CN103995918A - Method for analyzing influences of wing deformation and vibration on aircraft transfer alignment - Google Patents

Abstract

The invention belongs to the field of aircraft structural dynamics, and relates to a method for analyzing influences of wing deformation and vibration on aircraft transfer alignment. The method is characterized by including the following steps: (1) building a full-aircraft structure dynamical model, (2) building a full-aircraft unsteady aerodynamic model, (3) building a transfer relationship between plug-ins and inertial navigation, and (4) obtaining a primary-secondary inertial navigation error fitting empirical formula. The method has the advantages that speed errors and angular speed errors generated between the primary inertial navigation and the secondary inertial navigation of an aerial movable platform are analyzed through the built aircraft structure dynamical model, then the errors are eliminated through filtering, and in this way, the transfer alignment time of the primary inertial navigation and the secondary inertial navigation during guided missile launching can be shortened to a great degree.

Description

A kind of wing distortion and the analytical approach of vibration on the impact of aircraft Transfer Alignment

Technical field

This patent belongs to aircaft configuration dynamics field, relates to a kind of wing distortion and the analytical approach of vibration on the impact of aircraft Transfer Alignment.

Background technology

First the inertial navigation system of the guided weapon that aerial platform carries will carry out initial alignment before starting working, and its objective is and makes inertial navigation system can set up suitable navigation coordinate system.The precision of initial alignment is directly connected to the operational performance of armament systems, so initial alignment is one of most important gordian technique of inertial navigation system.In order to eliminate the error between main and sub inertial navigation, adopt the method for flight actual measurement before, this method is that aircraft is under the state of certain Mach number, height and overload, utilize high-speed camera to take displacement and the Vibration Condition of wing distortion, its advantage is to obtain the error of simultaneously considering the many factors such as fire control system airborne equipment, wing are out of shape, guided missile is installed in carrier aircraft, and precision is higher.But the method for this actual measurement is consuming time longer, costly, can not each model adopt.So, this patent adopts large-scale finite element procedure, take statics model as basis, set up model via dynamical response and unsteady aerodynamic model, consider elastic deformation and the uncertain vibration equal error of wing, obtain the response relation of main and sub inertial navigation, thereby simulate graph of errors, for shortening the time of MISSILE LAUNCHING front transfer aligning, provide prerequisite.

Summary of the invention

By the aircaft configuration kinetic model of setting up, the velocity error producing between the main and sub inertial navigation of aerial moving platform and angular velocity error are analyzed, by filtering, eliminate error again, the time of main and sub inertial navigation Transfer Alignment in the time of can shortening MISSILE LAUNCHING to a great extent like this, opportunity of combat is improved to fight capability and be significant.

Technical scheme of the present invention is: a kind of wing distortion and the analytical approach of uncertain vibration on the impact of aircraft inertial navigation system Transfer Alignment.It is characterized in that, comprise the steps:

The first, set up full machine model via dynamical response

With aerial moving platform model, take full machine statics model as basis, carry out mass distribution adjustment,

Each node lumped mass sum that equivalence obtains equals gross mass m;

$\underset{i=1}{\overset{N}{\mathrm{\Σ}}}(m_{\mathrm{Ai}}+m_i)=m---\left(1\right)$

The barycenter of each node lumped mass that equivalence obtains and the coordinate position (x of structure barycenter _m, y _m, z _m) overlap;

$\underset{i=1}{\overset{N}{\mathrm{\Σ}}}(m_{\mathrm{Ai}}+m_i)x_i/m=x_m---\left(2\right)$

$\underset{i=1}{\overset{N}{\mathrm{\Σ}}}(m_{\mathrm{Ai}}+m_i)y_i/m=y_m---\left(3\right)$

$\underset{i=1}{\overset{N}{\mathrm{\Σ}}}(m_{\mathrm{Ai}}+m_i)z_i/m=z_m---\left(4\right)$

The moments of inertia of each node lumped mass that equivalence obtains and equating with total moments of inertia of structure;

$I_x=\underset{i=1}{\overset{N}{\mathrm{\Σ}}}\left\{\right[{(y_i-y_m)}^2+{(z_i-z_m)}^2](m_i+m_{\mathrm{Ai}})+I_{\mathrm{Axi}}\}---\left(5\right)$

$I_y=\underset{i=1}{\overset{N}{\mathrm{\Σ}}}\left\{\right[{(y_i-x_m)}^2+{(z_i-z_m)}^2](m_i+m_{\mathrm{Ai}})+I_{\mathrm{Ayi}}\}---\left(6\right)$

$I_z=\underset{i=1}{\overset{N}{\mathrm{\Σ}}}\left\{\right[{(y_i-y_m)}^2+{(x_i-x_m)}^2](m_i+m_{\mathrm{Ai}})+I_{\mathrm{Azi}}\}---\left(7\right)$

Carry out Dynamic Modeling, according to calculation requirement optimization, revise and obtain full machine model via dynamical response.

