CN109367758B - Thermal protection assembly and thermal protection system - Google Patents
Thermal protection assembly and thermal protection system Download PDFInfo
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- CN109367758B CN109367758B CN201811427161.3A CN201811427161A CN109367758B CN 109367758 B CN109367758 B CN 109367758B CN 201811427161 A CN201811427161 A CN 201811427161A CN 109367758 B CN109367758 B CN 109367758B
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C1/40—Sound or heat insulation, e.g. using insulation blankets
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Abstract
The invention relates to a thermal protection assembly and a thermal protection system, wherein the thermal protection assembly comprises a cover plate (111), a cover plate support (112) arranged on the cover plate (111), and a heat insulation structure (113) connected with the cover plate support (112); the cover plate (111) is a ceramic matrix composite plate, and an anti-oxidation coating is arranged on one side of the cover plate (111) far away from the cover plate support (112). The thermal protection assembly has the advantages of reusability, modular structure, light weight, low cost and easy replacement, has wide application prospect in the field of hypersonic aircrafts, and meets the strict requirements of the aircrafts on heat insulation, light weight, bearing, impact resistance and the like. Further, the invention has wide application range and higher applicability.
Description
Technical Field
The invention relates to the field of design of thermal protection systems, in particular to a thermal protection assembly and a thermal protection system.
Background
The surface of the hypersonic aircraft is subjected to a severe pneumatic heating effect, and a Thermal Protection System (TPS) has the main function of solving the pneumatic heating problem in the hypersonic cruise or reentry return process, so that a bottom layer main body structure is maintained in an allowed temperature range, and the normal working environment of mechanisms, electric devices and control devices in a cabin is ensured. The reusable aircraft is subjected to high temperature of more than l670 ℃ in the processes of emission and return in the flying process, so that the noise of up to 167dB is isolated, the reusable aircraft is resistant to the washing of various particles in the atmosphere, and the radiation of various rays and the like is resisted. Therefore, the development and optimized design of lightweight, durable, easy to operate, reusable thermal protection systems is critical to the success and failure of advanced hypersonic aircraft and weapons systems.
The metal thermal protection system researched at present has the advantages of high reliability, good toughness and the like, has a certain bearing capacity while having a heat-proof function, but is limited by the performance of metal materials, and the metal TPS is generally only applied to an area with the use temperature of about 980 ℃. For hypersonic aircraft, however, it is often desirable to employ thermal protection systems in the higher temperature range of use. In addition, the metal thermal protection system adopts a high-temperature alloy-based composite structure, so that the system quality is obviously improved, and the problems of unfavorable gap sealing and thermal matching design due to large thermal expansion coefficient exist.
Disclosure of Invention
The invention aims to provide a thermal protection assembly and a thermal protection system, and solves the problems that the application temperature range of the existing thermal protection system is not high enough and the thermal expansion is serious.
In order to achieve the above object, the present invention provides a thermal protection assembly, which includes a cover plate, a cover plate support disposed on the cover plate, and a thermal insulation structure connected to the cover plate support;
the cover plate is a ceramic matrix composite plate, and an anti-oxidation coating is arranged on one side of the cover plate, which is far away from the cover plate support.
According to one aspect of the invention, the deck support comprises a support portion and a connection portion;
the supporting part is an annular surrounding wall, one end of the supporting part is perpendicular to the connecting part, and the other end of the supporting part is fixedly connected with the cover plate to form a groove structure.
According to one aspect of the invention, the cover plate is a trapezoidal plate having a first edge and a second edge parallel to each other;
the second edge extends to the direction far away from the supporting part to form a first convex eaves structure;
the first edge overlaps the end face of the support part to form a step structure.
According to one aspect of the invention, the connecting portion is an annular plate-like body;
the supporting part is fixedly connected with the annular edge at the inner side of the connecting part in a mutually perpendicular mode.
According to one aspect of the invention, the insulation structure comprises: the radiation shielding layer comprises a first heat insulation layer, a second heat insulation layer, a glass fiber layer and a radiation shielding layer;
the first heat insulation layer, the second heat insulation layer, the glass fiber layer and the radiation shielding layer are assembled by adopting an adhesive and a sizing tape;
the radiation shielding layer is arranged between the adjacent heat insulation layers.
