CN112329148A - Optimization method and system for thermal protection structure of hypersonic aircraft - Google Patents

Optimization method and system for thermal protection structure of hypersonic aircraft Download PDF

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CN112329148A
CN112329148A CN202011294917.9A CN202011294917A CN112329148A CN 112329148 A CN112329148 A CN 112329148A CN 202011294917 A CN202011294917 A CN 202011294917A CN 112329148 A CN112329148 A CN 112329148A
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谢丹
张世豪
冀春秀
车驰
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Northwestern Polytechnical University
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Abstract

The invention discloses a method and a system for optimizing a thermal protection structure of a hypersonic aircraft. The invention carries out thermoelastometry analysis on the TPS metal wall plate of the universal reusable carrier in hypersonic flow, applies a two-way coupling scheme considering the influence of elastic deformation on pneumatic heating, and forms a thermoelastometry coupling analysis technology by establishing a model of pneumatic elasticity, pneumatic heat, heat conduction and multi-physical field coupling, so that the temperature of the bottom wall plate is lower, the stress is smaller, and the stable boundary is widened, thereby providing beneficial guidance for the optimization design of the TPS structure.

Description

Optimization method and system for thermal protection structure of hypersonic aircraft
Technical Field
The invention relates to the field of a thermal protection structure of a hypersonic aircraft, in particular to a method and a system for optimizing the thermal protection structure of the hypersonic aircraft.
Background
Due to the high Mach number of the hypersonic aircraft and the prerequisite condition of flight in the atmosphere, severe pneumatic heating can cause material degradation, thermal stress and thermal bending moment under the condition of temperature change, and finally the problem of thermal-pneumatic-elastic multi-physical field coupling of the thin-wall structure of the hypersonic aircraft is caused. Taking a typical hypersonic aircraft as an example, the lower surface of the hypersonic aircraft is a part of a scramjet engine (fig. 1), and in order to enable the fuselage to work in a safe temperature range and protect the fuselage from the influence of pneumatic heating, a Thermal Protection System (TPS) is designed on the lower surface of the hypersonic aircraft.
The existing TPS concept optimization design research mainly considers the constraints in the aspects of weight, heat load, surface temperature, structural strain/stress and the like. NASA performs a parametric weight comparison of different TPS concepts using a one-dimensional transient TPS structure code, designs an advanced metallic TPS concept, and determines the dimensions of the prototype wallboard to maintain acceptable temperatures under the design loads (including ramp-up and ramp-in heating rates, pressures, acoustics and accelerations). Recent studies have considered thermomechanical constraints of low thermal conductivity and high mechanical properties in TPS optimization. Also, based on the concept of thermomechanical optimization, the purpose of designing a metallic TPS is to minimize structural weight and meet various performance constraints, including deformation of the top wall panel, stress around the lug holes, stress of the honeycomb core, and internal temperature. Culler reduces the TPS prototype for the NAAX-43 aircraft to a panel structure to reduce modeling complexity and computational cost due to comprehensive aeroelastic analysis of fluid-thermal-structure coupling, however, this is an overall study on panel flutter in hypersonic flow and not for TPS design. Hypersonic wallboard flutter analysis studies that consider fluid-thermal-structural coupling remain limited.
Since aerodynamic forces, aerodynamic heat and structural deformations are almost all calculated based on a single model of decoupling, but in practice these are strongly coupled, it is crucial to consider predictions of surface heat flow, transient temperature distribution and structural deformations/stresses under the flow-thermal-solid coupling effect, i.e. thermoaeroelastic analysis.
Disclosure of Invention
The invention aims to provide a method and a system for optimizing a thermal protection structure of a hypersonic aircraft, wherein a TPS metal wall plate of a general reusable carrier in hypersonic flow is subjected to thermoelasticity analysis, a bidirectional coupling scheme considering the influence of elastic deformation on pneumatic heating is applied, and a thermoelasticity coupling analysis technology is formed by establishing a model of pneumatic elasticity, pneumatic heat, heat conduction and multi-physical field coupling, so that the temperature of a bottom wall plate is lower, the stress is smaller, a stable boundary is widened, and beneficial guidance is provided for the optimization design of the TPS structure.
In order to achieve the purpose, the invention provides the following scheme:
a method for optimizing a thermal protection structure of a hypersonic aircraft comprises the following steps:
solving the aeroelasticity model by taking the initial displacement and the initial temperature as input signals of the aeroelasticity model to obtain pressure, Mach number and temperature;
solving the aerodynamic thermal model by taking the pressure, the Mach number and the temperature as input signals of the aerodynamic thermal model to obtain net heat flux;
solving the heat transfer model by taking the net heat flux as an input signal of a heat conduction model to obtain the spatial temperature distribution of the TPS metal plate structure;
updating the temperature according to the space temperature distribution, and calculating the temperature change material parameters of the TPS metal wall plate;
calculating thermal stress and thermal bending moment according to the material parameters and the temperature distribution of the metal wall plate;
calculating the wallboard displacement according to the thermal stress and the thermal bending moment;
and optimizing the thermal protection structure of the hypersonic aircraft through thermal-dynamic elastic coupling analysis according to the updated temperature, the thermal stress, the thermal bending moment and the wallboard displacement.
Optionally, the solving the aerodynamic thermal model by using the pressure, the mach number, and the temperature as input signals of the aerodynamic thermal model to obtain a net heat flux specifically includes:
solving the aerodynamic thermal model by taking the aerodynamic force, the Mach number and the temperature as input signals of an aerodynamic thermal model to obtain aerodynamic heat flux and radiant heat flux;
and (4) making the difference between the pneumatic heat flux and the radiant heat flux to obtain a net heat flux.
Optionally, the aeroelastic model comprises an aerodynamic model and a structural dynamic model.
Optionally, the building of the structural dynamics model comprises:
Figure BDA0002784964660000031
Figure BDA0002784964660000032
Figure BDA0002784964660000033
Figure BDA0002784964660000034
wherein w is the wall displacement, hi,ρiRespectively represents the thickness and the density of the ith laminate, v is Poisson's ratio, TrefIs a reference temperature, NxIs a film force, NTFor thermal stress, MTT (x, z) is the space temperature distribution of the heat protection metal plate, E (x), and alpha (x) is the Young modulus and the thermal expansion coefficient of the wall plate along the chord direction respectively. x and z represent the chordwise and transverse coordinates of the panel, respectively.
