CN106043739A - Airplane with intelligent monitoring function - Google Patents

Abstract

The invention discloses an airplane with an intelligent monitoring function. The airplane comprises an airplane body and an intelligent monitoring device. The intelligent monitoring device comprises a data collection and storage module, an airplane body crack input module, a fatigue test module, a crack expansion analysis module, a service life prediction module, a display module and an alarming module, wherein the crack expansion analysis module is used for carrying out crack expansion analysis on a random load spectrum, crack positions and dimensions of all practical cracks and fatigue crack expansion speed curves of various cracks and determining crack expansion service life circulation numbers corresponding to various cracks. The intelligent monitoring device is arranged on the airplane body, and thus the fatigue condition of the airplane body is monitored in real time.

Description

A kind of aircraft with intellectual monitoring function

Technical field

The present invention relates to Airplane detection field, be specifically related to a kind of aircraft with intellectual monitoring function.

Background technology

For preventing the fatigue of airframe from causing its structural failure, need to predict the fatigue life of materials for airframes.Phase In the technology of pass, aircraft does not arrange corresponding monitoring device, the tired situation of airframe can not be monitored at any time.

Summary of the invention

For solving the problems referred to above, it is desirable to provide a kind of aircraft with intellectual monitoring function.

The purpose of the present invention realizes by the following technical solutions:

A kind of aircraft with intellectual monitoring function, including airframe and intelligent monitoring device, described intellectual monitoring fills Put and include:

(1) data set storage module, is used for gathering aircraft real flight conditions, it is thus achieved that the random load spectrum in aircraft flight is also Store described random load spectrum；

(2) structure crack input module, for inputting the crack position of each actual crack, size on described airframe, And various crackles are carried out geometry simplification classification；

(3) fatigue test module, for the material of described airframe is carried out fatigue test, obtains described material corresponding Fatigue crack growth rate curve in various crackles；

(4) crackle expand analyze module: for described random load spectrum, the crack position of each actual crack, size with And the fatigue crack growth rate curve of various crackle carries out Crack growth analysis, determine the cracks can spread corresponding to various crackles Life Cycle number；

(5) biometry module: for determining the residual fatigue of corresponding crackle according to described crack propagation life period The estimated value in life-span and the estimated value of airframe remanent fatigue life.

The invention have the benefit that the present invention arranges intelligent monitoring device on airframe, it is achieved that airframe The real-time monitoring of tired situation, thus solve above-mentioned technical problem.

Accompanying drawing explanation

The invention will be further described to utilize accompanying drawing, but the application scenarios in accompanying drawing does not constitute any limit to the present invention System, for those of ordinary skill in the art, on the premise of not paying creative work, it is also possible to obtain according to the following drawings Other accompanying drawing.

Fig. 1 is the structural representation of intelligent monitoring device of the present invention；

Fig. 2 is the structural representation of the fatigue test module of the present invention.

Reference:

Data set storage module 1, structure crack input module 2, fatigue test module 3, crackle are expanded and are analyzed module 4, life-span Prediction module 5, display module 6, alarm module 7, parameter computation module 31, fatigue crack growth rate curve build submodule 32。

Detailed description of the invention

In conjunction with following application scenarios, the invention will be further described.

Application scenarios 1

Seeing Fig. 1, Fig. 2, the aircraft with intellectual monitoring function of an embodiment of this application scene, including aircraft machine Body and intelligent monitoring device, described intelligent monitoring device includes:

(1) data set storage module 1, is used for gathering aircraft real flight conditions, it is thus achieved that the random load spectrum in aircraft flight And store described random load spectrum；

(2) structure crack input module 2, for inputting the crackle position of each actual crack in the structure of described airframe Put, size, and various crackles are carried out geometry simplification classification；

(3) fatigue test module 3, for the material of described airframe is carried out fatigue test, obtains described material pair Should be in the fatigue crack growth rate curve of various crackles；

(4) crackle expand analyze module 4: for described random load spectrum, the crack position of each actual crack, size with And the fatigue crack growth rate curve of various crackle carries out Crack growth analysis, determine the cracks can spread corresponding to various crackles Life Cycle number；

(5) biometry module 5: for determining the residual fatigue of corresponding crackle according to described crack propagation life period The estimated value in life-span and the estimated value of airframe remanent fatigue life.

Preferably, described intelligent monitoring device also includes that display module 6, described display module 6 are used for showing fatigue crack Spreading rate curve and airframe remanent fatigue life.

The above embodiment of the present invention arranges intelligent monitoring device on airframe, it is achieved that the tired situation of airframe Real-time monitoring, thus solve above-mentioned technical problem.

Preferably, described intelligent monitoring device also includes alarm module 7, and described alarm module 7 is for remaining at airframe Report to the police when the estimated value of remaining fatigue life is more than the threshold value set.

