CN119568433A - Unmanned aerial vehicle actuator dynamic characteristic test and modeling method based on actual mounting state - Google Patents
Unmanned aerial vehicle actuator dynamic characteristic test and modeling method based on actual mounting state Download PDFInfo
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Abstract
本发明公开了基于实装状态的无人机作动器动态特性测试与建模方法，属于无人机技术领域，包括基于无人机的作动器和舵面的原位安装状态下构建的测试系统，所述测试系统包括固定在所述舵面表面且距离舵面转轴最近位置处的MEMS惯导，所述MEMS惯导通过电缆分别与测试计算机和测试电源连接；在距离所述MEMS惯导为L远处安装有位于所述舵面表面上的吸盘，所述吸盘通过绳索连接有弹簧，所述弹簧的另一端与固定的固定架连接；所述舵面偏转0°时，所述弹簧一端的所述绳索刚好处于绷直状态。本发明采用上述的基于实装状态的无人机作动器动态特性测试与建模方法，可以提高作动器建模精度，提高飞行控制律设计的可靠性和飞行仿真结果的置信度。
The present invention discloses a method for testing and modeling the dynamic characteristics of an actuator of a UAV based on an actual installation state, which belongs to the technical field of UAVs, and includes a test system constructed under an in-situ installation state of an actuator and a rudder of a UAV, wherein the test system includes a MEMS inertial navigation system fixed on the surface of the rudder and at the position closest to the rudder rotating shaft, and the MEMS inertial navigation system is respectively connected to a test computer and a test power supply through cables; a suction cup located on the surface of the rudder is installed at a distance of L from the MEMS inertial navigation system, and the suction cup is connected to a spring through 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 actuator of a UAV based on an actual installation state, which can improve the modeling accuracy of the actuator, the reliability of the flight control law design, and the confidence of the flight simulation results.
Description
Technical Field
The invention relates to the technical field of unmanned aerial vehicles, in particular to a method for testing and modeling dynamic characteristics of an unmanned aerial vehicle actuator based on a real-world state.
Background
The unmanned aerial vehicle actuator is one of unmanned aerial vehicle key airborne equipment, and the actuator receives the instruction of flight control computer, and output machinery rotates or linear motion drives corresponding unmanned aerial vehicle rudder face to deflect to the rudder face deflection leads to unmanned aerial vehicle aerodynamic force or the change of moment to change unmanned aerial vehicle's motion, reach the purpose of control unmanned aerial vehicle motion.
In the prior art, the actuator is composed of a controller, a steering engine and a cable between the steering engine and the controller, the controller receives a motion instruction of a flight control computer, comprehensive steering engine motion and electric feedback information are subjected to information calculation processing, a steering engine driving signal is generated and output to the steering engine, a motor in the steering engine receives the driving signal of the controller to rotate, an internal speed reducer of the steering engine is driven to enable an output shaft of the steering engine to move to a required position, and information such as the position of the output shaft, the speed of the output shaft, armature current and the like is fed back to the controller.
The flight control computer gives out an actuator motion instruction to the control surface to deflect, and the middle of the flight control computer goes through the controller, the steering engine and links of a transmission mechanism (transmission mechanism for short) between the steering engine output shaft and the control surface. The control surface of the unmanned plane is a driving object of an actuator, the control surface is subjected to aerodynamic force, and the aerodynamic force generates moment on a control surface rotating shaft and is called pneumatic hinge moment. The moment of the pneumatic hinge is transmitted to the steering engine through the control surface and the transmission mechanism to become moment/force load (force load for short) of an output shaft of the steering engine. The control surface, the transmission mechanism and the mechanical parts moving in the steering engine are all provided with mass or rotational inertia, and the inertial load of the steering engine output shaft is formed together. The force load and the inertia load acting on the output shaft of the steering engine have obvious influence on the motion of the actuator. In addition, the motor, the speed reducer, the output shaft, the transmission mechanism and the control surface of the steering engine have factors such as friction, clearance, nonlinearity of mechanical transmission (the transmission ratio of a rocker arm in the transmission mechanism changes along with the deflection angle of the rocker arm) and the like, and have non-negligible influence on the motion of the actuator.
