Tilt-rotor unmanned aerial vehicle with unpowered falling function and algorithm thereof

Abstract

The invention discloses a tilt rotor unmanned aerial vehicle with an unpowered falling function, and relates to the technical field of unmanned aerial vehicles, the tilt rotor unmanned aerial vehicle comprises a vehicle body, two symmetrically arranged front holes are formed in the front end of the vehicle body, two symmetrically arranged rear holes are formed in the rear end of the vehicle body, and the front holes and the rear holes are internally connected with rotor engines through rotatable ducts; a sliding rail is arranged on the center line of the aircraft body, a front rotor wing cover plate matched with the front hole structure is connected to the sliding rail in a sliding mode, telescopic wings with variable extension lengths are arranged on the two sides of the aircraft body, the top face and the bottom face of the aircraft body are each provided with two symmetrically-arranged storage grooves, and resistance rudders are rotationally connected into the storage grooves. According to the tilt-rotor unmanned aerial vehicle with the unpowered falling function, through special structural design and application of a control algorithm, stable and vertical take-off can be finally achieved, lifting force is provided through the fixed wings after the tilt-rotor unmanned aerial vehicle is converted into a level flight cruise state, and the unmanned aerial vehicle is more stable and reliable. And during landing, the unmanned aerial vehicle can land safely in a form of almost not consuming own power.

Description

Technical Field

[0001] The present invention relates to the technical field of unmanned aerial vehicles, and in particular to a tilt-rotor unmanned aerial vehicle with an unpowered landing function. Background Art

[0002] Rotary-wing UAVs have the advantages of being easy to miniaturize, capable of vertical take-off and landing, and having mature control schemes. In recent years, with the development and maturity of technology, rotary-wing UAVs have not only been widely used in commercial activities and scientific exploration, but also received increasing attention in military operations[1][3]. However, during flight, rotary-wing UAVs completely rely on the recoil force of the airflow pushed by the rotor to overcome gravity. Compared with fixed-wing aircraft, they consume more energy during flight and have a shorter endurance during military operations.

[0003] In order to combine the advantages of rotorcraft with fixed-wing aircraft, tilt-rotor UAVs that can switch the direction of the rotors during flight have been designed. After the tilt-rotor UAV is lifted vertically by the rotors, it will switch the direction of the rotors when entering the level flight cruising state, so that the lift in the subsequent flight process is provided by the fixed wings, so that the power output of the rotors basically only needs to overcome the air resistance, achieving a more efficient and energy-saving flight [5]. However, on the one hand, it is difficult for current tilt-rotor UAVs to smoothly transform the flight state. When the rotors are turning, the body often becomes unstable or even shakes significantly [6]. On the other hand, the direct addition of fixed-wing design makes the tilt-rotor UAV structure less compact than ordinary rotor UAVs, and the moment of inertia is greatly increased. As a result, the tilt-rotor UAV often requires more power output to control the rotors during the process of lifting and descending by the rotors, and the energy consumption increases. It is particularly worth noting that if the power system of the tilt-rotor UAV is insufficient during the descent process, it will often cause catastrophic accidents. In order to solve this problem, the common practice now is to pre-configure the use of energy and automatically start returning when the energy storage is low. This practice is to avoid accidents to the greatest extent through preset rules. However, the above method is difficult to deal with the accidental excessive energy consumption encountered by the drone during flight. It is acceptable in the civilian field, but when the tilt-rotor drone is used for military reconnaissance, strike and other tasks, there are more unexpected situations and it is more likely to consume too much energy. Even if the automatic return can ensure safety, it will bring many inconveniences.

[0004] In order to solve the above problems, the present application proposes a new type of tilt-rotor UAV, which relies on the effect of air resistance to control attitude stability and landing position during the descent process, and only briefly starts the rotor to straighten the UAV when the UAV approaches the ground, so that the energy saving and safety of the landing process are greatly improved. In this way, the new type of tilt-rotor UAV consumes almost no energy during the descent process, thereby improving the safety of the UAV landing process, making it adaptable to more working conditions, and avoiding accidents due to insufficient energy when performing reconnaissance, strike and other tasks. Since the drag rudder mechanism is simple and not easy to be damaged, even if the new type of UAV faces serious accidents such as flight system failure, it can still decelerate and land to ensure flight safety. Summary of the invention

[0005] The purpose of the present invention is to provide a tilt-rotor UAV with an unpowered landing function, which can ultimately achieve a smooth vertical take-off through a special structural design and the application of a control algorithm, rely on fixed wings to provide lift after switching to a level flight cruising state, and land safely without consuming almost any of its own power during landing.

[0006] To achieve the above-mentioned objectives, the present invention provides a tilt-rotor UAV with an unpowered landing function, comprising a fuselage with a flying wing layout, the front end of the fuselage is provided with two symmetrically arranged front holes, the rear end of the fuselage is provided with two symmetrically arranged rear holes, the front holes and the rear holes are both connected to the rotor engine through a rotatable duct, the power direction of the rotor engine is consistent with the axial direction of the rotatable duct, a slide rail is provided on the center line of the fuselage, a front rotor cover plate adapted to the front hole structure is slidably connected to the slide rail, telescopic wings with variable extension are provided on both sides of the fuselage, the top and bottom surfaces of the fuselage are provided with two symmetrically arranged storage grooves, and a resistance rudder is rotatably connected in the storage groove.

[0007] Therefore, the tilt-rotor UAV with the above-mentioned structure and the unpowered landing function has the following beneficial effects:

[0008] 1. Through structural design, the UAV can adjust its own structural distribution between vertical rise form, stable flight form, air deceleration form, and resistance reduction form through the conversion of structural form, making the structure more compact and making vertical take-off and landing more reliable and safe.

[0009] 2. In terms of control algorithm, the stability of the flight state is mainly ensured by fine-tuning the rotor during level flight and during landing, the rotor is only activated at the end of landing, and the drag rudder solution is used in the middle to control the drone's smooth landing and vertical speed control. This not only ensures energy saving during the landing process and greatly improves safety, but also because the rotor tilt only occurs during the take-off phase, the impact of the rotor tilt on the flight process is also reduced;

[0010] 3. With a variable length structure, the UAV can adjust the moment of inertia, aerodynamic force, and center of gravity by retracting the wings in different working conditions to better adapt to different flight modes. Using the wide and flat fuselage surface of the flying wing layout, the drag rudder of the new UAV is designed to be hidden. During normal flight, it fits into the pre-designed groove on the fuselage surface and constitutes a part of the fuselage. It will only start working when landing, which makes the structure of the UAV more compact, avoids the influence of the drag rudder structure on the aerodynamic force during flight, and makes the layout simpler.