Preferably, define corresponding to crackle i=1,2 ... the estimated value collection of the remanent fatigue life of m is { P₁,P₂,…, P_i, the estimated value P of airframe remanent fatigue life_ZIt is then:

P_Z=min_{I=1,2 ... m}{P₁,P₂,…,P_i}。

This preferred embodiment determines the fatigue of the remanent fatigue life of airframe and each actual crack of airframe Relation between life-span, the fatigue life of the actual crack of the airframe that employing is minimum is as the residual fatigue longevity of airframe Life, meets Law of Barrel, and accuracy is high.

Preferably, described fatigue test module 3 includes parameter computation module 31 and fatigue crack growth rate curve structure Build submodule 32, particularly as follows:

(1) parameter computation module 31: for calculating the stress intensive factor range of various crackle, it is considered to Crack Tip end points Plastically deforming area can have conclusive impact to the fatigue fracture of material, crack tip plastic zone is equivalent to one containing phase The homogenizing that allergic effect becomes is mingled with, and defines stress intensive factor range Δ K_pcComputing formula be:

${\mathrm{\ΔK}}_{pc}=\left\{\begin{array}{cc}{K}_{pc}^{\max }-K_{yc}-{\mathrm{\ΔK}}_{sc},& R\≤0\\ K_{pc}^{\max }-K_{pc}^{\min },& R>0\end{array}\right.$

In formula

$\begin{array}{c}{\mathrm{\ΔK}}_{sc}=\frac{1}{2\sqrt{2\mathrm{\π}}}{\∫}_A{r}^{-3/2}\[\frac{K_{yc}}{\sqrt{2\mathrm{\π}r}}(3{\sin }^2\mathrm{\α}\cos \mathrm{\α}+2\cos \frac{\mathrm{\α}}{2}\cos \frac{3\mathrm{\α}}{2})\\ +3({\mathrm{\σ}}_{11}-{\mathrm{\σ}}_{22})\sin \mathrm{\α}\sin \frac{5\mathrm{\α}}{2}-6{\mathrm{\σ}}_{12}\sin \mathrm{\α}\cos \frac{5\mathrm{\α}}{2}-({\mathrm{\σ}}_{11}+{\mathrm{\σ}}_{22})\cos \frac{3\mathrm{\α}}{2}\]dA\end{array}$

Wherein,For in fatigue and cyclic load by the calculated stress intensity factor through plastic correcting of maximum load Value,For in fatigue and cyclic load by the calculated stress intensity factor value through plastic correcting of minimum load, K_ycFor far Stress intensity factor under field action, crackle LOAD FOR when opening completely obtains, Δ K_scRepresent crack tip plastic zone The stress intensity factor increment caused, A is the area of the plastic zone around crack tip, and it includes being produced in crack propagation process Raw plastic deformation tail district, σ₁₁、σ₁₂、σ₂₂For the stress in crack tip plastic zone, by crack tip plastic zone stress field Finite element method (fem) analysis obtain, R is the ratio of tensile load and compressive load；

(2) fatigue crack growth rate curve builds submodule 32, for building the crack Propagation speed of various crackle Rate curve, based on Paris formula, it is considered to fatigue crack is expanded the impact of speed by temperature, defines described fatigue crack and expands The modified computing formulae of exhibition speed is:

T<0℃OR T>T_maxTime,

$\frac{da}{dN}=C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M$

0℃≤T≤T_maxTime,

$\frac{da}{dN}=C{\left({\mathrm{\&Delta;K}}_{pc}\right)}^M$

In formula, T is test temperature, T_maxFor the maximum temperature set, T_maxSpan be [35 DEG C, 40 DEG C], a is for splitting Stricture of vagina extension length, N is cycle-index, C and M is material constant, Δ K_TFor cracks can spread performance curved surface at the improper temperature of matching The improper temperature fracture threshold value that post analysis obtains, embodies the temperature impact on spreading rate, and Δ K_TSpan [0, Δ K need to be met_pc)。

The calculating of stress intensive factor range Δ K_pc defined in the parameter computation module 31 that this preferred embodiment is arranged Formula, and the plastically deforming area considering Crack Tip end points can have conclusive impact to the fatigue fracture of material, and will split Plastic zone, stricture of vagina tip is equivalent to a homogenizing containing phase transition strain and is mingled with, thus the stress intensive factor range Δ K_pc defined can Analyze the shadow of crack tip plastic zone counter stress intensity factor with carrying out quantification being work perfectly well as a rational mechanical parameter Ring；In the fatigue crack growth rate curve structure submodule 32 arranged based on Paris formula, it is contemplated that temperature is to fatigue The impact of crackle expansion speed, and define the modified computing formulae of fatigue crack growth rate, improve the precision of calculating, and Simple and practical.

Preferably, the computing formula of described crack propagation life period N is:

$N={\&Integral;}_{a_0}^{a_c}\frac{1}{C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M}$

This preferred embodiment determines the computing formula of crack propagation life period N, improves the speed of biometry.

The maximum temperature T of this application scene above-described embodiment_maxIt is set as 35 DEG C, the fatigue life prediction to airframe Precision relatively improve 15%.

