CN108845553B

CN108845553B - Servo elastic vibration suppression comprehensive inspection method for slender aircraft

Info

Publication number: CN108845553B
Application number: CN201810618582.8A
Authority: CN
Inventors: 奚勇; 冯昊; 陈光山; 唐德佳; 陶键; 霍红梅; 王璐
Original assignee: Shanghai Aerospace Control Technology Institute
Current assignee: Shanghai Aerospace Control Technology Institute
Priority date: 2018-06-15
Filing date: 2018-06-15
Publication date: 2021-03-09
Anticipated expiration: 2038-06-15
Also published as: CN108845553A

Abstract

The invention discloses a comprehensive test method for servo elastic vibration suppression of a slender aircraft, which comprises the following implementation steps: step one, obtaining the elastic modal frequency of the aircraft in the pitching and yawing directions through calculation or ground test, and obtaining the torsional modal frequency of the aircraft in the rolling direction. And step two, obtaining the resonant frequency and the magnitude of the gyroscope output signal of the inertia measurement assembly under a certain vibration condition through a ground test. And step three, obtaining the control surface modal frequency of the electric steering engine through a ground test. And step four, selecting the characteristic point with the maximum gain in the closed-loop system, developing a closed-loop system test aiming at the frequency, and checking the servo elastic vibration of the aircraft. According to the invention, the design of the control system is checked by a ground test method, so that the risk of servo elastic vibration caused by vibration, external impact and the like in a real mechanical environment in the flight process is reduced.

Description

Servo elastic vibration suppression comprehensive inspection method for slender aircraft

Technical Field

The invention relates to a ground test method suitable for a slender aircraft, which can comprehensively consider the self vibration frequency of the aircraft, an inertia measurement combination hardware product and an electric steering engine hardware product and comprehensively check the design of a control system aiming at elastic vibration suppression.

Background

With the demand of long-distance flight of an aircraft, a wingless aerodynamic profile with a large slenderness ratio is generally adopted to increase the engine loading and reduce aerodynamic resistance, and the range is also increased in a ballistic high-throwing mode, so that the gain of a control system needs to be increased to realize control. On the other hand, in order to meet the requirement of the aircraft on response speed improvement, the bandwidth of the sensitive element and the electric steering engine is also obviously increased.

The flight control principle is that sensitive flight motion information is measured through inertia, a control system is used for calculating a rudder instruction, a steering engine drives a control surface to deflect, and aerodynamic force and inertial force formed by deflection of the control surface act on an aircraft to change the attitude and overload of the aircraft. If in the flying process, the inertial measurement also senses other high-frequency vibration signals, a high-frequency rudder instruction is formed after the signals are resolved by the control system, the steered engine part responds to the high-frequency vibration of the steering surface, high-frequency aerodynamic force and high-frequency inertia force are formed and act on the aircraft body in return to cause the further vibration of the aircraft body, and therefore the closed-loop feedback loop of the 'body structure-control' shown in the figure 1 is formed.

The servo vibration is a typical nonlinear closed loop feedback, and besides the high-frequency signal enters a feedback loop due to the increase of the bandwidth of a sensitive element and an electric steering engine, the servo vibration also has an important reason that the gain of a control loop is larger. If servo vibration occurs in the flying process, the working current of the steering engine is increased, the performance of the steering engine is reduced, and great risk is brought to the flying test.

Disclosure of Invention

The invention solves the problem that the design of a control system aiming at elastic vibration suppression can be comprehensively checked by comprehensively considering the self vibration frequency of an aircraft, an inertia measurement combination hardware product and an electric steering engine hardware product.

The technical scheme adopted by the invention for solving the technical problem is that,

the method comprises the following steps of firstly, obtaining the elastic modal frequency and the torsional modal frequency of the aircraft through calculation or ground test;

step two, obtaining the resonance frequency and the magnitude of signals output by a gyroscope and an accelerometer of the inertia measurement combination under the vibration condition through a ground test;

step three, obtaining the control surface modal frequency of the electric steering engine through a ground test;

and step four, selecting the characteristic point with the maximum gain in the closed-loop system, developing a closed-loop system test aiming at the frequency, and testing the control capability of the servo elastic vibration of the aircraft.

Further, the frequency range needing attention is determined according to the sampling frequency of the computer, and the upper limit of the frequency range is half of the sampling frequency of the computer.

Further, in the first step, the modal frequencies of the orders in the frequency range of interest are obtained, including the elastic modal frequencies of the orders in the pitch channel, the yaw channel, and the torsional modal frequencies of the orders in the roll channel.

Further, in the second step, in the vibration test of the inertia measurement combination, the test method is to perform random vibration of a certain magnitude along the X, Y, Z axis of the inertia measurement combination, and in each vibration process, the resonant frequency appearing in the output signals of the three gyroscopes and the three accelerometers is recorded.

Further, in the third step, a control surface modal test under an idle condition is carried out, and the modal frequency of the control surface is recorded.

