CN101000294A

CN101000294A - Investigating method for impact loading spectrum of aircraft laminated structure and its investigating device

Info

Publication number: CN101000294A
Application number: CNA2007100193714A
Authority: CN
Inventors: 周晚林
Original assignee: Nanjing University of Aeronautics and Astronautics
Current assignee: Nanjing University of Aeronautics and Astronautics
Priority date: 2007-01-18
Filing date: 2007-01-18
Publication date: 2007-07-18

Abstract

The invention discloses aircraft laminate structure striking -load spectrum monitoring method. It includes the following steps: cling or pre-embedding many piezoelectricity sensing elements on the surface of the laminate structure as diverging piezoelectricity board unit; while suffering random striking, data collecting processing integration board collects the dynamic response of the sensor; calculating out structure modality response coordinate; using unconditional stability fine step-by-step integration method to find out modality dynamics differential equation; iteration calculating out striking load time course by modality coordinate. The monitoring device of the invention includes computer, diverging piezoelectricity sensor and is connected with charge amplifier which is connected with computer by data collecting machine. The computer main board expanding groove is inserted data collecting processing integrating board.

Description

Monitoring method and monitoring device for impact load spectrum of aircraft laminated structure

Technical Field

The invention relates to the technical field of aircraft structure damage monitoring, in particular to a method for monitoring the time history of impact load on the surface of an aircraft laminated structure. The invention also relates to a monitoring device for implementing the method.

Background

With the application of laminated structures to engineering components such as aircraft airfoils, the problem of delamination damage caused by uncertain impact energy has attracted much attention to ensure safe operation of the structure, because delamination damage occurs in the interior of the material and cannot be found from the surface, with the potential for catastrophic and catastrophic failure. The impact damage mechanism of the composite material laminated structure is quite complex, and the research on the aspect is still in the development stage at present. Experimental research shows that the dynamic response of the composite material laminated structure under impact has obvious correlation with the damage of the structure, and the delamination damage area has certain correlation with the impact energy. Therefore, identifying the time history of uncertain impact load in the service process of the laminated structure in real time and calculating the impact energy is one of effective ways for evaluating the damage type and degree of the structure. The key of the problem is how to monitor the impact load spectrum of the laminated structure subjected to random impact in the service process in real time, and the adopted device does not influence the function of the laminated structure and the dynamic performance of the laminated structure.

The monitoring of the impact load spectrum actually belongs to the problem of identifying the dynamic load of the structure. The identification of the dynamic load is based on the dynamic characteristics of the known system and the dynamic load of the measured dynamic response inversion structure, and is a relatively difficult inverse problem of dynamics. The general technology is that an acceleration sensor is arranged on a structure, and a multipoint acceleration vibration pickup based method is adopted to invert the dynamic load of the structure. However, this method is not suitable for the laminated structure on the aircraft, because it requires more measuring sensors, has a narrow frequency response range of the acceleration sensor and has certain mass and volume, and has a large influence on the dynamic characteristics of the structure.

Disclosure of Invention

The invention aims to solve the technical problem of providing a method for monitoring the impact load spectrum of an aircraft laminated structure in real time without influencing the functions of the laminated structure and the dynamic performance of the laminated structure. In addition, another technical problem to be solved by the present invention is to provide a device for implementing the monitoring method.

In order to solve the technical problem, the monitoring method of the impact load spectrum of the aircraft laminated structure comprises the following steps of firstly, sticking or embedding a plurality of piezoelectric sensitive elements serving as discrete piezoelectric plate units on the surface of the laminated structure to form a piezoelectric laminated intelligent structure with the function of a piezoelectric modal sensor; when the structure is subjected to random impact, acquiring dynamic response of each piezoelectric modal sensor caused by the random impact through a data acquisition and processing integrated board; and finally, iteratively solving the time history (impact load spectrum) of the impact load by the modal coordinate of the structure.

The monitoring device for realizing the monitoring method of the impact load spectrum of the aircraft laminated structure comprises a computer and is characterized in that: the discrete piezoelectric sensor adhered to or embedded in the surface of the laminated structure is electrically connected with a charge amplifier, and the charge amplifier is connected with a computer through a data collector; and a data acquisition and processing integrated board is inserted in the computer mainboard expansion slot.

