CN107944137A - The thermographic curve computing technique of hypersonic aircraft trajectory state multi- scenarios method - Google Patents
The thermographic curve computing technique of hypersonic aircraft trajectory state multi- scenarios method Download PDFInfo
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Abstract
The invention discloses a kind of thermographic curve computational methods of hypersonic aircraft trajectory state multi- scenarios method, belong to flight vehicle aerodynamic calculating field.In the engineering fields such as aerospace, with the increase of supersonic vehicle flying speed, Flight Vehicle Structure suffer from the Aerodynamic Heating effect of sternness.Simultaneously because the mission requirements to become increasingly complex, flight path is increasingly complex, and pneumatic thermal environment is also more difficult to predict.Therefore the thermographic curve problem under complicated trajectory merits attention.The present invention carries out high-speed aircraft calculation of aerodynamic heating using without viscous outflow solution with the method that engineering method is combined, and this method can save substantial amounts of memory compared with pure values method and calculate the time, reduce calculating cost.And its scope of application is wider compared with pure engineering method, complex shape is adapted to, while more accurately pneumatic pyrolysis can be provided again.
Description
Technical field
The invention belongs to Aerodynamic Characteristic Calculation Methods for Aircraft field, is specifically a kind of more couplings of hypersonic aircraft trajectory state
The thermographic curve computational methods of conjunction.
Background technology
In the engineering fields such as aerospace, with the increase of supersonic vehicle flying speed, Flight Vehicle Structure by
The Aerodynamic Heating effect of sternness.Simultaneously because the mission requirements to become increasingly complex, flight path is increasingly complex, pneumatic thermal environment
Also it is more difficult to predict.Aerodynamic Heating effect can make under structural material mechanical property high temperature caused by structure and temperature gradient
Drop, the non-uniform thermal deformation of inside configuration and thermal stress can change the stiffness characteristics of structure, while can also change the intrinsic of structure
Vibration characteristics, and these influences are all often negative.Therefore the thermographic curve problem under complicated trajectory merits attention.
At present, hypersonic aircraft surface heat exchanging computational methods mainly include whole flow field method for numerical simulation, engineering is estimated
Method that calculation method and Euler solutions are combined with engineering estimation etc..In general, the limitation of engineering estimating method is larger, and one
As be difficult to be extended to complex appearance, and simple numerical solution need to consume it is substantial amounts of calculate time and memory, it is non-to calculate cost
Chang Gao.The present invention carries out high-speed aircraft calculation of aerodynamic heating using without viscous outflow solution with the method that engineering method is combined, this
Kind method can save substantial amounts of memory compared with pure values method and calculate the time, reduce calculating cost.And with pure engineering side
Method is wider compared to its scope of application, adapts to complex shape, while can provide more accurately pneumatic pyrolysis again.The party
The main thought of method is：It is theoretical based on prandtl boundary layer, by whole flow field be divided into outside boundary layer without viscosity flow field and boundary layer
Interior Viscous Flow two parts, so that Aerodynamic Heating problem be become to solve following two problems：(1) nothing outside boundary layer is glued
The solution in flow field；(2) heat exchange in the leading region of viscosity solves in boundary layer.On the basis of hypersonic aircraft surface heat exchanging
On, the One-dimensional heat transfer calculating along structural thickness direction is coupled, forms the basic skills of whole calculation of aerodynamic heating.
Under complicated trajectory, carry out the costly of time domain dynamic aeroelastic analysis to whole trajectory, can be by ignoring
Simplified compared with the Pneumatic pressure of weak coupling and influence Aerodynamic Heating.Thermographic curve for giving transient state temperature field
Research is carried out more abundant.Chen Wenjun (influence modern defense technology of the Aerodynamic Heating to flight vehicle aerodynamic elastic characteristic,
1998 (3)) describe the adverse effect that Aerodynamic Heating may bring aircraft, outline aerothermoelasticity (including hot rigidity,
Thermal vibration and heat flutter) analysis thinking and step, summarize Aerodynamic Heating to Flight Vehicle Structure rigidity, vibration and the shadow of flutter
The trend of sound.Yellow generation it is brave et al. (Constructional Modal Analysis guided missiles and carrying space technology under the brave thermal environments of Huang Shiyong, Wang Zhi,
2009(5):50-52.) Varying Thickness Plates have been carried out with the Constructional Modal Analysis under transient state thermal environment.The three of tablet in calculating
Given a temperature value on a position, the variation tendency of heated rear structural natural frequencies have studied.The result shows that structure is intrinsic
Frequency is on a declining curve, and bandwidth narrows between frequency, and fuel factor is maximum to single order Torsion Coupling Effect of Mode.It is and complicated
Method in terms of the prediction of structure thermal environment under trajectory state is less.
The content of the invention
Above-mentioned the deficiencies in the prior art are directed to, it is an object of the invention to provide a kind of hypersonic aircraft trajectory shape
The thermographic curve meter method of state multi- scenarios method, to reduce under complicated trajectory, carries out time domain dynamic aeroelastic to whole trajectory
Calculate the great cost produced with analysis.
