CN212046202U - Phase-change heat-insulation composite thermal protection structure of hypersonic aircraft - Google Patents
Phase-change heat-insulation composite thermal protection structure of hypersonic aircraft Download PDFInfo
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- CN212046202U CN212046202U CN202020352859.XU CN202020352859U CN212046202U CN 212046202 U CN212046202 U CN 212046202U CN 202020352859 U CN202020352859 U CN 202020352859U CN 212046202 U CN212046202 U CN 212046202U
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Abstract
The embodiment of the utility model provides a hypersonic aircraft phase transition heat insulation composite thermal protection structure, which comprises a surface layer, a connecting layer, a heat insulation layer, an aerogel layer and an aircraft body; the connecting layer is arranged on one side of the surface layer, a heat insulation layer is arranged on one side of the connecting layer away from the surface layer, an aerogel layer is arranged on one side of the heat insulation layer away from the connecting layer, and an aircraft body is arranged on one side of the aerogel layer away from the heat insulation layer; the insulation layer includes a first plate, a second plate, and a honeycomb body disposed between the first plate and the second plate. The embodiment of the utility model provides a hypersonic aircraft phase transition compound thermal protection structure that insulates against heat helps improving thermal-insulated efficiency.
Description
Technical Field
The utility model belongs to the hot protective structure field of hypersonic aircraft multilayer, concretely relates to compound hot protective structure of hypersonic aircraft phase transition thermal-insulated.
Background
The hypersonic aircraft is a winged or wingless aircraft with flight speed more than 5 times of sonic speed, such as airplanes, missiles, shells and the like, has the characteristic of high penetration success rate, and has great military value and potential economic value. With the development of aerospace technology, the flight speed of the aerospace craft is continuously improved, the service environment is more and more severe, and the effective thermal protection system can provide enough protection for the aerospace craft when the aerospace craft faces severe pneumatic heating, so that the aerospace craft can be prevented from being damaged by severe pneumatic thermal environment and can keep safe flight for a longer time. The reliable thermal protection system is one of the key systems for safe flight of high-performance aircraft, and the design of the thermal protection structure and the selection of thermal protection materials are the key for the design and development of the thermal protection system. With the continuous increase of the flight speed of the aircraft, the thermal protection problem of the aircraft becomes a bottleneck limiting the development of the aircraft.
Currently, multilayer thermal protection structures are an important direction for the development of thermal protection systems. The multilayer thermal protection structure can greatly improve the heat insulation prevention efficiency of the thermal protection system, reduce the structural quality of the thermal protection system, increase the effective load of the aircraft and effectively reduce the manufacturing and maintenance cost of the aircraft. The requirement on the thermal protection structure is very strict, and the application requirement can not be well met by only utilizing the thermal insulation material.
SUMMERY OF THE UTILITY MODEL
An object of the utility model is to provide a hypersonic aircraft phase transition compound thermal protection structure that insulates against heat helps improving thermal-insulated efficiency.
The utility model discloses the technical scheme who adopts does:
the utility model provides a hypersonic aircraft phase-change heat insulation composite thermal protection structure, wherein the aircraft comprises an aircraft body, and the thermal protection structure comprises a surface layer, a connecting layer, a heat insulation layer and an aerogel layer;
an aerogel layer is arranged on the outer side of the aircraft body, a thermal insulation layer is arranged on one side, away from the aircraft body, of the aerogel layer, a connection layer is arranged on one side, away from the aerogel layer, of the thermal insulation layer, and the surface layer is arranged on one side, away from the thermal insulation layer, of the connection layer;
the insulation layer includes a first plate, a second plate, and a honeycomb body disposed between the first plate and the second plate.
Preferably, the honeycomb body comprises a plurality of honeycomb units, each honeycomb unit is formed by mutually enclosing five honeycomb sheets, and a honeycomb cavity is formed in each honeycomb unit.
Preferably, the honeycomb cavity is filled with a phase change material;
the first plate and the honeycomb body, and the second plate and the honeycomb body are welded to each other.
