CN105151292B - Distributive vectored thrust system - Google Patents

Abstract

A distributed vector propulsion system belongs to the fields of aviation technology and electromechanical control technology, and in particular relates to a distributed vector propulsion system. The invention provides a distributed vector propulsion system capable of taking off, landing and hovering at any attitude. The present invention includes a free thrust unit and a flight controller, the free thrust unit includes a first servo mechanism frame, a first steering gear is arranged on the first servo mechanism frame, the shaft of the first steering gear is connected with the second servo mechanism frame, The second servo mechanism frame is provided with a second steering gear, the second steering gear shaft is connected with the power duct frame, and the second steering gear shaft is perpendicular to the first steering gear shaft; the flight controller controls the first steering gear The rotation angle input of the steering gear and the second steering gear, and the speed of the power duct.

Description

Distributed vector propulsion system

technical field

The invention belongs to the field of aviation technology and electromechanical control technology, and in particular relates to a distributed vector propulsion system.

Background technique

Existing fixed-wing aircraft do not have vertical take-off and landing capabilities, and are more dependent on take-off and landing sites. Although the tiltrotor mechanism has both vertical take-off and landing and level flight, it also has high requirements for flight quality similar to the fixed wing, and cannot meet the maneuvering flight in complex environments. Although the newly popular multi-rotor aircraft has good vertical take-off and landing and hovering performance, it cannot perform efficient cruise flight due to institutional limitations, resulting in a short range and unable to perform long-distance missions. Moreover, the trajectory and attitude are still coupled, and it is impossible to fly with full degrees of freedom and full attitude in space, and the maneuverability can be improved.

Thrust vectoring technology can turn part of the engine thrust into the control force, replacing or partially replacing the control surface, thereby greatly reducing the radar reflection area; no matter how large the angle of attack is and how low the flight speed is, the aircraft can use this part of the control force to control, This increases the maneuverability of the aircraft. Since the control force is directly generated, and the magnitude and direction are variable, the agility of the aircraft is increased, so the vertical tail can be appropriately reduced or removed, and some other control surfaces can also be replaced; this is helpful for reducing the detectability of the aircraft. It is beneficial and can also reduce the drag of the aircraft.

Contents of the invention

The present invention aims at the above problems and provides a distributed vector propulsion system capable of taking off, landing and hovering at any attitude.

In order to achieve the above object, the present invention adopts the following technical solutions. The present invention includes a free thrust unit and a flight controller. The free thrust unit includes a first servo mechanism frame, and the first servo mechanism frame is provided with a first steering gear. The shaft of the steering gear is connected with the frame of the second servo mechanism, the frame of the second servo mechanism is provided with a second steering gear, the shaft of the second steering gear is connected with the power duct frame, and the shaft of the second steering gear is connected with the frame of the first steering gear The axis is vertical; the flight controller controls the rotation angle input of the first steering gear and the second steering gear, and the rotational speed of the power duct.

As a preferred solution, the control method of the flight controller in the present invention is as follows.

A Cartesian rectangular coordinate system is established with the center of gravity of the carrier where the free thrust unit is located as the origin, r is the distance from the center of gravity to the duct, l is the length of the corresponding point when it is stable, and the corresponding point refers to: the plane where the carrier is located The three points are "connected" to the three points corresponding to the other upper plane parallel to the plane where the aircraft is located, and the connection line of the corresponding points is perpendicular to the two planes. The connection line is "chain", and the "chain" is linear elasticity. It satisfies Zhengxuan-Hooke's law, and the elastic coefficient is μ; τ is the angle of the top of the carrier aircraft.

Case 1: Under external disturbance, the angular displacement of the carrier around the x-axis is Correspondence needs to happen At the same time, the increased force required by the duct is:

Situation 2: Under external disturbance, the angular displacement of the loader around the y-axis is δ _λ , and the duct needs to have an angular displacement of -δ _λ . At the same time, the increased force required by the duct is: F=μrsinτ*sin(δ _λ ).