To reach above-mentioned purpose, the technical solution adopted by the present invention is as follows：
A kind of thermographic curve computational methods of hypersonic aircraft trajectory state multi- scenarios method of the present invention, including step
It is rapid as follows：
A kind of 1. thermographic curve computational methods of hypersonic aircraft trajectory state multi- scenarios method, it is characterised in that
It is as follows including step：
Under complicated trajectory, carry out the costly of time domain dynamic aeroelastic analysis to whole trajectory, can generally lead to
Cross and ignore compared with the Pneumatic pressure of weak coupling and influence Aerodynamic Heating to be simplified.Thus can be by aerothermoelastic analysis
It is reduced to pneumatic a heat problem and a single Aeroelastic Problems.Fig. 1 gives general analysis flow.I.e. first according to bullet
Road state, predicts the Transient Aerodynamic thermal environment on each point on trajectory, then from the thermal environment, calculates structure temperature field,
Then hot-die state of the structure under heat effect is analyzed, finally carries out aerothermoelastic analysis.Specific steps include：
Calculation of aerodynamic heating technology when step 1, trajectory state are long；
Kinetics equation and its solution under step 2, modal coordinate.
2. the thermographic curve of hypersonic aircraft trajectory state multi- scenarios method according to claim 1 calculates skill
Art, it is characterised in that：
Further, step 1 specifically includes：
Step 1-1, calculation of aerodynamic heating technology is calculated and included when trajectory state is long：
It is theoretical based on prandtl boundary layer, by whole flow field be divided into outside boundary layer without the viscosity in viscosity flow field and boundary layer
Flow field two parts, so that Aerodynamic Heating problem be become to solve following two problems：(1) the asking without viscosity flow field outside boundary layer
Solution；(2) heat exchange in the leading region of viscosity solves in boundary layer.On the basis of hypersonic aircraft surface heat exchanging, edge is coupled
The One-dimensional heat transfer in structural thickness direction calculates, and forms the basic skills of whole calculation of aerodynamic heating.This method scope of application
It is wider, complex shape is adapted to, while more accurately pneumatic pyrolysis can be provided again.
Step 1-2, calculation of aerodynamic heating technology calculation process when trajectory state is long：
, it is necessary to according to given trajectory parameter before calculating starts, the nothing for analyzing needs glues outflow solution status list, without viscous
The principle that state determines is mainly：1. the rectangular domain without viscous outflow solution state composition can include whole trajectory；2. the angle of attack,
Mach number changes violent region and suitably encrypts without viscous outflow solution state.Status list is solved according to without viscous outflow, carries out CFD numbers
Value simulation, obtains a series of nothing and glues logistics solution, put it into CFD without spare in viscous result of calculation storehouse.Preparation is completed, bullet
Calculation of aerodynamic heating flow is divided into following several steps, such as Fig. 2 when road state is long：
Initialized 1. calculating, the initial temperature in setting structure, the ginseng such as structure thermal protection scheme, thermally protective materials attribute
Number；
2. determine current trajectory point parameter according to flight time and trajectory parameter, from CFD without being selected in viscous result of calculation storehouse
Bent relevant destination file, by appropriate interpolation method, obtains viscous boundary layer's outer rim parameter under current trajectory point parameter；
3. combining One-dimensional heat transfer in exchange heat in boundary layer engineering calculation and structure to calculate, Temperature Distribution in structure is obtained；
Whether completed 4. judging to calculate, do not completed if calculating, go to 2；Completed if calculating, export result.
Step 1-3, hot-fluid calculates in boundary layer：
Heat current densimeter point counting calculates respectively for stagnation region and non-stagnation region two parts in boundary layer.Stagnation region heat flow density
Calculating, existing a variety of ripe computational methods suitable for different condition, such as：SCALA methods, LEES methods, Fay-
Riddell methods, Kemp-Riddell methods etc..Fay-Riddell methods are used herein, and Fay-Riddell formula are mainly used
In the calculating of axisymmetric equilibrium boundary layer stationary point heat flow density：
Wherein ρw, μw, hwRespectively surface density, viscosity and enthalpy, ρs, μs, hsRespectively Stagnation density, viscosity
And enthalpy, hDFor average air dissociation enthalpy, Pr=0.71, Le=1.0.
The surface heat flow of non-stagnation region calculates, and also having local similar solution in engineering, (LEES blunt body laminar flow hot-fluids are close
Spend distribution formula) and the methods of reference enthalpy method, herein using flat late heat transfer model Numerical heat transfer coefficient and heat flow density, to compared with
The calculating that complex appearance streams carries out simplifying processing：
Wherein ρe, ue, ReeRespectively boundary layer outer rim density, speed and Reynolds number.ε is the compression for considering compressible influence
The factor, F are form factor.The calculation formula of ε is as follows：
Wherein ρ*, μ*The density and viscosity that respectively reference enthalpy method obtains.
Step 1-4, Aerodynamic Heating and structural thermal calculate：
In order to achieve the purpose that fast prediction, the distribution of heat shield structure temperature builds difference side according to One-dimensional Heat Conduction Equation
Journey solves.The differential form of heat transfer formula is as follows：
Tw|T=0=T0
ρ in formulaδFor density of material, cδFor material specific heat capacity, δ is material thickness.
In hypersonic flowing, Aerodynamic Heating process that aircraft is subject to is a lasting transient, the hot side of wall
Boundary's condition is being in moment change before reaching balance.The heat flow density q that air-flow outside boundary layer is passed to aircraft surfacewRemove
, temperature T also with object plane related with rudders pneumatic power parameterwIt is related, and TwCalculating again with heat flow density qwCorrelation, therefore,
Accurate simulation aircraft is wanted to be heated situation, Aerodynamic Heating and structural thermal must carry out coupling calculating.
Wherein TawFor adiabatic wall temperature, α is heat transfer coefficient, and ε is material radiation coefficient.