The method is based on the piezoelectric structure modal response inversion impact load spectrum, the adopted piezoelectric structure modal response is obtained by indirectly calculating the piezoelectric response of the piezoelectric element according to the vibration mode superposition and piezoelectric sensing principles, and a piezoelectric modal sensor is not additionally arranged physically, so that the mode identification theory is utilized, no additional modal sensor is arranged, and the function and the dynamic characteristic of the laminated structure are not influenced. Meanwhile, after the modal response of the structure is obtained through calculation, a step-by-step integral method which is unconditionally stable is adopted for inverting the structure impact load spectrum, and the method is simple in algorithm, high in calculation speed, free of matrix ill-condition and the like. And because the method adopts the analysis method of the structural finite element piezoelectric mode, the method is widely suitable for complex laminated structures with any irregular shapes and boundary conditions.

In the monitoring device, the piezoelectric sensitive element is very flexible and thin, has small volume and light weight, and has very little influence on the function and the dynamic performance of the laminated structure, and the piezoelectric sensitive element has the advantages of low cost, high sensitivity, wide frequency response, large dynamic range and the like.

Drawings

The present invention will be described in further detail with reference to the following drawings and specific examples.

FIG. 1 is a schematic view of a monitoring device for implementing a method for monitoring impact load spectra of an aircraft laminate structure;

fig. 2 is a schematic view of an embodiment of the impact load spectrum monitoring method of the aircraft laminate structure of the present invention.

Detailed Description

FIG. 1 is a schematic view of a monitoring device for implementing a method for monitoring an impact load spectrum of an aircraft laminate structure. In FIG. 1, 1 is an aircraft laminated structure, 2 is discrete piezoelectric sensors (P1-P6), 3 is a computer, 4 is a data collector, and 5 is a charge amplifier; k is the assumed impact load application position. The discrete piezoelectric sensor 2 adhered to or embedded in the surface of the aircraft laminated structure 1 is electrically connected with a charge amplifier 5, and the charge amplifier 5 is connected with a computer 3 through a data collector 4; the discrete piezoelectric sensor 2 comprises six piezoelectric sensing units (generally adopting PVDF films) of P1, P2, P3, P4, P5 and P6, wherein the six piezoelectric sensing units are adhered to the lower surface of the aircraft laminated structure, and the electrode surfaces of the piezoelectric sensing units are mutually insulated, so that when impact load acts on the aircraft laminated structure 1, piezoelectric responses received by the six piezoelectric sensing units are simultaneously transmitted to the data acquisition and processing integrator 4 through the charge amplifier 5, and the piezoelectric modal response of the structure under impact is obtained after data processing. Therefore, the six piezoelectric sensing chips are actually discrete piezoelectric sensors having a piezoelectric mode sensor function.

Fig. 2 is a schematic diagram of an embodiment of the invention relating to monitoring of the impact load spectrum of an aircraft laminate structure. In FIG. 2, the aircraft laminated substrate 6 is 0.6m long, 0.6m wide and 0.8mm thick, and is composed of four layers of graphite epoxy composite materials, and four sides of the substrate are clamped. A PVDF piezoelectric film 7 is adhered to the lower surface of the aircraft laminated substrate 6, and the thickness is 0.1 mm. The whole PVDF piezoelectric film 7 is divided into 9 piezoelectric units, and each unit is a piezoelectric sensing unit slice. The impact test of the plate is carried out on the double-guide-rail drop hammer type experiment table, and in order to facilitate verification, the impact waveform is a simple half sine wave.

The purpose of impact load spectrum identification is to show the impact load randomly acting on the surface of the intelligent piezoelectric laminated plate structure by using a load-time relation curve. The method adopted by the invention is as follows: according to the modal sensing principle of the piezoelectric intelligent structure, the piezoelectric intelligent structure and the PVDF piezoelectric film are divided into finite elements in an up-down corresponding mode, namely the laminated plate in the figure 2 is also divided into 9 units.

After the piezoelectric intelligent plate is divided by finite elements, the e-type linear positive piezoelectric equation can be expressed as

D＝eε+βE (1)

Where D is an electric displacement array, E is a piezoelectric constant matrix, ε is a strain array, β is a dielectric constant matrix, and E is an electric field strength array. Strain array to arbitrary point in piezoelectric laminate

ε＝{ε_xε_yγ_xy}^TCan be expressed as

ε＝zψ；ψ＝Bδ^e (2)