Situation 3: Under external disturbances, the carrier aircraft has a small angular displacement around the z-axis of δ _θ , and the channel needs to be adjusted as .

An angular displacement α occurs around the x-axis

Angular displacement β around the y-axis

At the same time, the increased force required by the channel is

Side hovering: Set to rotate the upper plane around the y-axis κ, keep the carrier parallel to the upper plane, and simultaneously rotate the three ducts around the y-axis.

It is set to rotate the upper plane ζ around the x-axis, keep the carrier plane parallel to the upper plane, and simultaneously rotate the three ducts around the x-axis.

As another preferred solution, there are three free thrust units in the present invention; one of the free thrust units is located on the symmetry axis of the fuselage behind the center of gravity, and the distance from the center of gravity is L3; the other two are symmetrically distributed before the center of gravity, to the center of gravity The distances are L1 and L2 respectively; the control forces of the three free thrust units are F1, F2 and F3 respectively, and the lift focus is adjusted to coincide with the center of gravity through the free thrust, the resultant control force is F=F1+F2+F3, and the resultant control moment M= F1×L1+F2×L2+F3×L3.

As another preferred solution, the flight controller of the present invention controls the rotational speed of the power duct through an electronic governor.

As another preferred solution, the flight controller of the present invention includes an integrated sensor and a flight control board, the integrated sensor includes an inertial measurement unit, a GPS navigation module and a three-axis magnetometer module, and the inertial measurement unit includes a three-axis angular velocity measurement part and a three-axis acceleration measurement part; the flight controller measures the three-axis angular velocity and the three-axis acceleration, corrects it with the direction data, measures the flight attitude angle of the carrier aircraft, and uses the cosine algorithm to obtain the attitude data of the aircraft flight.

As another preferred solution, the flight control board of the present invention adopts Atmega1280/2560 chip.

As another preferred solution, the flight control board of the present invention includes a first receiver, a second receiver, an APM1 chip, an APM2 chip, an Arithmetic unit, a MWC1 board, a MWC2 board and a MWC3 board, and the signal input ports of the Arithmetic unit are respectively Be connected with the signal output port of the second receiver, the signal output port of APM2 chip, the signal output port of the first receiver is connected with the signal input port of APM1 chip; The signal output port of Arithmetic unit is connected with the signal input port of APM1 chip, The signal input port of the MWC1 board, the signal input port of the MWC2 board, and the signal input port of the MWC3 board are connected; the signal output port of the MWC1 board is connected with the first servo control signal input port and the second servo control signal input port of one of the free thrust units respectively. The signal input ports of the MWC2 board are respectively connected to the first steering gear control signal input port and the second steering gear control signal input port of another free thrust unit, and the signal output ports of the MWC2 board are respectively connected to the third free thrust unit. The first steering gear control signal input port of the unit is connected to the second steering gear control signal input port; the signal input port of the APM2 chip is connected to the signal output port of the optical flow sensor and the signal output port of the GPS sensor respectively; The signal output ports are respectively connected to the power duct speed control signal input ports of the three free thrust units.

The first receiver accepts the attitude data sent by the ground controller to the carrier aircraft, and inputs the attitude signal into the APM1 chip for calculation. The rotational speed of the power duct; the second receiver receives the track control signal sent by the ground controller to the carrier aircraft, and inputs the track control signal into the ARITHMETIC unit; the APM2 chip collects the signal data of the optical flow sensor and the GPS sensor, and sends the signal data to the ARITHMETIC The unit inputs four control signals 1, 2, 3, Y(1); the arithmetic unit converts the signal into seven output signals P1(OUT), P2(OUT), P3(OUT), R1(OUT), R2 (OUT), R3(OUT), and T are used as the input signals of the three MWC boards, and the three MWC control boards control the tilting of the six steering gears respectively.

The P signal is copied three times to obtain three signals of P1, P2, and P3.

The R signal is copied three times to obtain three signals R1, R2, and R3.

After the 3 signal and the 2 signal are superimposed, the signal intensity is reduced to 1/2 of the original, and then the 1 signal is subtracted to obtain the PG signal.