Optionally, the establishing of the aerodynamic model comprises:
Figure BDA0002784964660000035
Figure BDA0002784964660000036
Figure BDA0002784964660000037
Figure BDA0002784964660000038
Figure BDA0002784964660000039
wherein, Ma1，ρ1，p1，T1Mach number, density, pressure, temperature, Ma, of free incoming flow3，ρ3，p3，T3The Mach number, density, pressure and temperature of airflow at the front edge of the wallboard are shown, gamma is the specific heat ratio of air, and theta and beta are airflow turning angle and shock wave angle respectively.
Optionally, the calculation formula of the aerodynamic heat flux and the radiant heat flux is as follows:
Qaero＝St*ρ*Ue(Haw-Hw)
Figure BDA0002784964660000041
Figure BDA0002784964660000042
H*＝He+0.50(Hw-He)+0.22(Haw-He)
Figure BDA0002784964660000043
H＝CPT
wherein Q isaeroFor aerodynamic heat flux, QradTo radiate heat flux, H*For reference enthalpy, He，UeEnthalpy, flow velocity, H, at the outer edge of the boundary layerwIs wall enthalpy, HawIs the adiabatic wall enthalpy, r is the coefficient of restitution, H0In the form of the total enthalpy,
Figure BDA0002784964660000044
St*，
Figure BDA0002784964660000045
Figure BDA0002784964660000046
ρ*，η*prandtl number, Stanton number, local surface friction coefficient, local Reynolds number, density, viscosity, C, respectively, calculated at a reference enthalpyPIs the specific heat at constant pressure of air, sigma is a Stefan-Boltzmann constant, epsilon is the surface radiation coefficient, and T is the specific heat at constant pressureenvIs ambient temperature, TwThe temperature of the upper wall surface of the metal plate is protected by heat.
The invention also provides a hypersonic aircraft thermal protection structure optimization system, which comprises:
the aeroelastic model solving module is used for solving the aeroelastic model by taking the initial displacement and the initial temperature as input signals of the aeroelastic model to obtain the pressure, the Mach number and the temperature;
the aerodynamic thermal model solving module is used for solving the aerodynamic thermal model by taking the pressure, the Mach number and the temperature as input signals of the aerodynamic thermal model to obtain net heat flux;
the heat transfer model solving module is used for solving the heat transfer model by taking the net heat flux as an input signal of the heat conduction model to obtain the spatial temperature distribution of the TPS metal plate structure;
the material parameter calculation module is used for updating the temperature according to the space temperature distribution and calculating the temperature change material parameters of the TPS metal wall plate;
the thermal stress and thermal bending moment calculation module is used for calculating thermal stress and thermal bending moment according to the material parameters and the temperature distribution of the metal wall plate;
the displacement calculation module is used for calculating the displacement of the wallboard according to the thermal stress and the thermal bending moment;
and the optimization module is used for optimizing the thermal protection structure of the hypersonic aircraft through thermal-dynamic elastic coupling analysis according to the updated temperature, the thermal stress, the thermal bending moment and the wallboard displacement.
Optionally, the pneumatic thermal model solving module specifically includes:
the solving unit is used for solving the aerodynamic thermal model by taking the aerodynamic force, the Mach number and the temperature as input signals of the aerodynamic thermal model to obtain aerodynamic heat flux and radiant heat flux;
and the difference unit is used for making difference between the pneumatic heat flux and the radiant heat flux to obtain net heat flux.
According to the specific embodiment provided by the invention, the invention discloses the following technical effects:
the invention discloses a method and a system for optimizing a thermal protection structure of a hypersonic aircraft, wherein the method comprises the following steps: solving the aeroelasticity model by taking the initial displacement and the initial temperature as input signals of the aeroelasticity model to obtain pressure, Mach number and temperature; solving the aerodynamic thermal model by taking the pressure, the Mach number and the temperature as input signals of the aerodynamic thermal model to obtain net heat flux; solving the heat transfer model by taking the net heat flux as an input signal of a heat conduction model to obtain the spatial temperature distribution of the TPS metal plate structure; updating the temperature according to the space temperature distribution, and calculating the temperature change material parameters of the TPS metal wall plate; calculating thermal stress and thermal bending moment according to the material parameters and the temperature distribution of the metal wall plate; calculating the wallboard displacement according to the thermal stress and the thermal bending moment; and optimizing the thermal protection structure of the hypersonic aircraft through thermal-dynamic elastic coupling analysis according to the updated temperature, the thermal stress, the thermal bending moment and the wallboard displacement. The invention carries out thermoelastometry analysis on the TPS metal wall plate of the universal reusable carrier in hypersonic flow, applies a two-way coupling scheme considering the influence of elastic deformation on pneumatic heating, and forms a thermoelastometry coupling analysis technology by establishing a model of pneumatic elasticity, pneumatic heat, heat conduction and multi-physical field coupling, so that the temperature of the bottom wall plate is lower, the stress is smaller, and the stable boundary is widened, thereby providing beneficial guidance for the optimization design of the TPS structure.
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In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings needed to be used in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art to obtain other drawings without inventive exercise.