According to one aspect of the invention, the second insulating layer is two layers;
the first heat insulation layer is an alumina fiber composite layer, and the second heat insulation layer is a nano-silica composite layer.
To achieve the above object, the present invention provides a thermal protection system of a thermal protection assembly, comprising: a plurality of thermal protection assemblies, seals and connection assemblies;
the first convex eave structure and the step structure between two adjacent thermal protection assemblies are mutually overlapped, and the connecting parts between two adjacent thermal protection assemblies are mutually overlapped to form an accommodating cavity;
the sealing member is arranged in the accommodating cavity;
the coupling assembling with connecting portion interconnect.
According to one aspect of the invention, the gap between the cover plates of two adjacent heat protection assemblies is 0.5 mm-2 mm;
and a gap between the cover plates of two adjacent thermal protection assemblies is filled with heat-proof putty.
According to one aspect of the invention, the connection assembly comprises: the heat insulation device comprises a plurality of supports, a connecting piece, a first heat insulation gasket, a second heat insulation gasket, an adjusting gasket, a locking gasket and a base;
the plurality of first connecting lugs, the first heat-insulating gasket, the base, the second heat-insulating gasket, the adjusting gasket and the anti-loosening gasket are sequentially arranged in series through the connecting piece;
according to one aspect of the invention, the bracket is provided with a first connecting lug and a second connecting lug which are arranged in a staggered mode;
first engaging lug with base interconnect, the second engaging lug with connecting portion interconnect.
The support is an elastic alloy piece.
According to one scheme of the invention, the thermal protection assembly has the advantages of reusability, modular structure, light weight, low cost and easy replacement, so that the thermal protection assembly has wide application prospect in the field of hypersonic aircrafts, and meets the strict requirements of the aircrafts on heat insulation, light weight, bearing, impact resistance and the like. Further, the invention has wide application range and higher applicability.
According to one scheme of the invention, the Ceramic Matrix Composite (CMC) is adopted to make the cover plate type thermal protection system have higher temperature resistance, stronger bearing capacity, capability of manufacturing large-size components and lower maintenance cost. The thermal protection system disclosed by the invention is designed in a block mode, and the ceramic matrix composite material with the thermal expansion coefficient far smaller than that of metal is adopted, and the thermal expansion coefficient is only 1/2-1/3 of metal, so that the problems of thermal matching and thermal stress in the design of a thermal protection scheme can be effectively solved, and the reliability of the design of the thermal protection system is improved. In addition, the main body of the invention adopts the ceramic matrix composite material with high temperature resistance, high specific strength and relatively lower density, the total weight of the whole body structural member can be reduced by 30-40 percent, and the oxidation resistance of the composite material is better than that of the C/C composite material; meanwhile, the structure of the invention is not easy to fall off, absorb water, prevent rain, is not easy to deform and is beneficial to sealing gaps.
According to one scheme of the invention, the whole cover plate is arranged into the trapezoidal panel, so that the defects of poor aerodynamic characteristics and local aerodynamic heat caused by the fact that gaps among different cover plates are communicated into a long longitudinal seam are effectively avoided, and the method is favorable for ensuring the normal work of an internal structure and components of an aircraft adopting the invention.
According to the scheme of the invention, the heat insulation structure is of a multilayer structure, and the heat insulation layers with different heat insulation effects are arranged, so that the excellent performances of different heat insulation layers are fully utilized, the gradient heat insulation is realized, and the cost of the whole heat insulation structure is reduced. The radiation shielding layers are arranged among different heat insulation layers, so that the heat transmitted by the previous heat insulation layer is further reduced, the performance of the next heat insulation layer is guaranteed, particularly, the radiation shielding layers are arranged at the connection positions of the heat insulation layers, the defect that the surface of the next heat insulation layer is damaged due to abrupt temperature change is overcome, and the radiation shielding layer is beneficial to guaranteeing the stable structure and the good heat insulation effect of a heat insulation structure.
