the hovering mode, the pitch attitude control reduces the power of the propulsion rotor power system by increasing the power of the embedded ducted rotor power unit forward of the ducted ground effect wing;

[0026] The lateral attitude control method includes rolling attitude control and heading turning control in ground effect cruise mode;

[0027] The roll attitude control makes the distance between the floats at the left and right wingtips of the ducted ground effect wing and the water surface or the ground greater than the safe height by compensating the height supplementary instruction of the aircraft related to the roll angle into the height control channel;

[0028] Alternatively, the rolling attitude control method comprises installing a position sensor for measuring the distance from the water surface or the ground on the buoy, measuring the distance change between the buoy and the water surface or the ground in real time during the rolling process of the aircraft through the position sensor, and compensating the distance change into the altitude control channel;

[0029] The heading turn control establishes a desired roll angle through the roll attitude control method so that the aircraft generates a suitable centripetal force on the horizontal plane of motion;

[0030] The method for controlling motion on the water surface includes heading control and forward motion control on the water surface;

[0031] The heading control on the water surface is performed by controlling the differential speed of the rotors on the plurality of embedded ducted rotor power units in the embedded ducted rotor power system, by increasing the speed of the counterclockwise rotating rotor and reducing the speed of all clockwise rotating rotors to achieve the clockwise rotation of the aircraft; or reducing the speed of all counterclockwise rotating rotors and increasing the speed of all clockwise rotating rotors to achieve the counterclockwise rotation of the aircraft;

[0032] The forward motion control on the water surface is achieved by the forward rotation of the propulsion rotor power system to achieve the forward motion of the aircraft; and the backward motion of the aircraft is achieved by the reverse rotation of the propulsion rotor power system;

[0033] The emergency power jump control method increases the power of the embedded ducted rotor system and simultaneously increases the flight angle of attack of the aircraft to increase the aerodynamic lift of the entire aircraft, thereby achieving the jump maneuverability of the aircraft.

[0034] Furthermore, in the pitch attitude control, the following steps are also included:

[0035] Step S101: Calculate the rotor center x tx ,

[0036]

[0037] Among them, S A , S B , S C , S D , S F are the projected areas of the rotors of the embedded ducted rotor power device (A-D) and the propulsion rotor power system (F) on the horizontal plane, respectively, and x A 、x B 、x C 、x D 、x F are the positions of the rotors of the embedded ducted rotor power unit (A-D) and the propulsion rotor power system (F) on the x-axis respectively;

[0038] Step S102: Calculate the pitch control energy efficiency coefficient f of each rotor in the embedded ducted rotor power unit (A-D) and the propulsion rotor power system (F) under the pitch attitude control A_z 、f B_z 、f C_z 、f D_z 、f F_z , the calculation formulas are:

[0039]

[0040] Furthermore, the pitch angle control method adopts cascade PID to complete the control of the pitch angle, and includes the following steps:

[0041] Step S1: The pitch angle outer loop (the pitch angle to pitch angle rate) adopts PD control, and the desired pitch angle θ is obtained according to the pitch angle control strategy. g As well as the current pitch angle θ and pitch rate ω z , calculate the desired pitch angle rate ω zgz ,

[0042] ω zgz =k p1 (θ g -θ)-k d1 × x ,

[0043] Among them, k p1 is the proportionality coefficient, k d1 Derivative control coefficient;

[0044] Step S2: According to the performance requirements of the aircraft, set the pitch rate limit value ω zmax , and the desired pitch angle rate ω calculated in step S1 zgz Limit the amplitude to get the expected value of the pitch angle rate ω zg ,

[0045]

[0046] Step S3: The pitch angle inner loop (the pitch angle rate to the pitch channel rotor motor throttle) adopts PID control, according to the pitch angle rate expected value ω zg 、Current angular velocity ω z and the current angular acceleration ε z , calculate the pitch channel rotor motor throttle signal, and according to the use requirements of the aircraft, it is necessary to limit the integral link of the pitch angle inner loop, and the set amplitude is d pitch_i_max ,

[0047]

[0048] Among them, k i2 Integral control coefficient;

[0049] Step S4: Based on step S3, calculate the pitch control signal value d pitch_z ,

[0050] d pitch_z =k p2(w zg -w z )-k d2 ×ε z +d pitch_i ,

[0051] Among them, k p2 is the proportionality coefficient, k d2 Derivative control coefficient;

[0052] Step S5: According to the use requirements of the aircraft, the total control signal of the pitch angle inner loop needs to be limited, and the amplitude is set to d pitch_max ,

[0053]

[0054] Step S6: According to the aircraft usage requirements, the pitch control efficiency coefficient (f A_z 、f B_z 、f C_z 、f D_z 、f F_z ), the total control signal d of the pitch angle inner loop pitch Converted into the throttle signal of each rotor motor in the pitch channel; the throttle signal of each rotor motor pitch channel is dth A_z 、dth B_z 、dth C_z 、dth D_z 、dth F_z ,

[0055]

[0056] Among them, ζ is the pitch channel control allocation coefficient, which is about 0.7 to 0.8, and f A_z 、f B_z 、f C_z 、f D_z 、f F_z They are respectively the rotor pitch control efficiency coefficients of the embedded ducted rotor power unit (A~D) and the propulsion rotor power system (F).

[0057] Furthermore, in the rolling attitude control, the following steps are also included:

[0058] Step S201: Calculate the rotor center z tx ,

[0059] In the roll attitude control, the projected areas of the rotors of the embedded ducted rotor power unit (A-D) and the propulsion rotor power system (F) on the horizontal plane are S and A , S B , SC , S D , S F , the positions on the z-axis are z A 、z B 、z C 、z D 、z F The embedded ducted rotor power system (A-D) and the propulsion rotor power system (F) are both arranged symmetrically, with the rotor center located on the symmetry plane, z tx =0;

[0060] Step S202: Calculate the roll operation efficiency coefficient f of each rotor of the embedded ducted rotor power system (A-D) and the propulsion rotor power system (F) A_x 、f B_x 、f C_x 、f D_x 、f F_x ,

[0061]

[0062] Furthermore, the roll angle control method uses a cascade PID to complete the control of the roll angle, and includes the following steps:

[0063] Step S1: The roll angle outer loop (the roll angle to roll angle rate) adopts PD control, and the desired roll angle γ obtained according to the roll angle control strategy g And the current roll angle γ and roll angle rate ω x , calculate the desired roll angular rate ω xgz ,

[0064] ω xgz =k p1 (γ g -γ)-k d1 × x ,

[0065] Among them, k p1 is the proportionality coefficient, k d1 Derivative control coefficient;

[0066] Step S2: According to the performance requirements of the aircraft, set the roll angle rate limit value ω xmax , and the roll angular rate ω calculated in step S1 xgz Limit the amplitude to get the expected value of the roll angular rate ω xg ,

[0067]

[0068] Step S3: The inner loop of the roll angle (the roll angle rate to the roll channel rotor motor throttle) adopts PID control, according to the expected roll angle rate value ω xg and the current angular rate ω x and the current angular acceleration ε x , calculate the roll channel rotor motor throttle signal dth x According to the use requirements of the aircraft, it is necessary to limit the integral link of the inner loop of the roll angle, and the set amplitude is d roll_i_max ,

