Active twist device for the wings of an ornithopter and method for controlling it

Abstract

The application provides a wing active torsion device of a flapping-wing aircraft and a control method thereof. The wing active torsion device is arranged on a wing of the flapping-wing aircraft, and is used for adjusting the aerodynamic force generated by a flapping-wing system by changing the wing shape of the aircraft during flight. According to the control method based on an extended observer, the disturbance of the aircraft pitching attitude caused by the unsteady aerodynamic force of the flapping-wing system is estimated, the disturbance is compensated by the wing active torsion device and the elevator of a tail plane, and the attitude control is completed. The technical scheme of the application can reduce the attitude oscillation amplitude caused by the aerodynamic force fluctuation in the attitude control process of the flapping-wing aircraft, effectively reduces the difficulty of the attitude control, and reduces the deflection amplitude of the conventional aerodynamic control surface, and improves the control efficiency.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aircraft technology, in particular to a wing active torsion device of flapping-wing aircraft and a control method thereof. BACKGROUND

[0002] Flapping-wing aircraft is a kind of aircraft developed by imitating the flight mode of insects or birds in nature. Flapping-wing aircraft designed based on the bionics principle has the advantages of small size, light weight and good concealment, and is a popular direction in the field of micro-aircraft in recent years. At the same time, flapping-wing movement can generate lift and thrust at the same time, which has many potential advantages compared with traditional fixed-wing and rotary-wing.

[0003] Most of the traditional flapping-wing aircrafts adopt the bird-like horizontal flight mode, and such flapping-wing aircrafts are usually influenced by traditional fixed-wing aircrafts, adopt the traditional fixed-wing layout form, and use the tail to control the attitude. However, the most significant and biggest control difficulty of flapping-wing aircraft is that the aerodynamic force generated by the flapping-wing system is not stable, but is unsteady and changes with the flapping process. Therefore, the bird-like flapping-wing aircraft usually accompanies the periodic oscillation of pitch and height during flight. Therefore, introducing control variables other than the tail is the current research hotspot of bird-like flapping-wing aircraft. The introduction of the wing active torsion device of flapping-wing aircraft makes it possible to fundamentally solve the longitudinal oscillation problem of bird-like flapping-wing aircraft.

[0004] The existing bird-like flapping-wing aircraft usually adopts manual control or PID control automatic control method for attitude control, although both methods can make the bird-like flapping-wing aircraft fly for a certain distance. However, since no special control rate design is made according to the characteristics of the periodic aerodynamic force of the flapping-wing aircraft, the aircraft attitude oscillation is still very common in the flapping period, especially when the flapping frequency is small. The flapping frequency of birds is very small when flying in the cruise state, which makes the bionic characteristics of the bird-like flapping-wing aircraft and the control rate design have a certain conflict. Therefore, how to design a control rate to make the bird-like flapping-wing aircraft have good periodic oscillation resistance and reduce the burden of the controller as much as possible is a subject worthy of further study. SUMMARY

[0005] In order to solve the problems of the prior art, the present application provides a wing active torsion device of flapping-wing aircraft and a control method thereof, which can reduce the burden of the control surface on the tail, reduce the deflection amplitude of the elevator, effectively reduce the possibility that the control surface loses control ability due to reaching the stall angle of attack, and improve the efficiency of the control surface.

[0006] This invention provides an active wing twisting device for a flapping-wing aircraft, comprising a support frame with a flexible skin. The support frame includes an inner wing rib, an outer wing rib, and a middle wing rib connected in series via a wing spars. The inner end of the wing spars is connected to a flapping mechanism within the fuselage of the flapping-wing aircraft. The portion of the skin covering the upper side of the wing rib extends to the rear end of the wing rib and exceeds the support range of the wing rib. The excess portion deflects vertically around the forward section of the wing. A servo motor is fixed to the middle wing rib and remains relatively stationary. The drive shaft of the servo motor is connected to a metal transmission rod via a servo motor rocker arm. The two ends of the metal transmission rod are bent and pass through holes at the ends of the servo motor rocker arm and the lug support, respectively. The lug support is bonded to the flexible skin. The transmission rod, the servo motor rocker arm, and the lug support constitute a planar four-bar linkage.

