Unmanned aerial vehicle actuator dynamic characteristic test and modeling method based on installation state

Abstract

The present invention discloses a method for testing and modeling the dynamic characteristics of an unmanned aerial vehicle actuator based on an actual installation state, belonging to the field of unmanned aerial vehicle technology. The method includes a test system constructed based on the in-situ installation state of the actuator and rudder of the unmanned aerial vehicle. The test system includes a MEMS inertial navigation system fixed on the surface of the rudder and at the position closest to the rudder rotation axis. The MEMS inertial navigation system is connected to a test computer and a test power supply via cables. A suction cup is installed on the surface of the rudder at a distance L from the MEMS inertial navigation system. The suction cup is connected to a spring via a rope, and the other end of the spring is connected to a fixed fixing frame. When the rudder deflects 0°, the rope at one end of the spring is just in a straight state. The present invention adopts the above-mentioned method for testing and modeling the dynamic characteristics of an unmanned aerial vehicle actuator based on an actual installation state, which can improve the accuracy of actuator modeling, the reliability of flight control law design, and the confidence level of flight simulation results.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of unmanned aerial vehicles, and particularly to a method for testing and modeling dynamic characteristics of an unmanned aerial vehicle actuator based on an installation state. BACKGROUND

[0002] An unmanned aerial vehicle actuator is one of the key on-board devices of an unmanned aerial vehicle. The actuator receives instructions from a flight control computer, outputs mechanical rotation or linear motion, drives corresponding unmanned aerial vehicle rudder deflection, and causes the unmanned aerial vehicle rudder deflection to result in changes in the aerodynamic force or torque of the unmanned aerial vehicle, thereby changing the motion of the unmanned aerial vehicle and achieving the purpose of controlling the motion of the unmanned aerial vehicle.

[0003] In the prior art, the actuator is composed of a controller, a rudder, and a cable therebetween. The controller receives motion instructions from the flight control computer, comprehensively processes information by synthesizing rudder motion and electrical feedback information, generates rudder driving signals, and outputs the driving signals to the rudder. The motor in the rudder receives the driving signals from the controller to rotate, drives the internal speed reducer of the rudder to move the output shaft of the rudder to the required position, and feeds back the output shaft position, output shaft speed, armature current, and other information to the controller.

[0004] From the flight control computer giving the actuator motion instructions to the rudder deflection, the transmission mechanism (referred to as the transmission mechanism) between the controller, the rudder, the output shaft of the rudder, and the rudder is experienced. The unmanned aerial vehicle rudder is the driven object of the actuator, and the rudder is subjected to aerodynamic force. The aerodynamic force generates a torque on the rudder shaft, which is referred to as the aerodynamic hinge torque. The aerodynamic hinge torque is transmitted to the rudder through the rudder and the transmission mechanism, and becomes the torque / force load (referred to as the force load) of the output shaft of the rudder. The rudder, the transmission mechanism, and the internal moving mechanical components of the rudder all have mass or rotational inertia, which together constitute the inertia load of the output shaft of the rudder. The force load and the inertia load acting on the output shaft of the rudder have a significant impact on the motion of the actuator. In addition, the motor, the speed reducer, the output shaft, the transmission mechanism, and the rudder of the rudder each have factors such as friction, clearance, and nonlinearity of mechanical transmission (the transmission ratio of the rocker arm in the transmission mechanism changes with the rocker arm deflection angle), which also have a non-negligible impact on the motion of the actuator.

[0005] The input of the actuator is the rudder deflection instruction from the flight control computer, and the output is the rudder deflection angle. The characteristic of the output responding to the input is referred to as the dynamic characteristic of the actuator. The key indicators of the dynamic characteristic of the actuator are the bandwidth and the maximum output force / torque. The bandwidth reflects the speed of the motion of the actuator, and the force / torque on the output shaft of the rudder of the actuator directly affects the speed of the motion of the actuator. Therefore, the maximum output force / torque of the actuator also indirectly reflects the speed of the motion of the actuator.

[0006] The bandwidth of actuator has great influence on flight control performance. If the bandwidth is small, the response to the flight control computer instruction is slow, which leads to large signal transmission phase delay, and directly leads to flight control oscillation, even divergence. The dynamic characteristics of actuator must be considered in flight control law design and flight control simulation, otherwise the designed control law may not be suitable for actual unmanned aerial vehicle, and flight control oscillation or even divergence may occur.

