Flight control system structure modal coupling test system and method

Abstract

The invention provides a flight control system structure modal coupling test system and method, and belongs to the field of flight control systems, and the system comprises a test aircraft which comprises a control surface and a servo actuation system used for controlling the control surface to move; the ground oil source and the ground power source are used for providing a hydraulic source and a power source for the test aircraft; the air spring is used for providing support for the test aircraft; the sensor is used for measuring the response of the test aircraft in the excitation state; the flight control computer is used for generating a control law instruction for controlling the servo actuation system to act according to the control law; the tester is arranged between the flight control computer and the servo actuation system and is used for measuring, cutting off or injecting a control law instruction sent by the flight control computer, a test instruction sent by the frequency response analyzer and a sensor output signal; and the frequency response analyzer is connected with the tester and is used for generating a sweep frequency signal or a pulse signal and collecting a control law instruction measured by the tester, and the sensor outputs the control law instruction and performs frequency response analysis.

Description

Technical Field

[0001] The present application belongs to the field of ground testing of flight control systems, and in particular, relates to a flight control system structural modal coupling testing system and method. Background Art

[0002] The fly-by-wire flight control system is a servo control system with wide bandwidth and high authority. Its sensors not only receive the aircraft's rigid body motion signals, but also receive the elastic vibration signals of the structure. These signals output the control commands of the control surfaces through the flight control computer to drive the deflection of the control surfaces. The aerodynamic force generated by the control surfaces in turn affects the rigid body motion and elastic vibration of the aircraft. The feedback coupling between the two main parts of control and structure forms the aerodynamic servoelastic stability problem of the aircraft. The flight control system structural modal coupling test is used to determine the servoelastic stability characteristics of the aircraft. Summary of the invention

[0003] The purpose of the present application is to provide a flight control system structural modal coupling test system to solve or alleviate at least one problem in the background technology.

[0004] The technical solution of the present application is: a flight control system structure modal coupling test system, comprising:

[0005] A test aircraft, the test aircraft comprising a control surface and a servo actuation system for controlling the movement of the control surface;

[0006] Ground oil source and ground power supply for providing hydraulic and electrical power to the test aircraft;

[0007] air springs used to support the test aircraft;

[0008] Sensors used to measure the test aircraft's response under excitation conditions;

[0009] A flight control computer, the flight control computer is used to generate control law instructions for controlling the servo actuator system to perform actions according to the control law;

[0010] A tester, which is arranged between the flight control computer and the servo actuation system and is used to measure, intercept or inject the control law instructions issued by the flight control computer, the test instructions issued by the frequency response analyzer and the sensor output signal;

[0011] The frequency response analyzer is connected to the tester and is used to generate a sweep frequency signal or a pulse signal, collect control law instructions and sensor output measured by the tester, and perform frequency response analysis.

[0012] Preferably, the sensor includes an angular velocity sensor and an overload sensor.

[0013] Preferably, it also includes a flight test interface device, which is connected to the flight control computer and is used to monitor the status of the flight control computer.

[0014] In addition, the present application also provides a method for performing a structural modal coupling test using any of the flight control system structural modal coupling test systems described above, wherein the structural modal coupling test includes an open-loop frequency response test and a closed-loop impulse response test, wherein:

[0015] When conducting an open-loop frequency response test of a flight control system, the flight control computer and the servo actuator system are disconnected, the flight control computer outputs a control law instruction to the servo actuator system through a tester, a control surface sinusoidal frequency sweep signal is output through a frequency response analyzer, the output signal of the control law instruction is measured, the total open-loop transfer function is obtained based on the sinusoidal frequency sweep signal and the output signal of the control law instruction, the body response signal of the test aircraft is measured through a sensor, and the aircraft link-related transfer function is obtained based on the body response signal and the sinusoidal frequency sweep signal;

[0016] When conducting a closed-loop pulse test of the flight control system, the flight control computer and the servo actuator system are connected, the flight control computer outputs a control law instruction to the servo actuator system, and at the same time a rated gain is applied to the servo actuator system through the tester, a pulse signal is input to the operating surface through a frequency response analyzer, the output signal of the control law instruction is measured, and the body response signal of the test aircraft and the response signal of the control surface are measured through sensors.

[0017] Preferably, during the open-loop frequency response test of the flight control system and / or the closed-loop pulse test of the flight control system, if a load-bearing state-related test of the take-off and landing control law is performed, the landing gear is used to support the test aircraft, and if a non-load-bearing state-related test is performed, an air spring is used to support the test aircraft.

[0018] Preferably, when an air spring is used to support the test aircraft, the air spring is inflated.

[0019] The flight control system structural modal coupling test system and method of the present application can simulate the servo-elasticity problem in the air on the ground, and make a quantitative judgment on the servo-elasticity stability problem of the aircraft. When the servo-elasticity does not meet the index requirements, a notch filter can be added according to the test results to solve the problem of insufficient servo-elasticity stability as soon as possible. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] In order to more clearly illustrate the technical solution provided by the present application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of the present application.

[0021] Figure 1This is a composition diagram of the flight control system structural modal coupling test system of this application.

