A Simulink-based test signal simulation method

Abstract

The application provides a Simulink-based test signal simulation method, which comprises the following steps: firstly, solving a high-order first-order differential signal and a second-order differential signal through integral feedback of Simulink; then, converting the signals into simplified overload signals, and obtaining ideal angle of attack commands according to an approximate proportional relationship between the overload and the angle of attack; then, introducing a simplified aircraft model to solve an angle of attack rate signal; through a nonlinear proportional integral differential of an angle of attack error signal, a pitch channel rudder deflection control signal is formed to realize stable tracking of the angle of attack and to solve a pitch angle velocity signal and a pitch angle signal; then, a rate gyroscope model and an angle gyroscope model are established by using Simulink, and the pitch angle velocity signal, the angle signal and random errors are input into the models respectively to simulate drift of the gyroscope, and finally, aircraft pitch angle velocity and angle measurement equivalent signals are obtained to realize the function of a test equivalent device. The method has the advantage that a simulated dynamic process is more real.

Description

Technical Field

[0001] This invention relates to the field of test equivalents, and more specifically, to a test signal simulation method based on Simulink. Background Technology

[0002] In recent years, research on test equivalence devices has attracted increasing interest from scholars. Test equivalence devices are mainly used in large-scale test systems to replace the object under test, reducing damage to the object while simulating it to verify the correctness of the test system's operation. For example, the launch of large weapon systems and rockets requires multiple tests. Furthermore, the operation process of test systems is quite complex, and these systems can detect faults in complex systems. However, training on the operation of test systems requires multiple experimental tests. Using actual rocket or weapon launch conditions for testing in these situations is clearly inappropriate. Once the test is completed, the rocket has already been propelled and cannot be deactivated; moreover, repeated power-on and power-off cycles on the real system will damage it and affect its lifespan.

[0003] Therefore, using a test equivalent to simulate the attitude change process of an aircraft, and to simulate the signals and states of measurement components such as the gyroscope system on the aircraft, is very meaningful. Traditional methods often only focus on the on / off state of the gyroscope, or perform a simple and coarse simulation of the angle signals during the flight process of the aircraft, but cannot reproduce the complex dynamic process of the aircraft's angle from disturbance to stability during high-speed motion; while using full ballistic simulation is too complex and completely unnecessary. Based on the above background reasons, this invention proposes a method that can accurately simulate the attitude angular velocity, i.e., pitch angular velocity, pitch angle, and even pitch acceleration during high-speed flight using only the aircraft's altitude trajectory data, thus having high engineering practical value.

[0004] It should be noted that the information in the background section above is only used to enhance the understanding of the background of the present invention, and therefore may include information that does not constitute prior art known to those skilled in the art. Summary of the Invention

[0005] The purpose of this invention is to provide a test signal simulation method based on Simulink, thereby overcoming the problem that the simulation of the dynamic process of the gyroscope signal provided to the test system is not realistic enough due to the limitations and defects of related technologies.

[0006] According to one aspect of the present invention, a Simulink-based test signal simulation method is provided, comprising the following steps: Step S10: First, based on the requirements of the aircraft test mission, generate the aircraft's altitude trajectory data using a function; or directly use the actual flight altitude trajectory data and horizontal trajectory data; whereby the altitude trajectory data is denoted as... Its representative is The aircraft's altitude trajectory data at any given time, including The time interval of the discrete data is defined; then, the integrator in Simulink is used, the initial value of the integrator is set, and it is compared with the aircraft altitude trajectory data to form a feedback to generate an altitude error signal, which is then amplified to obtain the first-order differential altitude signal; then it is integrated to obtain the altitude inertial hysteresis signal.

[0007] Step S20: Based on the first-order differential altitude signal, an integrator in Simulink is used. The initial value of the integrator is set and compared with the first-order differential altitude signal to obtain the differential altitude error signal. Then, the signal is amplified proportionally to obtain the second-order differential altitude signal. The signal is then integrated to obtain the differential altitude hysteresis signal. The longitudinal simplified overload signal is then calculated based on the second-order differential altitude signal. Finally, the ideal angle of attack signal of the aircraft is calculated based on the longitudinal simplified overload and angle of attack ratio of the aircraft.