Further, kinetics equation and its solution specifically include under step 2, modal coordinate：
Dynamics equations can be written as form：
In formula [M], [C], [K] represents mass matrix, damping matrix and stiffness matrix respectively.Assuming that x can pass through structure
Condition shape represented by linear superposition：
Wherein φiFor the i-th rank mode, the mode of structure can be tried to achieve by model analysis, qiFor the generalized displacement of i rank mode.
Comprehensive above formula, and premultiplication [Φ]TIt can obtain：
Wherein [Φ] is φiThe matrix of composition, q qiThe column vector of composition.The orthogonality of natural mode is as follows：
[Φ]T[M] [Φ]=1
[Φ]T[K] [Φ]=[ω]2
[Φ]T[C] [Φ]=2 [ξ] [ω]
Wherein ω, ξ are diagonal matrix, then can be write as the scalar equation of one group of decoupling：
For above-mentioned equation, solved herein using Newmark methods.Newmark step_by_step integrations formula can be written as
Form：
It can thus be concluded that：
α and β can suitably be adjusted according to the stability of calculating and required precision in formula.It can prove to work as α >=1/2, β >=(α
+1/2)2When, Integration Scheme unconditional stability.General to take α=1/2, β=1/4, Newmark methods are average at this time
Acceleration method.In addition, work as α=1/2, Newmark methods deteriorate to linear acceleration method during β=1/6.
Beneficial effects of the present invention：
The present invention carries out high-speed aircraft calculation of aerodynamic heating using without viscous outflow solution with the method that engineering method is combined,
This method can save substantial amounts of memory compared with pure values method and calculate the time, reduce calculating cost.And with pure engineering
Method is wider compared to its scope of application, adapts to complex shape, while can provide more accurately pneumatic pyrolysis again.Together
When, aerothermoelastic analysis is reduced to pneumatic a heat problem and a single Aeroelastic Problems.I.e. first according to trajectory
State, predicts the Transient Aerodynamic thermal environment on each point on trajectory, then from the thermal environment, calculates structure temperature field, connect
Hot-die state of the analysis structure under heat effect, finally carries out aerothermoelastic analysis, can be under fast prediction complexity trajectory state
Structure thermal environment.
Brief description of the drawings
Fig. 1 is analysis calculation process.
Fig. 2 is calculation of aerodynamic heating flow when trajectory state is long.
Fig. 3 is trajectory schematic diagram.
Fig. 4 a are wing monitoring point schematic diagram.
Fig. 4 b are vertical fin monitoring point schematic diagram.
Fig. 5 a change over time for surfaces externally and internally temperature at 1 monitoring point.
Fig. 5 b change over time for surfaces externally and internally temperature at 2 monitoring point.
Fig. 5 c change over time for surfaces externally and internally temperature at 3 monitoring point.
Fig. 5 d change over time for surfaces externally and internally temperature at 4 monitoring point.
Fig. 6 a are distributed for t=390s moment top airfoils hull-skin temperature.
Fig. 6 b are distributed for t=390s moment lower aerofoils hull-skin temperature.
Fig. 7 a are calculating grid.
Fig. 7 b are calculating grid.
Fig. 8 varies with temperature curve for preceding 6 order frequency.
Fig. 9 a are no the first rank of temperature rise natural mode of vibration.
Fig. 9 b are temperature rise 400K the first rank natural mode of vibrations.
Fig. 9 c are no temperature rise second-order natural mode of vibration.
Fig. 9 d are temperature rise 400K second-order natural mode of vibrations.
Fig. 9 e are no the 3rd rank natural mode of vibration of temperature rise.
Fig. 9 f are the 3rd rank natural mode of vibrations of temperature rise 400K.
Fig. 9 g are no temperature rise fourth order natural mode of vibration.
Fig. 9 h are temperature rise 400K fourth order natural mode of vibrations.
Fig. 9 i are temperature rise 600K the first rank natural mode of vibrations.
Fig. 9 j are temperature rise 1000K the first rank natural mode of vibrations.
Fig. 9 k are temperature rise 600K second-order natural mode of vibrations.
Fig. 9 l are temperature rise 1000K second-order natural mode of vibrations.
Fig. 9 m are the 3rd rank natural mode of vibrations of temperature rise 600K.
Fig. 9 n are the 3rd rank natural mode of vibrations of temperature rise 1000K.
Fig. 9 o are temperature rise 600K fourth order natural mode of vibrations.
Fig. 9 p are temperature rise 1000K fourth order natural mode of vibrations.
Figure 10 a are t=660s moment upper outer surface Temperature Distributions.
Figure 10 b are distributed for t=660s moment lower aerofoils hull-skin temperature.
Figure 10 c are distributed for t=660s moment empennages hull-skin temperature.
Figure 11 is without considering 20 rank generalized displacement before heat affecting.
Figure 12 is generalized displacement.
Embodiment
A kind of thermographic curve computational methods of hypersonic aircraft trajectory state multi- scenarios method, including step are as follows：
Under complicated trajectory, carry out the costly of time domain dynamic aeroelastic analysis to whole trajectory, can generally lead to
Cross and ignore compared with the Pneumatic pressure of weak coupling and influence Aerodynamic Heating to be simplified.Thus can be by aerothermoelastic analysis
It is reduced to pneumatic a heat problem and a single Aeroelastic Problems.Fig. 1 gives general analysis flow.I.e. first according to bullet
Road state, predicts the Transient Aerodynamic thermal environment on each point on trajectory, then from the thermal environment, calculates structure temperature field,
Then hot-die state of the structure under heat effect is analyzed, finally carries out aerothermoelastic analysis.Specific steps include：
Calculation of aerodynamic heating technology when step 1, trajectory state are long；
Kinetics equation and its solution under step 2, modal coordinate.