The 3 signal and the 2 signal are eliminated to obtain the inverted RG signal.

P1(OUT) is the superposition of P1 signal and PG signal and then cancels with Y and Y(1) signal.

P2(OUT) is obtained by superimposing the four groups of signals P2, Y, Y(1), and PG.

P3(OUT) is obtained by superimposing the P3 signal and the PG signal.

R1(OUT) is obtained by superimposing the R1 signal and the RG signal.

R2(OUT) is obtained by superimposing the R2 signal and the RG signal.

R3(OUT) is obtained by superimposing the R3 signal and the RG signal.

The P—pitch signal, R—roll signal, T—throttle signal, Y—yaw signal, 1, 2, 3—computing signal, (out)—output signal.

Secondly, the track control signal of the present invention includes front and rear, left and right, nose pointing, and throttle signals, and the attitude data includes pitch and roll data.

In addition, the first servo mechanism frame of the present invention includes a horizontal frame, the front end of the horizontal frame is provided with a front arc-shaped frame that is bent forward to the upper end, and the rear part of the horizontal frame is provided with a rear arc that is curved to the rear upper part corresponding to the front arc-shaped frame. shaped frame, the rear end of the horizontal frame is provided with the first steering gear, the axis of the first steering gear is parallel to the horizontal frame and passes through the upper through hole of the rear arc-shaped frame; the second servo mechanism frame is Most of the semicircular sealing frame, the lower contour of the second servo mechanism frame corresponds to the contour surrounded by the horizontal frame, the front arc frame and the rear arc frame; the second steering gear is arranged on the second servo mechanism machine At the upper end of the frame, the second steering gear shaft is connected vertically downward to the power duct frame; the horizontal end of the middle part of the second servo mechanism frame is connected with the first steering gear shaft, and the other horizontal end of the middle part of the second servo mechanism frame passes through The horizontal axis is connected with the upper part of the front arc frame.

The invention has beneficial effects.

The free thrust unit of the present invention can freely adjust the magnitude and direction of the thrust.

The flight controller of the present invention separately controls the angle input of the first and second steering gears of the free thrust unit and the rotational speed of the power duct to obtain full free thrust.

The distributed multi-element vector propulsion system of the present invention does not depend on site conditions for take-off, can realize take-off, landing and hovering at any attitude, is especially suitable for take-off and landing in complex and narrow urban terrains and special environments, and can perform exploration, surveillance and investigation tasks with a stable attitude. When performing missions, the distributed multivariate vector propulsion system can achieve stable flight at any flight attitude, fast and flexible attitude adjustment, and fast start-up and parking, so it can fly efficiently in small urban streets and even inside buildings. Environments such as towns and ruins can also complete tasks efficiently.

Description of drawings

The present invention will be further described below in conjunction with the accompanying drawings and specific embodiments. The scope of protection of the present invention is not limited to the following expressions.

Fig. 1 is a schematic diagram of the structure of the present invention.

Fig. 2 is a front view of the free thrust unit of the present invention.

Fig. 3 is a perspective view of the free thrust unit of the present invention.

Fig. 4 is a schematic circuit diagram of the present invention.

Figure 5 is a load spectrum diagram of the present invention. (Grayscale images cannot express clearly)

Fig. 6 is a diagram for establishing the coordinate system of the control method of the present invention.

Fig. 7 is a schematic diagram of situation 1 of the control method of the present invention.

Fig. 8 is a schematic diagram of situation 2 of the control method of the present invention.

Fig. 9 is a schematic diagram of situation 3 of the control method of the present invention.