According to one scheme of the invention, after the adjacent thermal protection assemblies are assembled, the first ledge structure and the step structure are mutually overlapped, so that a first sealing structure is formed at the joint position, the air flow on the surface of an aircraft is prevented from directly flowing into the thermal protection system of the invention from the joint position, the connecting parts between the two adjacent thermal protection assemblies are mutually overlapped, the strength of the connecting position is enhanced, an accommodating cavity for installing a sealing element is formed between the two adjacent thermal protection assemblies, the second sealing structure is further formed under the action of the sealing element, the sealing property and the high temperature resistance of the connecting position are further enhanced, and the sealing property and the high temperature resistance of the whole thermal protection system are favorably ensured.
According to one scheme of the invention, the gap between the cover plates of two adjacent thermal protection assemblies is set within the range of 0.5-2 mm, so that sufficient allowance is left for the cover plates to expand under heating under the condition of ensuring the sealing effect of the invention, the damage caused by the adjacent cover plates expanding under heating and extruding mutually is avoided, and the working stability of the invention is ensured. In addition, the joint position is filled with heat release putty, so that the permeation of high-temperature gas from the joint position in the flying process of the aircraft is further avoided, and the good sealing property of the invention is further ensured.
Drawings
FIG. 1 schematically illustrates a block diagram of a thermal shield assembly according to one embodiment of the present invention;
FIG. 2 schematically illustrates a connection of a cover plate and a cover plate support according to an embodiment of the present invention;
FIG. 3 schematically illustrates a bottom view of a cover plate and cover plate support according to one embodiment of the present invention;
FIG. 4 schematically illustrates a block diagram of an insulation structure according to an embodiment of the present invention;
FIG. 5 schematically illustrates a block diagram of a thermal protection system according to an embodiment of the present invention;
FIG. 6 schematically illustrates a connection structure pattern of a thermal protection system according to an embodiment of the present invention;
fig. 7 schematically shows a block diagram of a connection assembly according to an embodiment of the invention.
Detailed Description
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the embodiments will be briefly described below. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
In describing embodiments of the present invention, the terms "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," and the like are used in an orientation or positional relationship that is based on the orientation or positional relationship shown in the associated drawings, which is for convenience and simplicity of description only, and does not indicate or imply that the referenced device or element must have a particular orientation, be constructed and operated in a particular orientation, and thus, the above-described terms should not be construed as limiting the present invention.
The present invention is described in detail below with reference to the drawings and the specific embodiments, which are not repeated herein, but the embodiments of the present invention are not limited to the following embodiments.
Referring to fig. 1, 2 and 3, a thermal shield assembly according to an embodiment of the present invention includes a cover plate 111, a cover plate support 112 and an insulation structure 113. In the present embodiment, the cover 111, the cover support 112, and the heat insulating structure 113 are sequentially provided and fixedly connected to each other. In the present embodiment, the cover plate 111 is made of a material capable of withstanding a high temperature of 1700 ℃. In the present embodiment, the cover plate 111 is made of a ceramic matrix composite (e.g., C/S iC material), and the side thereof away from the cover plate support 112 is provided with an oxidation resistant coating. In the present embodiment, a rib structure 111a is further provided on the side where the cover plate 111 and the cover plate support 112 are connected to each other. In the present embodiment, the cover plate 111 and the cover plate support 112 are integrally formed, and have a stable structure and high strength. Of course, the cover plate 111 and the cover plate support 112 may be detachably connected. Through the arrangement, the cover plate 111 is made of the high-temperature-resistant ceramic matrix composite material with the reinforcing ribs 111a, so that the structural strength is high, the high-temperature-resistant performance of the aircraft in the high-speed flight process is better, the thermal expansion deformation is small, and the stable aerodynamic appearance of the aircraft in the high-speed flight process is ensured. Meanwhile, the high-temperature-resistant ceramic matrix composite material is low in heat conduction performance, heat conducted to the inside in the flying process of the aircraft is effectively blocked, and stable operation of internal components and structures of the aircraft is guaranteed. The anti-oxidation coating is arranged on the outer side of the cover plate 111, so that the surface quality of the cover plate 111 can be effectively guaranteed, particularly, the surface quality of the cover plate 111 is further guaranteed through the anti-oxidation coating in a high-temperature environment in a high-speed flight process, and the service life of the thermal protection assembly is guaranteed to be favorable.