[0007] The flexible skin wraps around the leading edge of the rib, and the lower side of the leading edge of the flexible skin extends to one-quarter of the length of the rib chord.

[0008] The parts of the skin that come into contact with the wing ribs are all glued together. At the wing root of the entire wing section, the skin is connected to the fuselage of the flapping-wing aircraft.

[0009] The present invention also provides a control method for the active twisting device of the wing of a flapping-wing aircraft, comprising the following steps:

[0010] Step 1) Construct a longitudinal dynamic model of the flapping-wing aircraft. Based on the longitudinal dynamic model, obtain the simplified longitudinal system dynamic equations, specifically:

[0011]

[0012] Where θ is the aircraft's pitch angle. The pitch rate of the aircraft. This is a fictitious quantity that can be observed using an observer.

[0013] Step 2) Construct an extended observer based on this dynamic model;

[0014] Step 3) The attitude data measured by the inertial measurement unit is used to obtain the real-time disturbance of the aircraft using the extended observer. Based on the obtained discrete extended observer dynamic equation, the total disturbance in the pitch angle of the aircraft is obtained.

[0015] Step 4) Combine the linear control law and real-time disturbance to form the control quantity of the aircraft, complete the attitude control of the flapping-wing aircraft, and allocate the actual control quantity between the two actuators.

[0016] In step 1), during the construction of the longitudinal dynamic model of the flapping-wing aircraft, the longitudinal attitude motion of the flapping-wing aircraft is simplified into a second-order differential equation with respect to the pitch angle:

[0017] in, Let b1 be the angular acceleration of the aircraft pitch angle, b2 be the proportional control coefficient of the aircraft tail for the pitch angle, u1 be the actual tail deflection angle, b2 be the proportional control coefficient of the wing active twist device for the pitch angle, u2 be the actual wing active twist device deflection angle, and f be the total disturbance.

[0018] Step 2) describes the process of building the extended observer as follows:

[0019]

[0020] Where z1 is the observer's estimate of the pitch angle θ, and z2 is the observer's estimate of the pitch angular velocity. The estimated value is z3, which is the observer's estimate of the total disturbance f. β1, β2, and β3 are adjustable parameters. When the parameters are well tuned, z1→θ can be achieved. z1, z2 and z3 respectively track and observe the pitch angle, pitch rate and total disturbance in the pitch angle of the aircraft.

[0021] Step 3) describes the method for obtaining real-time disturbances of the aircraft as follows:

[0022] 3.1) is the pitch angle of the aircraft, which is a physical quantity measured by inertial sensors in the flight control system;

[0023] 3.2) This is the control signal sent by the flight control system to the servo motor on the tail fin that controls the elevator. This signal determines the deflection angle of the elevator.

[0024] 3.3) The flight control system sends control signals to the servo motor of the wing active twist device. These signals determine the angle of active wing twist. The total disturbance of the system is observed in real time at each moment through discrete extended observer dynamic equations using three known quantities.

[0025] Step 4) describes the specific method for combining the linear control law and real-time disturbance to form the control quantity of the aircraft, as follows:

[0026] The linear control law is a linear superposition of the deviation between the expected and actual pitch angles, and the pitch rate, specifically:

[0027]

[0028] Where u0 is the linear control quantity, K p For the controller parameter proportional gain, K d The differential gain of the controller parameters, θ d Let θ be the desired pitch angle, and θ be the pitch angle measured by the sensor. The sensor measures the pitch angular velocity.