Wherein z is the coordinate of the point in the direction of the deflection w,

<math> <mrow> <mi>ψ</mi> <mo>=</mo> <mo>-</mo> <msup> <mfenced open='{' close='}'> <mtable> <mtr> <mtd> <mfrac> <mrow> <msup> <mo>&PartialD;</mo> <mn>2</mn> </msup> <mi>w</mi> </mrow> <msup> <mrow> <mo>&PartialD;</mo> <mi>x</mi> </mrow> <mn>2</mn> </msup> </mfrac> </mtd> <mtd> <mfrac> <mrow> <msup> <mo>&PartialD;</mo> <mn>2</mn> </msup> <mi>w</mi> </mrow> <msup> <mrow> <mo>&PartialD;</mo> <mi>y</mi> </mrow> <mn>2</mn> </msup> </mfrac> </mtd> <mtd> <mn>2</mn> <mfrac> <mrow> <msup> <mo>&PartialD;</mo> <mn>2</mn> </msup> <mi>w</mi> </mrow> <mrow> <mo>&PartialD;</mo> <mi>x</mi> <mo>&PartialD;</mo> <mi>y</mi> </mrow> </mfrac> </mtd> </mtr> </mtable> </mfenced> <mi>T</mi> </msup> </mrow> </math>

for curvature and deflection vectors, B is the element deformation matrix, δ^eIs a unit node displacement array. When the piezoelectric laminate is used as a sensor, the electric field strength E is 0, and the electric displacement of the piezoelectric element surface in the polarization direction z is obtained from the formula (1)

D_z＝e₃₁ε_x+e₃₂ε_y+e₃₆γ_xy (3)

M piezoelectric pieces are adhered on the surface of the structure, and the electric quantity on the jth piezoelectric piece is obtained from the formula

<math> <mrow> <msub> <mi>q</mi> <mi>j</mi> </msub> <mo>=</mo> <munderover> <mi>Σ</mi> <mi>j</mi> <mi>e</mi> </munderover> <mover> <mi>z</mi> <mo>&OverBar;</mo> </mover> <munder> <mo>&Integral;</mo> <msup> <mi>A</mi> <mi>e</mi> </msup> </munder> <msub> <mi>e</mi> <mi>z</mi> </msub> <mi>B</mi> <msup> <mi>δ</mi> <mi>e</mi> </msup> <msup> <mi>dA</mi> <mi>e</mi> </msup> <mo></mo> <mrow> <mo>(</mo> <mi>j</mi> <mo>=</mo> <mn>1</mn> <mo>,</mo> <mo>.</mo> <mo>.</mo> <mo>.</mo> <mo>,</mo> <mi>m</mi> <mo>)</mo> </mrow> <mo>-</mo> <mo>-</mo> <mo>-</mo> <mrow> <mo>(</mo> <mn>4</mn> <mo>)</mo> </mrow> </mrow> </math>

Wherein

Indicating summing the cells covered by the jth piezoelectric patch,

is the average z coordinate of the piezoelectric patch.

Let phi_eiIs the e element, x, of the ith order mode in the structural mode matrix phi_iIs the ith order modal coordinate, and shifts the node by delta^eExpanding according to the modal coordinate x and taking m times of modal truncation, then having

Substituting the above formula into formula (4), and writing into matrix form to obtain piezoelectric modal sensing equation

Q＝AX (6)

Wherein Q ═ { Q ═ Q₁，q₂，…，q_m}^TIs the output charge column vector of the piezoelectric sheet, X ═ X₁，x₂，…x_m}^TIn the form of a column vector of modal coordinates,

<math> <mrow> <msub> <mi>A</mi> <mi>ji</mi> </msub> <mo>=</mo> <munderover> <mi>Σ</mi> <mi>j</mi> <mi>e</mi> </munderover> <munder> <mo>&Integral;</mo> <msup> <mi>A</mi> <mi>e</mi> </msup> </munder> <msub> <mi>e</mi> <mi>z</mi> </msub> <mi>B</mi> <msub> <mi>φ</mi> <mi>ei</mi> </msub> <msup> <mi>dA</mi> <mi>e</mi> </msup> <mo>,</mo> <mrow> <mo>(</mo> <mi>i</mi> <mo>,</mo> <mi>j</mi> <mo>=</mo> <mn>1</mn> <mo>,</mo> <mo>.</mo> <mo>.</mo> <mo>.</mo> <mo>,</mo> <mi>m</mi> <mo>)</mo> </mrow> <mo>.</mo> </mrow> </math>

according to the above formula, the first m-order modal response of the structure can be measured by only laying m piezoelectric sheets, where m is 9 in this embodiment. According to the m-order modal response, the following modal dynamics differential equation with n degrees of freedom can be solved:

where M is the mass matrix, K is the stiffness matrix, and C is the damping matrix. Delta, delta,

Respectively, node displacement, velocity and acceleration arrays of the structure.