FIG. 1 is a schematic structural view of a typical hypersonic aircraft;
FIG. 2 is a schematic structural view of a TPS thermal protective wall panel on an inclined surface of a scramjet engine;
FIG. 3 is a flow chart of a method for optimizing a thermal protection structure of a hypersonic aerocraft in accordance with an embodiment of the present invention;
FIG. 4 is a bi-directional coupling flow chart of a thermoaeroelastic analysis in accordance with an embodiment of the present invention;
FIG. 5 is a schematic diagram showing the variation of the parameters of the materials of each layer of the TPS wall board with the temperature;
FIG. 6 is a TPS structure design with different layer thickness dimensions;
FIG. 7 is a scheme comparison diagram;
FIG. 8 shows the difference H and MaStability boundary schematic under conditions;
FIG. 9 shows different schemes at different H and MaStability boundary schematic under conditions;
FIG. 10 is a graph comparing transient temperature profiles for the upper surface of the TPS and the surface of the wall plate at 75% of the length;
FIG. 11 is a graph comparing the transient stress profile of the surface of the middle wall panel at 75% of the length;
FIG. 12 is a flow chart of a thermoelastic coupling analysis for ballistic conditions;
FIG. 13 is a schematic diagram of an exemplary calculated trajectory of X-43A;
FIG. 14 is a calculation of a typical trajectory for X-43A;
FIG. 15 is a block diagram of a hypersonic aircraft thermal protection structure optimization system according to an embodiment of the invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
The invention aims to provide a method and a system for optimizing a thermal protection structure of a hypersonic aircraft, wherein a TPS metal wall plate of a general reusable carrier in hypersonic flow is subjected to thermoelasticity analysis, a bidirectional coupling scheme considering the influence of elastic deformation on pneumatic heating is applied, and a thermoelasticity coupling analysis technology is formed by establishing a model of pneumatic elasticity, pneumatic heat, heat conduction and multi-physical field coupling, so that the temperature of a bottom wall plate is lower, the stress is smaller, a stable boundary is widened, and beneficial guidance is provided for the optimization design of the TPS structure.
The invention takes a TPS (thermoplastic polystyrene) wall plate on a part of the lower surface of a scramjet of a general hypersonic aircraft as a research object (as shown in figure 2), and the TPS wall plate is a multi-layer wall plate structure consisting of a TPS thermal protection layer and a metal wall plate at the bottom layer. The TPS comprises a radiation layer and a heat insulation layer which are respectively used for radiating heat to the environment and blocking the heat from being transmitted into the metal wall plate of the bottom layer. The bottom wall panel is considered to be a von Karman panel simply supported at both ends. And theta and beta are respectively an airflow turning angle and a shock wave angle and are used for calculating airflow parameters in the aeroelastic model.
And establishing a coordinate system for the von Karman plates simply supported at two ends, wherein the simple supporting point at the front end is taken as an original point, the chord direction of the wallboard is the positive direction of the x axis, and the direction of the z axis which is vertical to the x axis is positive.
The TPS panel layers are distributed in the z-direction. The bidirectional coupling process is shown in fig. 4, and comprises the solution of submodules, data transfer between modules and time domain iterative advancement. The initial parameters of the wall plate and the pneumatic parameters of the incoming flow are known, and the local pressure, the Mach number and the temperature are output by solving the aeroelastic model as the input quantity of the gas-heat model. Solving the aerodynamic thermal model yields an aerodynamic heat flux and a radiant heat flux, the latter representing the amount of heat radiated by the radiant layer into the environment. The difference between the two is the net heat flux Qaero-QradAnd is conducted to the TPS wall plate through the heat insulation layer. The heat conduction problem is solved by a two-dimensional transient heat transfer equation to obtain new space temperature distribution of the whole sandwich plate structure, so that the quantity changing along with the temperature is updated and used for the next calculation. The updated elastic modulus, thermal expansion coefficient, thermal stress and thermal bending moment are fed back to the aeroelastic module, the thermal conduction coefficient and specific heat are fed back to the thermal conduction module, and the temperature of the wall plate is fed back to a local parameter solving process to calculate new aerodynamic heat. In this way, the time-domain coupling iterations between aerothermal-aeroelastic are completed until the time of flight is as expected.
In order to make the aforementioned objects, features and advantages of the present invention comprehensible, embodiments accompanied with figures are described in further detail below.
As shown in fig. 3-4, a method for optimizing a thermal protection structure of a hypersonic aerocraft includes:
step 101: and solving the aeroelasticity model by taking the initial displacement and the initial temperature as input signals of the aeroelasticity model to obtain the pressure, the Mach number and the temperature.
Step 102: and solving the pneumatic thermal model by taking the pressure, the Mach number and the temperature as input signals of the pneumatic thermal model to obtain the net heat flux.
Step 103: and solving the heat transfer model by taking the net heat flux as an input signal of the heat conduction model to obtain the spatial temperature distribution of the TPS metal plate structure.
Step 104: and updating the temperature according to the spatial temperature distribution, and calculating the temperature change material parameters of the TPS metal wall plate.
Step 105: and calculating the thermal stress and the thermal bending moment according to the material parameters and the temperature distribution of the metal wall plate.
Step 106: and calculating the displacement of the wall plate according to the thermal stress and the thermal bending moment.
Step 107: and optimizing the thermal protection structure of the hypersonic aircraft through thermal-dynamic elastic coupling analysis according to the updated temperature, the thermal stress, the thermal bending moment and the wallboard displacement.
The present invention is described in detail below:
(1) initial iteration with given initial panel displacement W0And an initial temperature T0As input signals for the aeroelastic model, the aeroelastic model is solved (step 101), wherein the aerodynamic part obtains the local (wall position, point 4) pressure P4Mach number Ma4And temperature T4And the method is used for solving the aerodynamic heat. The displacement part obtains the wall plate displacement w by solving the equation of motion, and the pneumatic parameters given by the aerodynamic part are needed in the solving process. W is calculated in order to obtain the change situation of the w along with time, and the w is used for analyzing different TPS design schemes and drawing a stable boundary diagram in subsequent development.
(2) With P4、Ma4、T4Local air flow parameters, respectively calculating the aerodynamic heat flux QaeroAnd radiant heat flux QradThereby obtaining a net heat flux Qaero-QradTo the upper surface of the TPS wall (step 102).
(3) Net heat flux Q of metal wall plateaero-QradAs input, the thermal conduction model is solved (step 103) to obtain the spatial temperature distribution T (z, x, T) of the TPS metal plate structure. z and x are the thickness direction and chord direction positions of the wall plate in the coordinate system respectively, and t represents time. T is4Is the local temperature at the location 4 of the wall plate in the iteration step, and T (z, x, T) is the spatial temperature distribution of the entire TPS metal plate at a new time calculated after heat conduction.