Further, step 1 specifically includes：
Step 1-1, calculation of aerodynamic heating technology is calculated and included when trajectory state is long：
It is theoretical based on prandtl boundary layer, by whole flow field be divided into outside boundary layer without the viscosity in viscosity flow field and boundary layer
Flow field two parts, so that Aerodynamic Heating problem be become to solve following two problems：(1) the asking without viscosity flow field outside boundary layer
Solution；(2) heat exchange in the leading region of viscosity solves in boundary layer.On the basis of hypersonic aircraft surface heat exchanging, edge is coupled
The One-dimensional heat transfer in structural thickness direction calculates, and forms the basic skills of whole calculation of aerodynamic heating.This method scope of application
It is wider, complex shape is adapted to, while more accurately pneumatic pyrolysis can be provided again.
Step 1-2, calculation of aerodynamic heating technology calculation process when trajectory state is long：
, it is necessary to according to given trajectory parameter before calculating starts, the nothing for analyzing needs glues outflow solution status list, without viscous
The principle that state determines is mainly：1. the rectangular domain without viscous outflow solution state composition can include whole trajectory；2. the angle of attack,
Mach number changes violent region and suitably encrypts without viscous outflow solution state.Status list is solved according to without viscous outflow, carries out CFD numbers
Value simulation, obtains a series of nothing and glues logistics solution, put it into CFD without spare in viscous result of calculation storehouse.Preparation is completed, bullet
Calculation of aerodynamic heating flow is divided into following several steps, such as Fig. 2 when road state is long：
Initialized 1. calculating, the initial temperature in setting structure, the ginseng such as structure thermal protection scheme, thermally protective materials attribute
Number；
2. determine current trajectory point parameter according to flight time and trajectory parameter, from CFD without being selected in viscous result of calculation storehouse
Relevant destination file is taken, by appropriate interpolation method, obtains viscous boundary layer's outer rim parameter under current trajectory point parameter；
3. combining One-dimensional heat transfer in exchange heat in boundary layer engineering calculation and structure to calculate, Temperature Distribution in structure is obtained；
Whether completed 4. judging to calculate, do not completed if calculating, go to 2；Completed if calculating, export result.
Step 1-3, hot-fluid calculates in boundary layer：
Heat current densimeter point counting calculates respectively for stagnation region and non-stagnation region two parts in boundary layer.Stagnation region heat flow density
Calculating, existing a variety of ripe computational methods suitable for different condition, such as：SCALA methods, LEES methods, Fay-
Riddell methods, Kemp-Riddell methods etc..Fay-Riddell methods are used herein, and Fay-Riddell formula are mainly used
In the calculating of axisymmetric equilibrium boundary layer stationary point heat flow density：
Wherein ρw, μw, hwRespectively surface density, viscosity and enthalpy, ρs, μs, hsRespectively Stagnation density, viscosity
And enthalpy, hDFor average air dissociation enthalpy, Pr=0.71, Le=1.0.
The surface heat flow of non-stagnation region calculates, and also having local similar solution in engineering, (LEES blunt body laminar flow hot-fluids are close
Spend distribution formula) and the methods of reference enthalpy method, herein using flat late heat transfer model Numerical heat transfer coefficient and heat flow density, to compared with
The calculating that complex appearance streams carries out simplifying processing：
Wherein ρe, ue, ReeRespectively boundary layer outer rim density, speed and Reynolds number.ε is the compression for considering compressible influence
The factor, F are form factor.The calculation formula of ε is as follows：
Wherein ρ*, μ*The density and viscosity that respectively reference enthalpy method obtains.
Step 1-4, Aerodynamic Heating and structural thermal calculate：
In order to achieve the purpose that fast prediction, the distribution of heat shield structure temperature builds difference side according to One-dimensional Heat Conduction Equation
Journey solves.The differential form of heat transfer formula is as follows：
Tw|T=0=T0
ρ in formulaδFor density of material, cδFor material specific heat capacity, δ is material thickness.
In hypersonic flowing, Aerodynamic Heating process that aircraft is subject to is a lasting transient, the hot side of wall
Boundary's condition is being in moment change before reaching balance.The heat flow density q that air-flow outside boundary layer is passed to aircraft surfacewRemove
, temperature T also with object plane related with rudders pneumatic power parameterwIt is related, and TwCalculating again with heat flow density qwCorrelation, therefore,
Accurate simulation aircraft is wanted to be heated situation, Aerodynamic Heating and structural thermal must carry out coupling calculating.
Wherein TawFor adiabatic wall temperature, α is heat transfer coefficient, and ε is material radiation coefficient.
Further, kinetics equation and its solution specifically include under step 2, modal coordinate：
Dynamics equations can be written as form：
In formula [M], [C], [K] represents mass matrix, damping matrix and stiffness matrix respectively.Assuming that x can pass through structure
Condition shape represented by linear superposition：
Wherein φiFor the i-th rank mode, the mode of structure can be tried to achieve by model analysis, qiFor the generalized displacement of i rank mode.