The input of the actuator is a control surface deflection command from a flight control computer, the output is a control surface deflection angle, the characteristic of the output response input is called as the dynamic characteristic of the actuator, the key index of the dynamic characteristic of the actuator is bandwidth, and the key index of the static characteristic is maximum output/moment. The bandwidth reflects the speed of the motion of the actuator, and the force/moment received on the steering engine output shaft of the actuator directly affects the speed of the motion of the actuator, so that the maximum output force/moment of the actuator also indirectly reflects the speed of the motion of the actuator.
The bandwidth of the actuator has great influence on the flight control performance, the bandwidth is small, the response of the actuator to the instruction of the flight control computer is slow, the delay of a larger signal transmission phase is caused, and the flight control oscillation and even divergence can be directly caused. The dynamic characteristics of the actuators must be considered in the design of the flight control law and the simulation of the flight control, otherwise, the designed control law may not be suitable for an actual unmanned aerial vehicle, and the phenomenon of oscillation and even divergence of the flight control occurs.
The dynamic characteristics of an actuator can be represented by its mathematical model in the form of an output-to-input transfer function, typically the transfer function of an actuator is in the form of:
Where s is the Laplace operator, N(s), D(s) are polynomials about s, and N(s) order is smaller than D(s).
In view of the importance of the dynamic characteristics of the actuators, the dynamic characteristics of the actuators must be considered when designing the unmanned aerial vehicle flight control law and simulating the flight, and therefore, mathematical models of the actuators must be established, namely, N(s) and D(s) in the transfer function of the equation (1) are determined.
The actuator with the actuator output as the rotation angle is generally tested by adopting the system shown in fig. 1, and the transfer function of the actuator is obtained by analyzing test data. In fig. 1, an actuator is shown in a dashed frame, and the actuator comprises a controller and a steering engine, wherein the controller is connected with the steering engine through a cable. The steering engine is fixed on the test table surface, the controller is powered, the tail end of the steering engine rocker arm is connected with a mass block with a certain mass, the inertia load of each link from the steering engine rocker arm to the control surface (including) is simulated, a spring is connected between the mass block and the fixing frame, and the pneumatic hinge moment load of the control surface is simulated. The test computer is connected with the controller through a cable, the test computer outputs an actuator movement instruction signal to the controller, meanwhile, the controller feeds back a steering engine deflection angle to the test computer, and the steering engine deflection angle is measured by a sensor in the steering engine and is transmitted to the controller through the cable.
The pneumatic hinge moment load is related to the dynamic pressure of the unmanned aerial vehicle and the deflection angle of the control surface, the pneumatic hinge moment load borne by the steering engine under different dynamic pressures and deflection angles of the control surface is different, and the motion characteristics of the actuator are directly influenced, namely the transfer functions of the actuator are different under different dynamic pressures and deflection angles of the control surface. According to the dynamic pressure and rudder deflection angle state to be measured, the pneumatic hinge moment coefficient of the unmanned aerial vehicle is considered to determine the spring stiffness, a test computer sends a sweep frequency signal to a controller, meanwhile, the steering engine deflection angle fed back by the controller is received and recorded, and after that, the output signal (steering engine deflection angle) and the input command (sweep frequency signal) are subjected to numerical analysis to obtain the transfer function of the actuator.
For the actuator with linear motion output by the actuator, a test system similar to that in fig. 1 is adopted, except that the output shaft of the steering engine is in linear motion, and the other is the same.
The modeling method of the actuator has the following problems:
(1) The inertia load between the steering engine output shaft and the steering surface (including), the inertia load between the steering engine output shaft and the steering surface, and the pneumatic load between the steering surface and the steering engine output shaft, wherein the pneumatic hinge moment of the steering surface, the pneumatic load between the steering engine output shaft and the steering engine output shaft, are all changed along with the deflection angle of the steering surface, and the mass block and the spring stiffness are required to be adjusted according to the deflection angle of the steering surface, so that the steering engine is complicated and complex, and the error between the folded mass block and the spring stiffness is large.
(2) The method does not consider factors such as friction, clearance, nonlinearity of mechanical transmission (the transmission ratio of a rocker arm changes along with the deflection angle of the rocker arm) and the like existing in each link between an output shaft of a steering engine and a control surface (including the steering engine), and causes errors of test results.
(3) The signal fed back by the steering engine does not reflect the real deflection angle of the control surface.
Because of the above problems, the modeling of the actuator using the method of fig. 1 is cumbersome and has a large error.