Referring to fig. 2 and 3, according to an embodiment of the present invention, the cover plate support 112 includes a support portion 1121 and a connection portion 1122. In the present embodiment, the supporting portion 1121 is a circular surrounding wall, one end of which is connected to the connecting portion 1122 perpendicularly, and the other end of which is fixedly connected to the cover 111 to form a recessed structure. In the present embodiment, the connecting portion 1122 is an annular plate-like body. The supporting portion 1121 is fixedly connected to the inner annular edge of the connecting portion 1122 in a perpendicular manner.
Referring to fig. 2 and 3, according to an embodiment of the present invention, the cover plate 111 is a trapezoidal plate. In the present embodiment, the cover plate 111 has a first edge 1111 and a second edge 1112 parallel to each other, and the cover plate 111 has a third edge 1113 and a fourth edge 1114 inclined to each other. In the present embodiment, the cover 111 is fixedly supported at one end of the support portion 1121 remote from the connecting portion 1122. The second edge 1112 extends away from the supporting portion 1121 to form the first flange structure 114, that is, the second edge 1112 of the cover plate 111 extends beyond the side surface of the supporting portion 1121, and the protruding portion of the cover plate 111 forms the first flange structure 114. In the present embodiment, the first edge 1111 overlaps the end surface of the support portion 1121 to form the stepped structure 115, that is, the first edge 1111 does not extend beyond the side surface of the support portion 1121, and the first edge 1111 of the cover plate 111 and the end surface of the support portion 1121 form the stepped structure 115. In the present embodiment, the third edge 1113 and the fourth edge 1114 respectively extend away from the supporting portion 1121 to form a protruding eave structure, that is, the third edge 1113 and the fourth edge 1114 of the cover plate 111 respectively extend beyond the side surfaces of the supporting portion 1121, and the protruding portion of the cover plate 111 forms the protruding eave structure. Through the arrangement, the whole cover plate 111 is arranged into the trapezoidal panel, so that the defects of poor aerodynamic characteristics and local aerodynamic heat caused by the fact that gaps among different cover plates 111 are communicated into longer longitudinal seams are effectively avoided, and the method is favorable for ensuring the normal work of the internal structure and the components of the aircraft.
As shown in fig. 4, according to an embodiment of the present invention, the heat insulation structure 113 includes: first insulation layer 1131, second insulation layer 1132, fiberglass layer 1133, and radiation shield layer 1134. In this embodiment, the first thermal insulation layer 1131, the second thermal insulation layer 1132, the glass fiber layer 1133, and the radiation shield layer 1134 are assembled using an adhesive and a fixing tape. In the present embodiment, the heat insulating effect of the first heat insulating layer 1131 is greater than that of the second heat insulating layer 1132. In this embodiment, the first heat insulating layer 1131 is made of an alumina fiber composite material, and the second heat insulating layer 1132 is made of a nano silica composite material. Referring to fig. 4, in the present embodiment, the second thermal insulation layer 1132 has two layers. Radiation shield 1134 is disposed between adjacent thermal barriers, i.e., radiation shield 1134 is disposed between first thermal barrier 1131 and second thermal barrier 1132, and between second thermal barrier 1132 and second thermal barrier 1132. Of course, second insulating layer 1132 may also be provided in other numbers (e.g., three, four, or more). Likewise, the first thermal barrier layer 1131 can also be provided in multiple layers (e.g., two layers, three layers, etc.). Through the above arrangement, the heat insulation structure 113 is arranged into a multilayer structure, and the heat insulation layers with different heat insulation effects are arranged, so that the excellent performances of different heat insulation layers are fully utilized, the gradient separation of heat is realized, the heat insulation advantages and the heat insulation advantages of various materials are furthest exerted, the maximum heat insulation efficiency is obtained, and the cost of the whole heat insulation structure 113 is favorably reduced. The radiation shielding layer 1134 is arranged between different heat insulation layers, so that the heat transferred by the previous heat insulation layer is further reduced, the performance of the next heat insulation layer is ensured, particularly, the radiation shielding layer 1134 is arranged at the connecting position of each heat insulation layer, the defect that the surface of the next heat insulation layer is damaged due to abrupt temperature change is avoided, and the structure stability and the good heat insulation effect of the heat insulation structure 113 are ensured.