Comprehensive above formula, and premultiplication [Φ]TIt can obtain：
Wherein [Φ] is φiThe matrix of composition, q qiThe column vector of composition.The orthogonality of natural mode is as follows：
[Φ]T[M] [Φ]=1
[Φ]T[K] [Φ]=[ω]2
[Φ]T[C] [Φ]=2 [ξ] [ω]
Wherein ω, ξ are diagonal matrix, then can be write as the scalar equation of one group of decoupling：
For above-mentioned equation, solved herein using Newmark methods.Newmark step_by_step integrations formula can be written as
Form：
It can thus be concluded that：
α and β can suitably be adjusted according to the stability of calculating and required precision in formula.It can prove to work as α >=1/2, β >=(α
+1/2)2When, Integration Scheme unconditional stability.General to take α=1/2, β=1/4, Newmark methods are average add at this time
Tachometric method.In addition, work as α=1/2, Newmark methods deteriorate to linear acceleration method during β=1/6.
For the ease of the understanding of those skilled in the art, the present invention is made into one with reference to specific embodiment result of calculation
The explanation of step.
Trajectory state lower class X-37B aircraft thermographic curve calculates and analysis
Technical parameter
Carry out the unsteady thermographic curve of trajectory state for class X-37B aircraft to calculate and analyze.Fig. 3 gives trajectory
Schematic diagram, about 880 seconds ballistic flight total times, flying height is gradually reduced with the time, and Mach number is also gradually reduced.Flight course
In flown all the time with positive incidence, and be gradually reduced with the flight time.
Full machine uses same thermal protection scheme in calculating, i.e. top layer is 2mm titanium alloys, and inner layer is the titanium dioxide of 20mm
Silicon.The material properties of titanium alloy and silica are shown in Tables 1 and 2.Initial hull-skin temperature is 500K, and internal surface temperature is
300K, the temperature in structure intermediate layer are obtained by surfaces externally and internally temperature linearity interpolation.
1 titanium alloy material attribute of table
2 earth silicon material attribute of table
Result of calculation
Fig. 4 gives the monitoring point schematic diagram on wing and vertical fin.Fig. 5 gives the surfaces externally and internally temperature at these points
Change over time curve.PG is ideal gas result of calculation in figure, and CG is the result of calculation considered after chemical non-equilibrium effect.
As can be seen from the figure monitoring point hull-skin temperature is first raised with the time and reduced afterwards, and internal surface temperature gradually rises, and outer surface
Temperature change it is very violent, and inner surface then more relaxes.Monitoring point hull-skin temperature rise is primarily due to height
Reduction, atmospheric density becomes dense, and crosses monitoring point hull-skin temperature after maximum temperature and reduce and be primarily due to aircraft
Speed is more and more lower.Effects of chemical reaction all has an impact the surface temperature of wing and vertical fin, and influence mainly appears on 400~
600 second stage (height 60Km~30Km, Mach number 18~6).Main cause is probably that this altitude range air is denser, at the same time
Flying speed is very big, therefore air themperature is higher, and chemical reaction is more violent, and violent chemical reaction causes aircraft surface temperature
Decrease.The leading edge of a wing o'clock was to reach maximum 1444K at 450 seconds under the conditions of ideal gas, and vertical fin leading edge was reached at 453 seconds
To maximum 988K.Chemical non-equilibrium gas condition lower wing leading edge point was to reach maximum 1418K at 404.3 seconds, before vertical fin
Edge reached maximum 952K at 428.2 seconds.Than the leading edge of a wing point highest under the conditions of chemical non-equilibrium under the conditions of ideal gas
Temperature high 26K, the high 36K of vertical fin leading edge temperature, it is seen that effects of chemical reaction can a degree of highest for reducing leading edge point
Temperature.Influence of the effects of chemical reaction to wing is bigger than the influence of vertical fin, the reason is that due to band angles-of-attack, empennage is in leeward
Area, the temperature in flow field is not on the temperature height near the wing in area windward, thus wing is also compared in the influence of effects of chemical reaction
It is weak.It can also be seen that the temperature changing trend of two monitoring points on wing is more consistent in figure, and maximum temperature values are also almost
It is identical, and the temperature changing trend difference of two monitoring points on empennage is larger, the maximum temperature close to the monitoring point of fuselage is shown
Write and be less than the monitoring point away from fuselage.
Fig. 6 gives the t=390s Temperature Distributions of moment upper and lower surface, and curve is on the section at z=1.2m in figure
Temperature Distribution.Flying height H=60Km at this time, flight Mach number Ma=18, AOA=40 ° of flying drilling angle.Can from figure
Go out, under the state, the skewness of the upper and lower surface of aerofoil, the temperature difference of upper lower aerofoil is very big.Top airfoil temperature compares bottom wing
The high about 1200K in face, it can be seen that, during the big angles-of-attack of hypersonic aircraft, windward side and leeward temperature difference are very big,
During design, it is necessary to use different thermal protection schemes.
Thermographic curve calculating has been carried out to class X-37B shape aircraft, analyze first uniform temperature rise to intrinsic frequency and
The influence of natural mode of vibration, the thermographic curve that temperature distribution state when then choosing 660s on trajectory carries out full machine calculate.It is whole
The structural model of a aircraft uses shell unit, is specially：Vertical fin thickness 40mm, material SiC；Fuselage thickness is 70mm,
Material is aluminium alloy；Wing thickness is 30mm, and material is aluminium alloy.Grid is calculated as shown in fig. 7, aluminium alloy and SiC under room temperature
Material properties as shown in Table 3 and Table 4, it is assumed that the Young's modulus of aluminium alloy is uniformly reduced with the rise of temperature, poplar during 1300K
Family name's modulus is down to 36GPa, and the Young's modulus of SiC does not vary with temperature then.Mode is analyzed using Nastran, in calculating
Only consider influence of the heat to Material Physics attribute.