[0007] The dynamic characteristics of actuator can be represented by its mathematical model, which is the transfer function of output to input. The form of transfer function of general actuator is as follows:

[0008]

[0009] Wherein, s is Laplace operator, N(s) and D(s) are polynomials about s, and the order of N(s) is less than that of D(s).

[0010] In view of the importance of the dynamic characteristics of actuator, the dynamic characteristics of actuator must be considered in the design of unmanned aerial vehicle flight control law and flight simulation. Therefore, the mathematical model of actuator needs to be established, that is, N(s) and D(s) in formula (1) are determined.

[0011] For the actuator with angular output, the system shown in FIG. 1 is generally used for testing, and the transfer function of the actuator is obtained by analyzing the test data. Figure 1 The dashed box in FIG. 1 is the actuator, which includes a controller and a rudder. The controller and the rudder are connected by a cable. The rudder is fixed on the test table, and the controller is powered. The end of the rudder rocker arm is connected to a mass block to simulate the inertia load of each link between the rudder rocker arm and the rudder surface. A spring is connected between the mass block and the fixed frame to simulate the aerodynamic hinge moment load of the rudder surface. The test computer is connected to the controller through a cable, and the test computer outputs the actuator motion instruction signal to the controller. At the same time, the controller feeds back the rudder deflection angle to the test computer. The rudder deflection angle is measured by a sensor in the rudder and transmitted to the controller through a cable. Figure 1 The aerodynamic hinge moment load is related to the dynamic pressure and rudder deflection angle of the unmanned aerial vehicle. The aerodynamic hinge moment load received by the rudder is different under different dynamic pressure and rudder deflection angle, which directly affects the motion characteristics of the actuator, that is, the transfer function of the actuator is different under different dynamic pressure and rudder deflection angle. According to the desired dynamic pressure and rudder deflection angle state, the spring stiffness is determined by considering the aerodynamic hinge moment coefficient of the unmanned aerial vehicle. The test computer sends a sweep signal to the controller, receives the rudder deflection angle feedback by the controller and records it, and then performs numerical analysis on the output signal (rudder deflection angle) and input instruction (sweep signal) to obtain the transfer function of the actuator.

[0012]

[0013] ​For actuators whose output is linear motion, use Figure 1 A similar test system, except that the servo output shaft moves in a linear manner, and everything else is the same.

[0014] The above actuator modeling method has the following problems:

[0015] (1) The inertia load converted to the servo output shaft from the servo output shaft to the rudder surface (inclusive) and the aerodynamic load converted to the servo output shaft from the rudder surface aerodynamic hinge torque both change with the rudder surface deflection angle. It is necessary to adjust the mass block and spring stiffness according to the rudder surface deflection angle, which is relatively tedious and complicated, and the errors of the converted mass block and spring stiffness are large.

[0016] (2) The above method does not take into account factors such as friction, clearance, and nonlinearity of mechanical transmission (the transmission ratio of the rocker arm changes with the rocker arm deflection angle) in each link between the servo output shaft and the rudder surface (inclusive), which causes errors in the test results.

[0017] (3) The feedback signal from the servo does not reflect the actual deflection angle of the rudder.

[0018] Due to the above problems, the Figure 1 The method of testing and establishing the actuator model is cumbersome and has large errors. Summary of the Invention

[0019] The purpose of the present invention is to provide a dynamic characteristics testing and modeling method for UAV actuators based on the actual installation state, which can improve the accuracy of actuator modeling, ensure the safety of the control performance of the UAV full-envelope flight / actuator in the full motion range at the actuator model level, and improve the reliability of the flight control law design and the confidence of the flight simulation results.

[0020] To achieve the above-mentioned objectives, a test system is constructed based on the in-situ installation state of the actuator and rudder of a drone. The test system includes a MEMS inertial navigation system fixed on the surface of the rudder and at the position closest to the rudder rotation axis. The MEMS inertial navigation system is connected to a test computer and a test power supply via cables. A suction cup is installed on the surface of the rudder at a distance L from the MEMS inertial navigation system. The suction cup is connected to a spring via a rope, and the other end of the spring is connected to a fixed fixing frame. When the rudder deflects 0°, the rope at one end of the spring is just in a straight state.