The second, set up the non-permanent aerodynamic model of full machine

Adopt dull and stereotyped Aerodynamic Model.For the aeroelastic analysis in subsonic envelope, we have adopted the non-permanent aerodynamic model of ZONA6; For aeroelastic analysis in supersonic envelope, adopt the non-permanent aerodynamic model of ZONA7.These two kinds of models are all the non-permanent three-dimensional linearization microvariations potential equations based on separately.These two kinds of unsteady aerodynamic models are all frequency domain Aerodynamic Model, are reduced frequency k and Mach number M _∞function.

The 3rd, set up the transitive relation between plug-in and inertial navigation

According to full machine aeroelasticity equation, can obtain the aerodynamic loading under unsteady regime, put on model via dynamical response, can calculate frequency response function and transport function between plug-in in micro-of time and inertial navigation, it is carried out to depression of order processing, obtain the low order displacement transitive relation between plug-in and inertial navigation.For the structure of change in displacement, recalculate aerodynamic force and again calculate frequency response function and the transport function between plug-in in micro-of new time and inertial navigation, for the plug-in configuration of difference and state of flight, repeat above process, set up displacement transitive relation data between plug-in and inertial navigation.The aeroelasticity equation of system can be expressed as:

$M_{\mathrm{hh}}\stackrel{..}{q}\left(t\right)+D_{\mathrm{hh}}\stackrel{.}{q}\left(t\right)+K_{\mathrm{hh}}q\left(t\right)=\frac{1}{2}\mathrm{\ρ}{V}^2{Q}_{\mathrm{hh}}(k,M_{\∞})q\left(t\right)+\frac{1}{2}\mathrm{\ρ}{V}^2{Q}_{\mathrm{hg}}(k,M_{\∞})\frac{w_g\left(t\right)}{V}---\left(8\right)$

In formula, M _hh, D _hh, K _hhbe respectively modal mass, modal damping and modal stiffness matrix, q (t) is modal displacement.V represents speed of incoming flow.Matrix Q _hh(k, M _∞) be non-permanent aerodynamic force matrix, wherein, M _∞be Mach number, k is reduced frequency.Matrix Q _hg(k, M _∞) be prominent general mood dynamic matrix, w _g(t) be prominent wind speed degree.

The 4th, main and sub ins error matching experimental formula

According to the features of response of main inertial navigation and sub-inertial navigation, therefrom can seek an experimental formula, show as much as possible error propagation feature between main inertial navigation and sub-inertial navigation.According to the feature of error propagation, suppose that the form of experimental formula is

f ₁(t)＝f ₁(t)sin(2πf ₂(t)) （9）

F wherein ₁(t) be amplitude distracter, f ₂(t) be periodic disturbances item, according to the feature of error propagation amplitude, suppose:

$f_1\left(t\right)=a_1{e}^{-a_{3}t}+a_2{e}^{-a_{4}t}---\left(10\right)$

F wherein ₁(t) be amplitude distracter, f ₂(t) be periodic disturbances item, according to the feature of error propagation amplitude, suppose:

$f_1\left(t\right)=a_1{e}^{-a_{3}t}+a_2{e}^{-a_{4}t}---\left(11\right)$

Application least square method, makes error sum of squares minimum

$\begin{array}{c}{\left|\right|\mathrm{\δ}\left|\right|}_2^2=\underset{i=1}{\overset{8}{\mathrm{\Σ}}}{\mathrm{\δ}}_i^2=\underset{i=1}{\overset{8}{\mathrm{\Σ}}}(s_1\left(t_i\right)-f_1\left(t_i\right))\\ =\min \underset{i=1}{\overset{8}{\mathrm{\Σ}}}(s_1\left(t_i\right)-f_1\left(t_i\right))\end{array}---\left(12\right)$

In like manner suppose

$f_2\left(t\right)=b_1{e}^{-b_{3}t}+b_2{e}^{-{\mathrm{ba}}_{4}t}---\left(13\right)$

Thus, can estimate f ₁and f (t) ₂(t) correlation parameter.

