CN116798559A - Simulation design method and preparation process of heat protection material structure and heat protection material - Google Patents
Simulation design method and preparation process of heat protection material structure and heat protection material Download PDFInfo
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- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
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- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces
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Abstract
The invention provides a simulation design method of a heat protection material structure, a preparation process and a heat protection material, wherein the simulation design method comprises the following steps: acquiring unsteady aerodynamic heat data parameters of a full flight trajectory; preliminarily determining structural parameters of the heat protection material according to the unsteady aerodynamic heat data parameters, and selecting a combination mode of the heat protection material according to the structural parameters of the heat protection material; performing simulation analysis on the combination mode of the thermal protection materials to obtain the change relation of temperature data of the thermal protection materials along with time; judging whether the change relation of the temperature data of the heat protection material along with time meets the design requirement, and if the change relation of the temperature data of the heat protection material along with time meets the design requirement, outputting the structure and the temperature data of the heat protection material. The thermal protection material designed by the simulation design method can have the characteristics of flexibility and toughness at the same time.
Description
Technical Field
The invention relates to the technical field of composite materials, in particular to a simulation design method of a thermal protection material structure, a preparation process and a thermal protection material.
Background
With the gradual improvement of the human space exploration technology, the number of engines assembled by the ultra-high sound velocity aircraft is gradually increased, and the working conditions of high temperature environment and high airflow scouring are increasingly bad. The harsh bottom environment of ultra-high sound speed aircraft places higher demands on the thermal protection materials. The layout mode of the multiple engines of the ultra-high sound speed aircraft (the number of the existing 6 engines, 9 engines, 28 engines and the subsequent engines is rapidly increased) and the swinging requirement of the engines all put forward higher requirements on the flexibility of the thermal protection material, so that the thermal protection requirement is met, and the working environment for starting flexible swinging is adapted.
The number of engines assembled on the ultra-high sound speed aircraft is gradually increased, and the high-pressure and high-speed airflow formed at the outlet of the spray pipe is increasingly serious in scouring the heat protection material, so that higher requirements are put on the toughness of the heat protection material. The flight characteristics of the ultra-high sound velocity aircraft determine that the ultra-high sound velocity aircraft is very sensitive to the weight of the heat protection material, and the weight and design mode of the heat protection material also put forward higher requirements, and the design of the existing heat protection material cannot simultaneously have the characteristics of flexibility and toughness.
Disclosure of Invention
In view of the above, an object of the embodiments of the present invention is to provide a simulation design method, a preparation process and a thermal protection material for a thermal protection material structure, so as to solve the technical problem that the design of the thermal protection material in the prior art cannot have the characteristics of flexibility and toughness at the same time.
In order to achieve the above object, in a first aspect, an embodiment of the present invention provides a method for designing a thermal protection material structure, the method comprising:
acquiring unsteady aerodynamic heat data parameters of a full flight trajectory;
preliminarily determining structural parameters of the thermal protection material according to the unsteady aerodynamic thermal data parameters, and selecting a combination mode of the thermal protection material according to the structural parameters of the thermal protection material;
performing simulation analysis on the combination mode of the thermal protection materials to obtain the change relation of temperature data of the thermal protection materials along with time;
judging whether the change relation of the temperature data of the thermal protection material along with time meets the design requirement, and outputting the structure and the temperature data of the thermal protection material if the change relation of the temperature data of the thermal protection material along with time meets the design requirement.
In some possible embodiments, the method further comprises: and if the change relation of the temperature data of the heat protection material with time is insufficient to design requirements, the structural parameters of the heat protection material are redetermined, and the combination mode of the heat protection material is redelected according to the structural parameters of the heat protection material.
In some possible embodiments, the acquiring unsteady aerodynamic heat data parameters of the full-flight trajectory specifically includes:
acquiring full flight trajectory parameters;
performing full-flight trajectory aerodynamic heat simulation calculation according to the full-flight trajectory parameters to obtain unsteady aerodynamic heat data parameters of the full-flight trajectory, wherein the unsteady aerodynamic heat data parameters comprise the temperature and heat flow of an engine spray pipe; the full flight trajectory parameters include the flight speed, the flight environment, and the in-cabin environment of the aircraft.