[0069]

[0070] Among them, k i2 Integral control coefficient;

[0071] Step S4: Based on step S3, calculate the roll control signal value d roll_z ,

[0072] d roll_z =k p2 (w xg -w x )-k d2 ×ε x +d roll_i ,

[0073] Among them, k p2 is the proportionality coefficient, k d2 Derivative control coefficient;

[0074] Step S5: According to the use requirements of the aircraft, the total control signal of the inner loop of the roll angle needs to be limited, and the amplitude is set to d roll_max ,

[0075]

[0076] Step S6: According to the aircraft use requirements and the roll control efficiency coefficients of each rotor of the aircraft, the total control signal d of the inner loop of the roll angle is converted into roll Converted into the throttle signal of each rotor motor in the roll channel. The throttle signal of each rotor motor roll channel is dth A_x 、dth B_x 、dth C_x 、dth D_x 、dth F_x ,

[0077]

[0078] Where ζ is the roll channel control allocation coefficient, which is about 0.7 to 0.8, and f A_x 、f B_x 、fC_x 、f D_x 、f F_x are the rotor roll control efficiency coefficients of the embedded ducted rotor power unit (A-D) and the propulsion rotor power system (F);

[0079] Step S7: In order to ensure the flight safety of the aircraft's rolling maneuver, it is necessary to perform altitude compensation on the altitude control channel. The altitude compensation amount H of the rolling maneuver in the ground effect cruise mode is g With the desired roll angle γ g Related,

[0080] H g = k × L z ×γ g ,

[0081] Among them, L z The spanwise distance from the wingtip buoy of the ducted ground effect wing to the symmetry plane is k, a coefficient, and k is 0.9 to 1.1.

[0082] The beneficial effects of the present invention are:

[0083] 1. The vertical take-off and landing ground effect vehicle of the present invention adopts a vertical take-off / landing mode. Compared with traditional ground effect vehicles, it has no taxiing process, greatly improves the riding comfort of the passengers during the take-off and landing stages, and can fly seamlessly from flat ground to shallow water to water surface.

[0084] 2. The vertical take-off and landing ground-effect vehicle of the present invention adopts automatic control to assist driving, which reduces the risk of wingtips colliding with the water surface / ground during lateral maneuvers such as turning, thereby improving safety. At the same time, it has the ability of powered jump, which greatly improves flight safety; it can directly take off and land on buildings and islands without docks, greatly expanding the application field. BRIEF DESCRIPTION OF THE DRAWINGS

[0085] Figure 1 is a three-dimensional structural diagram of the aircraft of the present invention;

[0086] Figure 2 for Figure 1 Front view of

[0087] Figure 3 for Figure 2 Deformation state diagram of middle and outer wings;

[0088] Figure 4 for Figure 1 Side view of

[0089] Figure 5 for Figure 1 A top view of

[0090] Figure 6This is a structural diagram of an embodiment of a single vertical tail plus horizontal tail structure of the present invention;

[0091] Figure 7 This is a structural diagram of an embodiment of a double vertical tail plus horizontal tail structure of the present invention;

[0092] Figure 8 This is a structural diagram of an embodiment of a six-rotor layout of the present invention;

[0093] Fig. 9 The control method of the present invention;

[0094] Fig.10 is the equilibrium pitch angle variation curve of the aircraft as the speed changes;

[0095] Fig.11 The pitch control principle diagram of the aircraft of the present invention;

[0096] Fig.12 It is a schematic diagram of the lateral control related parameters of the aircraft of the present invention;

[0097] Fig.13 It is a schematic diagram of the distribution of forces and moments in the ground effect cruise mode of the aircraft of the present invention;

[0098] The markings in the figure are:

[0099] 1. Fuselage, 101. Door, 2. Ducted ground effect wing, 3. Outer wing, 301. Wing bar, 4. Vertical tail, 5. Tail pontoon, 6. Float, 7. Rotor power installation bracket, 8. Strut, 9. Horizontal tail;

[0100] A~D, embedded ducted rotor power unit, F, propulsion rotor power unit. DETAILED DESCRIPTION

[0101] In order to make the above contents, purposes and beneficial effects of the present invention more obvious and understandable, the specific embodiments of the present invention are described in detail below in conjunction with the accompanying drawings. In the following description, many specific details are set forth to facilitate a full understanding of the present invention. However, the present invention can be implemented in many other ways different from those described herein, and those skilled in the art can make similar improvements without violating the connotation of the present invention, so the present invention is not limited by the specific embodiments disclosed below.

[0102] It should be noted that, unless otherwise defined, all technical and scientific terms used herein have the same meaning as those generally understood by those skilled in the art of the present invention. The terms used herein in the specification of the present invention are only for the purpose of describing specific embodiments and are not intended to limit the present invention. The term "and / or" used herein includes any and all combinations of one or more related listed items.

[0103] like Figure 1-13 As shown, the present invention provides a vertical take-off and landing ground effect vehicle and a control method thereof.

[0104] like Figure 1 As shown, the present invention provides a vertical take-off and landing ground effect vehicle, including a fuselage 1, a ducted ground effect wing 2, an outer wing 3, a vertical tail 4, a tail fin buoy 5, a buoy 6, a rotor power mounting bracket 7, an embedded ducted rotor system and a propulsion rotor power system.

[0105] Among them, the ducted ground effect wing 2 is installed on the upper side of the fuselage 1 symmetrically. Floats 6 are arranged and installed at the left and right wing tips of the ducted ground effect wing 2 to form left and right floats. The outer wing 3 is arranged and installed on the outer sides of the left and right floats through a hinge assembly, tilted outward, to form left and right outer wings. The vertical tail 4 is arranged and installed on the upper side of the rear part of the ducted ground effect wing 2 through the tail fin buoy 5, or directly arranged and installed on the upper side of the tail of the fuselage 1. The embedded ducted rotor system is embedded and installed in the ducted ground effect wing 2, and the propulsion rotor power system is fixedly connected to the vertical tail 4 through the strut 8, fixedly connected between the double vertical tails, or fixedly connected to both sides of the vertical tail.

[0106] Further, such as Figure 1 and Figure 2 As shown, the fuselage 1 is a ship-shaped streamlined fuselage, but is not limited to a streamlined fuselage. An openable door 101 is installed at the upper front end of the fuselage 1. The interior of the fuselage 1 is used to arrange airborne electronic equipment, as well as other equipment or loads such as passengers or cargo. Figure 2 As shown, the ducted ground effect wing 2 is fixedly connected to the middle upper side of the boat-shaped fuselage 1, and the bottom shape of the boat-shaped fuselage 1 adopts a boat-shaped design. The boat-shaped design can meet the stability of the attitude during high-speed water entry and water skiing stages when the aircraft makes an emergency landing.

[0107] Furthermore, the ducted ground effect wing 2 is bilaterally symmetrical and fixed to the upper side of the middle and rear section of the fuselage 1. Figure 2 As shown, in one embodiment, the ducted ground effect wing 2 adopts a vaulted wing structure.

