CN103853890A - Aeroelastic tailoring method of hypersonic flight vehicle - Google Patents

Abstract

The invention relates to an aeroelastic tailoring method of a hypersonic flight vehicle, in particular to a method for improving the aeroelasticity property of the hypersonic flight vehicle under a thermal environment. The aeroelastic tailoring method comprises the steps of selecting a design variable-a skin ply angle needing to be optimized, a constraint condition of keeping the structure weight of a wing surface unchanged and an optimization target of maximum airplane flutter critical velocity; obtaining a critical flutter velocity value of the wing surface under the thermal environment by carrying out aeroelasticity property analysis of the wing surface under the thermal environment, judging whether the critical flutter velocity value meets the optimization target or not, exiting a loop if so, updating the value of the design variable-the skin ply angle through an optimization algorithm set in an FD ISIGHT optimization assembly if not, and then carrying out the aeroelasticity property analysis of the wing surface under the thermal environment again. According to the aeroelastic tailoring method of the hypersonic flight vehicle under an aerodynamic thermal environment, disclosed by the invention, the flutter velocity value of the wing surface can be increased through a tailoring design of a composite skin, and thus the aeroelastic performance of the hypersonic flight vehicle can be improved.

Description

A kind of hypersonic aircraft aeroelastic tailoring method

Technical field

The present invention relates to a kind of aeroelastic tailoring method of aerospace field, specifically, is a kind of vehicle aeroelastic behavior improvement method under hypersonic aircraft thermal environment.

Background technology

Compound substance aeroelastic tailoring designs this Optimization Design by thereby the design and cut-out of composite laminated plate is controlled to its stiffness characteristics and dynamic perfromance, finally obtains superior aeroelasticity performance.In concrete airplane design engineering practice, the aeroelasticity performance that conventionally can improve has aeroelastic stability, control stability, lift-drag ratio and maneuver load etc.Its optimum ideals has obtained huge success in application in practice, for example X-29 swept forward wing plane of the U.S. in 1984 and the typical example of Muscovite S-37.

Aeroelastic tailoring design is the multidisciplinary problem of a many-sided correlation theory of design, and its main theory relating to comprises Compound Material Engineering, aeroelastic analysis method and optimized algorithm theory.In essence, compound substance aeroelasticity optimal design still belongs to optimum structure design method, by the reasonable laying design to composite laminated plate, acquisition conforms with the rigidity of structure of configuration, on the basis of coupling effect that utilizes material itself, arrange and meet actual structure and the process constraint condition of design, under aerodynamic effect aircraft produce favourable elastic deformation finally in keeping vehicle mass constant to improve aeroelastic characteristic.The optimization of hypersonic aircraft compound substance aeroelastic tailoring has salient feature at structure and material, how to utilize the fundamental characteristics that meets material, consider under hypersonic condition, aerodynamic force, aerothermal combined action, inquiring into out a kind of optimization cut-out method that improves the comprehensive aeroelasticity performance of hypersonic aircraft is a very challenging multidisciplinary problem.

Summary of the invention

The present invention is directed to above-mentioned prior art situation and a kind of hypersonic aircraft aeroelastic tailoring method that provides is provided, its objective is and solve under Aerodynamic Heating environment, improve aerofoil flutter speed value by the design of cutting out to composite material skin, thus hypersonic aircraft aeroelasticity performance.

The technical solution adopted in the present invention comprises the following steps:

(1) selecting to need the design variable of optimizing is covering laying angle [alpha], and constraint condition is for keeping airfoil structure weight constant, and optimization aim is airplane flutter critical velocity maximum, and the mathematical model of this optimization problem can be expressed as:

max?FlutterVelocity

ΔMass=0

s.t.x∈S

In formula, FluterVelocity represents critical flutter Mach number, and Mass represents full machine model quality, and x represents design variable, and S represents design variable set, and according to the restriction of actual manufacturing process, value is series of discrete numerical value :-45 °, and 0 °, 45 °, 90 °;