Referring to fig. 1, 5 and 6, according to one embodiment of the present invention, a thermal protection system of the present invention comprises: a plurality of thermal shield assemblies 11, a seal 12 and a connecting assembly 13. In the present embodiment, a protection system for an aircraft is formed by splicing a plurality of thermal protection assemblies 11, and the cover plates 111 on the thermal protection assemblies 11 are spliced to form an aerodynamic shape of the aircraft. In the present embodiment, in the airflow direction, the first eaves structure 114 and the step structure 115 between two adjacent thermal protection assemblies 11 overlap with each other, and the connecting portions 1122 between two adjacent thermal protection assemblies 11 overlap with each other to form the receiving cavity 14. In the present embodiment, the sealing member 12 is an elongated body made of a high temperature resistant material, and is disposed in a manner matching the receiving cavity 14, that is, the sealing member 12 is installed in the receiving cavity 14 in cooperation with the receiving cavity 14. In the present embodiment, the connecting portion 1122 on the thermal protection component 11 and the connecting component 13 are connected to each other, thereby realizing the assembly of the thermal protection system of the present invention. Through the arrangement, after the adjacent thermal protection assemblies 11 are assembled, the first eave structure 114 and the step structure 115 are mutually overlapped, so that a first sealing structure is formed at a joint position, the condition that airflow on the surface of an aircraft directly flows into the thermal protection system of the invention from the joint position is avoided, the connecting parts 1122 between the two adjacent thermal protection assemblies 11 are mutually overlapped, the strength of the connecting position is enhanced, the accommodating cavity 14 for installing the sealing element 12 is formed between the two adjacent thermal protection assemblies 11, a second sealing structure is further formed through the action of the sealing element 12, the sealing property and the high temperature resistance of the connecting position are further enhanced, and the sealing property and the high temperature resistance of the whole invention are ensured.
According to an embodiment of the present invention, after the plurality of thermal protection assemblies 11 are connected to each other, the gap between the cover plates 111 of two adjacent thermal protection assemblies 11 is 0.5mm to 2 mm. In the present embodiment, the gap between the cover plates 111 of two adjacent thermal protection assemblies 11 is filled with thermal putty. Through the arrangement, the gap between the cover plates 111 of the two adjacent thermal protection assemblies 11 is set within the range of 0.5 mm-2 mm, so that sufficient allowance is reserved for the thermal expansion of the cover plates 111 under the condition of ensuring the sealing effect of the invention, the damage caused by the thermal expansion and mutual extrusion of the adjacent cover plates 111 is avoided, and the working stability of the invention is ensured. In addition, the joint position is filled with heat release putty, so that the permeation of high-temperature gas from the joint position in the flying process of the aircraft is further avoided, and the good sealing property of the invention is further ensured.
As shown in fig. 7, according to an embodiment of the present invention, the connection assembly 13 includes: a plurality of brackets 131, a connecting member 132, a first heat insulating gasket 133, a second heat insulating gasket 134, an adjustment gasket 135, a lock gasket 136, and a base 137. In this embodiment, the bracket 131 is provided with a first connecting lug 1311, a second connecting lug 1312, and an intermediate member 1313 for connecting the first connecting lug 1311 and the second connecting lug 1312, and the intermediate member 1313 is an elongated plate-shaped member, and has an arc-shaped cross section, so that the structural strength of the intermediate member 1313 is enhanced, and the connecting strength of the bracket 131 is ensured. The first coupling lug 1311 and the second coupling lug 1312 are respectively disposed at two opposite ends of the middle piece 1313, and the first coupling lug 1311 and the second coupling lug 1312 are respectively disposed at two sides of the middle piece 1313 in a staggered parallel arrangement. Through holes are respectively formed on the first connecting lug 1311 and the second connecting lug 1312 for the connection members to pass through. The first coupling lug 1311 is used to couple with the base 137, and the second coupling lug 1312 couples with the coupling portion 1122 through a coupling member.
As shown in fig. 7, in the present embodiment, a bracket 131, a first heat-insulating washer 133, a second heat-insulating washer 134, an adjustment washer 135, and a lock washer 136 are mounted on a base 137 through a coupling 132. In the present embodiment, the first coupling lugs 1311 of the plurality of brackets 131 are stacked on each other and coupled by the coupling member 132. In the present embodiment, the plurality of first coupling lugs 1311, the first heat-insulating washer 133, the seat 137, the second heat-insulating washer 134, the adjustment washer 135, and the lock washer 136 are sequentially connected in series by the coupling member 132 (e.g., a bolt and a nut). In the present embodiment, the base 137 has a substantially zigzag structure.