[0021] Preferably, the actuator includes a controller and a servo, the controller is connected to the test computer and the UAV power supply respectively through the cable, and the controller and the servo are also connected through the cable; the rudder rotating shaft is connected to a rudder rocker arm, the servo output shaft of the servo is connected to a servo rocker arm, and the rudder rocker arm and the servo rocker arm are connected by a connecting rod.

[0022] Preferably, the MEMS inertial navigation system is fixed on the control surface by double-sided tape.

[0023] Preferably, if the control surface is installed horizontally, the spring is installed below the control surface, and if the control surface is installed vertically, the spring is installed on either side of the control surface.

[0024] Preferably, the method for determining the size L is specifically as follows:

[0025] Let the spring stiffness be k, and the aerodynamic hinge moment coefficient derivative of the control surface of the unmanned aerial vehicle be The dynamic pressure of the unmanned aerial vehicle is q, and the aerodynamic hinge moment generated by the deflection δ of the control surface is Where S is the aerodynamic reference area, l is the span of the unmanned aerial vehicle, the distance x moved up and down or left and right of the suction cup when the control surface deflects δ is x = Lδπ / 180, the tension of the spring is F = kx = kLδπ / 180, and the moment of the spring on the control surface shaft is M = FL = kL 2 δπ / 180, the moment of the spring on the control surface shaft is equal to the aerodynamic hinge moment, i.e. M = M j , and we have:

[0026]

[0027] Preferably, the construction of the test system specifically includes the following steps:

[0028] S1, fix the MEMS inertial navigation system on the control surface and connect all cables;

[0029] S2, select a spring with a stiffness k, select the maximum dynamic pressure state of the unmanned aerial vehicle within the flight envelope, calculate L using the maximum dynamic pressure, and determine the installation position of the suction cup;

[0030] S3, set the deflection command of the control surface to 0 by the test computer, install the suction cup, rope, and spring at a distance L from the control surface shaft, and adjust the length of the rope or the position of the fixing frame so that the tension of the spring makes the rope straight;

[0031] S4, send a 0° deflection command by the test computer, and subsequently increase the command step by Δ every 5 seconds until the control surface deflects to the specified maximum deflection;

[0032] S5, set the transfer function structure of the actuator;

[0033] S6, for each step response data collected, use the system identification app of matlab to determine a set of parameters (n0, n1, d0, d1, d2) in the set actuator structure, so that the step response of the transfer function approximates the actual step response;

[0034] S7, for each set of parameters (n0, n1, d0, d1, d2) corresponding transfer function, using the frequency characteristics of matlab drawing function to draw the frequency characteristic curve of the transfer function, thus reading the bandwidth of the transfer function;

[0035] S8, taking the minimum bandwidth of the set of parameters (n0, n1, d0, d1, d2) as the final transfer function of the actuator.

[0036] Preferably, in step S4, Δ is 2° to 3°.

[0037] Preferably, in step S5, the structure of the transfer function is set as:

[0038]

[0039] In the formula, s is the Laplace operator, N(s) and D(s) are polynomials about s, the order of N(s) is less than that of D(s), n1 and n0 are the coefficients of the first order term and the constant term of the polynomial N(s) respectively, and d2, d1 and d0 are the coefficients of the second order term, the first order term and the constant term of the polynomial D(s) respectively.

[0040] Therefore, the beneficial effects of the unmanned aerial vehicle actuator dynamic characteristic test and modeling method based on the installation state are:

[0041] (1) Directly using the installed rudder, actuator and the transmission mechanism therebetween, avoiding the complex and tedious conversion process of the inertia load on the output shaft of the rudder and the aerodynamic hinge load, ensuring the accuracy of the test.

[0042] (2) The test takes into account the friction, gap, nonlinearity of mechanical transmission and other factors existing between the rudder output shaft and the rudder, and is the actual state, ensuring the accuracy of the test.

[0043] (3) Directly using the real deflection angle of the rudder, ensuring the accuracy of the test.