[0029] Step 4) involves allocating the actual control input between the two actuators, assigning the elevator deflection angle and the wing active twist angle the same value. Specifically:

[0030]

[0031] Step 4) describes the actual control quantity between the two actuators, using the elevator for aircraft pitch angle control, and the wing active twisting device for compensating for disturbances, specifically:

[0032]

[0033] The beneficial effects of this invention are as follows:

[0034] 1. An active wing twisting device is a mechanism that can freely adjust the wing shape in a flapping wing system as needed. Using a reasonable control strategy, the periodic fluctuations in aerodynamic forces generated by the flapping wing aircraft can be minimized, and attitude oscillations caused by unsteady aerodynamic forces can be effectively mitigated.

[0035] 2. The active wing twisting device is a supplement to the control actuators of traditional bird-like flapping-wing aircraft. When this device works together with the traditional control surfaces on the tail for aircraft attitude control, it can reduce the burden on the control surfaces on the tail, reduce the deflection of the elevator, effectively reduce the possibility of the control surfaces losing control capability due to reaching the stall angle of attack, and improve the efficiency of the control surfaces.

[0036] 3. The control method based on the extended observer introduced can further mitigate the attitude oscillation caused by unsteady aerodynamic forces by actively compensating for disturbances, while ensuring the normal flight of the flapping-wing aircraft. Attached Figure Description

[0037] To more clearly illustrate the technical solutions in the embodiments of the present invention, the accompanying drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0038] Figure 1 This is a schematic diagram of the active wing twisting device of the present invention.

[0039] Figure 2 This is a control block diagram of the active torsion device control method based on an extended observer according to the present invention.

[0040] Figure 3 This is a control block diagram of the active torsion device control method based on an extended observer according to the present invention.

[0041] In the diagram: 1. Flexible skin; 2. Inner wing rib; 3. Outer wing rib; 4. Middle wing rib; 5. Wing spars; 6. Servo; 7. Fixing screw; 8. Servo rocker arm; 9. Drive rod; 10. Lug support. Detailed Implementation

[0042] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0043] The invention provides an active twisting device for the wings of a flapping-wing aircraft, such as Figure 1 As shown, it includes a flexible skin 1, an inner wing rib 2, an outer wing rib 3, a middle wing rib 4, a wing spars 5, a servo motor 6, a fixing screw 7, a servo motor rocker arm 8, a transmission rod 9, and an ear plate support 10.

[0044] like Figure 1 As shown, the wing spars 5 pass through circular through-holes on the inner wing rib 2, outer wing rib 3, and middle wing rib 4, and are fixed to them by adhesive. The three wing ribs and the wing spars together form a relatively fixed whole, which will be used to support the flexible skin. One end of the wing spars is connected to a flapping mechanism (not shown) located inside the fuselage of the flapping-wing aircraft. During flight, the flapping mechanism drives the wing spars, which in turn drive the wing ribs, causing the entire wing to flap around the fuselage.

[0045] The wing skin, made of flexible material, wraps around the leading edge of the wing rib, extending its lower side to approximately one-quarter of the rib chord length, without interfering with the movement of the servo rocker arm 8. The portion of the skin covering the upper side of the wing rib extends to the rear end of the rib, exceeding its support range. The skin is glued to the wing rib, and the rib-supported portion of the skin constitutes the relatively undeformable part of the flapping wing, maintaining its shape consistency during flapping. At the wing root, the skin connects to the fuselage (not shown).

[0046] In the wing arrangement described above, the wing is divided into two parts. The leading section of the wing in the chord direction is supported by ribs, and its skin shape is relatively fixed, making it less prone to deformation. However, the trailing section of the chord direction lacks rib support and has the ability to deflect vertically around the leading section of the wing.