[0008] Step S30: Set the initial values ​​of the aircraft angle of attack signal and pitch rate signal to zero. Solve for the aircraft angle of attack rate signal based on the simplified linear model of the aircraft. Then integrate the aircraft angle of attack signal to obtain the aircraft angle of attack signal. Compare the aircraft angle of attack with the ideal aircraft angle of attack signal to obtain the angle of attack error signal. Then perform a nonlinear transformation on the angle of attack error signal to obtain the angle of attack error nonlinear signal. Then integrate the angle of attack error signal to obtain the angle of attack error integral signal. Finally, superimpose the angle of attack rate signal and the angle of attack error nonlinear signal from the simplified linear model to form the pitch rudder deflection control signal.

[0009] Step S40: Based on the simplified linear model of the aircraft, the pitch acceleration signal of the aircraft is solved by using the pitch control signal, the pitch velocity signal of the aircraft, and the angle of attack signal of the aircraft; then the aircraft pitch velocity signal is obtained by integration, and the aircraft pitch angle signal is obtained by integration based on the aircraft pitch velocity signal.

[0010] Step S50: Based on the model parameters of the aircraft rate gyroscope, the rate gyroscope damping coefficient and natural frequency are set. A second-order rate gyroscope system model is built using Simulink. The aircraft pitch angular velocity signal and velocity random error parameters are input to obtain the equivalent signal for aircraft pitch angular velocity measurement. Based on the model parameters of the aircraft angle gyroscope, the angle gyroscope damping coefficient and natural frequency are set. A second-order angle gyroscope system model is built using Simulink. The aircraft pitch angle signal is input and compared with the equivalent signal for aircraft pitch angle measurement to obtain the pitch angle error signal. Then, a first-order inertial system is designed based on the angle gyroscope damping coefficient and natural frequency, and the pitch angle error signal is input to this system to obtain the pitch angular rate measurement signal. The angle random error parameters are then superimposed to obtain the pitch angle gyroscope integral input signal. Finally, integration is performed to obtain the equivalent signal for aircraft pitch angle measurement, which is then provided to the test system to realize the function of the test equivalent unit.

[0011] In one exemplary embodiment of the present invention, an integrator from Simulink is used. An initial value is set for the integrator, and it is compared with the aircraft's altitude trajectory data to form a feedback altitude error signal. This signal is then amplified to obtain the first-order differential altitude signal. Further integration yields the altitude inertial hysteresis signal, including: ; ; ; in This is a height error signal. It is a highly first-order differential signal; This is a signal with high inertial hysteresis. It is a constant value and is the integration time parameter.

[0012] In one exemplary embodiment of the present invention, based on the first-order differential altitude signal, an integrator in Simulink is used. An initial value is set for the integrator and compared with the first-order differential altitude signal to obtain a differential altitude error signal. This error signal is then amplified to obtain a second-order differential altitude signal. Further integration yields a differential altitude hysteresis signal. The longitudinal simplified overload signal is then calculated based on the second-order differential altitude signal. Finally, the ideal angle-of-attack signal of the aircraft is calculated based on the relationship between the longitudinal simplified overload and the angle-of-attack ratio. ; ; ; ; ; in This is a highly differential error signal. It is a highly second-order differential signal; It is a highly differentially delayed signal; , is a constant, and is the integration time parameter, where It is the gravitational acceleration constant; To simplify the overload signal in the longitudinal direction, For the ideal angle of attack signal, To simplify the overload and angle of attack constant proportional parameters in the longitudinal direction, The initial value is the highly differentially delayed signal.