In the figure, 1 is the first steering gear, 2 is the second steering gear, 3 is the first servo mechanism rack, 4 is the second servo mechanism rack, 5 is the first steering gear shaft, and 6 is the second steering gear shaft , 7 is a power duct, 8 is a free thrust unit, 9 is a carrier, 10 is a rear arc frame, 11 is a horizontal frame, 12 is a front arc frame, and 13 is a horizontal axis.

detailed description

As shown in the figure, the present invention includes a free thrust unit 8 and a flight controller, and the free thrust unit 8 includes a first servo mechanism frame 3, on which the first servo mechanism frame 3 is provided with a first steering gear 1, the first steering gear The shaft 5 is connected with the second servo mechanism frame 4, the second servo mechanism frame 4 is provided with the second steering gear 2, the second steering gear shaft 6 is connected with the power duct 7, and the second steering gear shaft 6 It is perpendicular to the first steering gear axis 5; the flight controller controls the rotation angle input of the first steering gear 1 and the second steering gear 2, and the rotational speed of the power duct 7.

Electric power can be selected as the energy source of the power system to obtain faster response speed and make thrust adjustment more precise.

The control method of the flight controller is as follows.

Establish a Cartesian rectangular coordinate system with the center of gravity of the carrier aircraft 9 where the free thrust unit 8 is located as the origin, as shown in Figure 8, r is the distance from the center of gravity to the duct, l is the length of the corresponding point when it is stable, and the corresponding point refers to: "Tie" the three points on the plane where the carrier aircraft 9 is located on the three points corresponding to another upper plane parallel to the plane where the carrier aircraft 9 is located, and the connection line of the corresponding points is perpendicular to these two planes, and the connection line is " "chain", "chain" is linear elasticity, which satisfies Zhengxuan-Hooke's law, and the coefficient of elasticity is μ;

Situation 1: Under external disturbance, the angular displacement of carrier 9 around the x-axis is Correspondence needs to happen At the same time, the increased force required by the duct is:

Situation 2: Under external disturbance, the angular displacement of carrier 9 around the y-axis is δ _λ , and the duct needs to have an angular displacement of -δ _λ . At the same time, the increased force required by the duct is: F=μrsinτ*sin( _δλ ).

Situation 3: Under external disturbances, the carrier aircraft 9 has a small angular displacement around the z-axis of δ _θ , and the adjustment required for the channel is .

An angular displacement α occurs around the x-axis

Angular displacement β around the y-axis

At the same time, the increased force required by the channel is

Lateral hovering: set to rotate the upper plane around the y-axis by κ, keep the carrier 9 parallel to the upper plane, and simultaneously rotate the three ducts around the y-axis.

It is set to rotate the upper plane ζ around the x-axis, keep the carrier 9 parallel to the upper plane, and simultaneously rotate the three ducts around the x-axis.

The control method of the flight controller of the present invention adopts mathematical modeling of the relationship between the tension and direction of the suspension rope and the attitude of the object when the object is suspended from the ceiling, and uses each independent power unit to simulate the suspension rope, and the vector of each power unit The direction dynamically simulates the direction of each hanging rope, and the tension on the rope is simulated by the thrust of the power unit.

Objects suspended on the ceiling will tend to be stable under the action of gravity and resistance. Based on this, we have established a mathematical and mechanical model for the hanging rope of the hanging object; The direction of the unit vector and the magnitude of the thrust are linearly simulated, and the dynamic analysis is performed on the situation when the suspended object is disturbed, and then each power unit can be coordinated to complete the control of the aircraft.

The model consists of two parts: the first part is the dynamic equations based on the laws of dynamics, and the other part is the kinematic equations obtained through the coordinate transformation relationship.

Before building the aircraft model, make the following settings.

(1) The aircraft is an absolute rigid body, and the influence of structural elasticity is not considered.

(2) The mass and moment of inertia of the aircraft are constant.

(3) Neglect the airflow interference of the three ducted fans.

(4) The structure and quality of the same parts are the same.

(5) The structure is centrosymmetric.

Due to external disturbances, the balance of the aircraft will be affected to a certain extent. By changing the angle of the channel and increasing the thrust, the carrier aircraft 9 of the multi-element vector propulsion system can be restored to a stable attitude. The acceleration sensor is used on the aircraft, and the angular acceleration can be obtained from the device, and then the offset angle can be obtained.

The control method of the present invention can control the forward, backward and left-right movement of the carrier aircraft 9 of the multivariate vector propulsion system, and only needs to control the upper plane to move horizontally. It will be subjected to forward force, thereby realizing the horizontal movement of the carrier aircraft 9 of the multivariate vector propulsion system.