3 aluminum alloy materials attribute of table
Table 4SiC material properties
Table 5 gives the intrinsic frequency of the preceding 10 rank mode of complete machine under different temperatures, and it is intrinsic that Fig. 8 gives preceding 6 rank mode
Frequency variation with temperature curve.It can be seen from the graph that elevated reduction of the intrinsic frequency of structure with temperature, this master
If caused by young modulus of material is reduced with temperature.Fig. 9 gives the preceding 4 rank Mode Shape of complete machine under different temperatures.
It can be seen from the figure that, 4 rank Mode Shapes are basically identical before no temperature rise, temperature rise 400K, temperature rise 600K, illustrate that temperature raises, right
The vibration shape is substantially without influence.And the preceding 2 rank mode of temperature rise 1000K with other three kinds in the case of on the contrary, this is because wing and tail
The different reason of wing material.Without under temperature rising state, first-order modal is the bending of empennage, and second-order modal bends for wing, wing material
The Young's modulus of material is reduced with the rise of temperature, and structure softens, and vibration frequency reduces, and the Young's modulus of empennage is not with temperature
Degree change, vibration frequency are constant.
Preceding 10 rank intrinsic frequency (unit under the different wing temperatures of table 5：HZ)
The Temperature Distribution of wing and vertical fin when Figure 10 gives 660s.Simplified place has been done in calculating process to Temperature Distribution
Reason, takes the mean temperature on wing and vertical fin to be applied in structure and carries out hot-die state analysis respectively, then carries out dynamic hot gas and moves
Elastic calculation.Figure 11 is given without considering preceding 20 rank generalized displacement under the influence of temperature, and each rank mode is in constant amplitude under perturbation action
The trend of vibration, this is because the smaller influence to structure of unsteady aerodynamic force is also smaller, while it can also be seen that two before structure
Rank mode plays a major role.Figure 12, which gives, considers that temperature is influenced and without considering 2 rank generalized displacements before under the influence of temperature, considered
With the no generation flutter of state influenced without considering Aerodynamic Heating.After considering heat affecting, structure quake frequency reduces, it is easier to
Reach dynamic response border.
Concrete application approach of the present invention is very much, and the above is only the preferred embodiment of the present invention, it is noted that for
For those skilled in the art, without departing from the principle of the present invention, some improvement can also be made, this
A little improve also should be regarded as protection scope of the present invention.
Claims (3)
1. a kind of thermographic curve computational methods of hypersonic aircraft trajectory state multi- scenarios method, it is characterised in that by heat
Aeroelastic analysis is reduced to pneumatic a heat problem and a single Aeroelastic Problems, i.e., first according to trajectory state, in advance
The Transient Aerodynamic thermal environment on each point on trajectory is surveyed, then from the thermal environment, calculates structure temperature field, then analysis knot
Hot-die state of the structure under heat effect, finally carries out aerothermoelastic analysis, specific steps include：
Calculation of aerodynamic heating when step 1, trajectory state are long；
Kinetics equation and its solution under step 2, modal coordinate.
2. the thermographic curve computing technique of hypersonic aircraft trajectory state multi- scenarios method according to claim 1,
It is characterized in that：Step 1 specifically includes：
Step 1-1, calculation of aerodynamic heating includes when trajectory state is long：
It is theoretical based on prandtl boundary layer, by whole flow field be divided into outside boundary layer without the Viscous Flow in viscosity flow field and boundary layer
Two parts, so that Aerodynamic Heating problem be become to solve following two problems：(1) solution without viscosity flow field outside boundary layer；(2)
The heat exchange in the leading region of viscosity solves in boundary layer.On the basis of hypersonic aircraft surface heat exchanging, couple thick along structure
The One-dimensional heat transfer for spending direction calculates, and forms the basic skills of whole calculation of aerodynamic heating；
Step 1-2, calculation of aerodynamic heating flow when trajectory state is long：
, it is necessary to according to given trajectory parameter before calculating starts, the nothing for analyzing needs glues outflow solution status list, according to without viscous
Outflow solution status list, carries out CFD numerical simulations, obtains a series of nothing and glue logistics solution, put it into CFD without viscous result of calculation
Spare in storehouse, preparation is completed, and calculation of aerodynamic heating flow is divided into following four step when trajectory state is long：
Initialized 1. calculating, the initial temperature in setting structure, the parameter such as structure thermal protection scheme, thermally protective materials attribute；