[0044] (4) Using the maximum dynamic pressure state in the flight envelope, and the single-step deflection in the direction opposite to the torque direction generated by the aerodynamic hinge torque and the rudder gravity on the shaft, the model of the actuator under the most unfavorable state can be obtained, the bandwidth of the model is the smallest, and the reliability of the flight control law design and the confidence of the flight simulation result are improved.

[0045] The technical solutions of the present application will be further described in detail below through the drawings and examples. BRIEF DESCRIPTION OF DRAWINGS

[0046] Figure 1 is a structural schematic diagram of a conventional actuator modeling test system;

[0047] Figure 2 It is a structural schematic diagram of the test system in the dynamic characteristics testing and modeling method of the UAV actuator based on the actual installation state of the present invention.

[0048] Reference numerals

[0049] 1. Actuator; 2. Controller; 3. UAV power supply; 4. Servo; 5. Cable; 6. Test computer; 7. Test bench; 8. Servo rocker arm; 9. Mass block; 10. Spring; 11. Fixing bracket; 12. Rudder; 13. MEMS inertial navigation; 14. Test power supply; 15. Rudder shaft; 16. Suction cup; 17. Rope; 18. Rudder rocker arm; 19. Connecting rod. DETAILED DESCRIPTION

[0050] The technical solution of the present invention is further described below with reference to the accompanying drawings and embodiments.

[0051] Unless otherwise defined, the technical or scientific terms used in the present invention shall have the usual meanings understood by persons of ordinary skill in the field to which the present invention belongs. The words "first", "second" and similar terms used in the present invention do not indicate any order, quantity or importance, but are only used to distinguish different components. Words such as "include" or "comprise" mean that the elements or objects preceding the word include the elements or objects listed after the word and their equivalents, without excluding other elements or objects. Words such as "connect" or "connected" are not limited to physical or mechanical connections, but may include electrical connections, whether direct or indirect. "Up", "down", "left", "right" and the like are only used to indicate relative positional relationships. When the absolute position of the object being described changes, the relative positional relationship may also change accordingly.

[0052] Example 1

[0053] like Figure 2 As shown, the present invention provides a method for testing and modeling the dynamic characteristics of a UAV actuator based on the actual installation state, including a test system constructed based on the in-situ installation state of the actuator 1 and the rudder 12 of the UAV, that is, the test is carried out when the UAV is parked on the ground, and the rudder 12, the actuator 1, and the mechanical transmission mechanism from the servo output shaft to the rudder 12 are all in the actual installation state.

[0054] The actuator 1 includes a controller 2 and a steering gear 4. The controller 2 is connected to a test computer 6 and a drone power supply 3 via cables 5. The controller 2 and the steering gear 4 are also connected via cables 5. A rudder rocker arm 18 is connected to the rudder shaft 15. The rudder output shaft of the rudder 4 is connected to the rudder rocker arm 8. The rudder rocker arm 18 and the rudder rocker arm 8 are connected via a connecting rod 19.

[0055] The test system comprises a MEMS inertial navigation device 13 fixed on the surface of the rudder 12 at a position closest to the rudder rotation shaft 15, for measuring the rudder deflection angle. The position of the MEMS inertial navigation device 13 is as close as possible to the rudder rotation shaft 15, so that the influence of the moment of inertia of the rudder 12 relative to the rudder rotation shaft 15 can be ignored, thereby improving the test accuracy. Since the volume and weight of the MEMS inertial navigation device 13 are very small, in this embodiment, the MEMS inertial navigation device 13 is fixed on the surface of the rudder 12 by double-sided tape, and the MEMS inertial navigation device 13 is connected to the test computer 6 through the cable 5 to feed back the rudder deflection angle signal to the test computer 6. The MEMS inertial navigation device 13 is also connected to the test power supply 14 through the cable 5 for obtaining working power.

[0056] An adsorber 16 on the surface of the rudder 12 is installed at a distance L from the MEMS inertial navigation device 13, and the adsorber 16 is connected to the spring 10 through the rope 17, and the other end of the spring 10 is connected to the fixed frame 11, thereby simulating the aerodynamic hinge moment acting on the rudder 12 by the spring 10. When the rudder 12 is deflected by 0°, the rope 17 at one end of the spring 10 is just in a straight state, and the spring 10 has a very small tension. The placement mode of the rudder 12 is different, and the positions of other structures also need to be adjusted accordingly. If the rudder 12 is installed horizontally, the spring 10 is installed below the rudder 12, and if the rudder 12 is installed vertically, the spring 10 is installed on either side of the rudder 12.