In some possible embodiments, the selecting a combination of the heat protection materials according to the structural parameters of the heat protection materials specifically includes:
and calling a material database, and selecting a combination mode of the heat protection materials from the material database according to the structural parameters of the heat protection materials.
In some possible embodiments, the simulation analysis includes: floefd simulation analysis, fluent simulation analysis, or Abaqus simulation analysis.
In a second aspect, an embodiment of the present invention provides a process for preparing a thermal protection material, where the process is based on a method for designing a structure of the thermal protection material, and the process includes:
preparing a mold for preparing a heat protection material;
selecting a target material according to the structure and temperature data of the heat protection material, wherein the target material comprises an ablation material, a heat expansion material and a structure reinforcing material;
spraying the ablation material on the inner surface of the die, and placing the die in an oven for curing to fix the ablation material on the inner surface of the die;
and after the heat spreading material and the structural reinforcing material are sequentially paved on the ablation material, spraying the ablation material again on the heat spreading material and the structural reinforcing material, and placing the material in a baking oven for curing to obtain the heat protection material.
In some possible embodiments, the preparation process further comprises:
judging whether the quality of the heat protection material meets the requirement or not;
if not, adjusting the type and the structural parameters of the target material;
and if so, warehousing the heat protection material qualified and performing an ablation test.
In some possible embodiments, the ablative material is a silicone rubber composite coating, the heat spreading material is a boron nitride film, and the structural reinforcement material is a wire mesh; wherein,,
the preparation method of the silicone rubber composite coating comprises the following steps:
selecting phenolic resin hollow microspheres, high silica fibers, aluminum hydroxide, auxiliary materials and silicone rubber according to the mass ratio of 1:1:1:1:6, mixing and stirring the components in proportion to obtain the silicone rubber composite coating.
In some possible embodiments, the wire mesh is fixedly connected to the aircraft arrow body in the form of a weld.
In a third aspect, an embodiment of the present invention provides a thermal protection material, which is prepared according to the preparation process, and the thermal protection material includes: a first silicon rubber composite coating layer, a second silicon rubber composite coating layer, a boron nitride film and a silk screen, wherein,
the boron nitride film and the screen are disposed between the first and second silicone rubber composite coating layers.
The beneficial technical effects of the technical scheme are as follows:
the embodiment of the invention provides a simulation design method, a preparation process and a thermal protection material of a thermal protection material structure, wherein the simulation design method comprises the following steps: acquiring unsteady aerodynamic heat data parameters of a full flight trajectory; preliminarily determining structural parameters of the heat protection material according to the unsteady aerodynamic heat data parameters, and selecting a combination mode of the heat protection material according to the structural parameters of the heat protection material; performing simulation analysis on the combination mode of the thermal protection materials to obtain the change relation of temperature data of the thermal protection materials along with time; judging whether the temperature data of the thermal protection material changes with time to meet the design requirement, and outputting the structure and the temperature data of the thermal protection material if the temperature data of the thermal protection material changes with time to meet the design requirement. According to the embodiment of the invention, the structure and the temperature data of the heat protection material are analyzed through simulation, and the combination mode of the heat protection material is selected according to the structure and the temperature data, so that the heat protection material has the characteristics of flexibility and toughness.
Drawings
In order to more clearly illustrate the embodiments of the invention or the technical solutions in the prior art, the drawings that are required in the embodiments or the description of the prior art will be briefly described, it being obvious that the drawings in the following description are only some embodiments of the invention, and that other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.
FIG. 1 is a flow chart of a method for simulating design of a thermal protection material structure according to an embodiment of the present invention;
FIG. 2 is a schematic diagram of a method for simulating design of a thermal protection material structure according to an embodiment of the present invention;
FIG. 3 is a flow chart of a process for preparing a thermal protection material according to the present invention;
FIG. 4 is a schematic diagram of a process for preparing a thermal protection material according to an embodiment of the present invention;
fig. 5 is a schematic structural diagram of a thermal protection material according to an embodiment of the present invention.