Preferably, the connecting layer is used for connecting the surface layer and the heat insulation layer.
Preferably, the surface layer is made of a C/C composite material, and/or the connecting layer is made of nano phosphate glue.
Preferably, the phase-change material is prepared from a high-temperature phase-change heat-insulating material by a preparation method disclosed in CN 104591767A.
Preferably, the aerogel layer is made of SiC aerogel material and can resist 1400 ℃.
Preferably, the thickness of the surface layer is 5-8 mm, the thickness of the connecting layer is 2-3 mm, the thickness of the thermal insulation layer is 30-35 mm, the thickness of the aerogel layer can be 3-4mm, and the thickness of the aircraft fuselage is 8-12 mm.
Preferably, each adjacent two honeycomb units contact with each other to share one honeycomb sheet.
The beneficial effects of the utility model reside in that:
1. the utility model provides a hypersonic aircraft phase transition compound thermal protection structure that insulates against heat helps improving thermal-insulated efficiency.
2. The embodiment of the utility model provides a help improving thermal protection system's thermal-insulated efficiency of preventing, promote thermal protection system's available temperature range, reduce thermal protection system's density and structure weight, effectively reduce the manufacturing and the maintenance cost of aircraft.
3. The embodiment of the utility model provides an adopt metal honeycomb structure as the skeleton of insulating layer. The framework can provide certain supporting bearing for the heat protection structure, and can provide a fixed position space for the phase change material.
4. The phase-change material is filled in the honeycomb cavity in the metal honeycomb structure of the heat insulation layer, and the material absorbs a large amount of heat during phase change, so that the heat insulation performance of the heat insulation layer can be greatly improved.
5. After the phase change insulation by the phase change material, the insulation can be further insulated by the aerogel layer, so that only a small portion of the heat finally reaches the aircraft surface.
6. In this example, SiC aerogel materials (i.e., the ultra-light carbide ceramic foam disclosed in CN 110066175 a) were used, which have a large room for improvement in gas phase thermal barrier and long thermal conduction path due to their low density and large open-cell structure. The SiC aerogel material (namely the ultra-light carbide ceramic foam disclosed by CN 110066175A) has the characteristics of light weight, high specific strength and capability of resisting the high temperature of 1400 ℃.
7. The embodiment of the utility model utilizes the phase change of the material to absorb excessive heat, thereby preventing or delaying the heat from transmitting into the interior of the aircraft and ensuring that the control instrument works normally; and the aerogel layer is arranged on one side of the thermal insulation layer, so that the thermal insulation can be further realized, and the temperature finally reaching the aircraft aluminum alloy aircraft fuselage meets the requirement range of the aluminum alloy material.
Drawings
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 description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are only examples of the present invention, and for those skilled in the art, other drawings can be obtained according to the provided drawings without creative efforts.
Fig. 1 is a schematic structural view of a phase-change thermal insulation composite thermal protection junction of a hypersonic aircraft according to an embodiment of the present invention;
FIG. 2 is a schematic view of a three-dimensional structure of a metal honeycomb panel according to an embodiment of the present invention;
FIG. 3 is a schematic view of a partial structure of a metal honeycomb panel according to an embodiment of the present invention; (omitting the first plate)
Fig. 4 is a schematic diagram of partial discharge at a in fig. 3.
Description of the drawings:
1, surface layer; 2 a connecting layer; 3, a heat insulation layer; 31 a first plate; 32 a second plate; 33 a honeycomb body; 331 a honeycomb sheet; 332 honeycomb cavities; 4 an aerogel layer; 5 aircraft fuselage.
Detailed Description
The conception, specific structure, and technical effects of the present invention will be described clearly and completely with reference to the following embodiments, so that the objects, features, and effects of the present invention can be fully understood. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all embodiments, and other embodiments obtained by those skilled in the art without inventive labor based on the embodiments of the present invention all belong to the protection scope of the present invention.