[0022] Figure 2 This is a schematic diagram of the open-loop frequency response test principle in this application.

[0023] Figure 3 Schematic diagram of the closed-loop pulse test principle in this application.

[0024] Figure 4 This is the test result of an open-loop frequency response test of an embodiment of the present application.

[0025] Figure 5 This is a closed-loop pulse test result of an embodiment of the present application. DETAILED DESCRIPTION

[0026] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below in conjunction with the drawings in the embodiments of this application.

[0027] First, Figure 1 The figure shows a schematic diagram of a flight control system (or flight control system for short) structure modal coupling test system provided by the present application. The structure modal coupling test system 100 includes:

[0028] A test aircraft 101 includes a control surface 111 and a servo actuation system 112 for controlling the control surface 111 to move;

[0029] The ground oil source 102 and the ground power source 103 are used to provide hydraulic source and power source for the test aircraft 101;

[0030] An air spring 104 for providing support to the test aircraft 101;

[0031] Sensors, including an angular velocity sensor 105 and an overload sensor 106, for measuring the overload and angular velocity of the test aircraft 101 under control instructions or excitation;

[0032] A flight control computer (FCC) 107 is used to generate control law instructions for controlling the servo actuator system 112 to perform actions according to the control law, and drive the control surface to move;

[0033] The tester 108 (or disconnect box) is connected to the flight control computer 107 and the servo actuator system 112 through a test cable, and can measure, intercept or inject control law instructions issued by the flight control computer 107, test instructions issued by the frequency response analyzer 109, and angular velocity / overload sensor output signals;

[0034] The frequency response analyzer 109 is connected to the tester 108 via a cable, and is used to generate a frequency sweep signal or a pulse signal, and collect control law instructions and sensor output measured by the tester 108 and perform frequency response analysis;

[0035] The flight test interface device (FTI) 110 is connected to the flight control computer 107 and is used to monitor the status of the flight control computer 107 .

[0036] Based on the above-mentioned flight control system structural modal coupling test system, a flight control system structural modal coupling test is carried out. Usually, the structural modal coupling test includes an open-loop frequency response test and a closed-loop pulse response test. The open-loop frequency response test is used to test the open-loop transfer function of the servo-elastic system to quantitatively determine its stability, and the closed-loop pulse test is used to test the time domain response of the servo-elastic system to qualitatively verify its stability.

[0037] like Figure 2 The figure shows a schematic diagram of the open-loop frequency response test principle in the present application. When conducting an open-loop frequency response test of a flight control system, the flight control computer 107 and the servo actuator system 112 are disconnected, and the flight control computer 107 outputs a control law instruction to the servo actuator system 112 through the tester 108. At the same time, a sinusoidal frequency sweep signal X1 is input to the corresponding control surface (i.e., the control surface 111) through the frequency response analyzer 109, and the output signal Y of the control law instruction is measured to obtain a total open-loop transfer function Y / X1. The body response signal X2 of the test aircraft 101 is measured by the angular velocity / overload sensor to obtain the aircraft link-related transfer function X2 / X1.

[0038] like Figure 3 The figure shows a schematic diagram of the closed-loop pulse test principle in the present application. When conducting a closed-loop pulse test of the flight control system, the flight control computer 107 and the servo actuator system 112 are connected, and the flight control computer 107 outputs a control law instruction to the servo actuator system 112. At the same time, a rated gain G is given to the servo actuator system 112 through the tester 108, and a pulse signal X1 is input to the operating surface through the frequency response analyzer 109, and the output signal Y of the control law instruction is measured. The body response signal X2 of the test aircraft 101 and the response of the control surface are measured through the angular velocity / overload sensor.

[0039] The specific process of the flight control system structure modal coupling test of this application is as follows:

[0040] S1, check the support and protection status of the test aircraft 101, and determine the support form of the test aircraft 101 according to the test status, that is, use landing gear support or air spring 104 support, wherein, the landing gear is used to support the test aircraft 101 for the load-bearing state related tests of the take-off and landing control law, and the air spring 104 is used to support the test aircraft 101 for other non-load-bearing state (air state) related tests; when the air spring 104 is used for support, the air spring is inflated.

[0041] S2, power on the test aircraft 101 through the ground power supply 102, and supply pressure to the test aircraft 102 through the ground oil source 103.

[0042] S3, checking the integrity of the flight control computer 107 and the servo actuator system 112 through the flight test interface device 110.

[0043] S4: When conducting the frequency response test of the control law in the air, retract the landing gear and check whether the working state of the flight control system is in the air; when conducting the frequency response test of the control law in the takeoff and landing, lower the landing gear and check whether the working state of the flight control system is in the takeoff and landing state.

[0044] S5, connect the tester 108, and the open-loop frequency response test uses the tester 108 to disconnect from the D / A port of the flight control computer 107 and the servo actuator system 112, so that the servo elastic system becomes an open loop.

[0045] S6, connecting the frequency response analyzer 109 to the tester 108.