[0011] The technical solution of the present invention is further described in detail below through the accompanying drawings and embodiments. BRIEF DESCRIPTION OF THE DRAWINGS

[0012] Figure 1 The UAV in the rotor state of the present invention;

[0013] Figure 2 The UAV in the fixed-wing state of the present invention;

[0014] Figure 3 The UAV in the drag rudder state of the present invention;

[0015] Figure 4 is the variation of lift coefficient with span and angle of attack;

[0016] Figure 5 is the variation of pitch moment coefficient with span and angle of attack;

[0017] Figure 6 It is the fitting effect diagram of lift coefficient;

[0018] Figure 7 It is the fitting effect diagram of lift moment coefficient;

[0019] Figure 8 It is a schematic diagram of the longitudinal force of the UAV model during level flight cruising;

[0020] Fig. 9 It is a comparison chart of the UAV parameter values ​​in equilibrium state and the parameter values ​​when the uncontrolled longitudinal model is stable;

[0021] Fig.10 This is a simplified two-dimensional diagram of the landing phase of the drone;

[0022] Fig.11 This is the simulation result of the first set of initial values;

[0023] Fig.12 is the second set of initial value simulation results;

[0024] Fig.13 This is the simulation result of the third group of initial values;

[0025] Fig.14 The simulation results of the uncontrolled state of the UAV longitudinal dynamics model when the rotor engine deflection angle δ=0°, 30°, -30° in level flight cruise;

[0026] Fig.15 The simulation results of the uncontrolled longitudinal dynamics model of the UAV when the rotor engine deflection angle δ = 0°, 40°, 80° in level flight cruise;

[0027] Fig.16 is the rotor engine deflection angle control target θ in level flight cruise r =0°, 10°, -10° when the simulation results of the UAV longitudinal dynamics model under traditional PID control;

[0028] Fig.17 is the rotor engine deflection angle control target θ in level flight cruise r =0°, 30°, -30° when the simulation results of the UAV longitudinal dynamics model under traditional PID control.

[0029] Reference numerals

[0030] 1. Fuselage; 2. Front hole; 3. Rear hole; 4. Rotor engine; 5. Rotatable duct; 6. Connecting rod; 7. Slide rail; 8. Front rotor cover; 9. Retractable wing; 10. Storage slot; 11. Drag rudder. DETAILED DESCRIPTION

[0031] The technical solution of the present invention is further described below through the accompanying drawings and embodiments.

[0032] Example

[0033] like Figure 1-17 As shown, a tilt-rotor UAV with an unpowered landing function includes a body 1 with a flying wing layout, the front end of the body 1 is provided with two symmetrically arranged front holes 2, and the rear end of the body 1 is provided with two symmetrically arranged rear holes 3. The front holes 2 and the rear holes 3 are both connected to the rotor engine 4 through a rotatable duct 5, and the power direction of the rotor engine 4 is consistent with the axial direction of the rotatable duct 5. The two ends of the rotatable duct 5 are rotatably connected to the inner walls of the front hole 2 and the rear hole 3 through a connecting rod 6, and one of the connecting rods 6 is connected to the first rotating power device located in the body 1.

[0034] A slide rail 7 is provided on the midline of the body 1, and a front rotor cover plate 8 matching the structure of the front hole 2 is slidably connected to the slide rail 7. A screw is rotatably connected in the slide rail 7, one end of which is connected to the second rotary power device located in the body 1, and the screw is fixedly connected to the front rotor cover plate 8 through a nut seat.

[0035] The two sides of the fuselage 1 are provided with telescopic wings 9 with variable lengths. The front half of the telescopic wings 9 is a rectangular wing, and the rear half of the telescopic wings 9 is a delta wing that retracts forward. The sweep angle of the telescopic wings 9 is 25°. The top and bottom surfaces of the fuselage 1 are provided with two symmetrically arranged storage grooves 10, and a resistance rudder 11 is rotatably connected in the storage groove 10. The resistance rudder 11 is rotatably connected to the storage groove 10 through a rotating shaft, and the rotating shaft is connected to the third rotary power device located in the fuselage 1.

[0036] When in use, the drone can be divided into three flight states: rotor state, fixed-wing state, and drag rudder state. Figure 1 , wherein the power direction of the rotor engine 4 is consistent with the axial direction of the front hole 2 and the rear hole 3, and the telescopic wing 9 is retracted to the minimum span, and the whole can be basically regarded as a symmetrical X-shaped four-rotor drone; the fixed-wing drone is as follows Figure 2 , wherein the rotatable duct 5 in the rear hole 3 rotates 90 degrees relative to the fuselage, driving the internal rotor engine 4 to change the power output direction backward, and the rotatable duct 5 in the front hole 2 will also be converted to a downward suspension state to change the power output direction of the internal rotor engine 4, and at the same time, the telescopic wing 9 is fully extended, and the rotor cover on the fuselage will slide forward to cover the two front holes 2 left in the rotor state, thereby improving the aerodynamic shape during fixed-wing flight; the state of the drag rudder 11 is as follows Figure 3 , wherein the rotor cover is slid downward, the four rotatable ducts 5 are rotated to the same position as in the rotor state, the four drag rudders 11 are opened and corresponding control actions are performed, and the telescopic wings 9 are extended to half, ensuring that the center of gravity is configured below the drag pressure center while not forming an excessively large fuselage drag surface. At this time, the UAV descends with its nose facing the sky.