[0013] In one exemplary embodiment of the present invention, the angle-of-attack rate signal of the aircraft is solved based on a simplified linear model of the aircraft; then, the aircraft angle-of-attack signal is obtained by integration; the aircraft angle of attack is compared with the ideal angle-of-attack signal to obtain an angle-of-attack error signal; then, a nonlinear transformation is performed on the angle-of-attack error signal to obtain a nonlinear angle-of-attack error signal; then, the angle-of-attack error signal is integrated to obtain an integral angle-of-attack error signal; finally, the angle-of-attack rate signal of the simplified linear model and the nonlinear angle-of-attack error signal are superimposed to form a pitch rudder control signal, including: ; ; ; ; ; ; in , The force-related constant parameters for the simplified linear model of the aircraft are determined by the aerodynamic shape and aerodynamic characteristics of the aircraft. This is the aircraft's angle-of-attack rate signal; For the aircraft's angle of attack signal, This is the angle of attack error signal. This is a nonlinear signal representing the angle of attack error. , and These are constant parameters for nonlinear transformations; This is the integral signal of the angle of attack error; For pitch control; , , These are constant control parameters.

[0014] In one exemplary embodiment of the present invention, based on a simplified linear model of the aircraft, the pitch acceleration signal of the aircraft is solved using the pitch control rudder signal, the pitch rate signal of the aircraft, and the angle of attack signal of the aircraft; then, the aircraft pitch rate signal is obtained by integration; and finally, the pitch angle signal is obtained by integration based on the pitch rate signal. ; ; ; in , , The torque-related constant parameters for the simplified linear model of the aircraft are determined by the aerodynamic shape and aerodynamic characteristics of the aircraft. This is the pitch angle acceleration signal of the aircraft; For the aircraft's pitch angular velocity signal, This is the pitch angle signal for the aircraft.

[0015] In one exemplary embodiment of the present invention, based on the model parameters of the aircraft rate gyroscope, the rate gyroscope damping coefficient and natural frequency are set. A second-order rate gyroscope system model is built using Simulink. The aircraft pitch angular velocity signal and velocity random error parameters are input to obtain the equivalent signal for aircraft pitch angular velocity measurement. Based on the model parameters of the aircraft angle gyroscope, the angle gyroscope damping coefficient and natural frequency are set. A second-order angle gyroscope system model is built using Simulink. The aircraft pitch angle signal is input and compared with the equivalent signal for aircraft pitch angle measurement to obtain the pitch angle error signal. Then, based on the angle gyroscope damping coefficient and natural frequency, a first-order inertial system is designed and the pitch angle error signal is input to the system to obtain the pitch angular rate measurement signal. The angle random error parameters are then superimposed to obtain the pitch angle gyroscope integral input signal. Finally, integration is performed to obtain the equivalent signal for aircraft pitch angle measurement, including: ; ; ; ; ; ; in The differential operator for the model transfer function of the aircraft rate gyroscope. This refers to the rate gyroscope damping coefficient. The natural frequency of the rate gyroscope. The input signal is the second-order system model of the rate gyroscope. For speed random error parameters, Equivalent signal for measuring the pitch angular velocity of an aircraft; This refers to the damping coefficient of the angle gyroscope. The natural frequency parameter of the angle gyroscope. This is the pitch angle error signal; The pitch rate measurement signal; The input signal is used for the pitch angle gyroscope integration. Angular random error parameters, This provides an equivalent signal for measuring the pitch angle of an aircraft.

[0016] Beneficial effects This invention presents a Simulink-based test signal simulation method, with six main innovations: First, this method only requires aircraft altitude trajectory data to simulate attitude angles including pitch angular velocity signals, whereas conventional methods also require horizontal trajectory data; thus, it has a wider range of applications. Second, by introducing a simplified aircraft model, this method can more accurately simulate the dynamic characteristics of aircraft in flight engineering; other methods only simulate under the assumption of ideal rotation processes, thus failing to accurately simulate the complex dynamic changes in angular velocity during high-speed aircraft movement. Third, this method can generate pitch acceleration signals of the aircraft, thereby enabling the simulation of pitch acceleration signals and expanding the functionality of the test equivalent. Fourth, this method is based on the nonlinear proportional-integral-differential (PID) method of the angle-of-attack error. The introduction of the differential signal, in particular, simplifies the analysis and design of the control process. Furthermore, since it is an equivalent simulation, no actual control is required. This allows the introduction of the angle-of-attack rate signal to simplify the control design process while still meeting the needs of accurate simulation. Our numerous experiments have shown that this method closely matches the attitude angle changes during flight controlled by the attitude stabilization system, resulting in high accuracy and meeting the requirements for testing equivalents. Fifth, using Simulink to build gyroscope and other models offers advantages such as simple and convenient modeling, and facilitates parameter tuning. Simultaneously, the integral feedback method solves the problem of inconvenient initial state setting for inertial elements in Simulink. Sixth, the simple conversion relationship from overload to angle of attack avoids complex calculations and ensures clear physical meaning, making it very easy to implement in engineering.