The control method of the invention effectively solves the problem of how to use the vector unit to control the aircraft when outputting with multiple degrees of freedom. The closed-loop correction of the attitude of the aircraft can be performed through the three-axis accelerometer and the three-axis gyroscope, so as to achieve the purpose of coordinating the flight of the air vehicle with multiple degrees of freedom.

There are three free thrust units 8; one of them is located on the symmetry axis of the fuselage behind the center of gravity, and the distance from the center of gravity is L3; the other two are symmetrically distributed before the center of gravity, and the distances to the center of gravity are L1 and L2 respectively The control forces of the three free thrust units 8 are respectively F1, F2, and F3, and the focus of the lift force is adjusted to coincide with the center of gravity through the free thrust, the combined control force is F=F1+F2+F3, and the combined control torque M=F1×L1+F2× L2+F3×L3 (both are vector operations). Through the linkage of the power units distributed on the fuselage, multiple thrust vectors are synthesized into one control force and one force couple, so as to achieve independent control of the attitude and heading of the aircraft, so as to obtain the functions of good cruise capability, fixed-point hovering and maneuvering performance All-in-one aerial platform.

The flight controller controls the speed of the power duct 7 through an electronic governor.

The flight controller includes an integrated sensor and a flight control board, the integrated sensor includes an inertial measurement unit, a GPS navigation module and a three-axis magnetometer module, and the inertial measurement unit includes a three-axis angular velocity measurement part and a three-axis acceleration measurement part; The flight controller measures the three-axis angular velocity and the three-axis acceleration, corrects it with the direction data, measures the flight attitude angle of the carrier aircraft 9, and uses the cosine algorithm to obtain the flight attitude data of the aircraft. The GPS navigation module can measure the aircraft's current latitude and longitude, altitude, track direction (track), ground speed and other information. The three-axis magnetometer module can measure the current heading of the aircraft. The flight controller can also set the airspeed gauge, air pressure gauge, and AD chip.

The flight control board adopts Atmega1280/2560 chip. The Atmega1280/2560 chip has a PPM decoding chip, which is responsible for monitoring the PWM signal of the mode channel in order to switch between manual mode and other modes.

Described flight control board comprises first receiver, second receiver, APM1 chip, APM2 chip, Arithmetic unit, MWC1 board, MWC2 board and MWC3 board, and the signal input port of Arithmetic unit is connected with the signal output port of the second receiver respectively , the signal output port of the APM2 chip is connected, the signal output port of the first receiver is connected with the signal input port of the APM1 chip; the signal output port of the Arithmetic unit is connected with the signal input port of the APM1 chip, the signal input port of the MWC1 board, and the MWC2 board The signal input port of the MWC3 board is connected with the signal input port of the MWC3 board; the signal output port of the MWC1 board is respectively connected with the first steering gear 1 control signal input port and the second steering gear 2 control signal input port of one of the free thrust units 8, MWC2 The signal output port of the board is connected with the first steering gear 1 control signal input port and the second steering gear 2 control signal input port of another free thrust unit 8 respectively, and the signal output port of the MWC2 board is connected with the third free thrust unit 8 respectively. The first steering gear 1 control signal input port is connected to the second steering gear 2 control signal input ports; the signal input port of the APM2 chip is connected to the signal output port of the optical flow sensor and the signal output port of the GPS sensor respectively; The signal output ports are respectively connected to the power duct 7 rotational speed control signal input ports of the three free thrust units 8 .

The first receiver accepts the attitude data sent by the ground controller to the carrier aircraft 9, and inputs the attitude signal into the APM1 chip for calculation. The APM1 chip also accepts the throttle signal processed by the arithmetic unit, and outputs three throttle signals to control the The speed of the three power ducts 7; the second receiver receives the track control signal sent by the ground controller to the carrier aircraft 9, and inputs the track control signal into the Arithmetic unit; the APM2 chip collects the signal data of the optical flow sensor and the GPS sensor , input four control signals 1, 2, 3, Y(1) to the Arithmetic unit; the arithmetic unit converts the signals into seven output signals P1(out), P2(out), P3(out), R1(out ), R2(out), R3(out), and T are used as the input signals of the three MWC boards, and the three MWC control boards control the tilting of the six steering gears respectively.