[0037] By retracting the wings 9, the overall moment of inertia of the drone can be adjusted. When the drone retracts its wings 9 in the rotor state, the overall structure will be more compact, and the moment of inertia of the fuselage will be smaller, which will make it easier for the rotor to control the flight of the drone. Secondly, by changing the span, the lift of the fixed wing can be more smoothly "replaced" by the lift of the rotor. By adjusting the span at the same time during the rotor tilting process, the lift of the fixed wing of the drone can be changed more smoothly, thereby more effectively compensating for the rotor lift lost by the drone when the rotor tilts, making the transition from rotor to fixed wing smoother. Finally, by moving the retractable wings 9, the relative position of the center of gravity and the aerodynamic pressure center of the drone can be adjusted to make the drone stable in various states. For example, when the drone is cruising in level flight, fully unfolding the wings can further increase the distance between the lift pressure center and the center of gravity, and ensure that the center of gravity is in front of the lift pressure center, so that the drone has longitudinal static stability when cruising in level flight. For example, when a drone is performing an unpowered landing, by extending a portion of its wings, the drone's center of gravity can be kept below the center of pressure of air resistance, so that the drone's balance point when it is falling is itself a stable balance point, which is convenient for attitude control and vertical deceleration.

[0038] The wing type used in the UAV in this embodiment is N60R, and the parameters of the UAV structure are as follows:

[0039]

[0040] When the drone takes off, the telescopic wings 9 are folded, the front rotor cover 8 is pulled down, and the drone rises vertically in a similar rotor state. Then the rotor engine 4 gradually tilts, the telescopic wings 9 are extended, and the front rotor cover 8 is closed. The drone begins to cruise in fixed-wing state. The control technology of the rotor state is already very mature and will not be described in detail in this application. The stability control algorithm of the drone in cruise flight includes the following steps:

[0041] Step 1: Calculate the relationship between aerodynamic coefficient and span-angle of attack:

[0042] The aerodynamic coefficients include the lift coefficient and the pitch moment coefficient. The vortex lattice method is used in MATLAB to calculate the aerodynamic force when the wingspan length l is 0.2 to 2m and the angle of attack α is 0 to 10 degrees, and the changes of the lift coefficient and the pitch moment coefficient are given. The calculation results of the lift coefficient and the pitch moment coefficient are as follows: Figure 4 , Figure 5 Then, a binary quadratic function is used to fit the relationship between the aerodynamic coefficient and the span and angle of attack, and the following fitting formula is obtained:

[0043]

[0044] Among them, C L is the lift coefficient, m z is the pitching moment coefficient, l is the wingspan, α is the angle of attack, k1，L , k 2,L , k 3,L , k 4,L , k 5,L , k 6,L , k 1,m , k 2,m , k 3,m , k 4,m , k 5,m , k 6,m is the fitting coefficient in the fitting process;

[0045] The fitting results are shown in the following table:

[0046] Function fitting results of aerodynamic coefficients

[0047]

[0048] The fitting effect is as follows Figure 6 and Figure 7 The effects of quadratic function fitting and extrapolation are relatively good, so this form of function can be used to approximate the calculation of aerodynamic coefficients.

[0049] Step 2: Derive the UAV longitudinal dynamics model:

[0050] Assumptions: (1) The elastic factor is not considered and the object is defined as a rigid body. This is because the actual research problem is mainly the control problem of the UAV, and the influence of the elastic factor is ignored, so the study of the control problem is taken as the main research problem; (2) The displacement of the center of mass of the UAV, mass, and structural changes are not considered. This can simplify the complexity of the system to the greatest extent, but the defect is that it is easy to reduce the accuracy of the model to a certain extent; (3) The vertical take-off and landing UAV is a one-sided symmetrical structure; (4) Considering that the rotor engines on the left and right sides of the fuselage are in the same state at any time, the sum of the power of the four rotor engines is simplified to the state of a two-dimensional force system; (5) The ground coordinate system is stationary relative to the earth, that is, the ground coordinate system can be regarded as an inertial reference system; (6) By assuming that the sideslip angle, roll angle, track deviation angle, and roll angular velocity are zero, the dynamic model of the UAV is simplified to a longitudinal model that only considers movement in the longitudinal direction.

[0051] The force diagram of the UAV longitudinal model is as follows: Figure 8As shown. The ground coordinate system, body coordinate system and track coordinate system are selected in the figure. For the ground coordinate system, the center of mass A of the aircraft at takeoff is selected as the origin, the x-axis is the intersection of the track plane and the horizontal plane, and the direction pointing to the target is positive; the y-axis is upward along the vertical line, and the direction of the z-axis is determined by the right-hand rule. For the track coordinate system, the instantaneous center of mass O of the aircraft is the origin, and the x2 axis coincides with the velocity vector V of the aircraft; y2 is in the plumb plane containing the velocity vector V, perpendicular to the x2 axis, and pointing upward is positive; the direction of its z2 axis is determined according to the right-hand rule. For the body coordinate system, the center of mass O of the drone is also selected as the origin, the x3 axis is the longitudinal axis of the fuselage, and the direction pointing to the nose is positive; y3 is located on the symmetry plane of the aircraft, perpendicular to the y3 axis, and the direction of its z3 axis is determined according to the right-hand rule.

[0052] When the UAV is cruising, all the external forces acting on the UAV are mainly thrust P, drag D, lift L and gravity G. Since only the movement of the UAV in the longitudinal plane is considered, there is only one pitch moment M acting on its longitudinal model. z , this moment is about A z Axis direction.

[0053] According to the dynamic equation of the center of mass motion of the rigid body, the following equation is obtained:

[0054]

[0055]

[0056] in, and The algebraic representation of the components of the external forces (total aerodynamic force and gravity) on each axis of the track coordinate system. and It represents the thrust components on each axis of the track coordinate system.

[0057] From this, the final dynamic equation of the UAV center of mass motion can be obtained:

[0058]

[0059]

[0060] In order to determine the motion trajectory of the drone's center of mass relative to the ground coordinate system, it is necessary to establish the kinematic equation of the drone's center of mass relative to the ground coordinate system, as shown below:

[0061]

[0062]

[0063] According to the relationship between the Euler angles of the drone - pitch angle is equal to the sum of the angle of attack and the track inclination, and the attitude angle relationship can be obtained:

[0064]

[0065] According to the kinematic equation of the drone rotating around the center of mass, the following equation is obtained:

[0066]

[0067] According to the dynamic equation of the drone rotating around the center of mass, the following equation is obtained:

[0068]

[0069] In summary, the mathematical expression of the longitudinal dynamics model of the UAV is:

[0070]

[0071] Among them, V is the flight speed of the drone; t is time; P is thrust; α is the angle of attack; δ is the deflection angle of the rotor engine in the rear hole. In the fixed-wing state, the rotor engine in the rear hole is the main power output, and the rotor engine in the front hole is more used in the lifting state. It is auxiliary power in other stages; m is the mass of the drone; X is the drag; g is the gravity coefficient; θ is the track inclination; Y is the lift; y is the ground coordinate system A with the center of mass A of the drone at takeoff as the origin y The movement distance on the axis; x is the ground coordinate system with the center of mass A of the drone when it takes off as the origin. x Distance of movement on the axis; is the pitch angle; ω z is the pitch angular velocity; M z is the pitch moment; J z is the moment of inertia.