[0017] It should be understood that the above general description and the following detailed description are exemplary and explanatory only, and are not intended to limit the invention. Attached Figure Description

[0018] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and, together with the description, serve to explain the principles of the invention. It is obvious that the drawings described below are merely some embodiments of the invention, and those skilled in the art can obtain other drawings based on these drawings without any inventive effort.

[0019] Figure 1 This is a flowchart of a test signal simulation method based on Simulink provided by the present invention; Figure 2 This is a curve showing the change in altitude trajectory over time (unit: meters) provided by the method in this embodiment of the invention. Figure 3 This is the high-order second-differential signal curve (unit: meters per second squared) of the method provided in the embodiments of the present invention. Figure 4 This is the ideal angle-of-attack signal curve (unit: degrees) of the method provided in the embodiments of the present invention. Figure 5 This is the aircraft angle-of-attack error signal curve (unit: degrees) provided by the method in the embodiment of the present invention. Figure 6 This is the aircraft pitch angular velocity signal curve (unit: degrees per second) provided by the method in the embodiments of the present invention. Figure 7 This is the aircraft pitch angle signal curve (unit: degrees) of the method provided in the embodiments of the present invention. Figure 8 This is the equivalent signal curve (unit: degrees per second) for measuring the pitch angular velocity of an aircraft using the method provided in the embodiments of the present invention. Figure 9 This is the equivalent signal curve (unit: degrees) for measuring the pitch angle of an aircraft provided by the method in the embodiments of the present invention. Detailed Implementation

[0020] Exemplary embodiments will now be described more fully with reference to the accompanying drawings. However, these exemplary embodiments can be implemented in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided to make the invention more comprehensive and complete, and to fully convey the concept of the exemplary embodiments to those skilled in the art. The described features, structures, or characteristics can be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to give a full understanding of embodiments of the invention. However, those skilled in the art will recognize that the technical solutions of the invention may be practiced with one or more of these specific details omitted, or other methods, components, apparatus, steps, etc., may be employed. In other instances, well-known technical solutions are not shown or described in detail to avoid obscuring various aspects of the invention.

[0021] This invention provides a Simulink-based test signal simulation method that can accurately simulate pitch attitude information during flight using only aircraft altitude trajectory data, fulfilling the requirements of a test equivalent. First, it solves for the first and second differential altitude signals using Simulink's integral feedback. Then, it converts these signals into simplified overload signals. Based on the approximate proportional relationship between overload and angle of attack, it obtains the ideal angle of attack command. Next, it introduces a simplified aircraft model that does not consider non-minimum phase characteristics to solve for the angle of attack rate signal. Then, it constructs the pitch channel rudder deflection control signal using the nonlinear proportional-integral-differential method of the angle of attack error signal to achieve stable angle of attack tracking. Simultaneously, it calculates the pitch velocity and pitch angle signals based on the simplified aircraft model. Finally, it uses Simulink to establish rate gyroscope and angle gyroscope models, inputting the pitch velocity, pitch angle signals, and random errors of velocity and angle to simulate gyroscope drift. Ultimately, it obtains the equivalent signals for aircraft pitch velocity and angle measurements, providing them to the test system to fulfill the function of a test equivalent. The advantage of this method is that it only requires altitude trajectory data and does not require horizontal data, and it can reflect the changes in angle and angular velocity during the dynamic process of the aircraft, thus making it more realistic.