(2) set up hypersonic aircraft aerofoil geometric model and finite element model, comprise following content:

(a) in 3D sculpting software, set up hypersonic aircraft flying tail geometric model, adopted double-wedge thin airfoil, be connected by rotating shaft with fuselage;

(b) after geometric model is read in to MSC.Patran, set up finite element model and model is carried out to finite element grid division, horizontal tail surface stressed-skin construction part is selected shell unit modeling, and adopt the quadrilateral units in MSC.Patran automatically to divide, the inner girder construction that adopts of horizontal tail, and give beam element attribute, horizontal tail leading edge portion adopts solid element and uses the tetrahedron element in MSC.Patran automatically to divide;

(c) selecting at material, the covering of flying tail part is selected carbon fibre composite laminate (T300/5222), lay if covering is eight layers of symmetrical equilibrium, select four laying angles of each symmetrical individual layer as design variable, be set as respectively [angle ₁, angle ₂, angle ₃, angle ₄], initial laying order is [0/0/0/0] _s;

(3) carry out the vehicle aeroelastic behavior analysis of thermal environment lower aerofoil, it is characterized in that, comprise the following steps:

(a) the heat flow density field distribution that computation model surface is subject to:

The Aerodynamic Heating phenomenon that leading edge of a wing part is subject to is the most obvious, and the leading edge portion of the slim Wing-Body Configurations of hypersonic aircraft and airfoil portion can be ignored thickness factor, regards flat board as and calculates:

With reference to coefficient of viscosity μ ^*solve and draw by the blue expression formula in Saudi:

${\mathrm{\μ}}^{*}={\left(\frac{T^{*}}{288.15}\right)}^{1.5}\frac{398.55}{T^{*}+110.4}\×{1.7894\×10}^{-5}$

Reference density ρ ^*solve and draw by state equation expression formula:

${\mathrm{\ρ}}^{*}=\frac{p_{\∞}}{{\mathrm{RT}}^{*}}$

Can obtain with reference to Reynolds number simultaneously, as follows:

${\mathrm{Re}}_x^{*}=\frac{{\mathrm{\ρ}}^{*}{V}_{\∞}x}{{\mathrm{\μ}}^{*}}$

Margoulis number is solved and is drawn by Reynolds analogy formula:

${\mathrm{St}}^{*}=\frac{c_f^{*}}{2}\frac{1}{{\left(\mathrm{Pr}\right)}^{2/3}}$

Final hypersonic aircraft surface heat flow is tried to achieve by following formula:

Q _aero=St ^*ρ ^*V _∞c _ρ(T _r-T _ω)

Along x direction, horizontal tail is divided into some equal portions, supposes that every part of residing heat flow density is certain, heat flow field is that several mutual not continuous discrete temperature values form, and in original ambient temperature one timing, the heat flow field calculating distributes;

(b) heat flux density field is loaded on finite element model, calculates the steady temperature field distribution of its finite element model: consider to flow through the air-flow of aircraft surface, the τ of aircraft surface any point _ωof equal value and τ that accordingly can baric flow _∞, the temperature of incompressible air-flow is increased to a given reference temperature value T ^*:

T ^*=0.5T _ω+0.22T _r+0.28T _∞

Wherein:

T ^*represent aircraft surface temperature,

represent boundary layer outer rim temperature, can be obtained T by aerodynamic force calculating section parameter _rfor recovery temperature,

According to the heat flux density field calculating, in MSC.Patran, be carried in finite element model surface, add suitable radiation value balance in addition, solving sequence SOL153 by MSC.Nastran carries out Temperature Distribution calculating to it, obtains the steady temperature field distribution of its finite element model simultaneously;

(c) obtain the structural parameters after steady-state thermal stress distortion according to Steady-State Thermal Field distribution situation:

Making structure initial cell linear stiffness matrix is K ₀, the unnecessary rigidity generating for thermal stress effect is K _σ, structure is equivalent to by the total stress stiffness matrix after thermal effect effect