Example 1
The embodiment 1 of the utility model provides a hypersonic aircraft phase transition thermal insulation composite thermal protection structure, the aircraft comprises an aircraft body 5, and the thermal protection structure comprises a surface layer 1, a connecting layer 2, a thermal insulation layer 3 and an aerogel layer 4;
an aerogel layer 4 is arranged on the outer side of the aircraft body 5, a thermal insulation layer 3 is arranged on one side, away from the aircraft body 5, of the aerogel layer 4, a connecting layer 2 is arranged on one side, away from the aerogel layer 4, of the thermal insulation layer 3, and the surface layer 1 is arranged on one side, away from the thermal insulation layer 3, of the connecting layer 2;
the thermal insulation layer 3 comprises a first plate 31, a second plate 32 and a honeycomb body 33 arranged between the first plate 31 and the second plate 32.
Further, the thickness of insulating layer 3 is 30 ~ 35mm, insulating layer 3 whole thickness preferred is 30 mm.
Referring to fig. 2 to 4, the honeycomb body 33 includes a plurality of honeycomb units, each of which is formed by five honeycomb sheets 331 enclosed with each other, and a honeycomb cavity 332 is formed inside the honeycomb unit. Optionally, the shape of the top view of the honeycomb unit is a regular pentagon.
The heat insulating layer 3 adopts a plurality of metal honeycomb units and two metal plates (a first plate 31 and a second plate 32) as frameworks to form a metal honeycomb framework.
In this embodiment, each adjacent two honeycomb units contact each other to share one honeycomb sheet 331.
Further, the honeycomb cavity 332 is filled with a phase change material. The filling is preferably a high-temperature phase-change heat-insulating material.
Specifically, CN104591767A (patent No. 201510020733.6) discloses a method for preparing a high-temperature phase-change heat-insulating material; the high-temperature phase change heat insulation material of the embodiment is prepared by the preparation method disclosed by the CN 104591767A.
In the embodiment, the phase change temperature of the high-temperature phase change thermal insulation material is 1200-1300 ℃, the highest temperature resistance of the matrix reaches 1600 ℃, the preparation production time is short, the process is simple, the cost is low, and the large-scale industrial production is easy to realize.
The first plate 31 and the honeycomb body 33, and the second plate 32 and the honeycomb body 33 are welded to each other. Alternatively, the first plate 31, the second plate 32, and the honeycomb body 33 are made of nickel-based alloy.
In the embodiment, the surface layer 1 is made of a C/C composite material, and the thickness of the surface layer 1 can be 5-8 mm; the preferable thickness is 7mm, the heat conductivity coefficient is high, the specific heat capacity is large, and the material cannot be melted at high temperature.
The connecting layer 2 is made of nano phosphate glue, and the thickness of the connecting layer 2 can be 2-3 mm. The connecting layer 2 is used for connecting the surface layer 1 and the heat insulation layer 3, and the connecting layer 2 is preferably 3mm thick; the nano phosphate adhesive has high refractoriness, good thermal shock stability, good impermeability and strong impact resistance, higher load softening temperature and chemical stability, strong high-temperature bonding force and the like.
In this embodiment, the aerogel layer 4 is made of SiC aerogel material, and can resist 1400 ℃. The aerogel layer 4 may have a thickness of 3-4mm, preferably 3 mm.
Specifically, CN 110066175A (patent No. 201910391861X) discloses a method for preparing an ultra-light carbide ceramic foam; the SiC aerogel material of this embodiment is prepared by the preparation method disclosed in CN 110066175 a, and is the ultra-light carbide ceramic foam prepared by the method.
In the embodiment, the SiC aerogel material has the advantages of light weight, high specific strength, good processability and the like. The SiC aerogel material has the advantages of heat resistance temperature as high as 1400 ℃, wide and cheap raw material source, simple manufacturing process and environmental protection.
The thickness of the aircraft fuselage 5 is 8-12mm, and the thickness is preferably 8 mm; the aircraft fuselage 5 is made of an aluminum alloy material.