Detailed Description
Features and exemplary embodiments of various aspects of the invention are described in detail below. In the following detailed description, numerous specific details are set forth in order to provide a thorough understanding of the invention. It will be apparent, however, to one skilled in the art that the present invention may be practiced without some of these specific details. The following description of the embodiments is merely intended to provide a better understanding of the invention by showing examples of the invention. In the drawings and the following description, at least some well-known structures and techniques have not been shown in detail in order not to unnecessarily obscure the present invention; also, the dimensions of some of the structures may be exaggerated for clarity. Furthermore, the described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments.
Example 1
Fig. 1 is a flowchart of a simulation design method of a thermal protection material structure according to an embodiment of the present invention, as shown in fig. 1, the simulation design method includes the following steps:
step S11, obtaining unsteady aerodynamic heat data parameters of a full flight trajectory;
step S12, preliminarily determining structural parameters of the heat protection material according to the unsteady aerodynamic heat data parameters, and selecting a combination mode of the heat protection material according to the structural parameters of the heat protection material; in this embodiment, the unsteady aerodynamic thermal data parameters include the temperature and heat flow of the engine nozzle.
Step S13, performing simulation analysis on the combination mode of the thermal protection materials to obtain the change relation of temperature data of the thermal protection materials along with time;
in this embodiment, the aerodynamic heat data parameters determined in step S12 are used as input conditions for preliminary simulation design of the heat protection structure, for example, the design requirement that the cabin inner surface temperature is not higher than T1 ℃, and the combination mode of the heat protection materials and the thickness of each component part are selected according to the input conditions, which can be determined preliminarily according to the experience of the previous model. Then, developing detailed design, and adjusting until the detailed design is satisfied, for example, the selected heat protection materials are a combination mode of cork, honeycomb and aluminum plates, and the thickness of each heat protection material corresponds to the combination mode; and calculating the temperature of the inner surface of the heat protection material to be less than T1 ℃, namely, designing the temperature of the inner surface of the cabin to be not higher than T1 ℃ and obtaining the change relation of temperature data of the heat protection material along with time. In this embodiment, several iterations are performed through Floefd simulation analysis, fluent simulation analysis, abaqus simulation analysis, or the like according to the input conditions and the selected thermal protection materials and thicknesses thereof, if the input conditions are adjusted, the types and structures of the thermal protection materials are also adjusted, and the change relation between the temperatures of the thermal protection layer and the structural layer along with time is calculated by adopting a calculation method of convection, radiation and thermal conduction theory, so as to design the structure of the thermal protection material. As an illustration, thermal protection layer refers to thermal protection materials such as honeycomb paper, carbon skin or rubber layers, structural layers refer to aluminum skin, etc.
Step S14, judging whether the change relation of the temperature data of the heat protection material along with time meets the design requirement, and if the change relation of the temperature data of the heat protection material along with time meets the design requirement, outputting the structure and the temperature data of the heat protection material. The thermal protection material is prepared according to the time-dependent temperature data obtained by the simulation design method provided by the embodiment.
In particular, hypersonic aircraft experience the atmosphere in flight, and their surfaces can produce aerodynamic heating, causing the aircraft surface and interior temperatures to rise. When the temperature is too high, the instrumentation of the aircraft may not work properly. The key to solving the problem of 'thermal barrier' is to design a thermal protection system reasonably to ensure the normal operation of the aircraft. The aim of thermal protection of an aircraft is to ensure that the internal structure of the aircraft is within a safe temperature range, the design criteria being that the internal temperature of the structure does not exceed the permissible temperature.
The heat protection simulation design method provided by the embodiment of the invention mainly aims at:
1. calculating and analyzing the transient temperature of each structure in the interior, namely the temperature of three layers of heat protection materials, such as cork, paper honeycomb and aluminum plate;
2. calculating the internal temperature of the thermal protection structure to evaluate the performance of the thermal protection structure;
3. and combining the calculation result of the aerodynamic thermal simulation of the full flight trajectory, and developing an unsteady (considering time variation) simulation design of the thermal protection structure.
In some embodiments, the method further comprises the steps of: if the temperature data of the heat protection material does not meet the design requirement along with the change relation of time, the structural parameters of the heat protection material are redetermined, and the combination mode of the heat protection material is redetected according to the redetermined structural parameters of the heat protection material.