The P signal is copied three times to obtain three signals of P1, P2, and P3.

The R signal is copied three times to obtain three signals R1, R2, and R3.

After the 3 signal and the 2 signal are superimposed, the signal intensity is reduced to 1/2 of the original, and then the 1 signal is subtracted to obtain the Pg signal.

The 3 signal and the 2 signal are eliminated to obtain the inverted Rg signal.

P1(out) is the superposition of P1 signal and Pg signal and then cancels with Y and Y(1) signal.

P2(out) is obtained after the four groups of signals P2, Y, Y(1), and Pg are superimposed on each other.

P3(out) is obtained by superimposing the P3 signal and the Pg signal.

R1(out) is obtained by superimposing the R1 signal and the Rg signal.

R2(out) is obtained by superimposing the R2 signal and the Rg signal.

R3(out) is obtained by superimposing the R3 signal and the Rg signal.

The P—pitch signal, R—roll signal, T—throttle signal, Y—yaw signal, 1, 2, 3—computing signal, (out)—output signal.

The track control signal includes front and rear, left and right, nose pointing, and throttle signals, and the attitude data includes pitch and roll data.

The first servo mechanism frame 3 includes a horizontal frame 11, the front end of the horizontal frame 11 is provided with a front arc-shaped frame 12 that is bent forward to the upper end, and the rear portion of the horizontal frame 11 is provided with a rear upper curved frame corresponding to the front arc-shaped frame 12. The rear arc-shaped frame 10, the rear end of the horizontal frame 11 is provided with the first steering gear 1, the first steering gear shaft 5 is parallel to the horizontal frame 11 and passes through the upper through hole of the rear arc-shaped frame 10; The second servo mechanism frame 4 is a mostly semicircular sealing frame, and the lower contour of the second servo mechanism frame 4 corresponds to the contour surrounded by the horizontal frame 11, the front arc frame 12, and the rear arc frame 10; The second steering gear 2 is arranged on the upper end of the second servo mechanism frame 4, and the second steering gear shaft 6 is connected to the power duct 7 vertically downward; the horizontal end of the second servo mechanism frame 4 middle part is connected to the The first steering gear shaft 5 is connected, and the other lateral end of the middle part of the second servo mechanism frame 4 is connected with the upper part of the front arc frame 12 through a transverse shaft 13 . The invention synthesizes a plurality of thrust vectors into one control force and one force couple through the linkage of the power units distributed on the fuselage, so as to achieve independent control of the attitude and heading of the aircraft. The multiple dual-axis omnidirectional vector propulsion units in this system provide a guarantee for the full vector maneuverability of the aircraft. By adopting a newly built flight control platform and a new flight control method to control the rotation angle of two mutually perpendicular steering gears and the motor Accurately control the real-time interaction between each vector power unit and the power signal, which can precisely adjust the attitude and trajectory of the aircraft. The distributed multi-element vector system combines the characteristics of the distributed power layout to improve the control mode of the aircraft, overcome the problems of under-actuation and under-stability of the rotorcraft, and realize the high maneuverability of the aircraft's multi-attitude hovering in the air. The following table is an optimal table of hardware parameters of the power system of the present invention.

The following table is a list of circuit hardware parameters of the present invention.

It can be understood that the above specific descriptions of the present invention are only used to illustrate the present invention and are not limited to the technical solutions described in the embodiments of the present invention. Those of ordinary skill in the art should understand that the present invention can still be modified or Equivalent replacements to achieve the same technical effect; as long as they meet the needs of use, they are all within the protection scope of the present invention.