Taking the temperature distribution T (z, x, T) as an input signal, 3 data updates are to be done: calculating new T by using the updated temperature distribution T (z, x, T) as an input signal of the pneumatic thermal module4And updating the temperature boundary. Secondly, because the TPS heat conduction coefficient k (T) and the air specific heat c (T) are functions of temperature, k (T), c (T) and T which change along with the temperature are updated on the basis of temperature distribution and are used as input signals of a next time domain iteration heat conduction module. Thirdly, the Young modulus E (T) and the thermal expansion coefficient alpha (T) of the wall plate are related to the temperature, the wall plate Young modulus E (T) and the thermal expansion coefficient alpha (T) are fitted to a function of x when the fourth step is carried out non-dimensionalization, the material parameters E (T) and alpha (T) of the metal wall plate are updated, and the thermal stress N is calculatedTAnd hot bending moment MTThe related formula is shown in the step four. Thereby converting E (T), alpha (T), NT,MTAnd feeding back the signal as an input signal to the aeroelastic module to perform the next time domain iteration.
And (5) next iteration, namely replacing W with the wall plate displacement W obtained by solving the equation of motion in the last iteration0The displacement part uses the updated E (T), alpha (T), N as the input of the aeroelastic modelT,MTAnd solving a motion equation, calculating local parameters according to the inflow parameters and the updated T (z, x, T) by the aerodynamic part, and then performing calculation by an aerodynamic thermal model. The heat conduction module is to use the heat flux of the pneumatic heat output and the updated k (T), c (T) in the previous step to solve.
And continuously carrying out model solving and data transmission in the process until the expected flight time is reached, and stopping iteration. And finally obtaining information such as wallboard deformation, temperature distribution, thermal stress change data along with time and the like in the time domain, and using the information for analyzing different TPS design schemes and drawing stable boundary diagrams in subsequent expansion.
In a word, the two-way coupling analysis of the aeroelasticity-aeroelasticity is completed through independent solution and data transfer among the aeroelasticity, the aerothermal and the heat conduction models. The bidirectional coupling is embodied as: the aeroelastic module inputs displacement deformation to the aerothermal/heat conduction module, and the aerothermal/heat conduction module inputs temperature distribution and material temperature change and thermal load caused by the temperature distribution to the aeroelastic module, so that the bidirectional coupling is realized, as shown in fig. 4.
For the TPS wallboard geometry of fig. 2, first, models of three sub-problems in the bidirectional coupling flow are established, respectively.
1) Establishing a aeroelastic model
For the von Karman wallboard shown in FIG. 2, consider the membrane force N caused by geometric nonlinearityxInertia force, aerodynamic force qaAnd in-plane thermal stress N caused by aerodynamic heatTBending moment MT。
Establishing an aerodynamic model to obtain a free airflow parameter T1，P1，Ma1，ρ1The result of the last iteration T, W is obtained (initial iteration is the initial temperature T of the panel)0Initial displacement W0) Calculating aerodynamic force P of wall plate position according to aerodynamic model4Mach number Ma4Unsteady aerodynamic force qaAnd temperature T4Signal, establishing a displacement model formula (6), and establishing the unsteady aerodynamic force q according to the unsteady aerodynamic force qaObtaining the last iteration result E, alpha, NT，MTAnd calculating a new displacement value W according to the displacement model, wherein the new displacement value is used for the next iteration calculation of the aerodynamic model.
Establishing an aerodynamic model:
calculating the air flow parameters at the position 3 according to the oblique shock wave theory:
Figure BDA0002784964660000101
Figure BDA0002784964660000102
Figure BDA0002784964660000103
Figure BDA0002784964660000104
Figure BDA0002784964660000105
in the formula, Ma1，ρ1，p1，T1Mach number, density, pressure, temperature, Ma, of free incoming flow3，ρ3，p3，T3Mach number, density, pressure and temperature of airflow at the front edge of the wallboard are shown, gamma is the specific heat ratio of air, and beta represents a shock wave angle.
The hypersonic aerodynamic force calculation adopts a classical nonlinear third-order piston theory:
Figure BDA0002784964660000106
in addition to the panel flexibility, the in-plane stress in the panel is given by:
Figure BDA0002784964660000107
the overall parameters are calculated from the isentropic flow assumption.
Figure BDA0002784964660000108
Figure BDA0002784964660000109
Figure BDA0002784964660000111
Figure BDA0002784964660000112
In the formula Tt，ρ0，p0The total temperature, total density and total pressure.
The air flow parameter at the front edge (position 3) of the wall plate is taken as an environmental parameter, and according to the piston theory, the pneumatic pressure at the researched place (position 4) is
Figure BDA0002784964660000113
And calculating the temperature and the Mach number according to the piston theory and the isentropic flow relation.
Figure BDA0002784964660000114
Figure BDA0002784964660000115
Building a structural dynamics model:
assuming no bow in the spanwise direction, the y-direction deflection is ignored. The aero-elastic equation of motion for a two-dimensional panel is therefore:
Figure BDA0002784964660000116
wherein
Figure BDA0002784964660000117
Wherein w is the wall displacement, hi,ρiRespectively represents the thickness and the density of the ith laminate, v is Poisson's ratio, TrefIs a reference temperature, NxIs thinMembrane force, NTFor thermal stress, MTIs the hot bending moment, qaT (x, z) is the space temperature distribution of the heat protection metal plate, E (x), and alpha (x) is the Young modulus and the thermal expansion coefficient of the wall plate along the chord direction respectively. x, z represent the coordinates of the panel in the chord and transverse (thickness) directions, respectively.