In this embodiment, if the combination mode of the initially designed thermal protection material and the thickness of each material layer do not meet the design requirement, the structural parameters of the thermal protection material are readjusted, and the materials in the material database are called to redesign. For example, the combination mode of "cork-honeycomb-aluminum plate" can be adjusted to "cork-glass fiber reinforced plastic-aluminum plate" or "rubber-aluminum plate" and the like to perform various combination calculations so as to find the optimal combination.
In some embodiments, acquiring unsteady aerodynamic thermal data parameters of the full flight trajectory in step S11 may specifically include:
and carrying out full-flight trajectory aerodynamic heat simulation calculation according to the full-flight trajectory parameters to obtain unsteady aerodynamic heat data parameters of the full-flight trajectory, wherein the full-flight trajectory parameters comprise the flight speed, the flight environment and the cabin environment of the aircraft.
Specifically, in this embodiment, before the simulation design is performed, the unsteady aerodynamic heat data parameters need to be calculated with the flying speed, the flying environment, and the cabin environment of the aircraft as input conditions. The flying environment refers to parameters such as flying speed, altitude, atmospheric temperature and the like, and the cabin environment refers to the environment such as pressure and temperature requirements in the aircraft.
In some embodiments, the combination of the heat protection materials is selected according to the structural parameters of the heat protection materials in step S12, specifically including:
the material database is called, and the combination mode of the heat protection materials is selected from the material database according to the structural parameters of the heat protection materials, for example, the combination mode of cork-glass fiber reinforced plastic-aluminum plate, the combination mode of rubber-aluminum plate, and the like. Specifically, a database of thermal protection materials is firstly established, and the existing data in the database can be called for the optimal design of thermal protection according to the adjustment of input conditions.
In order to enable those skilled in the art to better understand the technical scheme provided by the embodiment of the invention, a simulation design method of the heat protection material structure provided by the embodiment of the invention is described in detail below. Fig. 2 is a schematic diagram of a method for simulating design of a thermal protection material structure according to an embodiment of the present invention, and as shown in fig. 2, the method includes two stages: in the simulation preparation stage, aerodynamic heat calculation is required to be carried out according to flight trajectory parameters, and the flight trajectory parameters refer to the flight speed, the flight environment and the cabin environment of the aircraft, so that all-flight trajectory aerodynamic heat unsteady calculation data are obtained.
In the simulation design stage, data in a material database is loaded according to all-flight trajectory aerodynamic heat unsteady calculation data (namely unsteady aerodynamic heat data parameters) calculated in the simulation preparation stage, a combination mode of a heat protection material and the thickness of each material layer are initially designed, temperature data of the heat protection material, namely the change relation of temperature along with time, is obtained through simulation analysis, whether the design requirement is met is judged, if the design requirement is not met, the structural parameters of the heat protection material are continuously adjusted, the material database is called to reselect the combination mode of the heat protection material and the thickness of each material layer, and iterative adjustment is repeated until the designed heat protection material meets the design requirement, and finally the structure and the temperature data of the heat protection material are output.
The hypersonic aircraft anisotropic ablation heat protection material structure and temperature data with the characteristics of flexibility and toughness provided by the embodiment of the invention enable the protection material to meet the use condition of 400-2000 ℃, and the flexibility of the hypersonic aircraft anisotropic ablation heat protection material can enable the hypersonic aircraft anisotropic ablation heat protection material to be applied to large-size and complex appearance and meet the structural movement requirement.
Example two
Fig. 3 is a flowchart of a process for preparing a thermal protection material according to an embodiment of the present invention, which is based on the simulation design method of the thermal protection material structure according to the first embodiment, as shown in fig. 3, and includes the following steps:
s21, preparing a mold for preparing a heat protection material;
after the mold for preparing the heat protection material is prepared, the surface of the mold can be cleaned by adopting ethanol to prevent pollution.