Advantage of the present invention is: by the aircaft configuration kinetic model of setting up, the velocity error producing between the main and sub inertial navigation of aerial moving platform and angular velocity error are analyzed, grasp the typical dynamic characteristics of aerial platform, disclose the action rule of dynamic disturbance to Transfer Alignment precision and time, for Transfer Alignment theoretical research and Project Realization provide interference characteristic analysis and the mechanism of action thereof.Thereby the time of main and sub inertial navigation Transfer Alignment while shortening MISSILE LAUNCHING, opportunity of combat is improved to fight capability and be significant.

Accompanying drawing explanation

Fig. 1 is the response schematic diagram between typical main and sub inertial navigation

Fig. 2 is error and matched curve schematic diagram

Embodiment

Below by concrete enforcement, this patent is described in further detail.

The first, set up full machine model via dynamical response

Aerial moving platform model be take the full machine statics of certain type model as basis,

Carry out mass distribution adjustment, obtain full machine Structural Dynamics computation model.

The operating mass unloden that theory calculates is 498.435kg; Centre of gravity place is X=3.7762m, Y=-0.1m, Z=0.Through check, full machine modal calculation model quality is 497.55kg; Centre of gravity place is X=3.77662m, Y=-0.1m, Z=1.0E-4m.

The second, the complete non-permanent aerodynamic model of machine

Non-permanent aerodynamic model, is divided into 30 aerodynamic force unit altogether.Wherein, 14 unit of fuselage, 8 unit of wing, 4 unit of wing tip carriage, 4 unit of empennage.

The 3rd, the transitive relation between plug-in and inertial navigation

According to full machine aeroelasticity equation, with H=5km, M _∞=0.8, the flat state of flight flying, and adopt the sin wind model of dashing forward, its expression formula is

$T=\left\{\begin{array}{cc}0,& \mathrm{\&tau;}<0\\ \sin \left(\frac{2\mathrm{\&pi;\&tau;}}{L_g}\right)& 0\&le;\mathrm{\&tau;}\&le;L_g\\ 0,& \mathrm{\&tau;}>L_g\end{array}\right.$

Obtain the response between typical main and sub inertial navigation, as shown in Figure 1.

The 4th, main and sub ins error matching experimental formula

According to the form of formula (11), in order to determine f ₁(t) coefficient, gets the relevant information s of graph of errors crest ₁it is (t) as shown in table 1,

Table 1 graph of errors crest

According to the form of formula (13), in order to determine f ₂(t) coefficient, gets the relevant information s in graph of errors cycle ₂it is (t) as shown in table 2,

The table 2 graph of errors cycle

Thus, can estimate f ₁and f (t) ₂(t) correlation parameter is as shown in table 3.

Table 3 amplitude error parameters of formula

So just obtained matched curve as shown in Figure 2.

Thereby obtained the action rule of aerial moving platform dynamic disturbance to Transfer Alignment precision and time, the time of eliminating main and sub inertial navigation Transfer Alignment error for shortening provides help.

Claims (1)

1. wing distortion and the analytical approach of vibration on the impact of aircraft Transfer Alignment, is characterized in that, comprises the steps:

The first, set up full machine Structural Dynamics finite element model

With aerial moving platform model, take full machine statics model as basis, carry out mass distribution adjustment, each node lumped mass sum that equivalence obtains equals gross mass m;

$\underset{i=1}{\overset{N}{\mathrm{\&Sigma;}}}(m_{\mathrm{Ai}}+m_i)=m---\left(1\right)$

The barycenter of each node lumped mass that equivalence obtains and the coordinate position (x of structure barycenter _m, y _m, z _m) overlap;

$\underset{i=1}{\overset{N}{\mathrm{\&Sigma;}}}(m_{\mathrm{Ai}}+m_i)x_i/m=x_m---\left(2\right)$

$\underset{i=1}{\overset{N}{\mathrm{\&Sigma;}}}(m_{\mathrm{Ai}}+m_i)y_i/m=y_m---\left(3\right)$

$\underset{i=1}{\overset{N}{\mathrm{\&Sigma;}}}(m_{\mathrm{Ai}}+m_i)z_i/m=z_m---\left(4\right)$

The moments of inertia of each node lumped mass that equivalence obtains and equating with total moments of inertia of structure;

$I_x=\underset{i=1}{\overset{N}{\mathrm{\&Sigma;}}}\left\{\right[{(y_i-y_m)}^2+{(z_i-z_m)}^2](m_i+m_{\mathrm{Ai}})+I_{\mathrm{Axi}}\}---\left(5\right)$