[0047] A through-hole is provided in the center of the central wing rib to accommodate the servo motor and a fixing screw. The servo motor 6 is fixed to the central wing rib 4 by the fixing screw 7. The servo motor 6 remains relatively stationary with respect to the central wing rib. The servo motor rocker arm 8 is nested into the servo motor drive shaft through the through-hole and is fixed by adhesive. The metal transmission rod 9 has bends at both ends and passes through the holes at the ends of the servo motor rocker arm 8 and the lug support 10, respectively. The transmission rod 9, the servo motor rocker arm 8, and the lug support 10 constitute a planar four-bar linkage. The lug support 10 is bonded to the flexible skin 1.

[0048] The rotational motion of the drive shaft of servo motor 6 is converted into the up-and-down deflection motion of the flexible skin at the trailing edge of the wing via the aforementioned four-bar linkage. Servo motor 6 receives control signals from the flight control system on the flapping-wing aircraft, ultimately controlling the shape of the wing's trailing edge skin. During normal cruise, the deflection of the wing's trailing edge alters its angle of attack relative to the incoming airflow, resulting in different aerodynamic characteristics.

[0049] When the wing trailing edge deflects downwards, the angle of attack of the trailing edge increases relative to the incoming flow. Therefore, the lift generated by the wing increases before this deflection angle reaches the stall angle of attack. Since the aerodynamic center of the flapping-wing aircraft's wing is located in front of the aircraft's center of mass, the increased lift will generate a pitching moment on the entire aircraft. If the wing trailing edge deflects upwards, a pitching moment will be generated. In summary, this active wing twisting device can adjust the longitudinal attitude of the flapping-wing aircraft by generating a longitudinal moment.

[0050] This invention also provides a flight control method for a bird-like flapping-wing aircraft with an active wing twisting device, such as... Figure 2 and Figure 3 As shown, it includes the following steps:

[0051] Step 1: Construct the longitudinal dynamic model of the flapping-wing aircraft. The longitudinal attitude motion of the flapping-wing aircraft can be described by a simplified second-order differential equation about the pitch angle, specifically:

[0052]

[0053] Among them, among them, Let b1 be the angular acceleration of the aircraft's pitch angle, u1 be the proportional control coefficient of the aircraft's tail fin on the pitch angle, b2 be the proportional control coefficient of the wing's active twist mechanism on the pitch angle, u2 be the actual wing's active twist mechanism deflection angle, and f be the total disturbance. For the purposes of this invention, it mainly refers to the pitch angle disturbance caused by the deviation of the aerodynamic force generated by the flapping wing system from the average value.

[0054] Based on the longitudinal dynamics model of the aircraft, a simplified longitudinal system dynamic equation can be obtained, specifically:

[0055]

[0056] Where θ is the aircraft's pitch angle. The pitch rate of the aircraft. This is a fictitious quantity that can be observed using an observer.

[0057] Step 2: Based on the obtained longitudinal dynamic equations of the aircraft, an extended observer can be constructed, specifically as follows:

[0058]

[0059] Where z1 is the observer's estimate of the pitch angle θ, and z2 is the observer's estimate of the pitch angular velocity. The estimated value is z3, which is the observer's estimate of the total disturbance f. β1, β2, and β3 are adjustable parameters. When the parameters are well tuned, z1→θ can be achieved. z1, z2 and z3 respectively track and observe the pitch angle, pitch rate and total disturbance in the pitch angle of the aircraft.

[0060] Currently, the observer uses continuous dynamic equations, while the flight control system requires discrete dynamic equations for calculation. The following are the discrete equations obtained using the Euler method for the flight control system calculation:

[0061]

[0062] Where the subscript k represents the current time, k+1 represents the next time, and t is the sampling period of the flight control system sensor.

[0063] Step 3: Based on the discrete extended observer dynamic equations obtained above, the total disturbance in the aircraft's pitch angle can be obtained. For the extended observer, the total disturbance needs to be calculated using values ​​from three aircraft systems. First, the aircraft's pitch angle, which can be measured by inertial sensors in the flight control system. Second, the control signal sent by the flight control system to the servo motor controlling the elevator on the tail, which determines the elevator deflection angle; this value can be automatically retrieved by the program within the controller. Finally, the control signal sent by the flight control system to the servo motor of the wing's active twist mechanism, which determines the wing's active twist angle; this value can also be retrieved by the controller. Using these three known quantities, the total disturbance of the system can be observed in real time at each moment through the discrete extended observer dynamic equations.