S22, selecting materials according to the structure and temperature data of the heat protection material, wherein the materials comprise ablation materials, heat expansion materials and structure reinforcing materials;
s23, spraying an ablative material on the inner surface of the die, and placing the die in an oven for curing to fix the ablative material on the inner surface of the die; in this embodiment, the high temperature baking temperature may be 160℃and the time may be 6 hours.
S24, after the heat spreading material and the structural reinforcing material are sequentially paved on the ablation material, spraying the ablation material again on the heat spreading material and the structural reinforcing material, and curing in a baking oven to obtain the heat protection material.
In the embodiment of the invention, the heat-spreading material and the structural reinforcing material are coated inside the ablation material.
Table 1 shows the main components and effects of the heat protection material according to the embodiment of the present invention, as shown in table 1, any combination of a wire mesh, rubber, boron nitride, phenolic resin hollow beads, high silica fiber and aluminum hydroxide may be used, and of course, other alternative components may be used, for example, the wire mesh may be used to increase toughness, carbon fiber may be used to replace, the boron nitride may be used to insulate heat by heat expansion, and graphene may be used to replace. May be according to the actual situation.
TABLE 1
| Sequence number | Component (A) | Action | Remarks |
| 1 | Silk screen | Increasing toughness | Carbon fibers may be used instead of |
| 2 | Rubber material | Heat absorption by pyrolysis | |
| 3 | Boron nitride | Heat-expansion heat-insulation | Graphene replacement may be employed |
| 4 | Phenolic resin hollow micro-bead | Thermal insulation | |
| 5 | High silica fiber | Heat-insulating flame-retardant material | |
| 6 | Aluminum hydroxide | Heat absorption |
In some embodiments, the preparation process further comprises: judging whether the quality of the heat protection material meets the requirement; if not, adjusting the types and the structural parameters of the selected materials; and if so, carrying out qualified warehousing treatment and ablation test on the prepared thermal protection material. Here, it is determined whether the quality of the thermal protection material satisfies the requirement, or whether the designed thermal protection material satisfies the thermal protection material target formulated at the initial stage of the design, which is formulated according to the model.
In some embodiments, the ablative material is a silicone rubber composite coating, the heat spreading material is a boron nitride film, and the structural reinforcement material is a wire mesh; the preparation method of the silicone rubber composite coating comprises the following steps: selecting phenolic resin hollow microspheres, high silica fibers, aluminum hydroxide, auxiliary materials and silicone rubber according to the mass ratio of 1:1:1:1:6, mixing and stirring the components in proportion to obtain the silicone rubber composite coating. In the embodiment of the invention, the mass ratio of the components is 1:1:1:1:6, mixing and stirring are carried out according to the proportion, the heat protection requirement can be met, and flowing can not occur after the heating.
Specifically, the silicone rubber composite coating in this embodiment is sprayed by a spraying process. Firstly, spraying an outer layer high-temperature-resistant flexible silicon rubber composite material on the outer surface of a die, and then placing the die in a high-temperature baking oven for curing. Then, a wire mesh is laid on the outer surface of the silicon rubber composite material, and a boron nitride film is laid. And finally, spraying the high-temperature-resistant flexible silicon rubber composite material again on the upper surface of the paved silk screen and the boron nitride film, and then placing the silk screen and the boron nitride film in a high-temperature baking oven for curing. In addition, the wire mesh can be made of stainless steel.
In some embodiments, the wire mesh is fixedly connected to the aircraft arrow body in the form of a weld. According to the embodiment of the invention, the strength of the connection between the thermal protection material and the aircraft rocket body is improved by welding the silk screen and the aircraft, and the weight of the material is not obviously increased due to the hollowed silk screen.