[0108] Furthermore, an embedded ducted rotor system is symmetrically arranged and installed on the ducted ground effect wing 2. Figure 1 As shown, in one embodiment, the embedded ducted rotor system is composed of four embedded ducted rotor power devices A to D. Figure 8 As shown, in other embodiments, the embedded ducted rotor system can also be composed of six embedded ducted rotor power devices. Of course, in other embodiments, it can also be a rotor system composed of eight embedded ducted rotor power devices in an even number.

[0109] Further, such as Figure 1 and Figure 2As shown, the ducted ground effect wing 2 is symmetrically arranged with a plurality of ducts in a regular pattern. A plurality of rotor power devices are embedded in a plurality of ducts on the ducted ground effect wing 2 through a rotor power mounting bracket 7, thereby forming embedded ducted rotor power devices A to D. A plurality of embedded ducted rotor power devices together constitute the embedded ducted rotor power system of the present invention.

[0110] The ducted ground effect wing 2 of the present invention can effectively compress the airflow flowing from the front to the back on the lower surface of the ducted ground effect wing in the ground effect cruise stage, greatly improve the ground effect of the ducted ground effect wing, reduce the flight resistance, and improve the flight efficiency. Combined with the slipstream of the embedded ducted rotor system, the ground effect of the ducted ground effect wing 2 is further increased, the flight resistance is further reduced, and the flight efficiency is increased. At the same time, the ground effect of the ducted ground effect wing 2 also increases the efficiency of the embedded ducted rotor system and improves the power efficiency of the embedded ducted rotor system.

[0111] Further, such as Figure 1 As shown, in one embodiment, four embedded ducted rotor power devices A to D are symmetrically arranged on the ducted ground effect wing 2. The embedded ducted rotor power devices A and B are symmetrically arranged on the fuselage 1 and are regularly arranged at the front section of the ducted ground effect wing 2; the embedded ducted rotor power devices C and D are also symmetrically arranged on the fuselage 1 and are regularly arranged at the rear end of the ducted ground effect wing 2.

[0112] The embedded ducted rotor power device includes a motor, an electric regulator and a propeller, which are installed on a rotor power mounting bracket 7 and embedded in the ducted ground effect wing 2 through the rotor power mounting bracket 7. The embedded ducted rotor power device can adopt a single rotor structure or a coaxial double rotor structure.

[0113] Preferably, the rotors of several embedded ducted rotor power units are arranged symmetrically.

[0114] Furthermore, the propeller disc areas of the embedded ducted rotor power units arranged symmetrically on the left and right sides of the fuselage 1 are the same, and the propeller disc areas of the embedded ducted rotor power units arranged front and rear can be the same or different. Preferably, the propeller disc area of ​​the embedded ducted rotor power unit at the front section of the ducted ground effect wing 2 is larger than the propeller disc area of ​​the embedded ducted rotor power unit at the rear section thereof.

[0115] Further, such as Figure 2 As shown, the embedded ducted rotor power units A~D and the corresponding ducts are arranged on the ducted ground effect wing 2 at a certain camber angle β.

[0116] The ducted ground effect wing 2 and the embedded ducted rotor system (A, B, C, D) are designed to integrate the lift generated by the wing and the pulling force of the embedded ducted rotor system (A, B, C, D) in the forward flight stage to meet the main lift requirements of the entire aircraft. In the ground effect cruise stage of the aircraft using the ducted ground effect wing of the present invention, the aerodynamic drag of the aircraft is reduced by 15%, the flight efficiency is greatly improved, and the power efficiency of the embedded ducted rotor system is improved by 10%.

[0117] Further, the buoys 6 are regularly arranged and installed at the left and right wing tips of the ducted ground effect wing 2 to form left and right buoys. The tail buoys 5 are regularly arranged and installed at the rear of the ducted ground effect wing, and the tail buoys are also symmetrically installed at the rear of the ducted ground effect wing 2 to form left and right tail buoys.

[0118] Further, the outer wings 3 are arranged symmetrically on the left and right sides of the fuselage 1 and tilted outward. The outer wings 3 are fixedly mounted on the outer sides of the left and right pontoons 6 to form left and right outer wings. Alternatively, in another embodiment, the outer wings 3 are rotatably mounted on the outer sides of the left and right pontoons 6 by hinge components to form foldable left and right outer wings. The outer wings 3 can be rotated around the axis on the outer sides of the pontoons 6 by hinge components to be folded, such as Figure 2 and Figure 3 As shown, when the aircraft is stopped, the outer wing 3 can be rotated vertically upward around the axis of the hinge assembly, thereby reducing the footprint of the aircraft.

[0119] Furthermore, the outer wing 3 is arranged at a larger dihedral angle in normal working state to reduce the risk of the outer wing 3 being inserted into the water during high-speed water entry and hydroplaning phases when the aircraft makes an emergency landing. At the same time, the addition of the outer wing 3 can increase the lift and flight efficiency of the entire aircraft and improve the lateral stability.

[0120] Further, such as Figure 5 As shown, a wing slat 301 is added to the root of the outer wing 3, but it is not limited to adding a wing slat. The benefit of adding the wing slat 301 is that in maneuvering flight at a large angle of attack, the airflow will roll up a stable wing vortex from the leading edge of the slat, and under its favorable interference, the flow separation on the wing surface of the outer wing 3 can be greatly improved, the lateral stability in maneuvering flight is guaranteed, and the resistance of maneuvering flight is reduced.

[0121] like Figure 1 As shown, it is a structural diagram of an embodiment of a double vertical tail. In this embodiment, the vertical tail 4 includes a left vertical tail. The left vertical tail 4 is arranged symmetrically on the left and right sides of the fuselage 1, and the left vertical tail 4 is respectively fixed to the upper side of the left and right tail buoys 5. The main purpose of the vertical tail 4 is to meet the heading stability.

[0122] Further, in the embodiment of the double vertical tail structure, the tail pontoon 5 includes left and right tail pontoons, which are arranged symmetrically on the left and right sides of the fuselage and fixedly connected to the rear of the ducted ground effect wing 2. The upper end of the tail pontoon 5 is fixedly connected to the vertical tail 4, and its bottom shape adopts a ship-shaped design. The tail pontoon with a ship-shaped design can prevent the tail from entering the water too deeply during emergency landing of the aircraft, while meeting the stability of the posture during high-speed water entry and water skiing stages.

[0123] Further, in this embodiment, the propulsion rotor power device F is arranged obliquely between the left and right vertical tails through the strut 8. The strut 8 connects the wing surface between the left and right vertical tails, and the propulsion rotor power device F is arranged and installed at the middle position of the strut 8.

[0124] like Figure 6 As shown, it is a structural diagram of an embodiment of a single vertical tail plus a horizontal tail. In this embodiment, the vertical tail 4 adopts a single vertical tail structure and is arranged and installed on the upper side of the tail of the fuselage 1. A horizontal tail 8 is fixedly connected to the top wing tip thereof, and the horizontal tail 8 is fixedly connected to the top of the vertical tail 4 symmetrically.

[0125] Furthermore, in the embodiment of the single vertical tail plus horizontal tail structure, a propulsion rotor power device F is symmetrically installed on the vertical tail 4 through the strut 9.