2) Establishing a pneumatic thermal model
P obtained according to the steps4、Ma4、T4The local air flow parameters are used as input signals, the pneumatic thermal model is solved, and the pneumatic thermal flux Q generated by pneumatic heating is respectively calculatedaeroAnd radiant heat flux Q dissipated by the radiant layerradThereby obtaining the net heat flux Q of the metal wall plateaero-Qrad。
Based on the assumption that the TPS metal plate is positioned in complete turbulence, the aerodynamic heat flux of a turbulent boundary layer is calculated by adopting a classical semi-empirical Eckert reference enthalpy method. First, a reference enthalpy is calculated:
H*＝He+0.50(Hw-He)+0.22(Haw-He) (17)
wherein:
Figure BDA0002784964660000121
the aerodynamic heat flux is then:
Qaero＝St*ρ*Ue(Haw-Hw) (19)
wherein:
Figure BDA0002784964660000122
based on the reference enthalpy method, the reference temperature method can also be obtained through a temperature-enthalpy relation:
H＝CPT (21)
in addition to the aerodynamic heat flux, the thermal radiation flux is obtained taking into account the thermal radiation generated by the TPS skin opacity, scattering conditions and enclosure by the thermostatic environment:
Figure BDA0002784964660000123
in the formula, H*For reference enthalpy, He，UeEnthalpy, flow velocity, H, at the outer edge of the boundary layerwIs wall enthalpy, HawIs the adiabatic wall enthalpy, r is the coefficient of restitution, H0In the form of the total enthalpy,
Figure BDA0002784964660000124
ρ*，η*prandtl number, Stanton number, local surface friction coefficient, local Reynolds number, density, viscosity, C, respectively, calculated at a reference enthalpyPIs the specific heat at constant pressure of air, sigma is a Stefan-Boltzmann constant, epsilon is the surface radiation coefficient, and T is the specific heat at constant pressureenvIs ambient temperature, TwThe temperature of the upper wall surface of the metal plate is protected by heat.
3) Establishing a heat conduction model
Obtaining the net heat flux Q of the previous stepaero-QradAnd (3) solving a heat conduction model by combining TPS heat conduction coefficient K (T) and air specific heat C (T) parameters (T is 300K in the first iteration) generated in the last iteration as input signals to obtain the spatial temperature distribution T (z, x, T) of the TPS metal plate structure.
Modeling the transient temperature field in the panel using a two-dimensional heat transfer equation:
Figure BDA0002784964660000131
and (3) calculating the transient temperature distribution of the TPS structure by adopting an explicit format with time forward and space center discretization to carry out a finite difference method:
Figure BDA0002784964660000132
wherein ρ is density, c is specific heat, and k isx，kzRespectively in the chord direction and the thickness directionUpward heat conduction coefficients, delta x and delta z are heat conduction space step lengths in the chord direction and the thickness direction respectively, delta t is heat conduction time step length, subscripts i and j are ith and jth differential units in the chord direction and the thickness direction respectively, and superscript n represents nth time iteration step. Therefore, the temperature of the molten metal is controlled,
Figure BDA0002784964660000133
the temperature value of the ith unit in the chord direction and the jth unit in the thickness direction of the metal wallboard at the nth time step and the like.
The radiation layer adopts a PM-2000 honeycomb sandwich structure, the heat insulation layer adopts an internal multi-layer heat insulation structure (IMI), and the metal wall plate adopts high-temperature titanium (Ti-6Al-2Sn-4Zr-2 Mo). The heat transfer coefficient and specific heat as a function of temperature are shown in FIG. 5.
Considering that the distance between the nodes is variable and the materials at the two sides of the nodes are different, the temperature obtained based on the new time step is
Figure BDA0002784964660000134
Wherein
Figure BDA0002784964660000141
Figure BDA0002784964660000142
Half-step index
Figure BDA0002784964660000143
And
Figure BDA0002784964660000144
representing nodes between adjacent layers, xj,i-1,xj,i,xj,i+1Respectively representing the x-coordinate, z, of three chordally adjacent pointsj-1,i,zj,i,zj+1,iRespectively represent z coordinates of three adjacent points in the thickness direction, then | xj,i-1-xj,i|,|xj,i+1-xj,iI represents the distance between three points adjacent in chord direction, and Zj-1,i-zj,i|,|zj+1,i-zj,iAnd | represents the distance between three adjacent points in the thickness direction. The boundary condition is the thermal flux boundary of the TPS structure surface, which is exposed to the aerodynamic heating and thermal radiation environment, so the last term of equation (25) can be replaced with:
Figure BDA0002784964660000145
the aeroelasticity, the aerodynamic heat and the heat conduction models of the TPS wall plate are respectively established, so that the thermoelasticity coupling analysis can be carried out, and a foundation is provided for further TPS structure optimization design, thermoelasticity response, stability domain prediction and the like.
4) Iterative data update
Calculating material parameters E (T), alpha (T) of the metal wall plate according to the space temperature distribution T (z, x, T) of the whole TPS metal plate structure obtained in the step, and calculating thermal stress NTAnd hot bending moment MT(ii) a Calculating an iterative heat conduction coefficient k (T), and an air specific heat parameter c (T) as the heat conduction model input in the next iterative process; inputting the aerodynamic model established in the first step according to the spatial temperature distribution of the whole TPS metal plate structure in the steps, and calculating new temperature T required in the next iteration process4(ii) a And carrying out the next iteration process.
Figure BDA0002784964660000146
Wherein
Figure BDA0002784964660000151
D0，E0,α0,v,σ0Respectively bending stiffness, Young's modulus, coefficient of thermal expansion, Poisson's ratio and in-plane stress of the panel at the initial temperature. RhopH, a are TPS base layers respectivelyDensity, thickness and chordwise length of the wallboard. N is a radical ofx,NT,MTRespectively the film stress, thermal stress and thermal bending moment of the wall plate. q. q.s3,M3，ρ3The incoming flow pressure, mach number and density at position 3 are shown, respectively. λ, μ are called dimensionless dynamic pressure, mass ratio (pneumatic/structural), respectively. All physical quantities with an upper line represent dimensionless values. Thus, a dimensionless control equation is obtained:
Figure BDA0002784964660000152
to simplify the solution, the thermal expansion coefficient, young's modulus and thermal bending moment at a particular temperature are assumed to be a quadratic polynomial of x based on experience:
Figure BDA0002784964660000153
Figure BDA0002784964660000154
Figure BDA0002784964660000155
wherein alpha is1,α2,α3；E1,E2,E3；MT1,MT2 MT3Respectively represent
Figure BDA0002784964660000156
To be determined coefficient of the quadratic polynomial.