[0022] The following will further explain and illustrate a Simulink-based test signal simulation method of the present invention with reference to the accompanying drawings. (Reference) Figure 1 As shown, this Simulink-based test signal simulation method may include the following steps: Step S10: First, based on the requirements of the aircraft test mission, generate the aircraft's altitude trajectory data using a function; or directly use the actual flight altitude trajectory data and horizontal trajectory data; whereby the altitude trajectory data is denoted as... Its representative is The aircraft's altitude trajectory data at any given time, including The time interval of the discrete data is defined; then, the integrator in Simulink is used, the initial value of the integrator is set, and it is compared with the aircraft altitude trajectory data to form a feedback to generate an altitude error signal, which is then amplified to obtain the first-order differential altitude signal; then it is integrated to obtain the altitude inertial hysteresis signal.

[0023] Specifically, the generation of the first-order differential signal of altitude can be decomposed into the following three steps. The first step is to use Simulink's integrator based on the aircraft's altitude trajectory data, setting the initial value of the integrator to... The altitude error signal is generated by comparing the data with the aircraft's altitude trajectory data and forming a feedback loop. ; in This is the altitude error signal.

[0024] The second step involves amplifying the height error signal to obtain the first-order differential height signal, as follows: ; in It is a highly first-order differential signal.

[0025] The third step involves integrating the first-order differential signal of the height to obtain the height inertial hysteresis signal. ; in This is a signal with high inertial hysteresis. It is a constant value and is the integration time parameter. The initial value of the highly inertial hysteresis signal is constant.

[0026] Step S20: Based on the first-order differential altitude signal, an integrator in Simulink is used. The initial value of the integrator is set and compared with the first-order differential altitude signal to obtain the differential altitude error signal. Then, the signal is amplified proportionally to obtain the second-order differential altitude signal. The signal is then integrated to obtain the differential altitude hysteresis signal. The longitudinal simplified overload signal is then calculated based on the second-order differential altitude signal. Finally, the ideal angle of attack signal of the aircraft is calculated based on the longitudinal simplified overload and angle of attack ratio of the aircraft.

[0027] Specifically, it can be broken down into the following five steps. The first step is to use Simulink's integrator, based on the aforementioned first-order differential signal, and set the initial value of the integrator to... The height differential error signal is obtained by comparing it with the first-order differential signal of height: ; in This is a highly differential error signal. It is a highly differentially delayed signal.

[0028] The second step is to amplify the height differential error signal further to obtain the second-order height differential signal as follows: ; in It is a highly second-order differential signal.

[0029] The third step is to integrate the second-order differential signal of height to obtain the differential hysteresis signal of height as follows: ; in It is a highly differentially delayed signal; It is a constant value and is the integration time parameter. The initial value is the highly differentially delayed signal.

[0030] The fourth step is to solve for the simplified longitudinal overload signal based on the aforementioned second-order differential signal, as follows: ; in To simplify the overload signal in the longitudinal direction, is the gravitational acceleration constant.

[0031] Fifth, based on the simplified longitudinal overload and angle of attack ratio of the aircraft, the ideal angle of attack signal of the aircraft is calculated as follows: ; in For the ideal angle of attack signal, The longitudinal overload and angle of attack constant ratio parameters are simplified.

[0032] Step S30: Then, set the initial values ​​of the aircraft angle of attack signal and pitch rate signal to zero. Solve for the aircraft angle of attack rate signal based on the simplified linear model of the aircraft. Then, integrate the aircraft angle of attack signal to obtain the aircraft angle of attack signal. Compare the aircraft angle of attack with the ideal aircraft angle of attack signal to obtain the angle of attack error signal. Then, perform a nonlinear transformation on the angle of attack error signal to obtain the angle of attack error nonlinear signal. Then, integrate the angle of attack error signal to obtain the angle of attack error integral signal. Finally, superimpose the angle of attack rate signal and the angle of attack error nonlinear signal from the simplified linear model to form the pitch rudder deflection control signal.

[0033] Specifically, this can be broken down into the following six steps. Step 1: Given that the initial values ​​of the aircraft's angle of attack signal and pitch rate signal are zero, and based on the simplified linear model of the aircraft, the aircraft's angle of attack rate signal is calculated as follows: ; in , The force-related constant parameters for the simplified linear model of the aircraft are determined by the aerodynamic shape and aerodynamic characteristics of the aircraft. This is the aircraft's angle-of-attack rate signal.