[0057] The determination method of the above size L is as follows:

[0058] Let the spring stiffness be k, and the derivative of the aerodynamic hinge moment coefficient of the rudder of the unmanned aerial vehicle be The dynamic pressure of the unmanned aerial vehicle is q (unit: N / m 2 ), and the aerodynamic hinge moment generated by the deflection δ (unit: °) of the rudder is where S (unit: m 2 ) is the aerodynamic reference area, and l (unit: m) is the span length of the unmanned aerial vehicle. When the rudder is deflected by δ, the distance x moved up and down or left and right of the adsorber is x = Lδπ / 180, the tension of the spring is F = kx = kLδπ / 180, and the moment of the rudder rotation shaft is M = FL = kL 2 δπ / 180. The moment of the spring on the rudder rotation shaft is the same as the aerodynamic hinge moment, i.e. M = M j , and the following is obtained:

[0059]

[0060] In this embodiment, the unmanned aerial vehicle actuator dynamic characteristic test and modeling method based on the actual installation state specifically comprises the following steps:

[0061] S1, fix the MEMS inertial navigation device 13 on the rudder 12 and connect all the cables 5, and install and wire as shown in Figure 2 .

[0062] S2, select the appropriate spring 10, the spring 10 stiffness k, select the maximum dynamic pressure state of the unmanned aerial vehicle within the flight envelope, use the maximum dynamic pressure to calculate L from equation (2), determine the installation position of the suction cup 16.

[0063] S3, the test computer 6 gives the rudder deflection instruction as 0, the suction cup 16, the rope 17 and the spring 10 are installed at a distance of L from the rudder shaft 15, the length of the rope 17 or the position of the fixed frame 11 is adjusted, so that the tension of the spring 10 makes the rope 17 just straight.

[0064] S4, the test computer 6 sends a 0° deflection instruction, and then every 5 seconds, the instruction step increases Δ, generally 2°-3°, until the rudder 12 deflects to the specified maximum deflection.

[0065] S5, set the transfer function structure of the actuator 1, in this embodiment, the transfer function structure is set as:

[0066]

[0067] In the formula, s is the Laplace operator, N(s) and D(s) are polynomials with respect to s, the order of N(s) is less than that of D(s), n1 and n0 are the coefficients of the first order term and the constant term of the polynomial N(s) respectively, and d2, d1 and d0 are the coefficients of the second order term, the first order term and the constant term of the polynomial D(s) respectively.

[0068] S6, for each step response data collected, use the system identification app of matlab to determine a set of parameters (n0, n1, d0, d1, d2) in the set actuator structure, so that the step response of the transfer function formula (3) approximates the actual step response.

[0069] S7, for each set of parameters (n0, n1, d0, d1, d2) corresponding to the transfer function, use the frequency characteristic drawing function of matlab to draw the frequency characteristic curve of the transfer function, and thus read the bandwidth of the transfer function.

[0070] S8, take the set of parameters (n0, n1, d0, d1, d2) with the smallest bandwidth to substitute into the transfer function formula (3) as the final transfer function of the actuator 1.

[0071] Therefore, the unmanned aerial vehicle actuator dynamic characteristic test and modeling method based on the actual installation state improves the actuator modeling accuracy, considers the minimum bandwidth state of the actuator, and thus the control law design based on the actuator model can ensure the safety of the control performance within the full envelope flight of the unmanned aerial vehicle and the full motion range of the actuator, and improves the reliability of the flight control law design and the confidence of the flight simulation results.

[0072] It should be pointed out finally that the above embodiments are only used to illustrate the technical solutions of the present application but not to limit it, and although the present application has been described in detail with reference to the preferred embodiments, it should be understood by those skilled in the art that the technical solutions of the present application can be modified or equivalently replaced, and these modifications or equivalent replacements cannot make the modified technical solutions deviate from the spirit and scope of the technical solutions of the present application.