Further, in the fourth step, the closed loop system of the test comprises an aircraft object, inertia measurement combination hardware and electric steering engine hardware (comprising a control surface). And (4) performing examination through a closed loop system test on the characteristic points, wherein the selected test state is the characteristic point with the maximum gain in the closed loop system, and the standard that the closed loop system does not generate servo elastic vibration under vibration excitation is taken as the passing standard of the examination.

The invention has the beneficial effects that:

(1) starting from the mechanism analysis of 'body structure-control', the effect of the servo elastic vibration suppression design is verified in a ground test mode.

(2) And a frequency range needing attention is provided, and a method for only paying attention to the low-order modal frequency in the conventional design is expanded.

(3) Besides the traditional elastic modal frequency, the torsional modal frequency, the inertial measurement combination and the frequency of the electric steering engine are added.

Drawings

The invention is further illustrated by the following figures and examples.

FIG. 1 is a schematic diagram of a servo vibration feedback loop in the prior art.

FIG. 2 is a schematic diagram of a closed loop system servo vibration test according to an embodiment of the present invention.

Fig. 3 is a control schematic block diagram of a pitch channel of an embodiment of the present invention.

Fig. 4 is a control schematic block diagram of a scroll channel of an embodiment of the present invention.

Detailed Description

The invention is further illustrated below with reference to the figures and examples.

The invention discloses a comprehensive test method for the suppression of servo elastic vibration of a slender aircraft, which comprises the following implementation steps:

the frequency range to be paid attention to is determined according to the sampling frequency of the computer, the upper limit of the frequency range is half of the sampling frequency of the computer, the sampling frequency of the computer is 400Hz in the example, and therefore frequencies within 200Hz are paid attention to.

1. Obtaining elastic modal frequencies of each order of an aircraft pitch channel and a yaw channel within 200Hz and torsional modal frequency of each order of a rolling channel within 200Hz through calculation or ground test;

2. carrying out vibration test of the inertia measurement combination, wherein the test method comprises the steps of carrying out random vibration along the X, Y, Z axis of the inertia measurement combination respectively, determining the vibration magnitude according to the actual use condition, and recording the resonance frequency within 200Hz appearing in the output signals of the three gyroscopes and the three accelerometers in each vibration process;

3. carrying out a control surface modal test under a no-load condition, and recording the modal frequency within 200Hz of the control surface;

4. the closed loop system servo vibration test is developed as shown in fig. 2, and the hardware such as the aircraft, the inertia measurement combination, the steering engine, the control plane and the like are all in real states.

And (3) recording Y, Z by the inertia measurement combination, outputting by an axial accelerometer, deducting Y, Z axial gravity component in a control loop, electrifying the inertia measurement combination to work in the test process, and outputting high-frequency signals of angular velocity and acceleration caused by excitation in real time.

The computer is electrified to work, the characteristic point with the maximum gain in the closed-loop system is selected, fixed parameters such as speed and dynamic pressure are bound, angular speed and acceleration information output by the inertial measurement combination are received, and a rudder instruction is resolved according to a control loop shown in fig. 3-4.

The steering engine is required to be provided with a control surface, and in the test process, the steering engine is powered to normally work, receives a steering instruction output by the computer and pays out corresponding steering deviation.

According to the diagram in fig. 2, the vibration exciter applies vibration to the aircraft head at the frequencies of the elastic modal frequency, the inertial measurement unit and the resonant frequency of the electric steering engine itself obtained by the test.

The test state is judged to be normal without the closed-loop system entering self-excited vibration under the vibration excitation. Judgment basis of self-excited vibration:

observing self-excited vibration if the response of the steering engine is in a divergent trend;

the vibration state can be maintained after the vibration exciter is removed, and the vibration is self-excited vibration.

Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications of the present invention without departing from the spirit and scope of the present invention by using the methods and technical contents disclosed above.

Claims

1. A comprehensive checking method for servo elastic vibration suppression of a slender aircraft is characterized by comprising the following steps:

selecting a characteristic point with the maximum gain in the closed-loop system, developing a closed-loop system test aiming at the frequency, and testing the control capability of the servo elastic vibration of the aircraft;

focusing on a certain range of elastic modal frequency, wherein the upper limit of the elastic modal frequency range is half of the sampling frequency of a computer;

the elastic modal frequencies in the first step comprise elastic modal frequencies of all orders of a pitching channel and a yawing channel in a concerned frequency range and torsional modal frequencies of all orders of a rolling channel;

in the vibration test of the inertia measurement combination in the second step, the test method is that vibration is carried out along the X, Y, Z axis respectively, the vibration magnitude is determined by using conditions, and the resonant frequency appearing in the output signals of the three gyroscopes and the three accelerometers is recorded in the vibration process;

in the test in the third step, the control surface modal frequency under the no-load condition is recorded;

in the closed-loop system test in the fourth step, the closed-loop system comprises an aircraft object, inertial measurement combination hardware and electric steering engine hardware;

in the step four, in the closed loop system test, the control system works in real time, and the characteristic point with the maximum gain in the closed loop system is selected as a test state;

in the closed-loop system test in the fourth step, the condition that the closed-loop system does not generate servo elastic vibration under vibration excitation is taken as a passing standard of examination.