K=K ₀+K _σ

The structural vibration expression formula under thermal effect effect is

In formula, M represents mass matrix,

represent the vibration shape,

ω is the frequency that shakes,

Kinetics equation has now been described as structure vibration characteristics under Thermal Load, be reduced to the generalized eigenvalue problem of asking above formula, according to the Temperature Distribution situation obtaining, further solve sequence SOL153 to structure is carried out to static analysis by MSC.Nastran, obtain the structural parameters after steady-state thermal stress distortion, comprise stiffness matrix, stiffness matrix is now the effective stiffness matrix of considering under thermal effect, the input file that structure thermal stress Parameter File output is now calculated as aerodynamic force;

(d) set up aerodynamic grid model: the aeroelasticity module in application MSC Flightloads, the pneumatic subregion that becomes number to be applicable to model partition, each pneumatic subregion can be divided into again the applicable pneumatic plate of number, obtain aerodynamic grid parameter, the aerodynamic grid parameter of the structural parameters such as the stiffness matrix of derived type structure and division from MSC Patran, is used three rank piston theories to carry out the non-Unsteady Flow of linear frequency domain and calculates;

(e) structural parameters and the non-Unsteady Flow of frequency domain (d) obtaining after the steady-state thermal stress distortion obtaining according to (c), use p-k method to carry out FLUTTER CALCULATION, obtains flutter speed:

(4) enter Optimizing Flow, in FD ISIGHT, realize optimization integrated, determine whether and meet optimization aim according to the result of FLUTTER CALCULATION, if meet, finish Optimizing Flow, if do not meet, by the value of the optimized algorithm Renewal Design variable laying angle that arranges in FD ISIGHT optimization component, then repeating step (3) carries out the vehicle aeroelastic behavior analysis of thermal environment lower aerofoil.

The optimized algorithm arranging in FD ISIGHT optimization component described in above-mentioned steps (4) can be chosen as archipelago genetic algorithm.

The invention has the beneficial effects as follows:

1) a kind of computing method that solve aerodynamic force, Aerodynamic Heating and flutter speed in hypersonic aircraft aerothermoelastic analysis process have been realized, and take finite element model as research object, use the method can obtain it and consider the FLUTTER CALCULATION result in thermal effect situation;

2) provide a kind of by the optimization method of cutting out design raising aerofoil flutter speed value to composite material skin, thus hypersonic aircraft aeroelasticity performance.

By following description also by reference to the accompanying drawings, it is more clear that the present invention will become, and these accompanying drawings are used for explaining embodiments of the invention.

Accompanying drawing explanation

Fig. 1 is the schematic diagram of the basic procedure of a kind of hypersonic aircraft aeroelastic tailoring of the present invention method;

Fig. 2 is certain type hypersonic aircraft flying tail finite element model figure that step of the present invention (2) is set up;

Fig. 3 is that step of the present invention (3) content (a) is carried out Aerodynamic Heating calculating to certain type hypersonic aircraft flying tail aerofoil, obtains heat flow field distribution plan;

Fig. 4 is that step of the present invention (3) content (b) is loaded into heat flux density field on certain type hypersonic aircraft flying tail aerofoil finite element model, the steady temperature field pattern calculating;

Fig. 5 and Fig. 6 are respectively before case process of the present invention is optimized, certain type hypersonic aircraft flying tail aerofoil flutter V-g and V-f curve map of calculating;

Fig. 7 and Fig. 8 are respectively the interfaces of resolving model.bdf file and FluVel.txt file in FD ISIGHT;

Fig. 9 is the integrated interface of Optimizing Flow of the present invention in FD ISIGHT;

Figure 10 and Figure 11 are respectively cases of the present invention after optimizing, certain type hypersonic aircraft flying tail aerofoil flutter V-g and V-f curve map of calculating;

Embodiment

Embodiment of the method: take the integrated analysis of certain hypersonic aircraft aerofoil aeroelasticity as example, and for a kind of schematic diagram of basic procedure of hypersonic aircraft aeroelastic tailoring method as shown in Figure 1, the specific implementation method of utilizing a kind of hypersonic aircraft aeroelastic tailoring method to improve hypersonic aircraft aeroelasticity performance is described.