[0072] The expressions of drag, lift and pitch moment on the aircraft are:

[0073]

[0074] Where S is the extended area of ​​the telescopic wing, b is the aerodynamic mean chord length, and the lift coefficient C is L , and the pitch moment coefficient m z The value of is calculated in real time at different angles of attack using the fitting formula obtained in step 1. The drag coefficient C D It is given by the following formula:

[0075] C D =0.0012+0.0045α 2 .

[0076] Different thrust deflection angles, i.e. rear rotor deflection angles, are selected to simulate the longitudinal dynamics model of the UAV and observe the system response curve under uncontrolled conditions. The physical parameters of the UAV are set as:

[0077] Mass: m = 2.3 kg;

[0078] Pitch moment of inertia: J = 8kg*m 2 ;

[0079] Lift coefficient and pitch moment coefficient: C L 、m z According to the fitting formula in the previous article, the real-time calculation at different angles of attack is performed, and the span is 2m;

[0080] Drag coefficient: C D ;

[0081] Air density: ρ = 1.29 kg / m 3 ;

[0082] Gravitational acceleration: g = 9.8 m / s 2

[0083] The initial value of the drone is set as follows

[0084] Initial height: y0 = 500m;

[0085] Initial horizontal position: x0 = 0m;

[0086] Initial velocity: v0 = 10 m / s;

[0087] Initial pitch angle: θ0 = 0 rad;

[0088] Initial track inclination:

[0089] Initial pitch angular velocity: ω0 = 0 rad / s.

[0090] First, if Fig.14 , select the rear rotor deflection angle δ=0°, 30°, -30°, perform simulation and compare the analysis results. As can be seen from the figure, the UAV is stable in a large range when it is uncontrolled, and the change of δ affects the speed and pitch angle of the UAV in the equilibrium state, and affects whether the UAV has fluctuations in speed and pitch angle in the system response. As can be seen from the trajectory diagram, the UAV will eventually have two states of flying up and flying down under these three rear rotor deflection angles, which confirms that it is feasible to control the longitudinal flight trajectory of the UAV by the change of δ. From the attitude angle and track inclination change diagram, it can be seen that the attitude angle and track inclination of the UAV will converge quickly, which also makes the angle of attack change always small. When δ=30°, the pitch angle and speed will experience a long period of fluctuations before reaching equilibrium, and the equilibrium speed is the smallest under the same thrust.

[0091] In order to further observe the response of the UAV when δ is large, the present patent application further performs dynamic simulation when δ = 0°, 40°, and 80° and plots the results as shown in the figure. Fig.15 ,from Fig.15 It can be seen that when δ=80° is further increased, the pitch angle of the UAV in equilibrium state becomes negative, and the speed in equilibrium is greatly reduced. This also reveals that the relationship between the rear rotor deflection angle and the pitch angle is not a simple linear relationship, but a more complex connection.

[0092] Next, consider the controlled UAV flight. When using traditional PID control, let the initial δ = 0°, set different pitch angle targets for the control system, and run the simulation. First, give the following set of control targets θ r =0°, 10°, -10°, the simulation results are as follows Fig.16 ,As can be seen from the figure, under such a set of targets, the control system can ,make the UAV reach a stable state and fly according to the target track ,angle.

[0093] Try the following control target θ again r =0°, 30°, -30°, the simulation results are as follows Fig.17 As can be seen from the figure, when the expected track inclination angle is 300, the traditional PID attitude controller has failed. The essential reason for the failure is the nonlinear behavior between the rear rotor angle and the equilibrium track inclination angle when δ is large as mentioned above. r =-30°, the controller does not fail completely because it still has an approximate linear relationship when δ is negative.

[0094] The above argument is Fig. 9 It can be shown more clearly in Fig. 9 The speed, track inclination and angle of attack of the UAV in a balanced state are compared with the parameter values ​​when the simulation results of the uncontrolled longitudinal model reach stability under different rotor engine deflection angles, that is, rear rotor deflection angles. As can be seen from the figure, when δ>-30°, the rear rotor deflection angle δ and the equilibrium track inclination angle are linear or approximately linear, whether it is the simulation result or the solution equilibrium point result, and when δ<-30°, the equilibrium track inclination angle drops sharply, showing strong nonlinearity. This is also the main reason for the failure of traditional PID controllers. Therefore, this project proposes step 3.

[0095] Step 3: Design the stability control system of the UAV in level flight cruise based on eigenvalue configuration:

[0096] (1) The longitudinal dynamics model of the UAV obtained in step 2 is a nonlinear dynamics model. The nonlinear dynamics model is converted into a set of linear models by the Jacobi linearization method. The specific method is as follows: in the longitudinal dynamics model of the UAV, let the flight speed V and pitch angle of the UAV be The derivative of the track inclination angle θ with respect to time is 0. The longitudinal dynamics model of the UAV is converted into an algebraic equation. By solving the algebraic equation, the rotor engine deflection angle δ0 corresponding to each target state q0 equilibrium can be solved. The longitudinal dynamics model of the UAV is Taylor expanded at the equilibrium point corresponding to q0 and δ0, and the Jacobian matrix of the UAV at the equilibrium point is calculated. According to the Jacobian matrix, the linear approximate model of the UAV near the equilibrium point is obtained, which is usually expressed in the state space form as follows:

[0097]

[0098] Where q is the state variable, and is the state space equation, and

[0099] q0 is the target state of control; δ is the control input; δ0 is the rotor engine deflection angle corresponding to the UAV when it is balanced at q0, which is solved by balancing the forces in the x and y directions and the torque in the z direction; A and B are Jacobian matrices, and f i is the i-th component of the vector-valued function f, q j is the jth component of the state variable q, u j is the jth component of the input variable u. For this problem, u has only one component δ;