If the first m-order characteristic pair (omega) of the system is obtained by actual measurement_m ²，{φ_mAnd then the regularized mode matrix phi is set as [ phi ]₁φ₂…φ_m]As a transformation matrix, node displacement can be expressed as

δ＝φX (8)

Substituting the above formula into equation (7) to obtain uncoupled modal differential equation set

In the formula,

E＝diag[ξ₁ξ₂…ξ_m]，Ω²＝diag[ω₁ ²ω₂ ²…ω_m ²]，ω_r、ξ_rthe natural frequency and damping ratio of the r-th order mode, respectively, and the characteristic load array

{p(t)}＝φ^T{F(t)}

(10)

Writing formula (9) into decomposed form

(r＝1，…，m) (11)

If the modal coordinate x is obtained from the formula (9)_r(t), then the characteristic load p can be obtained by solving the differential equation (11)_r(t) thereby obtaining a load array by inverting the equation (10)

{F(t)}＝[φ^T]^-1{p(t)}

(12)

In order to obtain an accurate solution to the modal differential equation of equation (11), the present invention employs the following unconditionally stable fine stepwise integral solution: equation (11) is first reduced to a first order differential equation

{\dot{v} (t)} = [H] {v (t)} + {r (t)} - - - (13)

Wherein

r (t) = \{\begin{matrix} 0 \\ p_{r} (t) \end{matrix}\} - - - (15)

At the integration step t ∈ τ' (t)_j，t_j+1) The general solution of inner equation (13) is

{v(t)}＝[T(τ)]({v(t_j)}-{v_p(t_j)})+{v_p(t)} (16)

Wherein,

τ＝t-t_j，{v_p(t) } is a special solution of the equation.

Subdividing the integration step τ' into s 2^NEqual parts (20 for general structure N), Δ t τ'/s 2^-NTau' during a time period of deltat, having

[T(τ)]＝[e^[H]Δt]^s

≈[I+HΔt+(HΔt)²/2！+(HΔt)³/3！+(HΔt)⁴/4！]^s

≡[I+T_a0]^s

(17)

Form iterative

[T_ai]＝2[T_ai-1]+[T_ai-1][T_ai-1]

(i＝1，2，…，N) (18)

Then there is

At integration step length (t)_j，t_j+1) The internal load changes linearly; { k } is a function of₀}、{k₁Is a time invariant vector, then

For { r (t) } ═ k₀}+{k₁}×(t-t_j) (21)

The particular solution of equation (13) is

{v_p(t)}＝([H]^-1+[I]×t)(-[H]^-1{k₁})-[H]^-1({k₀}-{k₁}×t_j)

＝-[H]^-1({k₀}+[H]^-1{k₁})+[H]^-1{k₁}(t_j-t) (22)

The formula (16) is substituted and arranged to obtain

{v(t_j+1)}＝[T(τ)]({v(t_j)}+[H]^-1({k₀}+[H]^-1{k₁}))-

[H]^-1({k₀}+[H]^-1{k₁}+{k₁}×τ)

＝[T(τ)]{v(t_j)}+[C(τ)]{k₀}+[D(τ)]{k₁} (23)

Wherein [ C (tau)]＝[T(τ)][H]^-1-[H]^-1

[D(τ)]＝[T(τ)IH]^-1[H]^-1-[H]^-1[H]^-1-[H]^-1τ (24)

Can be written by the following formulae (14, 15) and (23)

Which is decomposed into

According to the above formula, i.e. from the modal coordinate x_r(t) iteratively calculating the impact load spectrum p_r(t)。

Compared with the applied semi-sinusoidal waveform, the result calculated by the embodiment of the method is better in waveform matching, which shows that the method provided by the invention is higher in accuracy when used for monitoring the impact load spectrum.

Claims

1. The monitoring method of the impact load spectrum of the aircraft laminated structure is characterized by comprising the following steps: firstly, sticking or embedding a plurality of piezoelectric sensitive elements serving as discrete piezoelectric plate units on the surface of a laminated structure to form a piezoelectric laminated intelligent structure with the function of a piezoelectric modal sensor; when the structure is subjected to random impact, acquiring dynamic response of each piezoelectric modal sensor caused by the random impact through a data acquisition and processing integrated board; and finally, iteratively solving the time history of the impact load according to the modal coordinate of the structure.

2. A monitoring device for implementing the method for monitoring an impact load spectrum of an aircraft laminate structure according to claim 1, comprising a computer, wherein: the discrete piezoelectric sensor adhered to or embedded in the surface of the laminated structure is electrically connected with a charge amplifier, and the charge amplifier is connected with a computer through a data collector; and a data acquisition and processing integrated board is inserted in the computer mainboard expansion slot.