Disclosure of Invention
The invention aims to provide a method for testing and modeling the dynamic characteristics of an unmanned aerial vehicle actuator based on a mounting state, which can improve the modeling precision of the actuator, ensure the safety of the control performance of the unmanned aerial vehicle in the full envelope flight/actuator full motion range on the actuator model level, and improve the reliability of the flight control law design and the confidence of the flight simulation result.
The test system comprises an MEMS inertial navigation device which is fixed on the surface of the control surface and is positioned at the position closest to a rotating shaft of the control surface, wherein the MEMS inertial navigation device is connected with a test computer and a test power supply respectively through cables, a sucker which is positioned on the surface of the control surface and is connected with a spring through a rope is arranged far away from the MEMS inertial navigation device L, the other end of the spring is connected with a fixed fixing frame, and when the control surface deflects for 0 DEG, the rope at one end of the spring is just in a straight state.
Preferably, the actuator comprises a controller and a steering engine, the controller is respectively connected with the test computer and the unmanned aerial vehicle power supply through the cables, the controller is also connected with the steering engine through the cables, a control surface rocker arm is connected on a control surface rotating shaft, a steering engine rocker arm is connected on a steering engine output shaft of the steering engine, and the control surface rocker arm is connected with the steering engine rocker arm through a connecting rod.
Preferably, the MEMS inertial navigation is fixed on the control surface through double faced adhesive tape.
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.
Preferably, the method for determining the dimension L specifically includes:
the spring stiffness is k, and the derivative of the moment coefficient of the pneumatic hinge of the control surface of the unmanned aerial vehicle is The dynamic pressure of the unmanned plane is q, and the pneumatic hinge moment generated by the steering surface deflection delta isWherein S is a pneumatic reference area, L is the extension length of the unmanned aerial vehicle, when the control surface deflects delta, the distance x=lδpi/180 of the vertical or horizontal movement of the suction disc, the tension force f=kx=klδpi/180 of the spring, the moment to the control surface rotating shaft is m=fl=kl 2 δpi/180, and the moment to the control surface rotating shaft by the spring is the same as the moment of the pneumatic hinge, namely m=m j, so that the method is obtained:
Preferably, the construction of the test system specifically comprises the following steps:
S1, fixing MEMS inertial navigation on a control surface and connecting all cables;
S2, selecting a spring, wherein the spring stiffness is k, selecting a maximum dynamic pressure state of the unmanned aerial vehicle in a flight envelope, calculating L by adopting the maximum dynamic pressure, and determining the mounting position of the sucker;
s3, giving a control surface deflection instruction of 0 by a test computer, installing a sucker, a rope and a spring at a position away from a control surface rotating shaft L, and adjusting the length of the rope or the position of a fixing frame to enable the rope to be just straightened by the tensile force generated by the spring;
s4, a test computer sends a 0-degree deflection instruction, and the step increase delta is instructed every 5 seconds later until the control surface deflects to the specified maximum deflection;
s5, setting a transfer function structure of the actuator;
S6, determining a set of parameters (n 0、n1、d0、d1、d2) in the set actuator structure by utilizing a system identification app of matlab for each acquired step response data, so that the step response of the transfer function approaches to the actual step response;
S7, drawing a frequency characteristic curve of the transfer function by using a matlab frequency characteristic drawing function for the transfer function corresponding to each group of parameters (n 0、n1、d0、d1、d2), so as to read out the bandwidth of the transfer function;
s8, substituting a group of parameters (n 0、n1、d0、d1、d2) with the minimum bandwidth into the transfer function to serve as a final transfer function of the actuator.
Preferably, in step S4, delta is 2 DEG to 3 deg.
Preferably, in step S5, the transfer function is configured as follows:
Where s is a laplace operator, N(s) and D(s) are polynomials related to s, the order of N(s) is smaller than D(s), N 1、n0 is a first order term coefficient and a constant term of the polynomial N(s), and D 2、d1、d0 is a second order term coefficient, a first order term coefficient and a constant term of the polynomial D(s), respectively.
Therefore, the unmanned aerial vehicle actuator dynamic characteristic testing and modeling method based on the mounting state has the beneficial effects that:
(1) The real control surface, the actuator and a transmission mechanism between the real control surface and the actuator are directly utilized, so that the complex and complicated folding process of inertia load and pneumatic hinge load on an output shaft of the steering engine is avoided, and the testing precision is ensured.