[0064] Step 4: Based on the total system disturbance observed by the extended observer, and combined with the linear control law, the specific actual control quantity of the actuator can be obtained. The linear control law is a linear superposition of the deviation between the expected and actual pitch angle values ​​and the pitch velocity, specifically:

[0065]

[0066] Where u0 is the linear control quantity, K p For the controller parameter proportional gain, K d The differential gain of the controller parameters, θ d Let θ be the desired pitch angle, and θ be the pitch angle measured by the sensor. The sensor measures the pitch angular velocity.

[0067] Because this invention introduces an active wing twisting device, compared with traditional bird-like flapping-wing aircraft, it has an additional actuator in addition to the elevator on the traditional tail, and it is necessary to reasonably allocate the actual control quantity between the two actuators.

[0068] This invention provides two allocation methods.

[0069] Firstly: Assign the same value to the elevator deflection angle and the wing active twist angle, specifically:

[0070]

[0071] This allocation method ensures that both actuators have the same twist angle, simultaneously compensating for disturbances in the longitudinal passageway through the elevator and the wing's active twist mechanism, while also controlling the aircraft's pitch angle. The advantage is a significant reduction in the elevator's workload, allowing for a smaller horizontal stabilizer capacity and overall aircraft weight reduction. However, because it results in a relatively large active twist angle for the wing, the flapping-wing aircraft may experience excessive angles of attack at certain angles, potentially leading to a stall.

[0072] Secondly: The aircraft's pitch angle is controlled using elevators, while the active wing twist mechanism compensates for disturbances. Specifically:

[0073]

[0074] This allocation method separates the aircraft's pitch control from disturbance compensation. The longitudinal attitude oscillations caused by the unsteady, periodic aerodynamic fluctuations of the flapping wing system are compensated for by the active wing twist mechanism, effectively reducing the overall aerodynamic fluctuations within the flapping wing system itself. Simultaneously, since the active wing twist mechanism does not handle the aircraft's attitude control, the deflection angle is relatively small, making it less likely for the wing to enter a stall angle of attack. However, this allocation method does not reduce the elevator area, thus decreasing the overall mass of the aircraft.

[0075] The various embodiments in this specification are described in a progressive manner. Similar or identical parts between embodiments can be referred to interchangeably. Each embodiment focuses on its differences from other embodiments. In particular, for the device embodiments, the above descriptions are merely preferred embodiments of the present invention. Since they are fundamentally similar to the method embodiments, the descriptions are relatively simple, and relevant parts can be referred to the descriptions of the method embodiments. The above descriptions are merely specific embodiments of the present invention, but the scope of protection of the present invention is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention, without departing from the principle of the present invention, should be included within the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be determined by the scope of the claims.