[0100] At the same time, define the control quantity change equation:

[0101]

[0102] Where K is a proportional coefficient; C is a coefficient matrix with 1 row and 6 columns, which has the following form:

[0103] C = [C1C2C3C4C5C6];

[0104] The above linear approximate model and the control variable change equation are used to organize the expression of a linear equation group as follows:

[0105]

[0106] (2) Based on the above linear equations, the stability analysis of the system is carried out. The sensor of the feedback state variable is used to obtain the value of the state variable q. Then, by setting the values ​​of the seven coefficients C1~C6 and K, that is, the value of the last row of the matrix T, the real part of the eigenvalue of the matrix T is all negative real numbers, and the system is in a stable state. The process of the drone landing from high altitude to the ground is divided into the unpowered stage and the rotor righting stage. In the unpowered stage, the drag rudder is opened and the rotor engine is not started. The entire drone is controlled by aerodynamic drag and drag rudder deflection for angle and attitude-position control; the second stage is when the drone is basically close to the ground, the rotor engine is started, and the drag rudder is immediately retracted. Subsequently, under the action of the rotor engine, the drone is righted and further decelerated until it lands safely. This stage is referred to as the "rotor righting stage".

[0107] Simplify the drone during landing into a two-dimensional longitudinal model:

[0108] 1. When the drone starts to fall, it has no angular velocity around its own x-axis and y-axis, and the xoy plane of the fuselage is perpendicular to the horizontal plane.

[0109] 2. The local wind speed is 0, and its disturbance effect on the drone can be ignored

[0110] 3. There is no differential between the two speed brakes on the same side of the xoy plane of the fuselage connection system, that is, it can be assumed that the air resistance to the speed brake will not generate torque around the x-axis and y-axis.

[0111] In this way, the falling drone can be simplified to a two-dimensional dynamic problem in the xoy plane of the fuselage. However, even if it is simplified to a two-dimensional problem, the shape of the drone is not just a flat plate. In fact, although the actual shapes of the fuselage and the drag rudder are relatively complex, when the problem is simplified to Fig.10 After the two-dimensional analysis, we can analyze the UAV system in the falling process as equivalent to three mutually hinged plates with different resistance areas and lengths in the plane. Fig.10 In the figure, the larger plate in the middle is the equivalent plate of the drag surface of the fuselage, referred to as the fuselage plate; the small plates on both sides represent the equivalent plates of the drag surface of the drag rudder, referred to as the drag plate; the airflow deflection angle is γ (counterclockwise deflection is positive); the fuselage angle is The two drag fins are now opened at angles α1 and α2.

[0112] The unpowered landing attitude-position control algorithm specifically includes the following steps:

[0113] Step 1: Obtain the aerodynamic drag estimation equation:

[0114] (1) According to the semi-empirical formula based on the mean momentum equation, the total aerodynamic drag of the drag plate on the side of the UAV where the shielding effect occurs, the total aerodynamic drag on the side of the UAV where the pre-rotation effect occurs, and the fuselage drag are obtained:

[0115] The total aerodynamic drag f on the side of the drone where the shielding effect occurs a1 The calculation formula is:

[0116]

[0117] in, is the fuselage angle, ρv 2 is the momentum current density, S a is the area of ​​the equivalent flat plate of the two drag rudders on the side of the UAV with pre-rotation effect or the two drag rudders on the side of the UAV with shielding effect, α1 is the opening angle of the drag rudder on the side of the UAV with shielding effect, γ is the airflow deflection angle, is the attenuation coefficient caused by the shading effect;

[0118] The total aerodynamic drag f on the side of the drone where the pre-rotation effect occurs a2 The calculation formula is:

[0119]

[0120] k m =max{0,k};

[0121]

[0122] Among them, k is the attenuation coefficient caused by the pre-rotation effect, and α2 is the opening angle of the drag rudder on the side of the UAV where the pre-rotation effect occurs;

[0123] Fuselage resistance f b The calculation formula is:

[0124]

[0125] Among them, S b is the area of ​​the equivalent flat plate of the fuselage's drag acting surface;

[0126] (2) Simplifying the aerodynamic drag of the two speed brakes to the center of mass, the simplified expression of the aerodynamic drag on the UAV to the center of mass is as follows:

[0127]

[0128] Among them, F x ,F y and M zThey are the x-direction force, y-direction force and moment around the z-axis perpendicular to the paper plane on the center of mass of the drone after the force system is simplified. a is the distance from the center of gravity of the resistance on the equivalent flat plate of the resistance action surface of the resistance rudder to the hinge point, l b The distance between the drag pressure center and the center of mass of the equivalent flat plate of the drag action surface of the fuselage;

[0129] (3) Assuming that the air has no self-flow and ignoring the shielding effect and pre-rotation effect, the attenuation coefficient caused by the shielding effect and the attenuation coefficient caused by the pre-rotation effect are recorded as 1, and a simple aerodynamic drag estimation equation is given:

[0130]

[0131] Among them, v represents the velocity at the current moment, γ is the velocity deflection angle, which is also the airflow deflection angle;

[0132] Step 2: Obtain the linear equations about α1, α2:

[0133] After many simulation experiments, it was found that when the resistance plate in the positive direction of the z-axis opens at an angle α + Equal to the opening angle α of the resistance plate in the negative direction of the z-axis - equal When , the UAV has a relatively stable balance point. Therefore, this balance point is selected to linearize the above aerodynamic drag estimation equation to obtain:

[0134]

[0135] Among them, K c is the function of aerodynamic drag F c The Jacobian matrix of , Δ represents the difference between the current value of the physical quantity and the value at the equilibrium point; next, cross out the second row of the matrices on both sides, and further organize the formula into a linear equation system about α1, α2:

[0136]

[0137] in:

[0138]

[0139] After linearization, when the matrix A c When non-singular, for each set F x and M z , in the corresponding The corresponding α1 and α2 can be solved under v and γ;

[0140] Step 3: Perform unpowered landing attitude control:

[0141] (1) Generate the expected force system at each moment using the following formula: and

[0142]