2. current trajectory point parameter is determined according to flight time and trajectory parameter, from CFD without selected songs phase in viscous result of calculation storehouse
The destination file of pass, by appropriate interpolation method, obtains viscous boundary layer's outer rim parameter under current trajectory point parameter；
3. combining One-dimensional heat transfer in exchange heat in boundary layer engineering calculation and structure to calculate, Temperature Distribution in structure is obtained；
Whether completed 4. judging to calculate, do not completed if calculating, go to 2；Completed if calculating, export result；
Step 1-3, hot-fluid calculates in boundary layer：
Heat current densimeter point counting calculates respectively for stagnation region and non-stagnation region two parts in boundary layer, the meter of stagnation region heat flow density
Calculation uses Fay-Riddell methods, and Fay-Riddell formula are based on axisymmetric equilibrium boundary layer stationary point heat flow density
Calculate：
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<msub>
<mi>q</mi>
<mrow>
<mi>w</mi>
<mi>s</mi>
</mrow>
</msub>
<mo>=</mo>
<mn>0.763</mn>
<mo>*</mo>
<msup>
<mi>Pr</mi>
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<mn>0.6</mn>
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<mi>w</mi>
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<mi>s</mi>
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<msub>
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<mi>du</mi>
<mi>e</mi>
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<mn>0.52</mn>
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<mfrac>
<msub>
<mi>h</mi>
<mi>D</mi>
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<mi>h</mi>
<mi>s</mi>
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<mrow>
<mo>(</mo>
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<mi>h</mi>
<mi>s</mi>
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<mo>-</mo>
<msub>
<mi>h</mi>
<mi>w</mi>
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<mo>)</mo>
</mrow>
</mrow>
Wherein ρw, μw, hwRespectively surface density, viscosity and enthalpy, ρs, μs, hsRespectively Stagnation density, viscosity and
Enthalpy, hDFor average air dissociation enthalpy, Pr=0.71, Le=1.0；
The surface heat flow of non-stagnation region calculates, using flat late heat transfer model Numerical heat transfer coefficient and heat flow density, to more complicated outer
The calculating that shape is streamed carries out simplifying processing：
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<mi>q</mi>
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<mi>t</mi>
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<mi>Pr</mi>
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<mi>w</mi>
</msub>
<mrow>
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<mn>2</mn>
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<mi>&rho;</mi>
<mi>e</mi>
</msub>
<msub>
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<mi>e</mi>
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</mrow>
<mo>*</mo>
<mi>&epsiv;</mi>
<mo>*</mo>
<mi>F</mi>
</mrow>
Wherein ρe, ue, ReeRespectively boundary layer outer rim density, speed and Reynolds number.ε be consider the compression of compressible influence because
Son, F are form factor.The calculation formula of ε is as follows：
<mrow>
<mi>&epsiv;</mi>
<mo>=</mo>
<msup>
<mrow>
<mo>(</mo>
<mfrac>
<msup>
<mi>&mu;</mi>
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<mi>e</mi>
</msub>
</mfrac>
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</mrow>
<mn>0.8</mn>
</msup>
</mrow>
Wherein ρ*, μ*The density and viscosity that respectively reference enthalpy method obtains；
Step 1-4, Aerodynamic Heating and structural thermal calculate：
In order to achieve the purpose that fast prediction, the distribution of heat shield structure temperature builds difference equation according to One-dimensional Heat Conduction Equation and asks
Solution.The differential form of heat transfer formula is as follows：
<mrow>
<msub>
<mi>&rho;</mi>
<mi>&delta;</mi>
</msub>
<msub>
<mi>c</mi>
<mi>&delta;</mi>
</msub>
<mi>&delta;</mi>
<mfrac>
<mrow>
<mi>d</mi>
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</mrow>
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Tw|T=0=T0
ρ in formulaδFor density of material, cδFor material specific heat capacity, δ is material thickness；
In hypersonic flowing, Aerodynamic Heating process that aircraft is subject to is a lasting transient, wall thermal boundary bar
Part is being in moment change before reaching balance.The heat flow density q that air-flow outside boundary layer is passed to aircraft surfacewExcept with
Rudders pneumatic power parameter is related, also the temperature T with object planewIt is related, and TwCalculating again with heat flow density qwCorrelation, therefore, standard
True simulated flight device is heated situation, and Aerodynamic Heating and structural thermal must carry out coupling calculating：
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<msub>
<mi>q</mi>
<mi>w</mi>
</msub>
<mo>=</mo>
<mi>&alpha;</mi>
<mrow>
<mo>(</mo>
<msub>
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<mrow>
<mi>a</mi>
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</mrow>
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<msub>
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</mrow>
<mo>-</mo>
<msubsup>
<mi>&epsiv;&sigma;T</mi>
<mi>w</mi>
<mn>4</mn>
</msubsup>
</mrow>
Wherein TawFor adiabatic wall temperature, α is heat transfer coefficient, and ε is material radiation coefficient.