[0126] like Figure 7 The above is a structural diagram of an embodiment of a double vertical tail plus a horizontal tail. In this embodiment, the vertical tail 4 includes left and right vertical tails, and its vertical tail structure, tail pontoon structure, and propulsion rotor power device F are similar to the structures of the double vertical tail outward camber embodiment, and horizontal tails 9 are regularly arranged and installed at the top wing tips of the left and right vertical tails. The horizontal tail 9 is connected to the wing tips of the left and right vertical tails.

[0127] Furthermore, in the double vertical tail structure, the left and right vertical tails 4 are vertically arranged and mounted on the left and right tail pontoons 5 , or are arranged and mounted on the left and right tail pontoons 5 in an outwardly inclined manner.

[0128] like Figure 1 and Figure 2 As shown, the buoy 6 is fixedly connected to the wing tip of the ducted ground effect wing 2, and the buoy 6 also adopts a boat-shaped streamlined shape, but is not limited to the streamlined shape. The bottom shape of the buoy 6 adopts a boat-shaped design, so that when the aircraft makes an emergency landing, the stability of the posture during the high-speed water entry and water skiing stages is met.

[0129] Furthermore, several rotor power devices are installed and fixed in the duct of the ducted ground effect wing 2 through the rotor power mounting bracket 7, thereby forming an embedded ducted rotor power system. The rotor power mounting bracket 7 is symmetrically installed and fixed to the lower part of the ducted ground effect wing 2, forming left and right rotor power mounting brackets. Figure 1 and Figure 2As shown, the embedded ducted rotor power units A and D are arranged front and back and installed on the upper end of the right rotor power mounting bracket, and the embedded ducted rotor power units B and C are arranged front and back and installed on the upper end of the left rotor power mounting bracket. The rotor power mounting brackets on the left and right sides can increase the rigidity of the ducted ground effect wing 2.

[0130] Furthermore, the rotor power mounting bracket 7 can also be installed and fixed on the upper part of the ducted ground effect wing 2, and correspondingly, the embedded ducted rotor power device is arranged and installed on the lower end of the rotor power mounting bracket 7.

[0131] like Figure 5 The figure shows the rotor rotation direction of four embedded ducted rotor power units A to D in an embodiment. Rotor power unit A rotates clockwise, rotor power unit B rotates counterclockwise, rotor power unit C rotates clockwise, and rotor power unit D rotates counterclockwise. The functions of the embedded ducted rotor system (A, B, C, D) mainly meet the lift requirements of the aircraft in the hovering stage, partial lift in the cruising stage, and attitude control requirements of the aircraft in each flight stage.

[0132] like Figure 1 As shown, the propulsion rotor power device F is installed on the vertical tail 4 through a strut 8, and the propulsion rotor power device includes a motor, an electric regulator and a propeller.

[0133] Further, such as Figure 4 As shown, the propulsion rotor power unit F can be installed in a forward tilted arrangement at a fixed tilt angle α, or can be flexibly arranged and installed on the vertical tail 4 through a tilt mechanism, so that the tilt angle of the propulsion rotor power unit F is adjusted and controlled by the tilt mechanism.

[0134] Furthermore, the propulsion rotor power device can be a single-propeller rotor or a multi-propeller rotor. Figure 1 As shown, the propulsion rotor power unit F is arranged and installed between the left and right vertical tails through the support rod 8. When the vertical tail adopts a single vertical tail structure, as shown in FIG. Figure 6 As shown, the propulsion rotor power device F is arranged and installed on the left and right sides of the vertical tail through the support rod 8, forming a left and right propulsion rotor power device. The left and right rotor propulsion power devices are arranged symmetrically on the left and right sides of the fuselage.

[0135] like Figure 1As shown, in this embodiment, the propulsion rotor power unit F adopts a fixed tilt coaxial twin-propeller form, and the forward tilt angle is 25° to 35°. Among them, the upper propeller of the propulsion rotor power unit F rotates counterclockwise around the axis, and the lower propeller rotates clockwise around the axis. In the hovering mode, the propulsion rotor power system F operates at a relatively small power, providing a relatively small lift for the whole aircraft in the hovering mode; in the ground effect cruise mode, the propulsion rotor power system F operates at a relatively large power, providing partial lift for the whole aircraft in the ground effect cruise mode, and providing most of the forward flight thrust for the whole aircraft.

[0136] like Fig. 9 As shown, the present invention also provides a control method for a vertical take-off and landing ground effect vehicle, including a longitudinal attitude control method 210, a lateral attitude control method 220, a water surface movement control method 230, and an emergency power jump control method 240.

[0137] Furthermore, in the longitudinal attitude control method 210, it is necessary to satisfy the pitch attitude control in the hovering mode and the ground effect cruise mode of the vertical take-off and landing ground effect vehicle.

[0138] In hover mode, the aerodynamic force of the VTOL ground effect vehicle is very small, and the pitch moment of the aerodynamic force is very small. The pitch moment of the whole aircraft is mainly generated by the embedded ducted rotor power system. In ground effect cruise mode, the aerodynamic force of the whole aircraft is already very large, and the lift reaches 40-60% of the total weight of the aircraft. The pitch moment of the aerodynamic force is usually a large nose-up moment. The distribution of longitudinal force and moment in ground effect cruise mode is shown in the figure below. Fig.13 shown.

[0139] In order to meet the pitch attitude control requirements of the hovering mode, meet the pitch attitude control and forward flight thrust requirements in the ground effect cruise mode, while ensuring that the total power in the cruise forward flight mode is low and the flight time is long, and the pitch attitude change from the hovering mode to the ground effect cruise mode is small.

[0140] The present invention adopts a combined rotor system of an embedded ducted rotor system (A, B, C, D) and a propulsion rotor power system (F), and reasonably configures the center of gravity, aerodynamic focus, and rotor center of the aircraft to meet the different requirements of the hovering mode and the ground effect cruise mode.

[0141] Furthermore, the center of gravity of the aircraft is arranged near the aerodynamic focus of the ground effect cruise mode, ensuring that the longitudinal aerodynamic characteristics of the aircraft in the ground effect cruise mode are neutrally stable or slightly statically stable. The center of gravity of the aircraft is located a short distance in front of the rotor center of the combined rotor system, ensuring that in the cruise mode, each rotor of the combined rotor system operates in its own optimal design state and the total power of the aircraft is near the lowest state. The vertical component of the resultant force at the center of the combined rotor and the forward flight aerodynamic lift jointly meet the gravity balance requirements. The resultant force at the center of the combined rotor balances most of the nose-up and pitching moments of the aerodynamic force during forward flight, and the forward pull of the inclined propulsion rotor power system (F) balances the forward flight aerodynamic resistance of the aircraft.

[0142] Furthermore, the center of gravity of the whole aircraft is located 0 to 0.05 times the average aerodynamic chord length C in front of the aerodynamic focus. A In the ground effect cruise mode, the vertical take-off and landing ground effect vehicle is in a small static stability or neutral stability in the longitudinal direction. The rotor center is located 0.02 to 0.08 times the average aerodynamic chord length C behind the center of gravity of the whole aircraft. A The resultant force at the center of the rotor on the center of gravity produces a nose-down moment that offsets the aerodynamic nose-up moment in the ground effect cruise mode.