Application scenarios 2

Seeing Fig. 1, Fig. 2, the aircraft with intellectual monitoring function of an embodiment of this application scene, including aircraft machine Body and intelligent monitoring device, described intelligent monitoring device includes:

(1) data set storage module 1, is used for gathering aircraft real flight conditions, it is thus achieved that the random load spectrum in aircraft flight And store described random load spectrum；；

(2) structure crack input module 2, for inputting the crackle position of each actual crack in the structure of described airframe Put, size, and various crackles are carried out geometry simplification classification；

(3) fatigue test module 3, for the material of described airframe is carried out fatigue test, obtains described material pair Should be in the fatigue crack growth rate curve of various crackles；

(4) crackle expand analyze module 4: for described random load spectrum, the crack position of each actual crack, size with And the fatigue crack growth rate curve of various crackle carries out Crack growth analysis, determine the cracks can spread corresponding to various crackles Life Cycle number；

(5) biometry module 5: for determining the residual fatigue of corresponding crackle according to described crack propagation life period The estimated value in life-span and the estimated value of airframe remanent fatigue life.

Preferably, described intelligent monitoring device also includes that display module 6, described display module 6 are used for showing fatigue crack Spreading rate curve and airframe remanent fatigue life.

The above embodiment of the present invention arranges intelligent monitoring device on airframe, it is achieved that the tired situation of airframe Real-time monitoring, thus solve above-mentioned technical problem.

Preferably, described intelligent monitoring device also includes alarm module 7, and described alarm module 7 is for remaining at airframe Report to the police when the estimated value of remaining fatigue life is more than the threshold value set.

Preferably, define corresponding to crackle i=1,2 ... the estimated value collection of the remanent fatigue life of m is { P₁,P₂,…, P_i, the estimated value P of airframe remanent fatigue life_zIt is then:

P_z=min_{I=1,2 ... m}{P₁,P₂,…,P_i}。

This preferred embodiment determines the fatigue of the remanent fatigue life of airframe and each actual crack of airframe Relation between life-span, the fatigue life of the actual crack of the airframe that employing is minimum is as the residual fatigue longevity of airframe Life, meets Law of Barrel, and accuracy is high.

Preferably, described fatigue test module 3 includes parameter computation module 31 and fatigue crack growth rate curve structure Build submodule 32, particularly as follows:

(1) parameter computation module 31: for calculating the stress intensive factor range of various crackle, it is considered to Crack Tip end points Plastically deforming area can have conclusive impact to the fatigue fracture of material, crack tip plastic zone is equivalent to one containing phase The homogenizing that allergic effect becomes is mingled with, and defines stress intensive factor range Δ K_pcComputing formula be:

${\mathrm{\&Delta;K}}_{pc}=\left\{\begin{array}{cc}{K}_{pc}^{\max }-K_{yc}-{\mathrm{\&Delta;K}}_{sc},& R\&le;0\\ K_{pc}^{\max }-K_{pc}^{\min },& R>0\end{array}\right.$

In formula

$\begin{array}{c}{\mathrm{\&Delta;K}}_{sc}=\frac{1}{2\sqrt{2\mathrm{\&pi;}}}{\&Integral;}_A{r}^{-3/2}\&lsqb;\frac{K_{yc}}{\sqrt{2\mathrm{\&pi;}r}}(3{\sin }^2\mathrm{\&alpha;}\cos \mathrm{\&alpha;}+2\cos \frac{\mathrm{\&alpha;}}{2}\cos \frac{3\mathrm{\&alpha;}}{2})\\ +3({\mathrm{\&sigma;}}_{11}-{\mathrm{\&sigma;}}_{22})\sin \mathrm{\&alpha;}\sin \frac{5\mathrm{\&alpha;}}{2}-6{\mathrm{\&sigma;}}_{12}\sin \mathrm{\&alpha;}\cos \frac{5\mathrm{\&alpha;}}{2}-({\mathrm{\&sigma;}}_{11}+{\mathrm{\&sigma;}}_{22})\cos \frac{3\mathrm{\&alpha;}}{2}\&rsqb;dA\end{array}$

Wherein,For in fatigue and cyclic load by the calculated stress intensity factor through plastic correcting of maximum load Value,For in fatigue and cyclic load by the calculated stress intensity factor value through plastic correcting of minimum load, K_ycFor far Stress intensity factor under field action, crackle LOAD FOR when opening completely obtains, Δ K_scRepresent crack tip plastic zone The stress intensity factor increment caused, A is the area of the plastic zone around crack tip, and it includes being produced in crack propagation process Raw plastic deformation tail district, σ₁₁、σ₁₂、σ₂₂For the stress in crack tip plastic zone, by crack tip plastic zone stress field Finite element method (fem) analysis obtain, R is the ratio of tensile load and compressive load；