3. the thermographic curve computing technique of hypersonic aircraft trajectory state multi- scenarios method according to claim 1,
It is characterized in that：Kinetics equation and its solution specifically include under step 2, modal coordinate：
Dynamics equations can be written as form：
<mrow>
<mo>&lsqb;</mo>
<mi>M</mi>
<mo>&rsqb;</mo>
<mo>{</mo>
<mover>
<mi>x</mi>
<mo>&CenterDot;&CenterDot;</mo>
</mover>
<mo>}</mo>
<mo>+</mo>
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<mi>C</mi>
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<mo>(</mo>
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</mrow>
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</mrow>
In formula [M], [C], [K] represents mass matrix, damping matrix and stiffness matrix respectively.Assuming that x can pass through the mould of structure
State shape is represented by linear superposition：
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<mi>x</mi>
<mo>=</mo>
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<mo>&Sigma;</mo>
<mn>1</mn>
<mi>n</mi>
</munderover>
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<mi>&phi;</mi>
<mi>i</mi>
</msub>
<msub>
<mi>q</mi>
<mi>i</mi>
</msub>
</mrow>
Wherein φiFor the i-th rank mode, the mode of structure can be tried to achieve by model analysis, qiFor the generalized displacement of i rank mode.It is comprehensive
Above formula, and premultiplication [Φ]TIt can obtain：
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<mi>&Phi;</mi>
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<mi>T</mi>
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<mi>C</mi>
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<mi>T</mi>
</msup>
<mo>&lsqb;</mo>
<mi>K</mi>
<mo>&rsqb;</mo>
<mo>&lsqb;</mo>
<mi>&Phi;</mi>
<mo>&rsqb;</mo>
<mo>{</mo>
<mi>q</mi>
<mo>}</mo>
<mo>=</mo>
<msup>
<mrow>
<mo>&lsqb;</mo>
<mi>&Phi;</mi>
<mo>&rsqb;</mo>
</mrow>
<mi>T</mi>
</msup>
<mo>{</mo>
<mi>F</mi>
<mrow>
<mo>(</mo>
<mi>t</mi>
<mo>)</mo>
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<mo>}</mo>
</mrow>
Wherein [Φ] is φiThe matrix of composition, q qiThe column vector of composition.The orthogonality of natural mode is as follows：
[Φ]T[M] [Φ]=1
[Φ]T[K] [Φ]=[ω]2
[Φ]T[C] [Φ]=2 [ξ] [ω]
Wherein ω, ξ are diagonal matrix, then can be write as the scalar equation of one group of decoupling：
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<mi>i</mi>
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<mi>q</mi>
<mo>=</mo>
<msup>
<msub>
<mi>&phi;</mi>
<mi>i</mi>
</msub>
<mi>T</mi>
</msup>
<mo>{</mo>
<mi>F</mi>
<mrow>
<mo>(</mo>
<mi>t</mi>
<mo>)</mo>
</mrow>
<mo>}</mo>
<mo>,</mo>
<mi>i</mi>
<mo>=</mo>
<mn>1</mn>
<mo>,</mo>
<mn>2</mn>
<mo>,</mo>
<mo>...</mo>
<mo>,</mo>
<mi>n</mi>
</mrow>
For above-mentioned equation, solved herein using Newmark methods, Newmark step_by_step integrations formula can be written as form：
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<mn>1</mn>
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<mi>&alpha;</mi>
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<mi>&alpha;</mi>
<msub>
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</mrow>
<mrow>
<mi>t</mi>
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<mi>&Delta;</mi>
<mi>t</mi>
</mrow>
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</mrow>
<mi>t</mi>
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<mi>&Delta;</mi>
<mi>t</mi>
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<mo>&lsqb;</mo>
<mrow>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>2</mn>
</mfrac>
<mo>-</mo>
<mi>&beta;</mi>
<mo>)</mo>
</mrow>
<msub>
<mrow>
<mo>{</mo>
<mover>
<mi>q</mi>
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</mover>
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</mrow>
<mi>t</mi>
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<mo>+</mo>
<mi>&beta;</mi>
<msub>
<mrow>
<mo>{</mo>
<mover>
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</mover>
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</mrow>
<mrow>
<mi>t</mi>
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<mi>&Delta;</mi>
<mi>t</mi>
</mrow>
</msub>
<mo>&rsqb;</mo>
<msup>
<mi>&Delta;t</mi>
<mn>2</mn>
</msup>
</mrow>
It can thus be concluded that：
<mfenced open = "" close = "">
<mtable>
<mtr>
<mtd>
<mrow>
<mo>(</mo>
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<mi>i</mi>
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<mi>i</mi>
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<msub>
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</mtr>
<mtr>
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<msub>
<mi>&xi;</mi>
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<mi>t</mi>
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</mtd>
</mtr>
</mtable>
</mfenced>
α and β is suitably adjusted according to the stability of calculating and required precision in formula, as α >=1/2, β >=(α+1/2)2When, Integration Scheme
Unconditional stability, generally takes α=1/2, and β=1/4, Newmark methods are average acceleration method at this time, in addition, when α=
Newmark methods deteriorate to linear acceleration method when 1/2, β=1/6.
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| CN109029907A (en) * | 2018-07-18 | 2018-12-18 | 大连理工大学 | A kind of parameter similar method of pneumatic thermal environmental test simulated conditions |
| CN111780949A (en) * | 2020-07-10 | 2020-10-16 | 南京航空航天大学 | CFD analysis-based total pressure correction method for high-speed air inlet channel precursor wind tunnel experiment |
| CN112036039A (en) * | 2020-09-01 | 2020-12-04 | 内蒙古科技大学 | High-precision numerical analysis method for thermal performance of conjugated heat transfer material |
| CN112036039B (en) * | 2020-09-01 | 2022-12-23 | 内蒙古科技大学 | High-precision numerical analysis method for thermal performance of conjugated heat transfer material |
| CN112329148A (en) * | 2020-11-18 | 2021-02-05 | 西北工业大学 | Optimization method and system for thermal protection structure of hypersonic aircraft |
| CN113642267A (en) * | 2021-07-15 | 2021-11-12 | 中国空气动力研究与发展中心空天技术研究所 | Aircraft surface front edge attachment line region extraction method |
| CN113642267B (en) * | 2021-07-15 | 2023-09-26 | 中国空气动力研究与发展中心空天技术研究所 | Extraction method for aircraft surface leading edge attachment line region |
| CN116383974A (en) * | 2023-06-06 | 2023-07-04 | 西安现代控制技术研究所 | Aircraft axial force correction method considering pneumatic heating effect |
| CN116383974B (en) * | 2023-06-06 | 2023-09-01 | 西安现代控制技术研究所 | Aircraft axial force correction method considering pneumatic heating effect |
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