[0034] The second step involves integrating the aircraft's angle-of-attack rate signal to obtain the aircraft's angle-of-attack signal, as follows: ; in This is the angle-of-attack signal for the aircraft.

[0035] The third step is to compare the aircraft's angle of attack with the ideal angle of attack signal to obtain the angle of attack error signal as follows: ; in This is the angle of attack error signal.

[0036] The fourth step involves performing a nonlinear transformation on the angle-of-attack error signal to obtain the following nonlinear signal: ; in This is a nonlinear signal representing the angle of attack error. , and These are constant parameters for nonlinear transformations.

[0037] Fifth, the angle-of-attack error signal is integrated to obtain the integrated angle-of-attack error signal as follows: ; in This is the integral signal of the angle of attack error.

[0038] Step 6: Based on the angle-of-attack error signal and the integral signal of the angle-of-attack error, and superimposed with the angle-of-attack rate signal of the simplified linear model and the nonlinear signal of the angle-of-attack error, the pitch rudder deflection control signal is formed as follows: ; in For pitch control; , , These are constant control parameters.

[0039] In step S40, based on the simplified linear model of the aircraft, the pitch acceleration signal of the aircraft is solved by using the pitch control signal, the pitch velocity signal of the aircraft, and the angle of attack signal of the aircraft; then the aircraft pitch velocity signal is obtained by integration, and the aircraft pitch angle signal is obtained by integration based on the aircraft pitch velocity signal.

[0040] Specifically, this can be broken down into the following three steps. The first step, based on the simplified linear model of the aircraft, uses the pitch control signal, the aircraft's pitch velocity signal, and the aircraft's angle of attack signal to calculate the aircraft's pitch acceleration signal as follows: ; in , , The torque-related constant parameters for the simplified linear model of the aircraft are determined by the aircraft's aerodynamic shape and aerodynamic characteristics. This is the pitch angular velocity signal of the aircraft.

[0041] The second step involves integrating the aircraft's pitch acceleration to obtain the aircraft's pitch velocity signal, as follows: ; in This is the pitch angle acceleration signal of the aircraft.

[0042] The third step involves integrating the aircraft's pitch angular velocity signal to obtain the aircraft's pitch angle signal, as follows: ; in This is the pitch angle signal for the aircraft.

[0043] In step S50, based on the model parameters of the aircraft rate gyroscope, the rate gyroscope damping coefficient and natural frequency are set. A second-order rate gyroscope system model is built using Simulink. The aircraft pitch angular velocity signal and velocity random error parameters are input to obtain the equivalent signal for aircraft pitch angular velocity measurement. Based on the model parameters of the aircraft angle gyroscope, the angle gyroscope damping coefficient and natural frequency are set. A second-order angle gyroscope system model is built using Simulink. The aircraft pitch angle signal is input and compared with the equivalent signal for aircraft pitch angle measurement to obtain the pitch angle error signal. A first-order inertial system is designed based on the angle gyroscope damping coefficient and natural frequency, and the pitch angle error signal is input to this system to obtain the pitch angular rate measurement signal. The angle random error parameters are then superimposed to obtain the pitch angle gyroscope integral input signal. Integration is then performed to obtain the equivalent signal for aircraft pitch angle measurement, which is then provided to the test system to realize the function of the test equivalent unit.

[0044] Specifically, this can be broken down into the following five steps. First, based on the model parameters of the aircraft's rate gyroscope, the rate gyroscope damping coefficient and natural frequency are set. A second-order rate gyroscope system model is built using Simulink. The aircraft's pitch rate signal and random error parameters are input, and the equivalent signal for measuring the aircraft's pitch rate is obtained as follows: ; ; in The differential operator for the model transfer function of the aircraft rate gyroscope. This refers to the rate gyroscope damping coefficient. The natural frequency of the rate gyroscope. The input signal is the second-order system model of the rate gyroscope. For speed random error parameters, This is the equivalent signal for measuring the pitch angular velocity of an aircraft.