Claims

1. A control method for an active wing twisting device of a flapping-wing aircraft, wherein the active wing twisting device includes a support frame with a flexible skin, the support frame including an inner wing rib, an outer wing rib, and a middle wing rib connected in series via a wing sparb, the inner end of the wing sparb being connected to a flapping mechanism inside the fuselage of the flapping-wing aircraft; the portion of the skin covering the upper side of the wing rib extends to the rear end of the wing rib and exceeds the support range of the wing rib, the excess portion deflecting vertically around the leading section of the wing; a servo motor is fixed to the middle wing rib and remains relatively stationary with respect to the middle wing rib, the drive shaft of the servo motor is connected to a metal transmission rod via a servo motor rocker arm; the metal transmission rod has bends at both ends and passes through holes at the ends of the servo motor rocker arm and the lug support, respectively, the lug support being bonded to the flexible skin, the transmission rod, the servo motor rocker arm, and the lug support forming a planar four-bar linkage, characterized in that... Includes the following steps: Step 1) Construct a longitudinal dynamic model of the flapping-wing aircraft. Based on the longitudinal dynamic model, obtain the simplified longitudinal system dynamic equations, specifically: Where θ is the aircraft's pitch angle. The pitch rate of the aircraft. This is a fictitious quantity that can be observed using an observer. Step 2) Construct an extended observer based on this dynamic model; Step 3) The attitude data measured by the inertial measurement unit is used to obtain the real-time disturbance of the aircraft using the extended observer. Based on the obtained discrete extended observer dynamic equation, the total disturbance in the pitch angle of the aircraft is obtained. Step 4) Combine the linear control law and real-time disturbance to form the control quantity of the aircraft, complete the attitude control of the flapping wing aircraft, and allocate the actual control quantity between the two actuators of the wing active twist device and the elevator on the tail.

2. The control method for the active wing twisting device of the flapping-wing aircraft according to claim 1, characterized in that: In step 1), during the construction of the longitudinal dynamic model of the flapping-wing aircraft, the longitudinal attitude motion of the flapping-wing aircraft is simplified into a second-order differential equation with respect to the pitch angle: in, Let b1 be the angular acceleration of the aircraft pitch angle, b2 be the proportional control coefficient of the aircraft tail for the pitch angle, u1 be the actual tail deflection angle, b2 be the proportional control coefficient of the wing active twist device for the pitch angle, u2 be the actual wing active twist device deflection angle, and f be the total disturbance.

3. The control method for the active wing twisting device of the flapping-wing aircraft according to claim 1, characterized in that: Step 2) describes the process of constructing the extended observer as follows: Where z1 is the observer's estimate of the pitch angle θ, and z2 is the observer's estimate of the pitch angular velocity. The estimated value is z3, which is the observer's estimate of the total disturbance f. β1, β2, and β3 are adjustable parameters. When the parameters are well tuned, z1→θ can be achieved. z1, z2 and z3 respectively track and observe the pitch angle, pitch rate and total disturbance in the pitch angle of the aircraft.

4. The control method for the active wing twisting device of the flapping-wing aircraft according to claim 1, characterized in that: Step 3) The method for obtaining the real-time disturbance of the aircraft is as follows: 3.1) is the pitch angle of the aircraft, which is measured by inertial sensors in the flight control system; 3.2) This is the control signal sent by the flight control system to the servo motor on the tail that controls the elevator, which determines the deflection angle of the elevator; 3.3) The flight control system sends control signals to the servo motor of the wing active twist device. These signals determine the angle of active wing twist. The total disturbance of the system is observed in real time at each moment through discrete extended observer dynamic equations using three known quantities.

5. The control method for the active wing twisting device of the flapping-wing aircraft according to claim 1, characterized in that: Step 4) describes the specific method for combining the linear control law and real-time disturbance to form the control quantity of the aircraft, as follows: The linear control law is a linear superposition of the deviation between the expected and actual pitch angles, and the pitch rate, specifically: Where u0 is the linear control quantity, K p For the controller parameter proportional gain, K d The differential gain of the controller parameters, θ d Let θ be the desired pitch angle, and θ be the pitch angle measured by the sensor. The sensor measures the pitch angular velocity.

6. The control method for the active wing twisting device of the flapping-wing aircraft according to claim 5, characterized in that: Step 4) involves allocating the actual control inputs between the wing's active twist mechanism and the elevator on the tail, assigning the elevator deflection angle the same value as the wing's active twist angle. Specifically:

7. The control method for the active wing twisting device of the flapping-wing aircraft according to claim 5, characterized in that: Step 4) involves allocating the actual control input between the active wing twist device and the elevator on the tail. The elevator is used for aircraft pitch control, and the active wing twist device is responsible for compensating for disturbances. Specifically:

Images