Claims (8)

1. distributed vector propulsion system, including free thrust unit and flight controller, it is characterised in that free thrust unit Including the first servo control mechanism frame, the first steering wheel, the first steering wheel axle and the second servo are provided with the first servo control mechanism frame Structure frame is connected, and the second steering wheel is provided with the second servo control mechanism frame, and the second steering wheel axle is connected with power duct frame, and described the Two steering wheel axles are vertical with the first steering wheel axle；The flight controller controls the anglec of rotation input of the first steering wheel and the second steering wheel, And the rotating speed of power duct；

The control method of the flight controller is：

Set up Descartes's rectangular coordinate system by origin of the center of gravity of free thrust unit place carrier aircraft, r be center of gravity to duct away from It is the distance of a pair of corresponding point when steady from, l, corresponding point are referred to：By carrier aircraft be located plane on three points " being " with On the parallel upper plane of plane that carrier aircraft is located is corresponding 3 points, perpendicular to this two plane, line is for " even for the line of corresponding point Lock ", " chain " is linear elasticity, meets Zheng Xuan-Hooke's law, and coefficient of elasticity is μ；τ is the drift angle number of degrees of carrier aircraft；

Situation 1：Under extraneous disturbance, there is angular displacement and be in carrier aircraft around x-axisExhausting needs generationAngular displacement, together When, the power increased needed for duct is：

Situation 2：Under extraneous disturbance, it is δ that carrier aircraft occurs angular displacement around y-axis_λ, exhausting needs generation-δ_λAngular displacement, meanwhile, The power of increase is needed for duct：F=μ rsin τ * sin (δ_λ)；

Situation 3:Under extraneous disturbance, it is δ that carrier aircraft occurs less angular displacement around z-axis_θ, that what is made needed for exhausting is adjusted to：

There is angular displacement alpha around x-axis

$\mathrm{\α}=arctan\frac{2rsin\frac{{\mathrm{\δ}}_{\mathrm{\θ}}}{2}*cos\frac{\mathrm{\τ}}{2}}{l}$

There is angular displacement beta around y-axis

$\mathrm{\β}=arctan\frac{2rsin\frac{{\mathrm{\σ}}_{\mathrm{\θ}}}{2}*sin\frac{\mathrm{\τ}}{2}}{l}$

Meanwhile, the power increased needed for duct is

Lateral hovering：Arrange and the upper plane is rotated into κ around y-axis, keep carrier aircraft parallel with upper plane, while three ducts are simultaneously Rotate around y-axis；

Arrange and the upper plane is rotated into ζ around x-axis, keep carrier aircraft parallel with upper plane, while three ducts are simultaneously around x-axis rotation Turn.

2. distributed vector propulsion system according to claim 1, it is characterised in that the free thrust unit is three；Its In free thrust unit be located at after center of gravity on fuselage axis of symmetry, be L3 with centroidal distance；Two other is symmetrically distributed Before center of gravity, the distance to center of gravity is respectively L1, L2；The controling power of three free thrust units is respectively F1, F2, F3, passes through Free thrust adjusts lift focus and overlaps with center of gravity, conjunction controling power be F=F1+F2+F3, conjunction control moment M=F1 × L1+F2 × L2+F3×L3。

3. distributed vector propulsion system according to claim 1, it is characterised in that the flight controller is adjusted by electronics Fast device controls the rotating speed of power duct.

4. distributed vector propulsion system according to claim 2, it is characterised in that the flight controller includes integrated biography Sensor and winged control plate, the integrated sensor includes Inertial Measurement Unit, GPS navigation module and three axle magnetometer module, inertia Measuring unit includes that three axis angular rates measurement part and 3-axis acceleration measure part；The flight controller measures three shaft angles speed Degree, 3-axis acceleration coordinates bearing data to be corrected, and measures the flight attitude angle of carrier aircraft, and with cosine-algorithm load is drawn The attitude data of machine flight.

5. distributed vector propulsion system according to claim 4, it is characterised in that the winged control plate adopts Atmega1280/ 2560 chips.