(2) fatigue crack growth rate curve builds submodule 32, for building the crack Propagation speed of various crackle Rate curve, based on Paris formula, it is considered to fatigue crack is expanded the impact of speed by temperature, defines described fatigue crack and expands The modified computing formulae of exhibition speed is:

T<0℃OR T>T_maxTime,

$\frac{da}{dN}=C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M$

0℃≤T≤T_maxTime,

$\frac{da}{dN}=C{\left({\mathrm{\&Delta;K}}_{pc}\right)}^M$

In formula, T is test temperature, T_maxFor the maximum temperature set, T_maxSpan be [35 DEG C, 40 DEG C], a is for splitting Stricture of vagina extension length, N is cycle-index, C and M is material constant, Δ K_TFor cracks can spread performance curved surface at the improper temperature of matching The improper temperature fracture threshold value that post analysis obtains, embodies the temperature impact on spreading rate, and Δ K_TSpan [0, Δ K need to be met_pc)。

The calculating of stress intensive factor range Δ K_pc defined in the parameter computation module 31 that this preferred embodiment is arranged Formula, and the plastically deforming area considering Crack Tip end points can have conclusive impact to the fatigue fracture of material, and will split Plastic zone, stricture of vagina tip is equivalent to a homogenizing containing phase transition strain and is mingled with, thus the stress intensive factor range Δ K_pc defined can Analyze the shadow of crack tip plastic zone counter stress intensity factor with carrying out quantification being work perfectly well as a rational mechanical parameter Ring；In the fatigue crack growth rate curve structure submodule 32 arranged based on Paris formula, it is contemplated that temperature is to fatigue The impact of crackle expansion speed, and define the modified computing formulae of fatigue crack growth rate, improve the precision of calculating, and Simple and practical.

Preferably, the computing formula of described crack propagation life period N is:

$N={\&Integral;}_{a_0}^{a_c}\frac{1}{C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M}$

This preferred embodiment determines the computing formula of crack propagation life period N, improves the speed of biometry.

The maximum temperature T of this application scene above-described embodiment_maxIt is set as 36 DEG C, the fatigue life prediction to airframe Precision relatively improve 14%.

Application scenarios 3

Seeing Fig. 1, Fig. 2, the aircraft with intellectual monitoring function of an embodiment of this application scene, including aircraft machine Body and intelligent monitoring device, described intelligent monitoring device includes:

(1) data set storage module 1, is used for gathering aircraft real flight conditions, it is thus achieved that the random load spectrum in aircraft flight And store described random load spectrum；；

(2) structure crack input module 2, for inputting the crackle position of each actual crack in the structure of described airframe Put, size, and various crackles are carried out geometry simplification classification；

(3) fatigue test module 3, for the material of described airframe is carried out fatigue test, obtains described material pair Should be in the fatigue crack growth rate curve of various crackles；

(4) crackle expand analyze module 4: for described random load spectrum, the crack position of each actual crack, size with And the fatigue crack growth rate curve of various crackle carries out Crack growth analysis, determine the cracks can spread corresponding to various crackles Life Cycle number；

(5) biometry module 5: for determining the residual fatigue of corresponding crackle according to described crack propagation life period The estimated value in life-span and the estimated value of airframe remanent fatigue life.

Preferably, described intelligent monitoring device also includes that display module 6, described display module 6 are used for showing fatigue crack Spreading rate curve and airframe remanent fatigue life.

The above embodiment of the present invention arranges intelligent monitoring device on airframe, it is achieved that the tired situation of airframe Real-time monitoring, thus solve above-mentioned technical problem.

Preferably, described intelligent monitoring device also includes alarm module 7, and described alarm module 7 is for remaining at airframe Report to the police when the estimated value of remaining fatigue life is more than the threshold value set.

Preferably, define corresponding to crackle i=1,2 ... the estimated value collection of the remanent fatigue life of m is { P₁,P₂,…, P_i, the estimated value P of airframe remanent fatigue life_ZIt is then:

P_Z=min_{I=1,2 ... m}{P₁,P₂,…,P_i}。

This preferred embodiment determines the fatigue of the remanent fatigue life of airframe and each actual crack of airframe Relation between life-span, the fatigue life of the actual crack of the airframe that employing is minimum is as the residual fatigue longevity of airframe Life, meets Law of Barrel, and accuracy is high.