In the embodiment of the utility model, the aircraft body 5 is a structural metal, is made of an aluminum alloy material, has the thickness of 8mm, high strength, good plasticity, excellent electrical conductivity, thermal conductivity and corrosion resistance, and good heat resistance; the aircraft fuselage 5 is made of an aluminium alloy and has a maximum design temperature of 176 ℃.
In this embodiment, the surface layer 1, the connecting layer 2, the thermal insulation layer 3, the aerogel layer 4, and the aircraft fuselage 5 are arranged in this order along the aircraft from the outside to the inside.
When the aircraft is flying at a hypersonic velocity, the surfaces generate a lot of heat due to aerodynamic heating. The surface layer 1 is made of C/C composite materials and can resist high temperature; when heat is transferred to the connecting layer 2, the temperature reaching the insulating layer 3 is greatly reduced due to the insulating ability of the aerogel. After the phase change and heat absorption of the metal honeycomb framework (the first plate 31, the second plate 32 and the honeycomb body 3332) and the phase change material, the aerogel layer 4 is used for further insulating heat, so that the temperature finally reaching the aircraft fuselage 5 is only about 170 ℃, and the service temperature range of the high-strength aluminum alloy adopted on the surface of the aircraft is completely met.
Characteristics of phase change material
When the spacecraft runs on the orbit, the heat load of instrument equipment is changed greatly due to the change of the external heat flow of the orbit. Particularly, the working temperature range of some special equipment with smaller heat capacity is narrower and narrower, and meanwhile, the requirement on temperature fluctuation is high, which brings a plurality of technical problems to the thermal control design of the spacecraft.
The phase-change material has the advantages of isothermy or approximately isothermy and absorption/release of a large amount of latent heat in the phase-change process, so that the phase-change material is particularly suitable for instruments and equipment with periodic pulse type work; the phase-change material has the other characteristic that no moving part exists, the reversible work can be performed for infinite times in principle, and the reliability is high.
Isolating the controlled device from the external environment by using a phase-change material, wherein when the temperature of a contact interface between the controlled device and the external environment is increased to the melting point of the phase-change material, the phase-change material is melted and absorbs heat equivalent to latent heat of melting, so that the temperature of the interface is kept near the melting point temperature; when the interface temperature drops due to internal or external causes, the phase change material solidifies and releases latent heat, maintaining the interface temperature substantially constant.
Secondly, characteristics of aerogel materials
The aerogel is a highly dispersed solid material which is formed by mutually agglomerating colloidal particles or high polymer molecules to form a nano porous network structure and filling gaseous dispersion media into pores. The aerogel is a super heat insulation material with the most potential due to the blocking effect of the air as a main heat insulation medium and the unique microstructure thereof on gas phase heat transfer and solid phase heat transfer.
The SiC aerogel material (ultra-light carbide ceramic foam) of CN 110066175A has the characteristics of light weight, high specific strength (basically no need of second phase reinforcement), simple process, good processability (having good processing potential of special-shaped pieces), and capability of resisting the high temperature of 1400 ℃. The controllable preparation can be realized within the range of 9-20 mg cm < -3 >, the compression strength is enhanced along with the increase of the density, the range is 0.085 MPa-0.35 MPa, the yield deformation rate is 10% -15%, the compression modulus is increased along with the increase of the density, and meanwhile, the deformation rebound rate is 4%. The porosity is more than 99%, the material has a 3D network structure and extremely low thermal conductivity (less than 0.03W/m.K).
It should be noted that, throughout the specification, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.
The principles and embodiments of the present invention have been explained herein using specific examples, which are presented only to assist in understanding the methods and their core concepts. It should be noted that there are infinite trial ways due to the limited character expression, and those skilled in the art can make some improvements, decorations or changes without departing from the principle of the present invention, and can also combine the above technical features in a proper way; the application of these modifications, variations or combinations, or the application of the concepts and solutions of the present invention in other contexts without modification, is not intended to be considered as a limitation of the present invention.