Then there is
Figure BDA0002784964660000157
Figure BDA0002784964660000158
In addition to this, the present invention is,
Figure BDA0002784964660000161
Figure BDA0002784964660000162
Figure BDA0002784964660000163
Figure BDA0002784964660000164
Figure BDA0002784964660000165
the motion equation is spatially discretized using the classical Galerkin method, assuming the lateral displacements of the panels as follows:
Figure BDA0002784964660000166
wherein
Figure BDA0002784964660000167
Time-domain coefficients representing sinusoidal modes, C1,C2,C3,C4Representing the undetermined coefficients of a quadratic polynomial.
After dispersion, a group of modal coefficients A is obtainednAnd (tau) solving an M-term second-order ordinary differential equation by adopting a fourth-order Runge Kutta (RK4) numerical integration method to obtain the thermal elastic time-domain displacement response of the wallboard, and further calculating the thermal stress, the thermal bending moment and the like of the TPS wallboard.
The law of the material parameter of each layer of the TPS wallboard along with the change of temperature is shown in FIG. 5, which is a temperature change curve chart of specific heat capacity, heat conduction coefficient, Young modulus and thermal expansion coefficient. And solving the thermoelasticity problem according to the method given in the step four and the bidirectional coupling flow of the figure 3.
6) TPS structure optimization design
The TPS structure optimization design can reduce the structure quality, improve the flutter boundary and reduce the structure working temperature, and the consideration on the fluid-heat-structure bidirectional coupling is less at present. Based on the thermoelastic coupling analysis technology proposed and verified in the foregoing, the TPS structure quality is calculated for the conditions of different TPS layer thicknesses, and the flutter boundary is determined based on the obtained panel thermoelastic time domain response, so as to evaluate the influence of the TPS structure size on the thermoelastic stability.
One of the TPS structure designs was chosen as the standard and seven comparative designs of other different layer thicknesses were chosen (as shown in fig. 6). Schemes 1, 2, 4 halve the radiation layer thickness compared to the standard scheme; schemes 1, 3, 6 halve the thickness of the thermal insulation layer; schemes 3 and 5 double the thickness of the radiation layer; schemes 2 and 7 double the thickness of the insulating layer. The fixed flight conditions H-30 km and Ma-10.5, resulted in the structural mass and flutter boundaries for all scenarios (as shown in fig. 6).
Ranking the TPS structure mass in descending order (figure 7(a)), it is evident that the 24.07kg for case 1 and 24.44kg for case 4 are reduced by about 5.4% and 4%, respectively, compared to the standard case 25.45kg structure mass. The chatter occurrence times (chatter boundaries) are arranged in ascending order (fig. 7(b)), and it is apparent that the chatter occurrence times are delayed by about 100% by the case 2 and the case 7. Meanwhile, a multi-objective optimization design concept based on lighter weight and later flutter is introduced (fig. 7(c)), and compared with the standard scheme, the scheme 2 and the scheme 7 are two better TPS structure (layer thickness) design schemes.
7) Hot gas bomb stability boundary prediction
In addition to the optimization of the structural quality and the flutter boundary, the stable region is also an index which needs to be considered in the TPS optimization design. The stable region shows the occurrence time of static stability, flexion, quasiperiodic and chaotic motion, and has important guiding significance for design.
Based on the provided thermoaeroelastic coupling analysis technology, the thermoelastic problem is solved by a fixed variable method under the condition that the flying height and the Mach number are respectively changed, the distribution condition of each stable boundary time point along with the change of the flying height and the Mach number is obtained by changing the flying height and the Mach number respectively, the thermoelastic stable area of the TPS metal wall plate is drawn, the thermoelastic stable area comprises the static stability, the buckling, the quasi-period and the chaos and other movements, and the stable boundaries of the scheme 2, the scheme 7 and the standard scheme are compared.
First, the flight height is changed within a range of 25 to 33km with a fixed mach number Ma of 12, and the flight time and the wall plate stability region in the flight height parameter plane are plotted (fig. 8 (a)). Secondly, the fixed altitude H is 30km, the mach number increases uniformly in steps of 0.5 in the range of 7.5-12, and fig. 8(b) shows a similar stable region. Finally, scheme 2, scheme 7 were compared to the standard scheme (fig. 9). The results show that: 1) both scheme 2 and scheme 7 delay the onset time of the buckling/chaotic transient/quasiperiod. Specifically, as H (fig. 9(a)) and Ma (fig. 9(b)) increase, the deviation gradually increases. 2) Comparing scheme 2 with scheme 7, the dashed and dotted lines (flutter boundaries) marked with open squares are closely tied together, the difference is not large; 3) the curve marked with triangles represents the bifurcation point from the chaotic transient to the quasiperiodic motion, concluding that scheme 2 experiences a shorter chaotic transient before entering the quasiperiodic steady-state response than scheme 7.
The invention has the following beneficial results:
the method comprises the steps of simplifying a thin-wall structure of a typical hypersonic aircraft body lower surface engine position, taking a three-layer thermal protection metal wall plate as a physical model, establishing a pneumatic elasticity, pneumatic heat and heat conduction model, and performing coupling modeling among multiple physical fields to form a thermal-pneumatic elasticity coupling analysis technology of the metal wall plate structure;
in order to carry out structure optimization design on the thermal protection structure, various structural schemes (TPS layer thickness) are selected, and compared with a standard scheme, the invention finds two better schemes by aiming at lighter weight and later generation of flutter: case 2 and case 7 delayed the chatter occurrence time by about 100% with a lighter structural weight.
The transient temperature (shown in fig. 10) and the in-plane stress (shown in fig. 11) of the surface of the wallboard obtained by the scheme 2 and the scheme 7 are evaluated at the same time, so that the temperature of the bottom wallboard is lower, the stress is lower, and the effect of optimizing the design is encouraged.
In terms of stable boundaries, both optimization schemes delay the occurrence time of the buckling/chaotic transient/quasiperiod, and compared with scheme 7, scheme 2 undergoes a shorter chaotic transient process before entering the quasiperiod steady-state response.