6. distributed vector propulsion system according to claim 4, it is characterised in that the winged control plate include the first receiver, Second receiver, APM1 chips, APM2 chips, Arithmetic units, MWC1 plates, MWC2 plates and MWC3 plates, Arithmetic The signal input port of unit signal output port respectively with the second receiver, the signal output port of APM2 chips are connected, The signal output port of the first receiver is connected with the signal input port of APM1 chips；The signal output of Arithmetic units Port signal input port respectively with APM1 chips, the signal input port of MWC1 plates, the signal input port of MWC2 plates, The signal input port of MWC3 plates is connected；The signal output port of MWC1 plates is respectively with the first of one of free thrust unit Servos control signal input port, the second servos control signal input port are connected, the signal output port of MWC2 plates respectively with First servos control signal input port of another free thrust unit, the second servos control signal input port are connected, MWC3 The signal output port of plate respectively with the first servos control signal input port, second servos control of the 3rd free thrust unit Signal input port is connected；The signal input port of APM2 chips signal output port, GPS respectively with light flow sensor The signal output port of sensor is connected；The signal output port of APM1 chips is contained respectively with the power of three free thrust units Road speed controling signal input port is connected；

First receiver receives the attitude data that ground controller sends to carrier aircraft, and attitude signal is input in APM1 chips Resolved, APM1 chips also receive the throttle signal after processing arithmetical unit, three road binders gate signals of output control respectively three The rotating speed size of individual power duct；Second receiver receives the flight tracking control signal that ground controller sends to carrier aircraft, by flight path Control signal is input into Arithmetic units；APM2 chips gather the signal data of light flow sensor and GPS sensor, to Arithmetic units are input into four tunnel control signals 1,2,3, Y (1)；Arithmetical unit will be converted to seven tunnel output signals after signal processing P1 (out), P2 (out), P3 (out), R1 (out), R2 (out), R3 (out), T as three blocks of MWC plates input signal, three pieces MWC panels control respectively verting for six steering wheels；

P signal is replicated three times and obtains tri- signals of P1, P2, P3；

R signal is replicated three times and obtains tri- signals of R1, R2, R3；

Reduce after 3 signals and 2 Signal averagings signal intensity be original two/deduct 1 signal again and again and obtain Pg signals；

3 signals and 2 signal cancellations obtain the signal of falling Rg；

P1 (out) is again with Y and Y (1) signal cancellation after P1 signals and Pg Signal averagings；

P2 (out) is P2, Y, Y (1), tetra- groups of signals of Pg are obtained after being overlapped mutually；

P3 (out) is that P3 signals are overlapped mutually what is obtained with Pg signals；

R1 (out) is that R1 signals are overlapped mutually acquisition with Rg signals；

R2 (out) is that R2 signals are overlapped mutually acquisition with Rg signals；

R3 (out) is that R3 signals are overlapped mutually acquisition with Rg signals；

P-the Pitch signal, R-rolling signal, T-throttle signal, Y-off-course signal, 1,2,3-computing signal, (out)-output signal, Y (1) is the off-course signal of course transmitter feedback on APM2.

7. distributed vector propulsion system according to claim 6, it is characterised in that before and after the flight tracking control signal includes, Left and right, head are pointed to, throttle signal, and the attitude data includes pitching, rolling data.

8. distributed vector propulsion system according to claim 1, it is characterised in that the first servo control mechanism frame includes Horizontal frame, horizontal frame front end is provided with the front arc frame of forward upper end bending, and horizontal frame rear portion arranges oriented corresponding to front arc frame The rear arc frame of upper back bending, the rear end of horizontal frame is provided with first steering wheel, and the first steering wheel axle is parallel to the horizontal frame And through the rear arc frame upper through-hole；The second servo control mechanism frame is many semicircle edge banding frames, the second servo The profile of structure frame is corresponding with the profile that front arc frame, rear arc frame are surrounded with the horizontal frame；Second rudder Machine is arranged on the second servo control mechanism frame upper end, and the second steering wheel axle is connected vertically downward with power duct frame；Second servo The horizontal one end of mechanism's central rack is connected with the first steering wheel axle, and the horizontal other end of the second servo control mechanism central rack is by horizontal stroke Axle is connected with front arc frame top.