(1) selecting to need the design variable of optimizing is covering laying angle [alpha], and constraint condition is for keeping airfoil structure weight constant, and optimization aim is airplane flutter critical velocity maximum, and the mathematical model of this optimization problem can be expressed as:

max?FlutterVelocity

ΔMass=0

s.t.x∈S

In formula, FluterVelocity represents critical flutter Mach number, and Mass represents full machine model quality, and x represents design variable, and S represents design variable set, and according to the restriction of actual manufacturing process, value is series of discrete numerical value :-45 °, and 0 °, 45 °, 90 °;

(2) set up hypersonic aircraft aerofoil geometric model and finite element model, comprise following content:

(a) in 3D sculpting software, set up hypersonic aircraft flying tail geometric model, adopted double-wedge thin airfoil, be connected by rotating shaft with fuselage;

(b) after geometric model is read in to MSC.Patran, set up finite element model and model is carried out to finite element grid division, horizontal tail surface stressed-skin construction part is selected shell unit modeling, and adopt the quadrilateral units in MSC.Patran automatically to divide, the inner girder construction that adopts of horizontal tail, and give beam element attribute, horizontal tail leading edge portion adopts solid element and uses the tetrahedron element in MSC.Patran automatically to divide;

(c) selecting at material, the covering of flying tail part is selected carbon fibre composite laminate (T300/5222), lay if covering is eight layers of symmetrical equilibrium, select four laying angles of each symmetrical individual layer as design variable, be set as respectively [angle ₁, angle ₂, angle ₃, angle ₄], initial laying order is [0/0/0/0] _s;

(3) carry out the vehicle aeroelastic behavior analysis of thermal environment lower aerofoil, it is characterized in that, comprise the following steps:

(a) the heat flow density field distribution that computation model surface is subject to:

The Aerodynamic Heating phenomenon that leading edge of a wing part is subject to is the most obvious, and the leading edge portion of the slim Wing-Body Configurations of hypersonic aircraft and airfoil portion can be ignored thickness factor, regards flat board as and calculates:

With reference to coefficient of viscosity μ ^*solve and draw by the blue expression formula in Saudi:

${\mathrm{\μ}}^{*}={\left(\frac{T^{*}}{288.15}\right)}^{1.5}\frac{398.55}{T^{*}+110.4}\×{1.7894\×10}^{-5}$

Reference density ρ ^*solve and draw by state equation expression formula:

${\mathrm{\ρ}}^{*}=\frac{p_{\∞}}{{\mathrm{RT}}^{*}}$

Can obtain with reference to Reynolds number simultaneously, as follows:

${\mathrm{Re}}_x^{*}=\frac{{\mathrm{\ρ}}^{*}{V}_{\∞}x}{{\mathrm{\μ}}^{*}}$

Margoulis number is solved and is drawn by Reynolds analogy formula:

${\mathrm{St}}^{*}=\frac{c_f^{*}}{2}\frac{1}{{\left(\mathrm{Pr}\right)}^{2/3}}$

Final hypersonic aircraft surface heat flow is tried to achieve by following formula:

Q _aero=St ^*ρ ^*V _∞c _ρ(T _r-T _ω)

Along x direction, flying tail is divided into some equal portions, supposes that every part of residing heat flow density is certain, heat flow field is that several mutual not continuous discrete temperature values form, at original ambient temperature one timing, the heat flow field calculating.This example is divided into 20 equal portions along x direction by flying tail, suppose that every part of residing heat flow density is certain, heat flow field is that 20 mutual not continuous discrete temperature values form, in the time that original ambient temperature is 300K, the heat flow field calculating distributes as shown in Figure 5, and heat flow density is along x to from 183935J/ (m as can be seen from this figure ²is s) to 140000J/ (m ²s) reduces gradually, and flying tail head is the place of suffered heat flux maximum;