The thermoaeroelastic coupling analysis model of the hypersonic aerocraft established by the invention can be further used for thermoelastic coupling analysis under a typical trajectory, and the specific flow is shown in FIG. 12.
Example 1: thermoelastic coupling analysis under X-34A typical design trajectory
A typical calculated trajectory for an X-34A vehicle (fig. 13) includes a post-boost acceleration phase with a prolonged hypersonic cruise flight phase, and a speed, altitude descent phase. The calculation and analysis of the long-term aerodynamic heating state and the thermal aerodynamic elasticity of the wall plate are carried out, and fig. 14 shows the calculation results of the calculation example, which specifically comprises the change curves of the temperature of the TPS metal plate surfaces 1/4, 1/2 and 3/4 along with the trajectory, and the change curves of the temperature, the thermal stress and the thermal bending moment of the bottom wall plate 3/4 along with the trajectory.
As shown in fig. 15, the present invention further provides a hypersonic aircraft thermal protection structure optimization system, including:
and the aeroelastic model solving module 201 is used for solving the aeroelastic model by taking the initial displacement and the initial temperature as input signals of the aeroelastic model to obtain the pressure, the Mach number and the temperature.
And the aerodynamic thermal model solving module 202 is configured to solve the aerodynamic thermal model by using the pressure, the mach number, and the temperature as input signals of the aerodynamic thermal model to obtain a net heat flux.
The pneumatic thermal model solving module specifically comprises:
the solving unit is used for solving the aerodynamic thermal model by taking the aerodynamic force, the Mach number and the temperature as input signals of the aerodynamic thermal model to obtain aerodynamic heat flux and radiant heat flux;
and the difference unit is used for making difference between the pneumatic heat flux and the radiant heat flux to obtain net heat flux.
And the heat transfer model solving module 203 is used for solving the heat transfer model by taking the net heat flux as an input signal of the heat conduction model to obtain the spatial temperature distribution of the TPS metal plate structure.
And the material parameter calculation module 204 is configured to update the temperature according to the spatial temperature distribution, and calculate a temperature change material parameter of the TPS metal wall plate.
A thermal stress and thermal bending moment calculation module 205, configured to calculate a thermal stress and a thermal bending moment according to the material parameter and the temperature distribution of the metal wall plate.
And a displacement calculation module 206 for calculating the wallboard displacement according to the thermal stress and the thermal bending moment.
And the optimization module 207 is used for optimizing the thermal protection structure of the hypersonic aircraft through thermal-dynamic elastic coupling analysis according to the updated temperature, the thermal stress, the thermal bending moment and the wall plate displacement.
The embodiments in the present description are described in a progressive manner, each embodiment focuses on differences from other embodiments, and the same and similar parts among the embodiments are referred to each other. For the system disclosed by the embodiment, the description is relatively simple because the system corresponds to the method disclosed by the embodiment, and the relevant points can be referred to the method part for description.
The principles and embodiments of the present invention have been described herein using specific examples, which are provided only to help understand the method and the core concept of the present invention; meanwhile, for a person skilled in the art, according to the idea of the present invention, the specific embodiments and the application range may be changed. In view of the above, the present disclosure should not be construed as limiting the invention.

Claims (8)

1. A method for optimizing a thermal protection structure of a hypersonic aircraft is characterized by comprising the following steps:
solving the aeroelasticity model by taking the initial displacement and the initial temperature as input signals of the aeroelasticity model to obtain pressure, Mach number and temperature;
solving the aerodynamic thermal model by taking the pressure, the Mach number and the temperature as input signals of the aerodynamic thermal model to obtain net heat flux;
solving the heat transfer model by taking the net heat flux as an input signal of a heat conduction model to obtain the spatial temperature distribution of the TPS metal plate structure;
updating the temperature according to the space temperature distribution, and calculating the temperature change material parameters of the TPS metal wall plate;
calculating thermal stress and thermal bending moment according to the material parameters and the temperature distribution of the metal wall plate;
calculating the wallboard displacement according to the thermal stress and the thermal bending moment;
and optimizing the thermal protection structure of the hypersonic aircraft through thermal-dynamic elastic coupling analysis according to the updated temperature, the thermal stress, the thermal bending moment and the wallboard displacement.
2. The method for optimizing the thermal protection structure of the hypersonic aircraft according to claim 1, wherein the step of solving the aerodynamic thermal model by using the pressure, the mach number and the temperature as input signals of the aerodynamic thermal model to obtain a net heat flux comprises:
solving the aerodynamic thermal model by taking the aerodynamic force, the Mach number and the temperature as input signals of an aerodynamic thermal model to obtain aerodynamic heat flux and radiant heat flux;
and (4) making the difference between the pneumatic heat flux and the radiant heat flux to obtain a net heat flux.
3. The hypersonic aircraft thermal protection structure optimization method of claim 1, wherein the aeroelastic model comprises an aerodynamic model and a structural dynamics model.
4. The optimization method for the thermal protection structure of the hypersonic aircraft according to claim 3, characterized in that the establishment of the displacement model comprises:
Figure FDA0002784964650000011
Figure FDA0002784964650000021
Figure FDA0002784964650000022
Figure FDA0002784964650000023
wherein w is the wall displacement, hi,ρiRespectively represents the thickness and the density of the ith laminate, v is Poisson's ratio, TrefIs a reference temperature, NxIs a film force, NTFor thermal stress, MTT (x, z) is the space temperature distribution of the heat protection metal plate, E (x), and alpha (x) is the Young modulus and the thermal expansion coefficient of the wall plate along the chord direction respectively. x and z represent the chordwise and transverse coordinates of the panel, respectively.
5. The optimization method for the thermal protection structure of the hypersonic aircraft according to claim 3, characterized in that the establishment of the aerodynamic model comprises:
Figure FDA0002784964650000024
Figure FDA0002784964650000025
Figure FDA0002784964650000026
Figure FDA0002784964650000027
Figure FDA0002784964650000028
wherein, Ma1，ρ1，p1，T1Mach number, density, pressure, temperature, Ma, of free incoming flow3，ρ3，p3，T3The Mach number, density, pressure and temperature of airflow at the front edge of the wallboard are shown, gamma is the specific heat ratio of air, and theta and beta are airflow turning angle and shock wave angle respectively.