[0045] The second step involves setting the damping coefficient and natural frequency of the angle gyroscope based on its model parameters. A second-order system model of the angle gyroscope is then built using Simulink. The aircraft pitch angle signal is input and compared with the equivalent signal from the aircraft pitch angle measurement to obtain the following pitch angle error signal: ; in This refers to the damping coefficient of the angle gyroscope. The natural frequency parameter of the angle gyroscope. This is the pitch angle error signal.

[0046] The third step involves designing a first-order inertial system based on the gyroscope's damping coefficient and natural frequency, and inputting the pitch angle error signal into the system to obtain the pitch rate measurement signal as follows: ; in This is the pitch rate measurement signal.

[0047] The fourth step is to obtain the pitch angle gyro integral input signal by superimposing the pitch rate measurement signal with random error parameters, as follows: ; in The input signal is used for the pitch angle gyroscope integration. Angle random error parameter.

[0048] Fifth, the pitch angle gyroscope integral input signal is integrated to obtain the equivalent signal for measuring the aircraft pitch angle, as follows: ; in This provides an equivalent signal for measuring the pitch angle of an aircraft.

[0049] Finally, by providing the equivalent signals of the aircraft pitch rate measurement and the aircraft pitch angle measurement to the test system, the function of the test equivalent generator can be realized.

[0050] Case Implementation and Computer Simulation Results Analysis In step S10, the variation law of the altitude trajectory over time is set as follows: The altitude trajectory curve over time is shown in the figure. Figure 2 As shown.

[0051] In step S20, select The highly second-order differential signal is obtained as follows: Figure 3 As shown. Select , To obtain the ideal angle of attack signal, such as Figure 4 As shown.

[0052] In step S30, select , , , , , , , The angle of attack error signal is obtained as follows: Figure 5 As shown.

[0053] In step S40, select , , The pitch angular velocity signal of the aircraft is obtained as follows: Figure 6 As shown; the aircraft pitch angle signal is obtained as follows: Figure 7 As shown.

[0054] In step S50, select The signal is a uniformly distributed random signal; select The signal is a Gaussian-distributed random signal. , , , The equivalent signal of the aircraft pitch angular velocity measurement is obtained, such as Figure 8 As shown; the equivalent signal for measuring the aircraft pitch angle is obtained as follows: Figure 9 As shown.

[0055] Depend on Figure 4 It can be seen that the ideal angle of attack signal can be generated based on the altitude trajectory, and its magnitude does not exceed the 6-degree limit, meeting engineering requirements and conforming to the background of dynamic physical changes. Figure 5 It can be seen that the angle-of-attack error converges to 0, and its dynamic convergence process simulates the control process of the missile attitude stabilization system, thus matching the physical process and resulting in higher simulation accuracy. Figure 6 and Figure 7 It can be seen that both the pitch angular velocity and pitch angle fluctuate within the normal range, consistent with their physical motion background; Figure 8 and Figure 9 It can be seen that the output signal introduces small glitches, which is consistent with the output curve state of the gyroscope under the influence of error. This allows the entire scheme to simulate the physical relationship and dynamic process between the gyroscope and the aircraft motion in a very detailed and accurate manner, and also meets the requirements of the test equivalent.