[0143] In the formula, e x Indicates the difference between the current x position and the target position, e v Represents the current speed v and The difference, Indicates the current Difference from target fuselage angle, P1 1 ,I1 1 ,D1 1 ,P1 2 ,I1 2 ,D1 2 ,P2,I2,D2 are all PID control coefficients, which are selected according to actual control needs;

[0144] (2) The above generated and Substitute into the linear equations about α1, α2 and solve to get the expected drag rudder opening angle and

[0145] (3) Build a proportional controller as follows to control the drag rudder to the expected drag rudder opening angle and limit the rudder deflection speed to form an attitude controller to prevent instability;

[0146]

[0147] P α is the proportional coefficient of the attitude controller, which is selected according to the actual control requirements;

[0148] Step 4: Perform unpowered landing position control:

[0149] (1) Ignoring the inertia of the drag rudders on both sides, the dynamic equation of the UAV is written as:

[0150]

[0151] Where m is the total mass of the UAV, J z is the moment of inertia of the drone around the z-axis of the fuselage, g is the local gravity acceleration, and F x ,F y and M z They are all given by the simplified expression of the aerodynamic drag on the center of mass of the UAV in step 1 (2);

[0152] (2) The generated and Substitute into the above UAV dynamics equation and solve to get the expected x r and Complete the construction of the position controller;

[0153] Step 5: Use the attitude controller as the inner loop and the position controller as the outer loop to control and Limits are imposed to ensure that the linear equations about α1 and α2 are within the applicable range, thereby ensuring better attitude stability control and position control in the unpowered stage.

[0154] A landing attitude-position control algorithm for the rotor righting phase of a tilt-rotor UAV with an unpowered landing function: The process of a UAV landing from a high altitude to the ground is divided into an unpowered phase and a rotor righting phase. In the rotor righting phase, when the UAV is basically close to the ground, the rotor engine is started and the drag rudder is retracted. Subsequently, under the action of the rotor engine, the UAV is righted and further decelerated until it lands safely.

[0155] The landing attitude-position control algorithm during the rotor righting phase specifically includes the following steps:

[0156] Step 1: Generate the expected control force at each moment using the following formula and expected control torque

[0157] Among them, e y Indicates the difference between the current y position and the target position. Indicates the current The difference with the target fuselage angle, P3, P4, I3, I4, D3, D4 are all PID control coefficients, which are selected according to actual control needs;

[0158] Step 2: After the unpowered falling stage, when the UAV is very close to the ground, the rotor engine must be started to straighten the fuselage and further decelerate. At this time, the rotor engine tilts again, and the power output direction is perpendicular to the fuselage surface. The force is positive when it points from the bottom of the fuselage to the fuselage. The power configuration equations are as follows:

[0159]

[0160] This system of equations is a linear system of equations about F1 and F2, where the forces acting on the front and rear rotors are F1 and F2, respectively, and the distances from the center of mass are l1 and l2, respectively;

[0161] Step 3: Convert the expected control force from step 1 to and expected control torque Substitute this into F from step 2 y and M z By solving F1 and F2, the attitude righting and landing control of the drone can be achieved.

[0162] The above power configuration equations and control algorithms are simulated using MATLAB to verify the control effect. Several typical initial values ​​of deviation from equilibrium are selected to test whether the UAV can maintain stability during the falling process.

[0163] The PID parameters are set to

[0164]

[0165]

[0166] P2=3.4, I2=0.002, D2=0.3;

[0167] P3=300, I3=0, D3=200;

[0168] P4=700, I4=0, D4=280;

[0169] Total mass of the drone: m = 2.3 kg;

[0170] The moment of inertia of the drone around the z-axis of the fuselage: J z ;

[0171] Local gravity acceleration: g = 9.8 m / s 2 ;

[0172] The switching rule is to use rotor righting and deceleration when the aircraft is 50m above the ground.

[0173] The simulation results of the entire landing process are as follows. Three groups of typical initial values ​​are selected as shown in the following table:

[0174]

[0175] The simulation results of the above three groups of treatments are as follows: Figure 11-13 It can be seen that the UAV is able to complete the landing in the unpowered stage and the rotor righting stage under the action of the above control algorithm, which also confirms the feasibility of the control of the new UAV.

[0176] Therefore, the present invention adopts the above-mentioned structure of the tilt-rotor UAV with the function of unpowered landing, which can eventually achieve a smooth vertical take-off through special structural design and the application of control algorithms. After switching to a level flight cruising state, the fixed wing provides lift, and the UAV can land safely without consuming almost any of its own power during landing.

[0177] Finally, it should be noted that the above embodiments are only used to illustrate the technical solution of the present invention rather than to limit it. Although the present invention has been described in detail with reference to the preferred embodiments, those skilled in the art should understand that they can still modify or replace the technical solution of the present invention with equivalents, and these modifications or equivalent replacements cannot cause the modified technical solution to deviate from the spirit and scope of the technical solution of the present invention.

Claims

1. A tilt-rotor UAV with an unpowered landing function, characterized in that: The invention comprises a fuselage with a flying wing layout, wherein the front end of the fuselage is provided with two symmetrically arranged front holes, and the rear end of the fuselage is provided with two symmetrically arranged rear holes, wherein the front holes and the rear holes are connected to the rotor engine through a rotatable duct, and the power direction of the rotor engine is consistent with the axial direction of the rotatable duct, a slide rail is provided on the center line of the fuselage, and a front rotor cover plate adapted to the front hole structure is slidably connected to the slide rail, telescopic wings with variable extension are provided on both sides of the fuselage, and the top and bottom surfaces of the fuselage are provided with two symmetrically arranged storage grooves, and a drag rudder is rotatably connected in the storage groove.

2. The tilt-rotor UAV with unpowered landing function according to claim 1, characterized in that: The front half of the telescopic wing is a rectangular wing, the rear half of the telescopic wing is a forward-folded delta wing, and the sweep angle of the telescopic wing is 25°.

3. The tilt-rotor UAV with unpowered landing function according to claim 2, characterized in that: Both ends of the rotatable duct are rotatably connected to the inner walls of the front hole and the rear hole via connecting rods, and one of the connecting rods is connected to a first rotating power device located in the body.