[0143] Since the center of gravity of the aircraft is located a short distance in front of the rotor center of the combined rotor system, in the hovering mode, the power of the front rotor needs to be increased, while the power demand of the propulsion rotor power system (F) is reduced to balance the pitch moment of the whole aircraft. In the hovering mode, the power of the embedded ducted rotor system (A, B) at the front increases by 5% to 8%, while the power of the propulsion rotor power system (F) decreases by 5% to 7%, which is within an acceptable range.

[0144] During the transition from the hovering mode to the ground effect cruise mode, the transition flight can be achieved by a small change in the forward flight pitch angle, and the passengers' riding comfort is relatively good. Fig.10 As shown, in an embodiment of a propulsion rotor power system (F) that can be tilted, the pitch angle change during the transition from 0 to 120 km / h is within 2°, and in an embodiment of a propulsion rotor power system (F) with a fixed tilt angle, the pitch angle change during the transition is within 5°, while when a traditional multi-rotor solution is used, the pitch angle change during the transition reaches 10°.

[0145] The present invention also provides a method for calculating the rotor center and each rotor pitch control efficiency coefficient, comprising the following steps:

[0146] Step 1: In this embodiment, a combined rotor system of an embedded ducted rotor system (A, B, C, D) and a propulsion rotor power system (F) is used. The projected areas of the five rotors on the horizontal plane are S A , S B, S C , S D , S F , their positions on the x-axis are x A 、x B 、x C 、x D 、x F , rotor center x tx The calculation formula (1) is as follows:

[0147]

[0148] Step 2: The pitch control efficiency coefficients of each rotor are f A_z 、f B_z 、f C_z 、f D_z 、f F_z , the calculation formula is as follows:

[0149]

[0150] The method for controlling the pitch angle of an aircraft in the present invention uses a cascade PID to control the pitch angle. The control process is as follows: Fig.11 Said, comprising the following steps:

[0151] Step 1: The outer loop of the pitch angle (pitch angle to pitch angle rate) adopts PD control, and the desired pitch angle θ is obtained according to the pitch angle control strategy g As well as the current pitch angle θ and pitch rate ω z Calculate the desired pitch rate ω zgz , as shown in formula (3), where k p1 is the proportionality coefficient, k d1 Derivative control coefficient;

[0152] ω zgz =k p1 (θ g -θ)-k d1 × z (3)

[0153] Step 2: Set the pitch rate limit value ω according to the aircraft performance requirements zmax , and the ω calculated in step 1 zgz Limit the amplitude to get the expected value of the pitch angle rate ω zg , as shown in formula (4);

[0154]

[0155] Step 3: The inner loop of the pitch angle (pitch angle rate to pitch channel rotor motor throttle) adopts PID control, and then according to the desired pitch angle rate ω zgand the current angular rate ω z and the current angular acceleration ε z Calculate the pitch channel rotor motor throttle signal. According to the aircraft's use requirements, it is necessary to limit the integral link of the pitch angle inner loop, and the set amplitude is d pitch_i_max , limiting formula (5), where k i2 Integral control coefficient;

[0156]

[0157] Step 4: Based on step 3, calculate the pitch control signal value d pitch_z , as shown in formula (6), where k p2 is the proportionality coefficient, k d2 Derivative control coefficient;

[0158] d pitch_z =k p2 (w zg -w z )-k d2 ×ε z +d pitch_i (6)

[0159] Step 5: According to the use requirements of the aircraft, the total control signal of the pitch angle inner loop needs to be limited, and the set amplitude is d pitch_max , limiting formula (7), the amplitude is usually 100;

[0160]

[0161] Step 6: According to the aircraft usage requirements and the pitch control efficiency coefficients of each rotor of the aircraft, the total control signal d pitch Converted into the throttle signal of each rotor motor in the pitch channel. In the embodiment, the throttle signal of each rotor motor pitch channel is dth A_z 、dth B_z 、dth C_z 、dth D_z 、dth F_z , calculated by formula (8). Where ζ is the pitch channel control allocation coefficient, which is about 0.7 to 0.8, and f A_z 、f B_z 、f C_z 、f D_z 、f F_z are the pitch control efficiency coefficients of each rotor respectively.

[0162]

[0163] exist Fig.11Among them, electronic control A~electronic control F are the electronic controls of embedded ducted rotor power units A~D and propulsion rotor power unit F respectively.

[0164] The lateral attitude control method 220 in the control method of the present invention requires the vertical take-off and landing ground effect vehicle to meet the rolling attitude control 221 and heading turn control 222 in the ground effect cruise mode. In the ground effect cruise mode, the aircraft is very low to the water surface or the ground, and the ground effect height is about 1 to 3 meters. When the aircraft is performing lateral maneuvers, the flight safety of the aircraft must be ensured, that is, the distance between the wingtip buoy 6 of the ducted ground effect wing 2 and the water surface or the ground is greater than the safety height H. a . Lateral control related parameters such as Fig.12 shown.

[0165] The roll attitude control method 221 can increase the flight altitude of the aircraft through the height compensation command related to the roll command angle, thereby ensuring that the distance between the wingtip buoy 6 of the ducted ground effect wing 2 and the water surface or the ground is greater than the safe altitude. g The relevant height supplementary instruction is H g In the roll control of the ground effect cruise mode of the vertical take-off and landing ground effect vehicle, the altitude supplementary instruction H g Compensate to the altitude control channel to ensure that the aircraft's flight altitude can be effectively increased during the roll control process of the ground effect cruise phase, ensuring that the lowest point of the aircraft is higher than the safe altitude H a , ensuring the flight safety of the aircraft.

[0166] The rolling attitude control method 222 can also measure the change in the distance between the buoy 6 and the ground or water surface during the rolling process by using a sensor installed on the surface of the buoy 6 to measure the distance from the ground or water surface, and use this to compensate the altitude control channel of the aircraft. In the embodiment, the sensor installed on the buoy 6 measures the distance H1 between the buoy 6 and the water surface / ground, and compares it with the safety distance H a Make the difference and get the height compensation command H g =H a -H1. Add the height to the command H g Compensate to the altitude control channel to ensure that the aircraft's flight altitude can be effectively increased during the roll control process of the ground effect cruise phase, ensuring that the lowest point of the aircraft is higher than the safe altitude H a , ensuring the flight safety of the aircraft.

[0167] The heading turn control 222 can establish a desired roll angle through the roll attitude control method 221, so that the aircraft generates a suitable centripetal force (the component of aerodynamic lift and rotor pull on the horizontal plane) in the horizontal plane of motion, thereby achieving a yaw turn of the aircraft.

[0168] The present invention also provides a method for calculating the roll control efficiency coefficient of each rotor, comprising the following steps:

[0169] Step 1: In this embodiment, an embedded ducted rotor system (A, B, C, D) and a propulsion rotor power system (F) are used. The projected areas of the five rotors on the horizontal plane are S A , S B , S C , S D , S F , the positions on the z-axis are z A 、z B 、z C 、z D 、z F Since the embedded ducted rotor system (A, B, C, D) and the propulsion rotor power system (F) are all arranged symmetrically, the rotor center is located on the symmetry plane, z tx =0.