Preferably, described fatigue test module 3 includes parameter computation module 31 and fatigue crack growth rate curve structure Build submodule 32, particularly as follows:

(1) parameter computation module 31: for calculating the stress intensive factor range of various crackle, it is considered to Crack Tip end points Plastically deforming area can have conclusive impact to the fatigue fracture of material, crack tip plastic zone is equivalent to one containing phase The homogenizing that allergic effect becomes is mingled with, and defines stress intensive factor range Δ K_pcComputing formula be:

${\mathrm{\&Delta;K}}_{pc}=\left\{\begin{array}{cc}{K}_{pc}^{\max }-K_{yc}-{\mathrm{\&Delta;K}}_{sc},& R\&le;0\\ K_{pc}^{\max }-K_{pc}^{\min },& R>0\end{array}\right.$

In formula

$\begin{array}{c}{\mathrm{\&Delta;K}}_{sc}=\frac{1}{2\sqrt{2\mathrm{\&pi;}}}{\&Integral;}_A{r}^{-3/2}\&lsqb;\frac{K_{yc}}{\sqrt{2\mathrm{\&pi;}r}}(3{\sin }^2\mathrm{\&alpha;}\cos \mathrm{\&alpha;}+2\cos \frac{\mathrm{\&alpha;}}{2}\cos \frac{3\mathrm{\&alpha;}}{2})\\ +3({\mathrm{\&sigma;}}_{11}-{\mathrm{\&sigma;}}_{22})\sin \mathrm{\&alpha;}\sin \frac{5\mathrm{\&alpha;}}{2}-6{\mathrm{\&sigma;}}_{12}\sin \mathrm{\&alpha;}\cos \frac{5\mathrm{\&alpha;}}{2}-({\mathrm{\&sigma;}}_{11}+{\mathrm{\&sigma;}}_{22})\cos \frac{3\mathrm{\&alpha;}}{2}\&rsqb;dA\end{array}$

Wherein,For in fatigue and cyclic load by the calculated stress intensity factor through plastic correcting of maximum load Value,For in fatigue and cyclic load by the calculated stress intensity factor value through plastic correcting of minimum load, K_ycFor far Stress intensity factor under field action, crackle LOAD FOR when opening completely obtains, Δ K_scRepresent crack tip plastic zone The stress intensity factor increment caused, A is the area of the plastic zone around crack tip, and it includes being produced in crack propagation process Raw plastic deformation tail district, σ₁₁、σ₁₂、σ₂₂For the stress in crack tip plastic zone, by crack tip plastic zone stress field Finite element method (fem) analysis obtain, R is the ratio of tensile load and compressive load；

(2) fatigue crack growth rate curve builds submodule 32, for building the crack Propagation speed of various crackle Rate curve, based on Paris formula, it is considered to fatigue crack is expanded the impact of speed by temperature, defines described fatigue crack and expands The modified computing formulae of exhibition speed is:

T<0℃OR T>T_maxTime,

$\frac{da}{dN}=C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M$

0℃≤T≤T_maxTime,

$\frac{da}{dN}=C{\left({\mathrm{\&Delta;K}}_{pc}\right)}^M$

In formula, T is test temperature, T_maxFor the maximum temperature set, T_maxSpan be [35 DEG C, 40 DEG C], a is for splitting Stricture of vagina extension length, N is cycle-index, C and M is material constant, Δ K_TFor cracks can spread performance curved surface at the improper temperature of matching The improper temperature fracture threshold value that post analysis obtains, embodies the temperature impact on spreading rate, and Δ K_TSpan [0, Δ K need to be met_pc)。

The calculating of stress intensive factor range Δ K_pc defined in the parameter computation module 31 that this preferred embodiment is arranged Formula, and the plastically deforming area considering Crack Tip end points can have conclusive impact to the fatigue fracture of material, and will split Plastic zone, stricture of vagina tip is equivalent to a homogenizing containing phase transition strain and is mingled with, thus the stress intensive factor range Δ K_pc defined can Analyze the shadow of crack tip plastic zone counter stress intensity factor with carrying out quantification being work perfectly well as a rational mechanical parameter Ring；In the fatigue crack growth rate curve structure submodule 32 arranged based on Paris formula, it is contemplated that temperature is to fatigue The impact of crackle expansion speed, and define the modified computing formulae of fatigue crack growth rate, improve the precision of calculating, and Simple and practical.

Preferably, the computing formula of described crack propagation life period N is:

$N={\&Integral;}_{a_0}^{a_c}\frac{1}{C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M}$

This preferred embodiment determines the computing formula of crack propagation life period N, improves the speed of biometry.

The maximum temperature T of this application scene above-described embodiment_maxIt is set as 38 DEG C, the fatigue life prediction to airframe Precision relatively improve 12%.

Application scenarios 4

Seeing Fig. 1, Fig. 2, the aircraft with intellectual monitoring function of an embodiment of this application scene, including aircraft machine Body and intelligent monitoring device, described intelligent monitoring device includes:

(1) data set storage module 1, is used for gathering aircraft real flight conditions, it is thus achieved that the random load spectrum in aircraft flight And store described random load spectrum；；

(2) structure crack input module 2, for inputting the crackle position of each actual crack in the structure of described airframe Put, size, and various crackles are carried out geometry simplification classification；

(3) fatigue test module 3, for the material of described airframe is carried out fatigue test, obtains described material pair Should be in the fatigue crack growth rate curve of various crackles；

(4) crackle expand analyze module 4: for described random load spectrum, the crack position of each actual crack, size with And the fatigue crack growth rate curve of various crackle carries out Crack growth analysis, determine the cracks can spread corresponding to various crackles Life Cycle number；

(5) biometry module 5: for determining the residual fatigue of corresponding crackle according to described crack propagation life period The estimated value in life-span and the estimated value of airframe remanent fatigue life.