4. The tilt-rotor UAV with unpowered landing function according to claim 3, characterized in that: A screw is rotatably connected in the slide rail, one end of the screw is connected to a second rotary power device located in the body, and the screw is fixedly connected to the front rotor cover plate via a nut seat.

5. The tilt-rotor UAV with unpowered landing function according to claim 4, characterized in that: The resistance rudder is rotationally connected to the storage groove via a rotating shaft, and the rotating shaft is connected to a third rotating power device located in the body.

6. The stability control algorithm of a tilt-rotor UAV with unpowered landing function in level flight cruise according to any one of claims 1 to 5, characterized in that: The following steps are involved: Step 1: Calculate the relationship between aerodynamic coefficient and span-angle of attack: The aerodynamic coefficients include the lift coefficient and the pitch moment coefficient. The vortex lattice method is used in MATLAB to calculate the aerodynamic force when the wingspan length l is 0.2 to 2 m and the angle of attack α is 0 to 10 degrees, and the changes of the lift coefficient and the pitch moment coefficient are given. Then, the relationship between the aerodynamic coefficients and the span length and the angle of attack is fitted with a binary quadratic function. The following fitting formula is used: C L =k 1,L ·l 2 +k 2,L ·l+k 3,L ·α 2 +k 4,L ·α+k 5,L ·lα+k 6,L m z =k 1,m ·l 2 +k 2,m ·l+k 3,m ·α 2 +k 4,m ·α+k 5,m ·lα+k 6,m ; Among them, C L is the lift coefficient, m z is the pitching moment coefficient, l is the wingspan, α is the angle of attack, k 1，L , k 2,L , k 3,L , k 4,L , k 5,L , k 6,L , k 1,m , k 2,m , k 3,m , k 4,m , k 5,m , k 6,m is the fitting coefficient in the fitting process; Step 2: Derive the UAV longitudinal dynamics model: The mathematical expression of the UAV longitudinal dynamics model is: Among them, V is the flight speed of the drone, t is time, P is thrust, α is the angle of attack, δ is the deflection angle of the rotor engine in the rear hole, m is the mass of the drone, X is the drag, g is the gravity coefficient, θ is the track inclination, Y is the lift, and y is the ground coordinate system A with the center of mass A of the drone at takeoff as the origin. y The movement distance on the axis, x is the ground coordinate system A with the center of mass A of the drone at takeoff as the origin x The movement distance on the axis, θ is the pitch angle, ω z is the pitch angular velocity, M z is the pitching moment, J z is the moment of inertia; The expressions of drag, lift and pitch moment on the aircraft are: Where S is the extended area of ​​the telescopic wing, b is the aerodynamic mean chord length, and the lift coefficient C is L , and the pitch moment coefficient m z The value of is calculated in real time at different angles of attack using the fitting formula obtained in step 1. The drag coefficient C D Given by the following formula: C D ＝0.0012+0.0045α 2 ； Step 3: Design the stability control system of the UAV in level flight cruise based on eigenvalue configuration: (1) The UAV longitudinal dynamics model obtained in step 2 is a nonlinear dynamics model. The nonlinear dynamics model is converted into a set of linear models by the Jacobian linearization method. The specific method is as follows: first, in the UAV longitudinal dynamics model, the derivatives of the UAV flight speed V, pitch angle θ, and track inclination angle θ with respect to time are all set to 0, and the UAV longitudinal dynamics model is converted into an algebraic equation. By solving the algebraic equation, the rotor engine deflection angle δ0 corresponding to each target state q0 equilibrium can be solved. The UAV longitudinal dynamics model is Taylor expanded at the equilibrium point corresponding to q0 and δ0, and the Jacobian matrix of the UAV at the equilibrium point is calculated. According to the Jacobian matrix, the linear approximate model of the UAV near the equilibrium point is obtained, which is usually expressed in the state space form as follows: Where q is the state variable, and is the state space equation, and q0 is the target state of control; δ is the control input; δ0 is the rotor engine deflection angle corresponding to the UAV when it is balanced at q0, which is solved by balancing the forces in the x and y directions and the torque in the z direction; A and B are Jacobian matrices, and f i is the i-th component of the vector-valued function f, q j is the jth component of the state variable q, u j is the jth component of the input variable u. For this problem, u has only one component δ; At the same time, define the control quantity change equation: Where K is a proportional coefficient; C is a coefficient matrix with 1 row and 6 columns, which has the following form: C = [C1 C2 C3 C4 C5 C6]; The above linear approximate model and the control variable change equation are used to organize the expression of a linear equation group as follows: (2) Based on the above linear equations, the system stability analysis is performed. The value of the state variable q is obtained by using the sensor of the feedback state variable. Then, by setting the values ​​of the seven coefficients C1 to C6 and K, that is, the values ​​of the last row of the matrix T, the real parts of the eigenvalues ​​of the matrix T are all negative real numbers, and the system is in a stable state.

7. The control algorithm of the tilt-rotor UAV with unpowered landing function according to claim 6 is characterized in that: Assumptions in step 2: (1) The elastic factor is not considered and the object is defined as a rigid body. This is because the actual research problem is mainly the control problem of the UAV, and the influence of the elastic factor is ignored, so the study of the control problem is taken as the main research problem; (2) The displacement of the center of mass of the UAV, mass, and structural changes are not considered. This can simplify the complexity of the system to the greatest extent, but the defect is that it is easy to reduce the accuracy of the model to a certain extent; (3) The vertical take-off and landing UAV is a one-sided symmetrical structure; (4) Considering that the rotor engines on the left and right sides of the fuselage are in the same state at any time, the sum of the power of the four rotor engines is simplified to the state of a two-dimensional force system; (5) The ground coordinate system is stationary relative to the earth, that is, the ground coordinate system can be regarded as an inertial reference system; (6) By assuming that the sideslip angle, roll angle, track deviation angle, and roll angular velocity are zero, the dynamic model of the UAV is simplified to a longitudinal model that only considers movement in the longitudinal direction.