(2) In the test, the factors such as friction, clearance, nonlinearity of mechanical transmission and the like existing in each link between the steering engine output shaft and the control surface are considered, and the test precision is ensured in an actual state.
(3) The real deflection angle of the control surface is directly adopted, so that the testing precision is ensured.
(4) The maximum dynamic pressure state in the flight envelope and unidirectional step deflection opposite to the moment direction generated by the pneumatic hinge moment and the control surface gravity on the rotating shaft of the flight envelope are adopted, so that a model of the actuator in the most unfavorable state can be obtained, the bandwidth of the model is minimum, and the reliability of the flight control law design and the confidence of the flight simulation result are improved.
The technical scheme of the invention is further described in detail through the drawings and the embodiments.
Drawings
FIG. 1 is a schematic diagram of a conventional actuator modeling test system;
Fig. 2 is a schematic structural diagram of a test system in the method for testing and modeling dynamic characteristics of an unmanned aerial vehicle actuator based on a mounting state.
Reference numerals
1. An actuator; 2, a controller, 3, an unmanned aerial vehicle power supply, 4, a steering engine, 5, a cable, 6, a test computer, 7, a test table, 8, a steering engine rocker arm, 9, a mass block, 10, a spring, 11, a fixing frame, 12, a control surface, 13, MEMS inertial navigation, 14, a test power supply, 15, a control surface rotating shaft, 16, a sucker, 17, a rope, 18, a control surface rocker arm, 19 and a connecting rod.
Detailed Description
The technical scheme of the invention is further described below through the attached drawings and the embodiments.
Unless defined otherwise, technical or scientific terms used herein should be given the ordinary meaning as understood by one of ordinary skill in the art to which this invention belongs. The terms "first," "second," and the like, as used herein, do not denote any order, quantity, or importance, but rather are used to distinguish one element from another. The word "comprising" or "comprises", and the like, means that elements or items preceding the word are included in the element or item listed after the word and equivalents thereof, but does not exclude other elements or items. The terms "connected" or "connected," and the like, are not limited to physical or mechanical connections, but may include electrical connections, whether direct or indirect. "upper", "lower", "left", "right", etc. are used merely to indicate relative positional relationships, which may also be changed when the absolute position of the object to be described is changed.
Example 1
As shown in fig. 2, the invention provides a method for testing and modeling dynamic characteristics of an unmanned aerial vehicle actuator based on a mounting state, which comprises a test system constructed under an in-situ mounting state of the unmanned aerial vehicle actuator 1 and a control surface 12, namely, the test is carried out under a ground shutdown state of the unmanned aerial vehicle, and mechanical transmission mechanisms from the control surface 12, the actuator 1 and a steering engine output shaft to the control surface 12 are all in a real mounting state.
The actuator 1 comprises a controller 2 and a steering engine 4, wherein the controller 2 is respectively connected with a test computer 6 and a power supply 3 through cables 5, and the controller 2 is also connected with the steering engine 4 through the cables 5. The control surface rotating shaft 15 is connected with a control surface rocker arm 18, a steering engine output shaft of the steering engine 4 is connected with a steering engine rocker arm 8, and the control surface rocker arm 18 is connected with the steering engine rocker arm 8 through a connecting rod 19.
The test system includes a MEMS inertial navigation 13 fixed to the surface of the control surface 12 and located closest to the control surface rotation axis 15 for measuring the control surface deflection angle. The MEMS inertial navigation 13 is located as close to the control surface rotation shaft 15 as possible, so that the influence of the MEMS inertial navigation 13 on the rotational inertia of the control surface 12 relative to the control surface rotation shaft 15 is negligible, and the test precision is improved. Since the volume and the weight of the MEMS inertial navigation 13 are small, in this embodiment, the MEMS inertial navigation 13 is fixed on the surface of the control surface 12 through a double sided adhesive tape, and the MEMS inertial navigation 13 is connected with the test computer 6 through the cable 5, and feeds back the control surface deflection angle signal to the test computer 6. The MEMS inertial navigation 13 is also connected to a test power supply 14 through a cable 5 for obtaining operating power.