(b) heat flux density field is loaded on finite element model, calculates the steady temperature field distribution of its finite element model: consider to flow through the air-flow of aircraft surface, the τ of aircraft surface any point _ωof equal value and τ that accordingly can baric flow _ω, the temperature of incompressible air-flow is increased to a given reference temperature value T ^*:

T ^*=0.5T _ω+0.22T _r+0.28T _∞

Wherein:

T ^*represent aircraft surface temperature,

represent boundary layer outer rim temperature, can be obtained T by aerodynamic force calculating section parameter _rfor recovery temperature,

According to the heat flux density field calculating, in MSC.Patran, be carried in finite element model surface, add suitable radiation value balance in addition simultaneously, solve sequence SOL153 by MSC.Nastran it is carried out to Temperature Distribution calculating, obtain the steady temperature field distribution of its finite element model, flying tail head temperature is the highest, is about 1600 ℃;

(c) obtain the structural parameters after steady-state thermal stress distortion according to Steady-State Thermal Field distribution situation:

Making structure initial cell linear stiffness matrix is K ₀, the unnecessary rigidity generating for thermal stress effect is K _σ, structure is equivalent to by the total stress stiffness matrix after thermal effect effect

K=K ₀+K _σ

The structural vibration expression formula under thermal effect effect is

In formula, M represents mass matrix,

represent the vibration shape,

ω is the frequency that shakes,

Kinetics equation has now been described as structure vibration characteristics under Thermal Load, be reduced to the generalized eigenvalue problem of asking above formula, according to the Temperature Distribution situation obtaining, further solve sequence SOL153 to structure is carried out to static analysis by MSC.Nastran, obtain the structural parameters after steady-state thermal stress distortion, comprise stiffness matrix, stiffness matrix is now the effective stiffness matrix of considering under thermal effect, the input file that structure thermal stress Parameter File output is now calculated as aerodynamic force, the first six rank natural vibration frequency is now as shown in the table:

(d) set up aerodynamic grid model: the aeroelasticity module in application MSC Flightloads, the pneumatic subregion that becomes number to be applicable to model partition, each pneumatic subregion can be divided into again the applicable pneumatic plate of number, obtain aerodynamic grid parameter, the aerodynamic grid parameter of the structural parameters such as the stiffness matrix of derived type structure and division from MSC Patran, is used three rank piston theories to carry out the non-Unsteady Flow of linear frequency domain and calculates;

(e) structural parameters and the non-Unsteady Flow of frequency domain (d) obtaining after the steady-state thermal stress distortion obtaining according to (c), use p-k method to carry out FLUTTER CALCULATION, the flutter critical Mach number obtaining before flying tail optimization is 4.16, flutter frequency is 287.5Hz, its coupled mode is that horizontal tail aerofoil reverses the vibration shape and the coupling of waving the vibration shape, and flutter V-g and V-f curve are distinguished as shown in Figure 5 and Figure 6:

(4) enter Optimizing Flow, realize optimization integrated in FD ISIGHT, first will resolve input file, document analysis is the key link of building Optimization Platform, comprises input file and output file parsing.Input file is resolved and is told the data (being generally design variable value) that ISIGHT need to replace, i.e. covering laying angle [alpha], and value is series of discrete numerical value :-45 °, 0 °, 45 °, 90 °; Output file is resolved and is told the data (being generally binding occurrence and target function value) that ISIGHT will extract, i.e. flutter speed value.Fig. 7 and Fig. 8 have shown respectively the interface of resolving model.bdf file and FluVel.txt file, determine whether and meet optimization aim according to the result of FLUTTER CALCULATION, if meet, finish Optimizing Flow, if do not meet, by FD ISIGHT optimization component, optimized algorithm being set, rule of thumb select archipelago genetic algorithm, the value of Renewal Design variable laying angle, then repeating step (3) carries out the vehicle aeroelastic behavior analysis of thermal environment lower aerofoil.