8. The unpowered landing attitude-position control algorithm of a tilt-rotor UAV with an unpowered landing function according to any one of claims 1 to 5, characterized in that: The process of a drone landing from high altitude to the ground is divided into an unpowered stage and a rotor righting stage. In the unpowered stage, the drag rudder is opened, the rotor engine is not started, and the entire drone relies only on aerodynamic drag and drag rudder deflection for angle and attitude-position control; The second stage is when the UAV is basically close to the ground, the rotor engine is started, and the drag rudder is retracted immediately. Then, under the action of the rotor engine, the UAV is straightened and further decelerated until it lands safely. This stage is referred to as the "rotor straightening stage"; The unpowered landing attitude-position control algorithm specifically includes the following steps: Step 1: Obtain the aerodynamic drag estimation equation: (1) According to the semi-empirical formula based on the mean momentum equation, the total aerodynamic drag of the drag plate on the side of the UAV where the shielding effect occurs, the total aerodynamic drag on the side of the UAV where the pre-rotation effect occurs, and the fuselage drag are obtained: The total aerodynamic drag f on the side of the drone where the shielding effect occurs a1 The calculation formula is: in, is the fuselage angle, ρv 2 is the momentum current density, S a is the area of ​​the equivalent flat plate of the two drag rudders on the side of the UAV with pre-rotation effect or the two drag rudders on the side of the UAV with shielding effect, α1 is the opening angle of the drag rudder on the side of the UAV with shielding effect, γ is the airflow deflection angle, is the attenuation coefficient caused by the shading effect; The total aerodynamic drag f on the side of the drone where the pre-rotation effect occurs a2 The calculation formula is: k m =max{0,k}; Among them, k is the attenuation coefficient caused by the pre-rotation effect, and α2 is the opening angle of the drag rudder on the side of the UAV where the pre-rotation effect occurs; Fuselage resistance f b The calculation formula is: Among them, S b is the area of ​​the equivalent flat plate of the fuselage's drag acting surface; (2) Simplifying the aerodynamic drag of the two speed brakes to the center of mass, the simplified expression of the aerodynamic drag on the UAV to the center of mass is as follows: Among them, F x ,F y and M z They are the x-direction force, y-direction force and moment around the z-axis perpendicular to the paper plane on the center of mass of the drone after the force system is simplified. a is the distance from the center of gravity of the resistance on the equivalent flat plate of the resistance action surface of the resistance rudder to the hinge point, l b The distance between the drag pressure center and the center of mass of the equivalent flat plate of the drag action surface of the fuselage; (3) Assuming that the air has no self-flow and ignoring the shielding effect and pre-rotation effect, the attenuation coefficient caused by the shielding effect and the attenuation coefficient caused by the pre-rotation effect are recorded as 1, and a simple aerodynamic drag estimation equation is given: Among them, v represents the velocity at the current moment, γ is the velocity deflection angle, which is also the airflow deflection angle; Step 2: Obtain the linear equations about α1, α2: After many simulation experiments, it was found that when the resistance plate in the positive direction of the z-axis opens at an angle α + Equal to the opening angle α of the resistance plate in the negative direction of the z-axis - equal When , the UAV has a relatively stable balance point. Therefore, this balance point is selected to linearize the above aerodynamic drag estimation equation to obtain: Among them, K c is the function of aerodynamic drag F c The Jacobian matrix of , Δ represents the difference between the current value of the physical quantity and the value at the equilibrium point; next, cross out the second row of the matrices on both sides, and further organize the formula into a linear equation system about α1, α2: in: After linearization, when the matrix A c When non-singular, for each set F x and M z , in the corresponding The corresponding α1 and α2 can be solved under v and γ; Step 3: Perform unpowered landing attitude control: (1) Generate the expected force system at each moment using the following formula: and In the formula, e x Indicates the difference between the current x position and the target position, e v Represents the current speed v and The difference, Indicates the current Difference from target fuselage angle, P1 1 ,I1 1 ,D1 1 ,P1 2 ,I1 2 ,D1 2 ,P2,I2,D2 are all PID control coefficients, which are selected according to actual control needs; (2) The above generated and Substitute into the linear equations about α1, α2 and solve to get the expected drag rudder opening angle and (3) Build a proportional controller as follows to control the drag rudder to the expected drag rudder opening angle and limit the rudder deflection speed to form an attitude controller to prevent instability; P α is the proportional coefficient of the attitude controller, which is selected according to the actual control requirements; Step 4: Perform unpowered landing position control: (1) Ignoring the inertia of the drag rudders on both sides, the dynamic equation of the UAV is written as: Where m is the total mass of the UAV, J z is the moment of inertia of the drone around the z-axis of the fuselage, g is the local gravity acceleration, and F x ,F y and M z They are all given by the simplified expression of the aerodynamic drag on the center of mass of the UAV in step 1 (2); (2) The generated and Substitute into the above UAV dynamics equation and solve to get the expected x r and Complete the construction of the position controller; Step 5: Use the attitude controller as the inner loop and the position controller as the outer loop to control and Limits are imposed to ensure that the linear equations about α1 and α2 are within the applicable range, thereby ensuring better attitude stability control and position control in the unpowered stage.

9. The landing attitude-position control algorithm of the tilt-rotor UAV with unpowered landing function in the rotor righting phase according to any one of claims 1 to 5, characterized in that: The process of a drone landing from a high altitude to the ground is divided into an unpowered stage and a rotor righting stage. In the rotor righting stage, when the drone is basically close to the ground, the rotor engine is started and the drag rudder is retracted. Subsequently, under the action of the rotor engine, the drone is righted and further decelerated until it lands safely. The landing attitude-position control algorithm during the rotor righting phase specifically includes the following steps: Step 1: Generate the expected control force at each moment using the following formula and expected control torque Among them, e y Indicates the difference between the current y position and the target position. Indicates the current The difference with the target fuselage angle, P3, P4, I3, I4, D3, D4 are all PID control coefficients, which are selected according to actual control needs; Step 2: After the unpowered falling stage, when the UAV is very close to the ground, the rotor engine must be started to straighten the fuselage and further decelerate. At this time, the rotor engine tilts again, and the power output direction is perpendicular to the fuselage surface. The force is positive when it points from the bottom of the fuselage to the fuselage. The power configuration equations are as follows: This system of equations is a linear system of equations about F1 and F2, where the forces acting on the front and rear rotors are F1 and F2, respectively, and the distances from the center of mass are l1 and l2, respectively; Step 3: Convert the expected control force from step 1 to and expected control torque Substitute this into F from step 2 y and M z By solving F1 and F2, the attitude righting and landing control of the drone can be achieved.

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