Preferably, described intelligent monitoring device also includes that display module 6, described display module 6 are used for showing fatigue crack Spreading rate curve and airframe remanent fatigue life.

The above embodiment of the present invention arranges intelligent monitoring device on airframe, it is achieved that the tired situation of airframe Real-time monitoring, thus solve above-mentioned technical problem.

Preferably, described intelligent monitoring device also includes alarm module 7, and described alarm module 7 is for remaining at airframe Report to the police when the estimated value of remaining fatigue life is more than the threshold value set.

Preferably, define corresponding to crackle i=1,2 ... the estimated value collection of the remanent fatigue life of m is { P₁,P₂,…, P_i, the estimated value P of airframe remanent fatigue life_zIt is then:

P_z=min_{I=1,2 ... m}{P₁,P₂,…,P_i}。

This preferred embodiment determines the fatigue of the remanent fatigue life of airframe and each actual crack of airframe Relation between life-span, the fatigue life of the actual crack of the airframe that employing is minimum is as the residual fatigue longevity of airframe Life, meets Law of Barrel, and accuracy is high.

Preferably, described fatigue test module 3 includes parameter computation module 31 and fatigue crack growth rate curve structure Build submodule 32, particularly as follows:

(1) parameter computation module 31: for calculating the stress intensive factor range of various crackle, it is considered to Crack Tip end points Plastically deforming area can have conclusive impact to the fatigue fracture of material, crack tip plastic zone is equivalent to one containing phase The homogenizing that allergic effect becomes is mingled with, and defines stress intensive factor range Δ K_pcComputing formula be:

${\mathrm{\&Delta;K}}_{pc}=\left\{\begin{array}{cc}{K}_{pc}^{\max }-K_{yc}-{\mathrm{\&Delta;K}}_{sc},& R\&le;0\\ K_{pc}^{\max }-K_{pc}^{\min },& R>0\end{array}\right.$

In formula

$\begin{array}{c}{\mathrm{\&Delta;K}}_{sc}=\frac{1}{2\sqrt{2\mathrm{\&pi;}}}{\&Integral;}_A{r}^{-3/2}\&lsqb;\frac{K_{yc}}{\sqrt{2\mathrm{\&pi;}r}}(3{\sin }^2\mathrm{\&alpha;}\cos \mathrm{\&alpha;}+2\cos \frac{\mathrm{\&alpha;}}{2}\cos \frac{3\mathrm{\&alpha;}}{2})\\ +3({\mathrm{\&sigma;}}_{11}-{\mathrm{\&sigma;}}_{22})\sin \mathrm{\&alpha;}\sin \frac{5\mathrm{\&alpha;}}{2}-6{\mathrm{\&sigma;}}_{12}\sin \mathrm{\&alpha;}\cos \frac{5\mathrm{\&alpha;}}{2}-({\mathrm{\&sigma;}}_{11}+{\mathrm{\&sigma;}}_{22})\cos \frac{3\mathrm{\&alpha;}}{2}\&rsqb;dA\end{array}$

Wherein,For in fatigue and cyclic load by the calculated stress intensity factor through plastic correcting of maximum load Value,For in fatigue and cyclic load by the calculated stress intensity factor value through plastic correcting of minimum load, K_ycFor far Stress intensity factor under field action, crackle LOAD FOR when opening completely obtains, Δ K_scRepresent crack tip plastic zone The stress intensity factor increment caused, A is the area of the plastic zone around crack tip, and it includes being produced in crack propagation process Raw plastic deformation tail district, σ₁₁、σ₁₂、σ₂₂For the stress in crack tip plastic zone, by crack tip plastic zone stress field Finite element method (fem) analysis obtain, R is the ratio of tensile load and compressive load；

(2) fatigue crack growth rate curve builds submodule 32, for building the crack Propagation speed of various crackle Rate curve, based on Paris formula, it is considered to fatigue crack is expanded the impact of speed by temperature, defines described fatigue crack and expands The modified computing formulae of exhibition speed is:

T<0℃OR T>T_maxTime,

$\frac{da}{dN}=C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M$

0℃≤T≤T_maxTime,

$\frac{da}{dN}=C{\left({\mathrm{\&Delta;K}}_{pc}\right)}^M$

In formula, T is test temperature, T_maxFor the maximum temperature set, T_maxSpan be [35 DEG C, 40 DEG C], a is for splitting Stricture of vagina extension length, N is cycle-index, C and M is material constant, Δ K_TFor cracks can spread performance curved surface at the improper temperature of matching The improper temperature fracture threshold value that post analysis obtains, embodies the temperature impact on spreading rate, and Δ K_TSpan [0, Δ K need to be met_pc)。

The calculating of stress intensive factor range Δ K_pc defined in the parameter computation module 31 that this preferred embodiment is arranged Formula, and the plastically deforming area considering Crack Tip end points can have conclusive impact to the fatigue fracture of material, and will split Plastic zone, stricture of vagina tip is equivalent to a homogenizing containing phase transition strain and is mingled with, thus the stress intensive factor range Δ K_pc defined can Analyze the shadow of crack tip plastic zone counter stress intensity factor with carrying out quantification being work perfectly well as a rational mechanical parameter Ring；In the fatigue crack growth rate curve structure submodule 32 arranged based on Paris formula, it is contemplated that temperature is to fatigue The impact of crackle expansion speed, and define the modified computing formulae of fatigue crack growth rate, improve the precision of calculating, and Simple and practical.