[0170] Step 2: The roll control efficiency coefficients of each rotor are f A_x 、f B_x 、f C_x 、f D_x 、f F_x , the calculation formulas are as follows: formula (9).

[0171]

[0172] The roll angle control method of the present invention uses cascade PID to complete the roll angle control, which is achieved by the following steps:

[0173] Step 1: The outer loop of the roll angle (roll angle to roll angle rate) adopts PD control, and the desired roll angle γ is obtained according to the roll angle control strategy g And the current roll angle γ and roll angle rate ω x Calculate the desired roll rate ω xgz , as shown in formula (10), where k p1 is the proportionality coefficient, k d1 Derivative control coefficient;

[0174] ω xgz =k p1 (γ g -γ)-k d1 × x (10)

[0175] Step 2: Set the roll rate limit value ω according to the aircraft performance requirements xmax , and the ω calculated in step 1 xgz Limit the amplitude to get the expected value of the roll angular rate ω xg, as shown in formula (11);

[0176]

[0177] Step 3: The inner loop of the roll angle (roll angle rate to the roll channel rotor motor throttle) adopts PID control, and then according to the desired roll angle rate ω xg and the current angular rate ω x and the current angular acceleration ε x Calculate the roll channel rotor motor throttle signal dth x ; According to the use requirements of the aircraft, it is necessary to limit the integral link of the inner loop of the roll angle, and the set amplitude is d roll_i_max , limiting formula (12), where k i2 Integral control coefficient;

[0178]

[0179] Step 4: Based on step 3, calculate the roll control signal value d roll_z , as shown in formula (13), where k p2 is the proportionality coefficient, k d2 Derivative control coefficient;

[0180] d roll_z =k p2 (w xg -w x )-k d2 ×ε x +d roll_i (13)

[0181] Step 5: According to the use requirements of the aircraft, the total control signal of the inner loop of the roll angle needs to be limited, and the set amplitude is d roll_max , limiting formula (14), the amplitude is usually 100;

[0182]

[0183] Step 6: According to the aircraft usage requirements and the roll control efficiency coefficients of each rotor of the aircraft, the total control signal d roll Converted into throttle signals of each rotor motor in the roll channel. In the embodiment, the throttle signals of each rotor motor roll channel are dth A_x 、dth B_x 、dth C_x 、dth D_x 、dth F_x , calculated by formula (15). Where ζ is the roll channel control allocation coefficient, which is about 0.7 to 0.8, and f A_x 、f B_x 、fC_x 、f D_x 、f F_x are the roll control efficiency coefficients of each rotor respectively.

[0184]

[0185] Step 7: In order to ensure the flight safety of the aircraft's rolling maneuver, it is necessary to compensate the altitude control channel. The altitude compensation amount H for the rolling maneuver in the ground effect cruise mode g With the desired roll angle γ g Related, can be calculated by formula (17), where L z k is a coefficient, which is 0.9 to 1.1 in this embodiment, and is the spanwise distance from the wingtip of the ducted ground effect wing (2) to the symmetry plane.

[0186] h g = k × L z ×γ g (17)

[0187] The motion control method 230 on the water surface of the present invention needs to meet the heading control and forward motion control of the vertical take-off and landing ground effect vehicle on the water surface. The heading control of the vertical take-off and landing ground effect vehicle on the water surface is not achieved by traditional rudders, but by the differential speed of the rotors of different rotation directions of the embedded ducted rotor system (increasing the rotor speeds of all counterclockwise rotations and reducing the rotor speeds of all clockwise rotations to achieve the clockwise rotation of the vertical take-off and landing ground effect vehicle; or reducing the rotor speeds of all counterclockwise rotations and increasing the rotor speeds of all clockwise rotations to achieve the counterclockwise rotation of the vertical take-off and landing ground effect vehicle), combined with the camber angle of the embedded ducted rotor system, to increase the torque of heading control.

[0188] In an embodiment, the rotation speed of the rotors of the embedded ducted rotor systems (A, C) can be increased, while the rotation speed of the rotors of the embedded ducted rotor systems (B, D) can be reduced, so as to achieve a counterclockwise turn of the vertical take-off and landing ground effect vehicle. The rotation speed of the rotors of the embedded ducted rotor systems (A, C) can be reduced, while the rotation speed of the rotors of the embedded ducted rotor systems (B, D) can be increased, so as to achieve a clockwise turn of the vertical take-off and landing ground effect vehicle. The camber angle β of the embedded ducted rotor systems (A, B, C, D) is 5°, and the steering torque of the heading is increased by about 1 times.

[0189] The water surface heading control of the water surface motion control method 230 can also be achieved by using a dual-axis dual-propeller scheme ( Figure 6The heading control of the vertical take-off and landing ground effect vehicle when moving on the water surface can be achieved by the differential rotation speed of the left propulsion rotor and the right propulsion rotor, such as making the left propulsion rotor rotation speed higher than the right propulsion rotor rotation speed to achieve a right turn, and making the left propulsion rotor rotation speed lower than the right propulsion rotor rotation speed to achieve a right turn.

[0190] The forward motion control of the motion control method 230 on the water surface is realized by the propulsion rotor power system (F). The forward rotation of the propulsion rotor power system (F) realizes the forward motion of the vertical take-off and landing ground effect vehicle, and the reverse rotation of the rotor of the propulsion rotor power system (F) realizes the backward motion of the vertical take-off and landing ground effect vehicle.

[0191] The emergency power jump control method 240 of the present invention needs to meet the requirements that the vertical take-off and landing ground effect vehicle can quickly increase its flight altitude and cross the obstacle when encountering special operating environments such as obstacles.

[0192] When a traditional ground effect vehicle encounters an obstacle, in order to increase its flight altitude, it is necessary to increase the throttle of the power system and manipulate the elevator to increase the angle of attack of the vehicle, thereby increasing the lift of the entire aircraft and thus increasing the flight altitude. As the ground effect altitude increases, the ground effect vehicle gradually leaves the ground effect area, and the lift of the entire aircraft tends to decrease. Therefore, the emergency jump capability of the traditional ground effect vehicle is relatively weak.

[0193] When the vertical take-off and landing ground effect vehicle of the present invention encounters an obstacle, the throttle of the embedded ducted rotor system can be directly increased, and the flight angle of attack can be increased at the same time, which can not only increase the aerodynamic lift of the entire aircraft, but also increase the dynamic lift of the embedded ducted rotor system at the same time, so that the total lift of the entire aircraft is greatly improved, the jumping maneuverability of the vertical take-off and landing ground effect vehicle is greatly improved, and the obstacle crossing capability of the aircraft is improved.