Claims

1. A method for simulating equivalent test signals based on Simulink, characterized by the following steps: Step S10: First, based on the requirements of the aircraft test mission, generate the aircraft's altitude trajectory data using a function; or directly use the actual flight altitude trajectory data and horizontal trajectory data; whereby the altitude trajectory data is denoted as... Its representative is The aircraft's altitude trajectory data at any given time, including The time interval for discrete data is defined. Then, an integrator in Simulink is used. The initial value of the integrator is set, and it is compared with the aircraft altitude trajectory data to form a feedback altitude error signal. This signal is then amplified to obtain the first-order differential altitude signal. Finally, integration yields the altitude inertial hysteresis signal as follows: ； ； ； in This is a height error signal. It is a highly first-order differential signal; This is a signal with high inertial hysteresis. It is a constant value and is the integration time parameter. The initial value of the highly inertial hysteresis signal is constant. Step S20: Based on the first-order differential altitude signal, an integrator in Simulink is used. The initial value of the integrator is set and compared with the first-order differential altitude signal to obtain the altitude differential error signal. This error signal is then amplified to obtain the second-order differential altitude signal. Further integration yields the altitude differential hysteresis signal. The simplified longitudinal overload signal is then calculated based on the second-order differential altitude signal. Finally, the ideal angle-of-attack signal of the aircraft is calculated based on the relationship between the simplified longitudinal overload and the angle-of-attack ratio, as follows: ； ； ； ； ； in This is a highly differential error signal. It is a highly second-order differential signal; It is a highly differentially delayed signal; is a constant value, which is the integration time parameter, where It is the gravitational acceleration constant; To simplify the overload signal in the longitudinal direction, For the ideal angle of attack signal, To simplify the overload and angle of attack constant proportional parameters in the longitudinal direction, The initial value is for the highly differentially delayed signal; Step S30: Set the initial values ​​of the aircraft's angle of attack signal and pitch rate signal to zero. Solve for the aircraft's angle of attack rate signal based on the simplified linear model of the aircraft. Integrate the result to obtain the aircraft's angle of attack signal. Compare the aircraft's angle of attack with the ideal angle of attack signal to obtain the angle of attack error signal. Perform a nonlinear transformation on the angle of attack error signal to obtain the angle of attack error nonlinear signal. Integrate the angle of attack error signal to obtain the angle of attack error integral signal. Finally, superimpose the angle of attack rate signal from the simplified linear model and the angle of attack error nonlinear signal to form the pitch rudder control signal as follows: ； ； ； ； ； ； in , The force-related constant parameters for the simplified linear model of the aircraft are determined by the aerodynamic shape and aerodynamic characteristics of the aircraft. This is the aircraft's angle-of-attack rate signal; For the aircraft's angle of attack signal, This is the angle of attack error signal. This is a nonlinear signal representing the angle of attack error. , and These are constant parameters for nonlinear transformations; This is the integral signal of the angle of attack error; For pitch control; , , These are constant control parameters; Step S40: Based on the simplified linear model of the aircraft, the pitch acceleration signal is calculated using the pitch control signal, the pitch rate signal, and the angle of attack signal. Then, the pitch rate signal is obtained by integration, and finally, the pitch angle signal is obtained by integration of the pitch rate signal as follows: ； ； ； in , , The torque-related constant parameters for the simplified linear model of the aircraft are determined by the aerodynamic shape and aerodynamic characteristics of the aircraft. This is the pitch angle acceleration signal of the aircraft; For the aircraft's pitch angular velocity signal, For the aircraft's pitch angle signal; Step S50: Based on the model parameters of the aircraft rate gyroscope, the rate gyroscope damping coefficient and natural frequency are set. A second-order rate gyroscope system model is built using Simulink. The aircraft pitch rate signal and velocity random error parameters are input to obtain the equivalent signal for aircraft pitch rate measurement. Based on the model parameters of the aircraft angle gyroscope, the angle gyroscope damping coefficient and natural frequency are set. A second-order angle gyroscope system model is built using Simulink. The aircraft pitch angle signal is input and compared with the equivalent signal for aircraft pitch angle measurement to obtain the pitch angle error signal. Then, a first-order inertial system is designed based on the angle gyroscope damping coefficient and natural frequency, and the pitch angle error signal is input to this system to obtain the pitch rate measurement signal. The angle random error parameters are then superimposed to obtain the pitch angle gyroscope integral input signal. Integration is then performed to obtain the equivalent signal for aircraft pitch angle measurement, which is then provided to the test system. The function of the test equivalent unit is as follows: ； ； ； ； ； ； in The differential operator for the model transfer function of the aircraft rate gyroscope. This refers to the rate gyroscope damping coefficient. The natural frequency of the rate gyroscope. The input signal is the second-order system model of the rate gyroscope. For speed random error parameters, Equivalent signal for measuring the pitch angular velocity of an aircraft; This refers to the damping coefficient of the angle gyroscope. The natural frequency parameter of the angle gyroscope. This is the pitch angle error signal; The pitch rate measurement signal; The input signal is used for the pitch angle gyroscope integration. Angular random error parameters, This provides an equivalent signal for measuring the pitch angle of an aircraft.

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