The suction cup 16 positioned on the surface of the control surface 12 is arranged far away from the MEMS inertial navigation 13, the suction cup 16 is connected with the spring 10 through a rope 17, the other end of the spring 10 is connected with the fixed fixing frame 11, and the spring 10 simulates the pneumatic hinge moment applied to the control surface 12. When the control surface 12 deflects 0 degrees, the rope 17 at one end of the spring 10 is ensured to be just in a straightened state, and the spring 10 has small tension. The control surface 12 is placed in different ways, and the positions of other structures need to be adjusted correspondingly. If the control surface 12 is mounted horizontally, the springs 10 are mounted below the control surface 12, and if the control surface 12 is mounted vertically, the springs 10 are mounted on either side of the control surface 12.
The method for determining the dimension L specifically comprises the following steps:
the spring stiffness is k, and the derivative of the moment coefficient of the pneumatic hinge of the control surface of the unmanned aerial vehicle is The dynamic pressure of the unmanned plane is q (unit N/m 2), and the pneumatic hinge moment generated by the deflection delta (unit degree) of the control surface isWherein S (unit M 2) is a pneumatic reference area, L (unit M) is the extension length of the unmanned aerial vehicle, when the control surface deflects delta, the suction cup moves up and down or left and right by a distance x=lδpi/180, the tension force f=kx=klδpi/180 of the spring, the moment on the control surface rotating shaft is m=fl=kl 2 δpi/180, and the moment on the control surface rotating shaft by the spring is the same as the moment of the pneumatic hinge, namely m=m j, so that the method is as follows:
in this embodiment, the method for testing and modeling the dynamic characteristics of the unmanned aerial vehicle actuator based on the mounting state specifically includes the following steps:
S1, fixing the MEMS inertial navigation device 13 on the control surface 12 and connecting all cables 5, and installing and wiring in a mode shown in FIG. 2.
S2, selecting a proper spring 10, wherein the stiffness of the spring 10 is k, selecting the maximum dynamic pressure state of the unmanned aerial vehicle in the flight envelope, calculating L by the formula (2) by adopting the maximum dynamic pressure, and determining the installation position of the sucker 16.
S3, giving a control surface deflection instruction of 0 by the test computer 6, installing a sucker 16, a rope 17 and a spring 10 at a position which is L away from the control surface rotating shaft 15, and adjusting the length of the rope 17 or the position of the fixing frame 11 to enable the rope 17 to be just straightened by the tensile force generated by the spring 10.
S4, a 0-degree deflection command is sent by the test computer 6, and the subsequent command step increment delta is increased every 5 seconds, wherein delta is generally 2-3 degrees until the control surface 12 deflects to the specified maximum deflection degree.
S5, setting a transfer function structure of the actuator 1, wherein in the implementation, the transfer function structure is set as follows:
Where s is a laplace operator, N(s) and D(s) are polynomials related to s, the order of N(s) is smaller than D(s), N 1、n0 is a first order term coefficient and a constant term of the polynomial N(s), and D 2、d1、d0 is a second order term coefficient, a first order term coefficient and a constant term of the polynomial D(s), respectively.
S6, determining a set of parameters (n 0、n1、d0、d1、d2) in the set actuator structure by using a system identification app of matlab for each acquired step response data, so that the step response of the transfer function formula (3) approximates to the actual step response.
S7, drawing a frequency characteristic curve of the transfer function by using a matlab frequency characteristic drawing function for the transfer function corresponding to each group of parameters (n 0、n1、d0、d1、d2), and reading out the bandwidth of the transfer function.
S8, substituting a group of parameters (n 0、n1、d0、d1、d2) with the minimum bandwidth into the transfer function formula (3) to serve as a final transfer function of the actuator 1.
Therefore, the method for testing and modeling the dynamic characteristics of the unmanned aerial vehicle actuator based on the mounting state improves the modeling precision of the actuator, considers the minimum bandwidth state of the actuator, so that the control law is designed according to the actuator model, the safety of the control performance of the unmanned aerial vehicle in the whole envelope flight/actuator in the whole movement range can be ensured on the actuator model level, and the reliability of the flight control law design and the confidence of the flight simulation result are improved.
It should be noted that the above-mentioned embodiments are merely for illustrating the technical solution of the present invention and not for limiting the same, and although the present invention has been described in detail with reference to the preferred embodiments, it should be understood by those skilled in the art that the technical solution of the present invention may be modified or substituted by the same, and the modified or substituted technical solution may not deviate from the spirit and scope of the technical solution of the present invention.
Claims (8)
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