In order to enable those skilled in the art to better understand the technical scheme provided by the embodiment of the invention, a preparation process of the heat protection material provided by the embodiment of the invention is described in detail below. Fig. 4 is a schematic diagram of a process for preparing a thermal protection material according to an embodiment of the present invention, and as shown in fig. 4, the process includes three stages: a production preparation stage, a preparation stage and a quality inspection stage;
selecting materials according to the structure and temperature data of the heat protection material in the production preparation stage, wherein the materials comprise an ablation material, a heat expansion material and a structure reinforcing material, and in the embodiment, the ablation material is phenolic resin hollow microspheres, high silica fibers, aluminum hydroxide, an auxiliary material and silicon rubber, the heat expansion material is a boron nitride film, and the structure reinforcing material is a silk screen;
in the preparation stage, firstly, phenolic resin hollow microspheres, high silica fibers, aluminum hydroxide, auxiliary materials and silicone rubber are mixed according to the mass ratio of 1:1:1:1:6, spraying a silica gel composite material on the bottom layer of the die, then sequentially paving a boron nitride film and a steel wire, wherein the paving sequence can be adjusted, for example, firstly paving a silk screen, then paving a boron nitride film, then spraying the silica gel composite material again, and then placing the coated material into an oven for baking, wherein the baking temperature is 160 ℃ generally, and the time is preferably 6 hours, so that the flexible protective material is obtained;
in the quality inspection stage, the flexible heat protection material prepared in the preparation stage is subjected to quality inspection, and if the quality is not satisfied, the type of the material prepared in the production preparation stage and the thickness of the laid material are readjusted, and the flexible heat protection material is prepared again. And if the quality meets the requirements, the materials are qualified and put in storage, and an ablation test of the full-flight ballistic rail thermal protection material is prepared.
The hypersonic aircraft anisotropic ablation heat protection material production process with the characteristics of flexibility and toughness provided by the embodiment of the invention has the characteristics of simplicity in operation, low cost and wide application range.
In addition, the special working environment characteristics of high-temperature, high-pressure and high-speed air flow flushing for the multi-jet pipe work of the ultra-high sound speed aircraft are effectively designed, and the working requirements of the thermal protection material at 400-2000 ℃ are realized through the selection and the proportioning of the preparation process formula materials;
meanwhile, the structure greatly improves the toughness of the material by arranging the silk screen layer on the premise of meeting the ablation heat protection requirement, and can be very well suitable for the characteristics of flushing and flexibility of high-temperature, high-pressure and high-speed air flow and meet the requirement of engine swing.
Example III
Fig. 5 is a schematic structural view of a thermal protection material according to an embodiment of the present invention, and as shown in fig. 5, the thermal protection material is disposed at a tail section of an arrow body, and includes: the coating comprises a first silicon rubber composite coating layer, a second silicon rubber composite coating layer, a boron nitride film and a silk screen, wherein the boron nitride film and the silk screen are arranged between the first silicon rubber composite coating layer and the second silicon rubber composite coating layer. Namely, the boron nitride film and the silk screen are wrapped inside the silicon rubber composite coating layer.
According to the heat protection material provided by the embodiment of the invention, the boron nitride film layer is arranged, so that on the premise of meeting the ablation heat protection requirement, according to the characteristic of an unsteady aerodynamic heat calculation result, namely, the heat of the outer surface of the heat protection material is unevenly loaded to the heat protection material, the high thermal conductivity 2400W/(mK) of boron nitride is fully utilized for redistributing the heat in the material, the effective utilization rate of the heat protection material is greatly improved by the increase of the material, and the thickness and the quality of the heat protection material are reduced. By arranging the silk screen layer, the toughness of the material is greatly improved on the premise of meeting the ablation heat protection requirement, and the material can be excellently suitable for scouring of high-temperature, high-pressure and high-speed air flow; meanwhile, excessive design allowance of the heat protection material caused by adopting the design method of the least adverse working condition is avoided.
In the description of the embodiments of the present invention, it should be noted that the orientation or positional relationship indicated by "upper, lower, inner and outer", etc. in terms are based on the orientation or positional relationship shown in the drawings, and are merely for convenience of describing the present invention and simplifying the description, rather than indicating or suggesting that the device or element in question must have a specific orientation, be configured and operated in a specific orientation, and thus should not be construed as limiting the present invention. Furthermore, the terms "first, second, or third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
The terms "mounted, connected, and coupled" in embodiments of the invention are to be construed broadly, unless otherwise specifically indicated and defined, for example: can be fixed connection, detachable connection or integral connection; it may also be a mechanical connection, an electrical connection, or a direct connection, or may be indirectly connected through an intermediate medium, or may be a communication between two elements. The specific meaning of the above terms in the present invention will be understood in specific cases by those of ordinary skill in the art.