$I_y=\underset{i=1}{\overset{N}{\mathrm{\&Sigma;}}}\left\{\right[{(y_i-x_m)}^2+{(z_i-z_m)}^2](m_i+m_{\mathrm{Ai}})+I_{\mathrm{Ayi}}\}---\left(6\right)$

$I_z=\underset{i=1}{\overset{N}{\mathrm{\&Sigma;}}}\left\{\right[{(y_i-y_m)}^2+{(x_i-x_m)}^2](m_i+m_{\mathrm{Ai}})+I_{\mathrm{Azi}}\}---\left(7\right)$

Carry out Dynamic Modeling, according to calculation requirement optimization, revise and obtain full machine model via dynamical response;

The second, set up the non-permanent aerodynamic model of full machine

Adopt dull and stereotyped Aerodynamic Model, for the aeroelastic analysis in subsonic envelope, adopt the non-permanent aerodynamic model of ZONA6; For aeroelastic analysis in supersonic envelope, adopt the non-permanent aerodynamic model of ZONA7, these two kinds of models are all the non-permanent three-dimensional linearization microvariations potential equations based on separately, and these two kinds of unsteady aerodynamic models are all frequency domain Aerodynamic Model, are reduced frequency k and Mach number M _∞function;

The 3rd, set up the transitive relation between plug-in and inertial navigation

According to full machine aeroelasticity equation, obtain the aerodynamic loading under unsteady regime, put on model via dynamical response, calculate frequency response function and transport function between plug-in in micro-of time and inertial navigation, it is carried out to depression of order processing, obtain the low order displacement transitive relation between plug-in and inertial navigation, for the structure of change in displacement, recalculate aerodynamic force and again calculate frequency response function and the transport function between plug-in in micro-of new time and inertial navigation, for the plug-in configuration of difference and state of flight, repeat above process, set up displacement transitive relation data between plug-in and inertial navigation, the aeroelasticity the Representation Equation of system is:

$M_{\mathrm{hh}}\stackrel{..}{q}\left(t\right)+D_{\mathrm{hh}}\stackrel{.}{q}\left(t\right)+K_{\mathrm{hh}}q\left(t\right)=\frac{1}{2}\mathrm{\&rho;}{V}^2{Q}_{\mathrm{hh}}(k,M_{\&infin;})q\left(t\right)+\frac{1}{2}\mathrm{\&rho;}{V}^2{Q}_{\mathrm{hg}}(k,M_{\&infin;})\frac{w_g\left(t\right)}{V}---\left(8\right)$

In formula, M _hh, D _hh, K _hhbe respectively modal mass, modal damping and modal stiffness matrix, q (t) is modal displacement, and V represents speed of incoming flow, matrix Q _hh(k, M _∞) be non-permanent aerodynamic force matrix, wherein, M _∞be Mach number, k is reduced frequency, matrix Q _hg(k, M _∞) be prominent general mood dynamic matrix, w _g(t) be prominent wind speed degree;

The 4th, main and sub ins error matching experimental formula

According to the features of response of main inertial navigation and sub-inertial navigation and error propagation feature, suppose that the form of experimental formula is

f ₁(t)＝f ₁(t)sin(2πf ₂(t)) （9）

F wherein ₁(t) be amplitude distracter, f ₂(t) be periodic disturbances item, according to the feature of error propagation amplitude, suppose:

$f_1\left(t\right)=a_1{e}^{-a_{3}t}+a_2{e}^{-a_{4}t}---\left(10\right)$

F wherein ₁(t) be amplitude distracter, f ₂(t) be periodic disturbances item, according to the feature of error propagation amplitude, suppose:

$f_1\left(t\right)=a_1{e}^{-a_{3}t}+a_2{e}^{-a_{4}t}---\left(11\right)$

Application least square method, makes error sum of squares minimum

$\begin{array}{c}{\left|\right|\mathrm{\&delta;}\left|\right|}_2^2=\underset{i=1}{\overset{8}{\mathrm{\&Sigma;}}}{\mathrm{\&delta;}}_i^2=\underset{i=1}{\overset{8}{\mathrm{\&Sigma;}}}(s_1\left(t_i\right)-f_1\left(t_i\right))\\ =\min \underset{i=1}{\overset{8}{\mathrm{\&Sigma;}}}(s_1\left(t_i\right)-f_1\left(t_i\right))\end{array}---\left(12\right)$

In like manner suppose

$f_2\left(t\right)=b_1{e}^{-b_{3}t}+b_2{e}^{-{\mathrm{ba}}_{4}t}---\left(13\right)$

Thus, can estimate f ₁and f (t) ₂(t) correlation parameter.