Preferably, the computing formula of described crack propagation life period N is:

$N={\&Integral;}_{a_0}^{a_c}\frac{1}{C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M}$

This preferred embodiment determines the computing formula of crack propagation life period N, improves the speed of biometry.

The maximum temperature T of this application scene above-described embodiment_maxIt is set as 39 DEG C, the fatigue life prediction to airframe Precision relatively improve 11%.

Application scenarios 5

Seeing Fig. 1, Fig. 2, the aircraft with intellectual monitoring function of an embodiment of this application scene, including aircraft machine Body and intelligent monitoring device, described intelligent monitoring device includes:

(1) data set storage module 1, is used for gathering aircraft real flight conditions, it is thus achieved that the random load spectrum in aircraft flight And store described random load spectrum；；

(2) structure crack input module 2, for inputting the crackle position of each actual crack in the structure of described airframe Put, size, and various crackles are carried out geometry simplification classification；

(3) fatigue test module 3, for the material of described airframe is carried out fatigue test, obtains described material pair Should be in the fatigue crack growth rate curve of various crackles；

(4) crackle expand analyze module 4: for described random load spectrum, the crack position of each actual crack, size with And the fatigue crack growth rate curve of various crackle carries out Crack growth analysis, determine the cracks can spread corresponding to various crackles Life Cycle number；

(5) biometry module 5: for determining the residual fatigue of corresponding crackle according to described crack propagation life period The estimated value in life-span and the estimated value of airframe remanent fatigue life.

Preferably, described intelligent monitoring device also includes that display module 6, described display module 6 are used for showing fatigue crack Spreading rate curve and airframe remanent fatigue life.

The above embodiment of the present invention arranges intelligent monitoring device on airframe, it is achieved that the tired situation of airframe Real-time monitoring, thus solve above-mentioned technical problem.

Preferably, described intelligent monitoring device also includes alarm module 7, and described alarm module 7 is for remaining at airframe Report to the police when the estimated value of remaining fatigue life is more than the threshold value set.

Preferably, define corresponding to crackle i=1,2 ... the estimated value collection of the remanent fatigue life of m is { P₁,P₂,…, P_i, the estimated value P of airframe remanent fatigue life_ZIt is then:

P_Z=min_{I=1,2 ... m}{P₁,P₂,…,P_i}。

This preferred embodiment determines the fatigue of the remanent fatigue life of airframe and each actual crack of airframe Relation between life-span, the fatigue life of the actual crack of the airframe that employing is minimum is as the residual fatigue longevity of airframe Life, meets Law of Barrel, and accuracy is high.

Preferably, described fatigue test module 3 includes parameter computation module 31 and fatigue crack growth rate curve structure Build submodule 32, particularly as follows:

(1) parameter computation module 31: for calculating the stress intensive factor range of various crackle, it is considered to Crack Tip end points Plastically deforming area can have conclusive impact to the fatigue fracture of material, crack tip plastic zone is equivalent to one containing phase The homogenizing that allergic effect becomes is mingled with, and defines stress intensive factor range Δ K_pcComputing formula be:

${\mathrm{\&Delta;K}}_{pc}=\left\{\begin{array}{cc}{K}_{pc}^{\max }-K_{yc}-{\mathrm{\&Delta;K}}_{sc},& R\&le;0\\ K_{pc}^{\max }-K_{pc}^{\min },& R>0\end{array}\right.$

In formula

$\begin{array}{c}{\mathrm{\&Delta;K}}_{sc}=\frac{1}{2\sqrt{2\mathrm{\&pi;}}}{\&Integral;}_A{r}^{-3/2}\&lsqb;\frac{K_{yc}}{\sqrt{2\mathrm{\&pi;}r}}(3{\sin }^2\mathrm{\&alpha;}\cos \mathrm{\&alpha;}+2\cos \frac{\mathrm{\&alpha;}}{2}\cos \frac{3\mathrm{\&alpha;}}{2})\\ +3({\mathrm{\&sigma;}}_{11}-{\mathrm{\&sigma;}}_{22})\sin \mathrm{\&alpha;}\sin \frac{5\mathrm{\&alpha;}}{2}-6{\mathrm{\&sigma;}}_{12}\sin \mathrm{\&alpha;}\cos \frac{5\mathrm{\&alpha;}}{2}-({\mathrm{\&sigma;}}_{11}+{\mathrm{\&sigma;}}_{22})\cos \frac{3\mathrm{\&alpha;}}{2}\&rsqb;dA\end{array}$