[0194] When the flight speed of a traditional ground effect vehicle is lower than the ground effect cruise speed, the lift of the whole aircraft gradually decreases. When the speed is lower than the stall speed, the total aerodynamic lift of the whole aircraft is less than the gravity, and the aircraft does not have any obstacle-crossing capability. However, a vertical take-off and landing ground effect vehicle has the ability to jump with emergency power at any transition flight speed from hovering mode to ground effect cruise mode, which greatly increases the safety of the transition flight phase. Therefore, the deployment flexibility of vertical take-off and landing ground effect vehicles is greatly increased, and vertical take-off and landing ground effect vehicles can also be operated in special application scenarios such as cliff islands where there are no conditions for building docks.

[0195] The specific embodiments described above further illustrate the objectives, technical solutions and beneficial effects of the present invention in detail. It should be understood that the above description is only a specific embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims

1. A vertical take-off and landing ground effect vehicle, characterized in that: include: Fuselage, ducted ground effect wings, outer wings, vertical tail, floats, embedded ducted rotor system and propulsion rotor power system; The interior of the fuselage is used to arrange airborne electronic equipment, and other equipment or loads such as passengers or cargo; The ducted ground effect wing is symmetrically arranged on the left and right and fixedly connected to the upper side of the middle and rear section of the fuselage, and a plurality of ducts are symmetrically and regularly arranged on the ducted ground effect wing; The buoys are arranged symmetrically on the left and right sides of the fuselage and are installed at the left and right wing tips of the ducted ground effect wing to form left and right buoys; The outer wings are arranged symmetrically on the left and right sides of the fuselage and tilted outwards, and are installed on the outer sides of the left and right floats to form left and right outer wings; or, the outer wings are rotatably installed on the outer sides of the left and right floats through hinge components to form foldable left and right outer wings; The vertical tail is arranged symmetrically on the left and right sides of the fuselage to form left and right vertical tails, and the left and right vertical tails are arranged and installed on the upper side of the rear part of the ducted ground effect wing through tail fin buoys, and the tail fin buoys are arranged symmetrically on the left and right sides of the fuselage and installed on the rear part of the ducted ground effect wing; the propulsion rotor power system is arranged tilted forward through a strut and installed between the left and right vertical tails; Alternatively, the vertical tail is directly arranged and installed on the upper side of the tail of the fuselage, and the propulsion rotor power system is arranged forwardly tilted through the strut and installed on the left and right sides of the vertical tail to form a left and right propulsion rotor power system; The embedded ducted rotor power system includes a plurality of rotor power devices, which are embedded in a plurality of ducts of the ducted ground effect wing through a rotor power mounting bracket to form a plurality of embedded ducted rotor power devices; the rotor power mounting bracket is fixedly connected to the lower part or the upper part of the ducted ground effect wing, and the plurality of rotor power devices are mounted on the rotor power mounting bracket and embedded in the duct; The plurality of embedded ducted rotor power units are arranged to be tilted outwards, and the angle at which the embedded ducted rotor power units are tilted outwards is 5 to 10 degrees.

2. A vertical take-off and landing ground effect vehicle according to claim 1, characterized in that: The ducted ground effect wing adopts an arched wing structure, the left and right pontoons are symmetrically and horizontally arranged through the wing tips of the ducted ground effect wing, and the tail pontoon is also symmetrically and horizontally arranged and installed at the rear of the ducted ground effect wing.

3. A vertical take-off and landing ground effect vehicle according to claim 1, characterized in that: The propulsion rotor power system is arranged to be tilted forward through a strut and fixedly installed between the left and right vertical tails or on the left and right sides of the vertical tail, and the forward tilt angle of the propulsion rotor power system is 25°~35°; or it is arranged between the left and right vertical tails or on the left and right sides of the vertical tail through a strut and a tilt mechanism, and the propulsion rotor power system is tilted through the tilt mechanism, and its tilt angle range is -5°~100°.

4. A vertical take-off and landing ground effect vehicle according to claim 1, characterized in that: The propeller disc area of ​​the embedded ducted rotor power device at the front section of the ducted ground effect wing is greater than or equal to the propeller disc area of ​​the embedded ducted rotor power device at the rear section thereof.

5. A vertical take-off and landing ground effect vehicle according to any one of claims 1 to 4, characterized in that: The embedded ducted rotor power system includes four, six or eight embedded ducted rotor power devices in an even order, and are installed in the duct of the ducted ground effect wing through a left-right symmetrical arrangement of rotor power mounting brackets.

6. A vertical take-off and landing ground effect vehicle according to any one of claims 1 to 4, characterized in that: The center of gravity of the vertical take-off and landing ground effect vehicle is set 0 to 0.05 times the average aerodynamic chord length in front of the aerodynamic focus, and the rotor center is set 0.02 to 0.08 times the average aerodynamic chord length behind the center of gravity of the vertical take-off and landing ground effect vehicle.

7. The control method of a vertical take-off and landing ground effect vehicle according to claim 6, characterized in that: It includes a longitudinal attitude control method (210), a lateral attitude control method (220), a surface motion control method (230) and an emergency power jump control method (240); The longitudinal attitude control method (210) includes pitch attitude control in a hovering mode and a ground effect cruise mode, wherein the pitch attitude control is performed by a combined rotor system composed of the embedded ducted rotor power system and the propulsion rotor power system; The lateral attitude control method (220) includes rolling attitude control and heading turning control in a ground effect cruise mode; The on-water surface motion control method (230) includes on-water surface heading control and forward motion control; The emergency power jump control method (240) increases the power of the embedded ducted rotor system and simultaneously increases the flight angle of attack of the aircraft to increase the aerodynamic lift of the entire aircraft, thereby achieving the jump maneuverability of the aircraft.

8. The control method of a vertical take-off and landing ground effect vehicle according to claim 7, characterized in that: The pitch attitude control, in the ground effect cruise mode, the resultant nose-down moment of the combined rotor system located behind the center of gravity balances most of the aerodynamic nose-up moment, and each rotor works near the optimal state; in the flight from the hovering mode to the ground effect cruise mode, by increasing the fixed tilt angle of the propulsion rotor power system or increasing the power of the tiltable propulsion rotor power system, the thrust for forward flight is provided while suppressing the nose-up aerodynamic moment of the aircraft, so that the pitch angle of the aircraft is controlled within 5°, so as to realize the attitude conversion control of the aircraft from the hovering mode to the ground effect cruise mode; In the hovering mode, the pitch attitude control reduces the power of the propulsion rotor power system by increasing the power of the embedded ducted rotor power unit forward of the ducted ground effect wing; The roll attitude control makes the distance between the floats at the left and right wingtips of the ducted ground effect wing and the water surface or the ground greater than the safe height by compensating the height supplementary instruction of the aircraft related to the roll angle into the height control channel; Alternatively, the rolling attitude control method comprises installing a position sensor for measuring the distance from the water surface or the ground on the buoy, measuring the distance change between the buoy and the water surface or the ground in real time during the rolling process of the aircraft through the position sensor, and compensating the distance change into the altitude control channel; The heading turn control establishes a desired roll angle through the roll attitude control method so that the aircraft generates a suitable centripetal force on the horizontal plane of motion; The heading control on the water surface is performed by controlling the differential speed of the rotors on the plurality of embedded ducted rotor power units in the embedded ducted rotor power system, by increasing the speed of the counterclockwise rotating rotor and reducing the speed of all clockwise rotating rotors to achieve the clockwise rotation of the aircraft; or reducing the speed of all counterclockwise rotating rotors and increasing the speed of all clockwise rotating rotors to achieve the counterclockwise rotation of the aircraft; The forward motion control on the water surface is achieved by the positive rotation of the propulsion rotor power system to realize the forward motion of the aircraft; and the backward motion of the aircraft is achieved by the reverse rotation of the propulsion rotor power system.