6. The method for optimizing the thermal protection structure of the hypersonic aircraft according to claim 2, characterized in that the calculation formula of the aerodynamic heat flux and the radiant heat flux is as follows:
Qaero＝St*ρ*Ue(Haw-Hw)
Figure FDA0002784964650000031
Haw＝r(H0-He)+He,
Figure FDA0002784964650000032
H*＝He+0.50(Hw-He)+0.22(Haw-He)
Figure FDA0002784964650000033
H＝CPT
wherein Q isaeroFor aerodynamic heat flux, QradTo radiate heat flux, H*For reference enthalpy, He，UeEnthalpy, flow velocity, H, at the outer edge of the boundary layerwIs wall enthalpy, HawIs the adiabatic wall enthalpy, r is the coefficient of restitution, H0In the form of the total enthalpy,
Figure FDA0002784964650000034
St*，
Figure FDA0002784964650000035
ρ*，η*prandtl number, Stanton number, local surface friction coefficient, local Reynolds number, density, viscosity, C, respectively, calculated at a reference enthalpyPIs the specific heat at constant pressure of air, sigma is a Stefan-Boltzmann constant, epsilon is the surface radiation coefficient, and T is the specific heat at constant pressureenvIs ambient temperature, TwThe temperature of the upper wall surface of the metal plate is protected by heat.
7. A hypersonic aircraft thermal protection configuration optimization system, comprising:
the aeroelastic model solving module is used for solving the aeroelastic model by taking the initial displacement and the initial temperature as input signals of the aeroelastic model to obtain the pressure, the Mach number and the temperature;
the aerodynamic thermal model solving module is used for solving the aerodynamic thermal model by taking the pressure, the Mach number and the temperature as input signals of the aerodynamic thermal model to obtain net heat flux;
the heat transfer model solving module is used for solving the heat transfer model by taking the net heat flux as an input signal of the heat conduction model to obtain the spatial temperature distribution of the TPS metal plate structure;
the material parameter calculation module is used for updating the temperature according to the space temperature distribution and calculating the temperature change material parameters of the TPS metal wall plate;
the thermal stress and thermal bending moment calculation module is used for calculating thermal stress and thermal bending moment according to the material parameters and the temperature distribution of the metal wall plate;
the displacement calculation module is used for calculating the displacement of the wallboard according to the thermal stress and the thermal bending moment;
and the optimization module is used for optimizing the thermal protection structure of the hypersonic aircraft through thermal-dynamic elastic coupling analysis according to the updated temperature, the thermal stress, the thermal bending moment and the wallboard displacement.
8. The optimization system of the thermal protection structure of the hypersonic aircraft according to claim 7, wherein the aerodynamic thermal model solving module specifically comprises:
the solving unit is used for solving the aerodynamic thermal model by taking the aerodynamic force, the Mach number and the temperature as input signals of the aerodynamic thermal model to obtain aerodynamic heat flux and radiant heat flux;
and the difference unit is used for making difference between the pneumatic heat flux and the radiant heat flux to obtain net heat flux.
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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113626932A (en) * 2021-08-11 2021-11-09 长春理工大学 Assembly method for eliminating array heat-proof structure paving accumulated error

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0618133A1 (en) * 1993-04-01 1994-10-05 Societe Europeenne De Propulsion Thermal protection device, particularly for spacecraft
US7281688B1 (en) * 2006-04-27 2007-10-16 The Boeing Company Materials for self-transpiring hot skins for hypersonic vehicles or reusable space vehicles
CN103970957A (en) * 2014-05-20 2014-08-06 北京航空航天大学 Simulation method for elastic waverider hypersonic flight vehicle
CN107368661A (en) * 2017-07-27 2017-11-21 中国空气动力研究与发展中心计算空气动力研究所 A kind of coupling calculation of hypersonic aircraft thermographic curve characteristic
CN107808065A (en) * 2017-11-23 2018-03-16 南京航空航天大学 The solid hot quick calculation method of 3 D complex profile high-speed aircraft stream
CN107944137A (en) * 2017-11-23 2018-04-20 南京航空航天大学 The thermographic curve computing technique of hypersonic aircraft trajectory state multi- scenarios method
CN108228921A (en) * 2016-12-13 2018-06-29 北京空天技术研究所 Flow tunnel testing device and its design method

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0618133A1 (en) * 1993-04-01 1994-10-05 Societe Europeenne De Propulsion Thermal protection device, particularly for spacecraft
US7281688B1 (en) * 2006-04-27 2007-10-16 The Boeing Company Materials for self-transpiring hot skins for hypersonic vehicles or reusable space vehicles
CN103970957A (en) * 2014-05-20 2014-08-06 北京航空航天大学 Simulation method for elastic waverider hypersonic flight vehicle
CN108228921A (en) * 2016-12-13 2018-06-29 北京空天技术研究所 Flow tunnel testing device and its design method
CN107368661A (en) * 2017-07-27 2017-11-21 中国空气动力研究与发展中心计算空气动力研究所 A kind of coupling calculation of hypersonic aircraft thermographic curve characteristic
CN107808065A (en) * 2017-11-23 2018-03-16 南京航空航天大学 The solid hot quick calculation method of 3 D complex profile high-speed aircraft stream
CN107944137A (en) * 2017-11-23 2018-04-20 南京航空航天大学 The thermographic curve computing technique of hypersonic aircraft trajectory state multi- scenarios method

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
DAN XIE等: "Effect of thermal protection system size on aerothermoelastic stability of the hypersonic panel", 《AEROSPACE SCIENCE AND TECHNOLOGY》 *

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113626932A (en) * 2021-08-11 2021-11-09 长春理工大学 Assembly method for eliminating array heat-proof structure paving accumulated error
CN113626932B (en) * 2021-08-11 2022-06-14 长春理工大学 Assembly method for eliminating array heat-proof structure paving accumulated error

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