Wherein,For in fatigue and cyclic load by the calculated stress intensity factor through plastic correcting of maximum load Value,For in fatigue and cyclic load by the calculated stress intensity factor value through plastic correcting of minimum load, K_ycFor far Stress intensity factor under field action, crackle LOAD FOR when opening completely obtains, Δ K_scRepresent crack tip plastic zone The stress intensity factor increment caused, A is the area of the plastic zone around crack tip, and it includes being produced in crack propagation process Raw plastic deformation tail district, σ₁₁、σ₁₂、σ₂₂For the stress in crack tip plastic zone, by crack tip plastic zone stress field Finite element method (fem) analysis obtain, R is the ratio of tensile load and compressive load；

(2) fatigue crack growth rate curve builds submodule 32, for building the crack Propagation speed of various crackle Rate curve, based on Paris formula, it is considered to fatigue crack is expanded the impact of speed by temperature, defines described fatigue crack and expands The modified computing formulae of exhibition speed is:

T<0℃OR T>T_maxTime,

$\frac{da}{dN}=C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M$

0℃≤T≤T_maxTime,

$\frac{da}{dN}=C{\left({\mathrm{\&Delta;K}}_{pc}\right)}^M$

In formula, T is test temperature, T_maxFor the maximum temperature set, T_maxSpan be [35 DEG C, 40 DEG C], a is for splitting Stricture of vagina extension length, N is cycle-index, C and M is material constant, Δ K_TFor cracks can spread performance curved surface at the improper temperature of matching The improper temperature fracture threshold value that post analysis obtains, embodies the temperature impact on spreading rate, and Δ K_TSpan [0, Δ K need to be met_pc)。

The calculating of stress intensive factor range Δ K_pc defined in the parameter computation module 31 that this preferred embodiment is arranged Formula, and the plastically deforming area considering Crack Tip end points can have conclusive impact to the fatigue fracture of material, and will split Plastic zone, stricture of vagina tip is equivalent to a homogenizing containing phase transition strain and is mingled with, thus the stress intensive factor range Δ K_pc defined can Analyze the shadow of crack tip plastic zone counter stress intensity factor with carrying out quantification being work perfectly well as a rational mechanical parameter Ring；In the fatigue crack growth rate curve structure submodule 32 arranged based on Paris formula, it is contemplated that temperature is to fatigue The impact of crackle expansion speed, and define the modified computing formulae of fatigue crack growth rate, improve the precision of calculating, and Simple and practical.

Preferably, the computing formula of described crack propagation life period N is:

$N={\&Integral;}_{a_0}^{a_c}\frac{1}{C{({\mathrm{\&Delta;K}}_{pc}-{\mathrm{\&Delta;K}}_T)}^M}$

This preferred embodiment determines the computing formula of crack propagation life period N, improves the speed of biometry.

The maximum temperature T of this application scene above-described embodiment_maxIt is set as 40 DEG C, the fatigue life prediction to airframe Precision relatively improve 10%.

Last it should be noted that, use above scene is only in order to illustrate technical scheme, rather than to the present invention The restriction of protection domain, although having made to explain to the present invention with reference to preferred application scene, the ordinary skill people of this area Member should be appreciated that and can modify technical scheme or equivalent, without deviating from technical solution of the present invention Spirit and scope.

Claims (3)

1. there is an aircraft for intellectual monitoring function, it is characterized in that, including airframe and the intelligence being arranged in airframe Energy monitoring device, described intelligent monitoring device includes:

(1) data set storage module, is used for gathering aircraft real flight conditions, it is thus achieved that random load spectrum in aircraft flight also stores Described random load spectrum；

(2) structure crack input module is for inputting the crack position of each actual crack, size on described airframe and right Various crackles carry out geometry simplification classification；

(3) fatigue test module, for the material of described airframe is carried out fatigue test, obtains described material corresponding to each Plant the fatigue crack growth rate curve of crackle；

(4) crackle is expanded and is analyzed module: for described random load spectrum, the crack position of each actual crack, size and each The fatigue crack growth rate curve planting crackle carries out Crack growth analysis, determines the crack propagation life corresponding to various crackles Period；

(5) biometry module: for determining the remanent fatigue life of corresponding crackle according to described crack propagation life period Estimated value and the estimated value of airframe remanent fatigue life.

A kind of aircraft with intellectual monitoring function the most according to claim 1, is characterized in that, described intelligent monitoring device Also including alarm module, described alarm module is for when the estimated value of airframe remanent fatigue life is more than the threshold value set Report to the police.

A kind of aircraft with intellectual monitoring function the most according to claim 2, is characterized in that, described intelligent monitoring device Also include that display module, described display module are used for showing fatigue crack growth rate curve and airframe residual fatigue longevity Life.