9. The control method of a vertical take-off and landing ground effect vehicle according to claim 8, characterized in that: In the pitch attitude control, the following steps are also included: Step S101: Calculate the rotor center x tx , Among them, S A , S B , S C , S D , S F are the projected areas of the rotors of the embedded ducted rotor power device (A-D) and the propulsion rotor power system (F) on the horizontal plane, respectively, and x A 、x B 、x C 、x D 、x F are the positions of the rotors of the embedded ducted rotor power unit (A-D) and the propulsion rotor power system (F) on the x-axis respectively; Step S102: Calculate the pitch control energy efficiency coefficient f of each rotor in the embedded ducted rotor power unit (A-D) and the propulsion rotor power system (F) under the pitch attitude control A_z 、f B_z 、f C_z 、f D_z 、f F_z , the calculation formulas are:

10. The control method of a vertical take-off and landing ground effect vehicle according to claim 9, characterized in that: The pitch angle control method adopts cascade PID to complete the control of the pitch angle, and includes the following steps: Step S1: The pitch angle outer loop (the pitch angle to pitch angle rate) adopts PD control, and the desired pitch angle θ is obtained according to the pitch angle control strategy. g As well as the current pitch angle θ and pitch rate ω z , calculate the desired pitch angle rate ω zgz , oh zgz =k p1 (i g -θ)-k d1 ×w z , Among them, k p1 is the proportionality coefficient, k d1 Derivative control coefficient; Step S2: According to the performance requirements of the aircraft, set the pitch rate limit value ω zmax , and the desired pitch angle rate ω calculated in step S1 zgz Limit the amplitude to get the expected value of the pitch angle rate ω zg , Step S3: The pitch angle inner loop (the pitch angle rate to the pitch channel rotor motor throttle) adopts PID control, according to the pitch angle rate expected value ω zg 、Current angular velocity ω z and the current angular acceleration ε z , calculate the pitch channel rotor motor throttle signal, and according to the use requirements of the aircraft, it is necessary to limit the integral link of the pitch angle inner loop, and the set amplitude is d pitch_i_max , Among them, k i2 Integral control coefficient; Step S4: Based on step S3, calculate the pitch control signal value d pitch_z , d pitch_z =k p2 (w zg -w z )-k d2 ×ε z +d pitch_i , Among them, k p2 is the proportionality coefficient, k d2 Derivative control coefficient; Step S5: According to the use requirements of the aircraft, the total control signal of the pitch angle inner loop needs to be limited, and the amplitude is set to d pitch_max , Step S6: According to the aircraft usage requirements, the pitch control efficiency coefficient (f A_z 、f B_z 、f C_z 、f D_z 、f F_z ), the total control signal d of the pitch angle inner loop pitch Converted into the throttle signal of each rotor motor in the pitch channel; the throttle signal of each rotor motor pitch channel is dth A_z 、dth B_z 、dth C_z 、dth D_z 、dth F_z , Among them, ζ is the pitch channel control allocation coefficient, which is about 0.7 to 0.8, and f A_z 、f B_z 、f C_z 、f D_z 、f F_z They are respectively the rotor pitch control efficiency coefficients of the embedded ducted rotor power unit (A~D) and the propulsion rotor power system (F).

11. The control method of a vertical take-off and landing ground effect vehicle according to claim 8, characterized in that: In the rolling attitude control, the following steps are also included: Step S201: Calculate the rotor center z tx , In the roll attitude control, the projected areas of the rotors of the embedded ducted rotor power unit (A-D) and the propulsion rotor power system (F) on the horizontal plane are S and A , S B , S C , S D , S F , the positions on the z-axis are z A 、z B 、z C 、z D 、z F The embedded ducted rotor power system (A-D) and the propulsion rotor power system (F) are both arranged symmetrically, with the rotor center located on the symmetry plane, z tx =0; Step S202: Calculate the roll operation efficiency coefficient f of each rotor of the embedded ducted rotor power system (A-D) and the propulsion rotor power system (F) A_x 、f B_x 、f C_x 、f D_x 、f F_x , 12. The control method of a vertical take-off and landing ground effect vehicle according to claim 9, characterized in that: The roll angle control method uses a cascade PID to complete the control of the roll angle, and includes the following steps: Step S1: The roll angle outer loop (the roll angle to roll angle rate) adopts PD control, and the desired roll angle γ obtained according to the roll angle control strategy g And the current roll angle γ and roll angle rate ω x , calculate the desired roll angular rate ω xgz , oh xgz =k p1 (c g -c)-k d1 ×w x , Among them, k p1 is the proportionality coefficient, k d1 Derivative control coefficient; Step S2: According to the performance requirements of the aircraft, set the roll angle rate limit value ω xmax , and the roll angular rate ω calculated in step S1 xgz Limit the amplitude to get the expected value of the roll angular rate ω xg , Step S3: The inner loop of the roll angle (the roll angle rate to the roll channel rotor motor throttle) adopts PID control, according to the expected roll angle rate value ω xg and the current angular rate ω x and the current angular acceleration ε x , calculate the roll channel rotor motor throttle signal dth x According to the use requirements of the aircraft, it is necessary to limit the integral link of the inner loop of the roll angle, and the set amplitude is d roll_i_max , Among them, k i2 Integral control coefficient; Step S4: Based on step S3, calculate the roll control signal value d roll_z , d roll_z =k p2 (w xg -w x )-k d2 ×ε x +d roll_i , Among them, k p2 is the proportionality coefficient, k d2 Derivative control coefficient; Step S5: According to the use requirements of the aircraft, the total control signal of the inner loop of the roll angle needs to be limited, and the amplitude is set to d roll_max , Step S6: According to the aircraft use requirements and the roll control efficiency coefficients of each rotor of the aircraft, the total control signal d of the inner loop of the roll angle is converted into roll Converted into the throttle signal of each rotor motor in the roll channel. The throttle signal of each rotor motor roll channel is dth A_x 、dth B_x 、dth C_x 、dth D_x 、dth F_x , Where ζ is the roll channel control allocation coefficient, which is about 0.7 to 0.8, and f A_x 、f B_x 、f C_x 、f D_x 、f F_x are the rotor roll control efficiency coefficients of the embedded ducted rotor power unit (A-D) and the propulsion rotor power system (F); Step S7: In order to ensure the flight safety of the aircraft's rolling maneuver, it is necessary to perform altitude compensation on the altitude control channel. The altitude compensation amount H of the rolling maneuver in the ground effect cruise mode is g With the desired roll angle γ g Related, H g =k×L z ×γ g , Among them, L z The spanwise distance from the wingtip buoy of the ducted ground effect wing to the symmetry plane is k, a coefficient, and k is 0.9 to 1.1.

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