As shown in Figure 9 integrated in FD ISIGHT of this Optimizing Flow, keeps model gross mass constant, only flutter speed is optimized.After having moved 12 steps, convergence obtains optimum solution.Result of calculation demonstration, the optimum laying order of covering compound substance is [0/45/-45/90] _s critical flutter Mach 2 ship 6.12 after optimization, flutter frequency is 345.7Hz, each order frequency after optimization all presents ascendant trend in various degree, particularly the torsion frequency of horizontal tail has larger raising, result shows that laying angle now of covering compound substance and laying order can effectively increase the rigidity of structure, postpone the generation of chatter phenomenon.Flutter V-g after its optimization and V-f curve are as shown in Figure 10 and Figure 11.

From above optimization is analyzed, we can draw such conclusion, covering compound substance laying mode, particularly laying angle and laying order etc., can produce impact to a certain degree to the aeroelastic characteristic of hypersonic aircraft, on the basis of hypersonic aircraft steady state thermal flutter speed method for solving, utilize compound substance aeroelastic tailoring technology can effectively improve the critical flutter speed of aircraft, hypersonic aircraft aeroelastic analysis method provided by the invention is carried out aeroelastic tailoring design on the multidisciplinary Optimization Software platform of FD ISIGHT, for practical engineering application provides simple and effective disposal route, can in engineering, be well used.

Claims (2)

1. a hypersonic aircraft aeroelastic tailoring method, is characterized in that: the step of the method is:

(1) selecting to need the design variable of optimizing is covering laying angle [alpha], and constraint condition is for keeping airfoil structure weight constant, and optimization aim is airplane flutter critical velocity maximum, and the mathematical model of this optimization problem can be expressed as:

max?FlutterVelocity

ΔMass=0

s.t.x∈S

In formula, FluterVelocity represents critical flutter Mach number, and Mass represents full machine model quality, and x represents design variable, and S represents design variable set, and according to the restriction of actual manufacturing process, value is series of discrete numerical value :-45 °, and 0 °, 45 °, 90 °;

(2) set up hypersonic aircraft aerofoil geometric model and finite element model, comprise following content:

(a) in 3D sculpting software, set up hypersonic aircraft flying tail geometric model, adopted double-wedge thin airfoil, be connected by rotating shaft with fuselage;

(b) after geometric model is read in to MSC.Patran, set up finite element model and model is carried out to finite element grid division, horizontal tail surface stressed-skin construction part is selected shell unit modeling, and adopt the quadrilateral units in MSC.Patran automatically to divide, the inner girder construction that adopts of horizontal tail, and give beam element attribute, horizontal tail leading edge portion adopts solid element and uses the tetrahedron element in MSC.Patran automatically to divide;

(c) selecting at material, the covering of flying tail part is selected carbon fibre composite laminate (T300/5222), lay if covering is eight layers of symmetrical equilibrium, select four laying angles of each symmetrical individual layer as design variable, be set as respectively [angle ₁, angle ₂, angle ₃, angle ₄], initial laying order is [0/0/0/0] _s;

(3) carry out the vehicle aeroelastic behavior analysis of thermal environment lower aerofoil, it is characterized in that, comprise the following steps:

(a) the heat flow density field distribution that computation model surface is subject to:

The Aerodynamic Heating phenomenon that leading edge of a wing part is subject to is the most obvious, and the leading edge portion of the slim Wing-Body Configurations of hypersonic aircraft and airfoil portion can be ignored thickness factor, regards flat board as and calculates:

With reference to coefficient of viscosity μ ^*solve and draw by the blue expression formula in Saudi:

${\mathrm{\μ}}^{*}={\left(\frac{T^{*}}{288.15}\right)}^{1.5}\frac{398.55}{T^{*}+110.4}\×{1.7894\×10}^{-5}$

Reference density ρ ^*solve and draw by state equation expression formula:

${\mathrm{\ρ}}^{*}=\frac{p_{\∞}}{{\mathrm{RT}}^{*}}$

Can obtain with reference to Reynolds number simultaneously, as follows:

${\mathrm{Re}}_x^{*}=\frac{{\mathrm{\ρ}}^{*}{V}_{\∞}x}{{\mathrm{\μ}}^{*}}$

Margoulis number is solved and is drawn by Reynolds analogy formula:

${\mathrm{St}}^{*}=\frac{c_f^{*}}{2}\frac{1}{{\left(\mathrm{Pr}\right)}^{2/3}}$

Final hypersonic aircraft surface heat flow is tried to achieve by following formula:

Q _aero=St ^*ρ ^*V _∞c _ρ(T _r-T _ω)

Along fuselage x direction, fuselage is divided into some equal portions, supposes that every part of residing heat flow density is certain, heat flow field is that several mutual not continuous discrete temperature values form, and in original ambient temperature one timing, the heat flow field calculating distributes;

(b) heat flux density field is loaded on finite element model, calculates the steady temperature field distribution of its finite element model: consider to flow through the air-flow of aircraft surface, the τ of aircraft surface any point _ωof equal value and τ that accordingly can baric flow _ω, the temperature of incompressible air-flow is increased to a given reference temperature value T ^*:

T ^*=0.5T _ω+0.22T _r+0.28T _∞

Wherein:

T ^*represent aircraft surface temperature,

represent boundary layer outer rim temperature, can be obtained T by aerodynamic force calculating section parameter _rfor recovery temperature,

According to the heat flux density field calculating, in MSC.Patran, be carried in finite element model surface, add suitable radiation value balance in addition, solving sequence SOL153 by MSC.Nastran carries out Temperature Distribution calculating to it, obtains the steady temperature field distribution of its finite element model simultaneously;

(c) obtain the structural parameters after steady-state thermal stress distortion according to Steady-State Thermal Field distribution situation:

Making structure initial cell linear stiffness matrix is K ₀, the unnecessary rigidity generating for thermal stress effect is k _σ, structure is equivalent to by the total stress stiffness matrix after thermal effect effect

K=K ₀+K _σ

The structural vibration expression formula under thermal effect effect is

In formula, M represents mass matrix,

represent the vibration shape,

ω is the frequency that shakes,

Kinetics equation has now been described as structure vibration characteristics under Thermal Load, be reduced to the generalized eigenvalue problem of asking above formula, according to the Temperature Distribution situation obtaining, further solve sequence SOL153 to structure is carried out to static analysis by MSC.Nastran, obtain the structural parameters after steady-state thermal stress distortion, comprise stiffness matrix, stiffness matrix is now the effective stiffness matrix of considering under thermal effect, the input file that structure thermal stress Parameter File output is now calculated as aerodynamic force;

(d) set up aerodynamic grid model: the aeroelasticity module in application MSC Flightloads, the pneumatic subregion that becomes number to be applicable to model partition, each pneumatic subregion can be divided into again the applicable pneumatic plate of number, obtain aerodynamic grid parameter, the aerodynamic grid parameter of the structural parameters such as the stiffness matrix of derived type structure and division from MSC Patran, is used three rank piston theories to carry out the non-Unsteady Flow of linear frequency domain and calculates;

(e) structural parameters and the non-Unsteady Flow of frequency domain (d) obtaining after the steady-state thermal stress distortion obtaining according to (c), use p-k method to carry out FLUTTER CALCULATION, obtains flutter speed:

(4) enter Optimizing Flow, in FD ISIGHT, realize optimization integrated, determine whether and meet optimization aim according to the result of FLUTTER CALCULATION, if meet, finish Optimizing Flow, if do not meet, by the value of the optimized algorithm Renewal Design variable laying angle that arranges in FD ISIGHT optimization component, then repeating step (3) carries out the vehicle aeroelastic behavior analysis of thermal environment lower aerofoil.

2. a kind of hypersonic aircraft aeroelastic tailoring method according to claim 1, is characterized in that: the optimized algorithm arranging in the FD ISIGHT optimization component described in above-mentioned steps (4) can be chosen as archipelago genetic algorithm.

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