While the invention has been described with reference to a preferred embodiment, various modifications may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In particular, the technical features mentioned in the respective embodiments may be combined in any manner as long as there is no structural conflict. The present invention is not limited to the specific embodiments disclosed herein, but encompasses all technical solutions falling within the scope of the claims.
Claims (10)
1. The simulation design method of the heat protection material structure is characterized by comprising the following steps of:
acquiring unsteady aerodynamic heat data parameters of a full flight trajectory;
preliminarily determining structural parameters of the thermal protection material according to the unsteady aerodynamic thermal data parameters, and selecting a combination mode of the thermal protection material according to the structural parameters of the thermal protection material;
performing simulation analysis on the combination mode of the thermal protection materials to obtain the change relation of temperature data of the thermal protection materials along with time;
judging whether the change relation of the temperature data of the thermal protection material along with time meets the design requirement, and outputting the structure and the temperature data of the thermal protection material if the change relation of the temperature data of the thermal protection material along with time meets the design requirement.
2. The simulated design method of claim 1, further comprising:
and if the temperature data of the heat protection material is not enough in design requirements along with the change relation of time, the structural parameters of the heat protection material are redetermined, and the combination mode of the heat protection material is redetected according to the redetermined structural parameters of the heat protection material.
3. The simulation design method according to claim 1, wherein the obtaining the unsteady aerodynamic heat data parameter of the full flight trajectory specifically comprises:
acquiring full flight trajectory parameters;
carrying out full-flight trajectory aerodynamic heat simulation calculation according to full-flight trajectory parameters to obtain unsteady aerodynamic heat data parameters of the full-flight trajectory, wherein the unsteady aerodynamic heat data parameters comprise the temperature and heat flow of an engine spray pipe; the full flight trajectory parameters include the flight speed, the flight environment, and the in-cabin environment of the aircraft.
4. The method for simulating design according to claim 1, wherein the selecting a combination of thermal protection materials according to the structural parameters of the thermal protection materials specifically comprises:
and calling a material database, and selecting a combination mode of the heat protection materials from the material database according to the structural parameters of the heat protection materials.
5. The simulated design method of claim 1, wherein the simulated analysis comprises: floefd simulation analysis, fluent simulation analysis, or Abaqus simulation analysis.
6. A process for preparing a thermal protection material based on the simulation design method of the thermal protection material structure according to any one of claims 1 to 5, characterized in that the process comprises:
preparing a mold for preparing a heat protection material;
selecting a target material according to the structure and temperature data of the thermal protection material, wherein the target material comprises an ablation material, a heat expansion material and a structure reinforcing material;
spraying the ablation material on the inner surface of the die, and placing the die in an oven for curing to fix the ablation material on the inner surface of the die;
and after the heat spreading material and the structural reinforcing material are sequentially paved on the ablation material, spraying the ablation material again on the heat spreading material and the structural reinforcing material, and placing the material in a baking oven for curing to obtain the heat protection material.
7. The manufacturing process according to claim 6, further comprising:
judging whether the quality of the heat protection material meets the requirement or not;
if not, adjusting the type and the structural parameters of the target material;
and if so, warehousing the heat protection material qualified and performing an ablation test.
8. The process of claim 6, wherein the ablative material is a silicone rubber composite coating, the heat spreading material is a boron nitride film, and the structural reinforcement material is a wire mesh; wherein,,
the preparation method of the silicone rubber composite coating comprises the following steps:
selecting phenolic resin hollow microspheres, high silica fibers, aluminum hydroxide, auxiliary materials and silicone rubber according to the mass ratio of 1:1:1:1:6, mixing and stirring the components in proportion to obtain the silicone rubber composite coating.
9. The manufacturing process according to claim 8, characterized in that the wire mesh is fixedly connected to the aircraft arrow body in the form of a weld.
10. A thermal protection material prepared according to the preparation process of claim 9, comprising: a first silicon rubber composite coating layer, a second silicon rubber composite coating layer, a boron nitride film and a silk screen, wherein,
the boron nitride film and the screen are disposed between the first and second silicone rubber composite coating layers.
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