Claims

1. A method for testing and modeling the dynamic characteristics of UAV actuators based on actual installation, characterized by: The invention relates to a test system constructed in an in-situ installation state of an actuator and a rudder based on a drone. The test system includes a MEMS inertial navigation system fixed on the surface of the rudder and at a position closest to the rudder rotation axis. The MEMS inertial navigation system is connected to a test computer and a test power supply via cables. A suction cup is installed on the rudder surface at a distance L from the MEMS inertial navigation system. The suction cup is connected to a spring via a rope. The other end of the spring is connected to a fixed bracket. When the rudder is deflected by 0°, the rope at one end of the spring is just in a straight state. The specific method for determining size L is: Assuming the spring stiffness is k, the derivative of the aerodynamic hinge torque coefficient of the UAV control surface is The dynamic pressure of the UAV is q, and the aerodynamic hinge torque generated by the rudder deflection δ is Where S is the aerodynamic reference area, l is the length of the drone, and when the rudder deflects δ, the distance the suction cup moves up and down or left and right is x = Lδπ / 180, the spring tension F = kx = kLδπ / 180, and the torque on the rudder axis is M = FL = kL 2 δπ / 180, the torque of the spring on the rudder shaft is the same as the aerodynamic hinge torque, that is, M=M j ,have to:

2. The method for testing and modeling the dynamic characteristics of a UAV actuator based on an installed state according to claim 1 is characterized by: The actuator includes a controller and a servo, the controller is connected to the test computer and the UAV power supply respectively through the cable, and the controller and the servo are also connected through the cable; the rudder rotating shaft is connected to a rudder rocker arm, and the servo output shaft of the servo is connected to a servo rocker arm, and the rudder rocker arm and the servo rocker arm are connected by a connecting rod.

3. The method for testing and modeling the dynamic characteristics of a UAV actuator based on an installed state according to claim 1 is characterized in that: The MEMS inertial navigation system is fixed to the rudder surface via double-sided tape.

4. The method for testing and modeling the dynamic characteristics of a UAV actuator based on an installed state according to claim 1 is characterized in that: If the rudder surface is installed horizontally, the spring is installed below the rudder surface; if the rudder surface is installed vertically, the spring is installed on either side of the rudder surface.

5. The method for testing and modeling the dynamic characteristics of a UAV actuator based on an installed state according to claim 1 is characterized in that: The construction of the test system specifically includes the following steps: S1. Fix the MEMS inertial navigation system on the rudder and connect all cables. S2. Select a spring with a spring stiffness of k, select the maximum dynamic pressure state of the drone within the flight envelope, calculate L using the maximum dynamic pressure, and determine the installation position of the suction cup; S3. The test computer sets the rudder deflection command to 0, installs the suction cup, rope, and spring at a distance L from the rudder axis, and adjusts the rope length or the position of the fixing bracket so that the tension generated by the spring just straightens the rope; S4. The test computer sends a 0° deflection command, and the command is then incremented by Δ every 5 seconds until the rudder deflects to the specified maximum deflection. S5. Setting the transfer function structure of the actuator; S6. For each collected step response data, use the system identification app of MATLAB to determine a set of parameters (n0, n1, d0, d1, d2) in the set actuator structure so that the step response of the transfer function approaches the actual step response; S7. For each set of transfer functions (n0, n1, d0, d1, d2), use the frequency characteristic plotting function of MATLAB to draw the frequency characteristic curve of the transfer function, thereby reading the bandwidth of the transfer function; S8. Take a set of parameters with the smallest bandwidth (n0, n1, d0, d1, d2) and substitute them into the transfer function as the final transfer function of the actuator.

6. The method for testing and modeling the dynamic characteristics of a UAV actuator based on an installed state according to claim 5 is characterized by: In step S4, Δ is set to 2° to 3°.

7. The method for testing and modeling the dynamic characteristics of a UAV actuator based on an installed state according to claim 5 is characterized by: In step S5, the structure of the transfer function is set as: Where s is the Laplace operator, N(s) and D(s) are polynomials with respect to s, the order of N(s) is less than that of D(s), n1 and n0 are the coefficient of the linear term and the constant term of the polynomial N(s), respectively, and d2, d1, and d0 are the coefficient of the quadratic term, the coefficient of the linear term, and the constant term of the polynomial D(s), respectively.

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