[0046] S7, given the dynamic and static pressures, setting the sensors to be normal or failed according to the test content, so that the test aircraft 101 is in a corresponding test state.

[0047] S8, using the tester 108 to input an excitation signal to the servo actuator system 112, and changing the input signal amplitude, and determining a reasonable input signal amplitude by observing the response of the control surface acceleration sensor and visually observing the control surface response.

[0048] S9. Reasonably determine the frequency range of the open-loop frequency response test based on the theoretical analysis results and the servo characteristics.

[0049] S10, use the frequency response analyzer 109 to sweep the frequency of the control surface 112, obtain the frequency characteristic curve, and test, analyze, record and save the data.

[0050] S11, test the remaining test states.

[0051] Table 1 shows the take-off and landing state test contents of the embodiment of the present application, and Table 2 shows the air state test contents of the embodiment of the present application. After the open-loop test is completed, a closed-loop pulse test is performed, and the control law gain is given in combination with the open-loop frequency response test results, and a pulse excitation signal is input to the servo loop, and the necessary time domain response data is tested, analyzed, recorded and saved.

[0052] Table 1 Take-off and landing test contents

[0053]

[0054]

[0055] Table 2 Test contents in air state

[0056]

[0057]

[0058]

[0059] like Figure 4 The figure shows the test results of the open-loop frequency response test of an embodiment of the present application. The frequency characteristic curve of the sine frequency sweep signal X1 and the output signal Y of the control law command measured in the test is shown in the figure. The amplitude margin of this state is 8.26dB, and the phase margin is 180°, which meets the requirements of the national military standard that the amplitude margin is greater than 6dB and the phase margin is greater than 60°. When the servo elasticity does not meet the requirements, a corresponding structural notch filter can be designed to meet the requirements.

[0060] like Figure 5 The figure shows the test results of a closed-loop pulse test of an embodiment of the present application. It can be seen from the test curve that when the flight control system gain is 2.5 times, the response of the test aircraft is attenuated after being subjected to pulse excitation, and the test aircraft is servo-elastic and stable.

[0061] The flight control system structural modal coupling test system and method of the present application can simulate the servo-elasticity problem in the air on the ground, and make a quantitative judgment on the servo-elasticity stability problem of the aircraft. When the servo-elasticity does not meet the index requirements, a notch filter can be added according to the test results to solve the problem of insufficient servo-elasticity stability as soon as possible.

[0062] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A flight control system structural modal coupling test system, characterized in that: include: A test aircraft, the test aircraft comprising a control surface and a servo actuation system for controlling the movement of the control surface; Ground oil source and ground power supply for providing hydraulic and electrical power to the test aircraft; air springs used to support the test aircraft; Sensors used to measure the test aircraft's response under excitation conditions; A flight control computer, the flight control computer is used to generate control law instructions for controlling the servo actuator system to perform actions according to the control law; A tester, which is arranged between the flight control computer and the servo actuation system and is used to measure, intercept or inject the control law instructions issued by the flight control computer, the test instructions issued by the frequency response analyzer and the sensor output signal; The frequency response analyzer is connected to the tester and is used to generate a sweep frequency signal or a pulse signal, collect control law instructions and sensor output measured by the tester, and perform frequency response analysis.

2. The flight control system structural modal coupling test system according to claim 1, characterized in that: The sensors include an angular velocity sensor and an overload sensor.

3. The flight control system structural modal coupling test system according to claim 1, characterized in that: It also includes a flight test interface device, which is connected to the flight control computer and is used to monitor the status of the flight control computer.

4. A method for performing a structural modal coupling test using the flight control system structural modal coupling test system according to any one of claims 1 to 3, characterized in that: The structural modal coupling test includes an open-loop frequency response test and a closed-loop pulse response test, wherein: When conducting an open-loop frequency response test of a flight control system, the flight control computer and the servo actuator system are disconnected, the flight control computer outputs a control law instruction to the servo actuator system through a tester, a control surface sinusoidal frequency sweep signal is output through a frequency response analyzer, the output signal of the control law instruction is measured, the total open-loop transfer function is obtained based on the sinusoidal frequency sweep signal and the output signal of the control law instruction, the body response signal of the test aircraft is measured through a sensor, and the aircraft link-related transfer function is obtained based on the body response signal and the sinusoidal frequency sweep signal; When conducting a closed-loop pulse test of the flight control system, the flight control computer and the servo actuator system are connected, the flight control computer outputs a control law instruction to the servo actuator system, and at the same time a rated gain is applied to the servo actuator system through the tester, a pulse signal is input to the operating surface through a frequency response analyzer, the output signal of the control law instruction is measured, and the body response signal of the test aircraft and the response signal of the control surface are measured through sensors.

5. The method according to claim 4, characterized in that During the open-loop frequency response test of the flight control system and / or the closed-loop pulse test of the flight control system, if the take-off and landing control law load-bearing state related test is performed, the test aircraft is supported by the landing gear, and if the non-load-bearing state related test is performed, the test aircraft is supported by the air spring.

6. The method according to claim 5, characterized in that When the test aircraft is supported by air springs, the air springs are inflated.

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