Claims (9)
1. A phase-change thermal insulation composite thermal protection structure of a hypersonic aircraft comprises an aircraft body (5), and is characterized in that the thermal protection structure comprises a surface layer (1), a connecting layer (2), a thermal insulation layer (3) and an aerogel layer (4);
an aerogel layer (4) is arranged on the outer side of the aircraft body (5), a heat insulation layer (3) is arranged on one side, away from the aircraft body (5), of the aerogel layer (4), a connecting layer (2) is arranged on one side, away from the aerogel layer (4), of the heat insulation layer (3), and the surface layer (1) is arranged on one side, away from the heat insulation layer (3), of the connecting layer (2);
the thermal insulation layer (3) comprises a first plate (31), a second plate (32) and a honeycomb body (33) arranged between the first plate (31) and the second plate (32).
2. The thermal protection structure according to claim 1, characterized in that said honeycomb body (33) comprises a plurality of honeycomb cells, each of said honeycomb cells being formed by five honeycomb sheets (331) mutually enclosing and having honeycomb cavities (332) formed therein.
3. The thermal protection structure according to claim 2, characterized in that said honeycomb cavities (332) are filled with a phase-change material;
the first plate (31) and the honeycomb body (33) and the second plate (32) and the honeycomb body (33) are welded to each other.
4. Thermal protection structure according to claim 1, characterized in that said connection layer (2) is intended to connect said skin layer (1) and said insulating layer (3).
5. The thermal protection structure according to claim 1, characterized in that said skin layer (1) is a C/C composite; and/or the presence of a gas in the gas,
the connecting layer (2) is made of nano phosphate glue.
6. The thermal protection structure of claim 3, wherein the phase-change material is made of a high-temperature phase-change heat-insulating material by a preparation method disclosed in CN 104591767A.
7. The thermal protection structure according to claim 1, characterized in that said aerogel layer (4) is of SiC aerogel material, able to withstand 1400 ℃.
8. Thermal protection structure according to claim 1, characterized in that said skin layer (1) has a thickness of 5-8 mm, said connecting layer (2) has a thickness of 2-3 mm, said insulating layer (3) has a thickness of 30-35 mm, said aerogel layer (4) may have a thickness of 3-4mm, and said aircraft fuselage (5) has a thickness of 8-12 mm.
9. The thermal protection structure of claim 2, wherein each adjacent two honeycomb units contact each other to share a honeycomb sheet (331).
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113955077A (en) * | 2021-11-04 | 2022-01-21 | 江苏大学 | Super high sound speed aircraft head surface heat insulation structure |
| CN114180026A (en) * | 2021-12-28 | 2022-03-15 | 中南大学 | Dredging phase change composite flexible thermal protection structure and application thereof in deformable aircraft |
| CN115230284A (en) * | 2022-08-11 | 2022-10-25 | 北京理工大学 | Codeable patterned heat-insulation/bearing/broadband stealth multifunctional integrated structure |
-
2020
- 2020-03-19 CN CN202020352859.XU patent/CN212046202U/en not_active Expired - Fee Related
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113955077A (en) * | 2021-11-04 | 2022-01-21 | 江苏大学 | Super high sound speed aircraft head surface heat insulation structure |
| CN113955077B (en) * | 2021-11-04 | 2024-03-19 | 江苏大学 | Head surface heat insulation structure of hypersonic aircraft |
| CN114180026A (en) * | 2021-12-28 | 2022-03-15 | 中南大学 | Dredging phase change composite flexible thermal protection structure and application thereof in deformable aircraft |
| CN114180026B (en) * | 2021-12-28 | 2023-12-01 | 中南大学 | Composite flexible heat protection structure for dredging phase change and application of composite flexible heat protection structure in deformable aircraft |
| CN115230284A (en) * | 2022-08-11 | 2022-10-25 | 北京理工大学 | Codeable patterned heat-insulation/bearing/broadband stealth multifunctional integrated structure |
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Granted publication date: 20201201 |