In the present embodiment, the holder 131 is made of a high temperature alloy material and has elasticity. The same is true. The connecting member 132 is also made of a high temperature alloy material. The support 131 is made of elastic alloy material, so that the defect of unbalanced stress caused by the difference of thermal expansion between the upper cover plate 111 of the heat protection component 11 and the aircraft body structure is relieved, and the structural stability of the heat protection system is effectively improved. Meanwhile, the connecting member 132 is made of alloy material, which is beneficial to eliminating the defect of thermal mismatch between the connecting member 132 and the bracket 131.
The foregoing is merely exemplary of particular aspects of the present invention and devices and structures not specifically described herein are understood to be those of ordinary skill in the art and are intended to be implemented in such conventional ways.
The above description is only one embodiment of the present invention, and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (9)
1. A thermal protection assembly, characterized by comprising a cover plate (111), a cover plate support (112) arranged on the cover plate (111), and an insulating structure (113) connected with the cover plate support (112);
the cover plate (111) is a ceramic matrix composite plate, and an anti-oxidation coating is arranged on one side of the cover plate (111) far away from the cover plate support (112);
the cover plate support (112) comprises a support portion (1121);
the cover plate (111) is a trapezoidal plate having a first edge (1111) and a second edge (1112) which are parallel to each other;
the second edge (1112) extends to a direction far away from the supporting part (1121) to form a first convex eaves structure (114);
the first edge (1111) overlaps the end face of the bearing part (1121) to form a step structure (115).
2. The thermal shield assembly of claim 1 wherein said cover plate support (112) further comprises a connecting portion (1122);
the supporting part (1121) is a circular surrounding wall, one end of the supporting part is mutually and vertically connected with the connecting part (1122), and the other end of the supporting part is mutually and fixedly connected with the cover plate (111) to form a groove structure.
3. The thermal shield assembly of claim 2 wherein said connecting portion (1122) is an annular plate;
the supporting part (1121) is fixedly connected with the annular edge at the inner side of the connecting part (1122) in a mutually perpendicular mode.
4. The thermal protection assembly according to claim 1, wherein said thermal insulation structure (113) comprises: a first thermal insulation layer (1131), a second thermal insulation layer (1132), a fiberglass layer (1133), and a radiation shield layer (1134);
the first heat insulation layer (1131), the second heat insulation layer (1132), the glass fiber layer (1133) and the radiation shielding layer (1134) are assembled by adopting an adhesive and a sizing tape;
the radiation shield (1134) is disposed between adjacent thermal insulation layers.
5. The thermal shield assembly of claim 4, wherein said second insulating layer (1132) is two layers;
the first heat insulation layer (1131) is an alumina fiber composite layer, and the second heat insulation layer (1132) is a nano-silica composite layer.
6. A thermal protection system employing the thermal protection assembly of any one of claims 1 to 5, comprising: a plurality of heat shield assemblies (11), seals (12) and connection assemblies (13);
the first eave structure (114) and the step structure (115) between two adjacent heat protection assemblies (11) are mutually overlapped, and the connecting parts (1122) between two adjacent heat protection assemblies (11) are mutually overlapped to form a containing cavity (14);
the seal (12) being arranged in the accommodation cavity (14);
the connecting member (13) and the connecting portion (1122) are connected to each other.
7. The thermal protection system according to claim 6, wherein the gap between the cover plates (111) of two adjacent thermal protection assemblies (11) is 0.5mm to 2 mm;
and a gap between the cover plates (111) of two adjacent heat protection assemblies (11) is filled with heat-proof putty.
8. Thermal protection system according to claim 6 or 7, characterized in that said connection assembly (13) comprises: a plurality of brackets (131), a connecting member (132), a first heat-insulating gasket (133), a second heat-insulating gasket (134), an adjusting gasket (135), a check gasket (136) and a base (137);
the bracket (131) is provided with a first connecting lug (1311) and a second connecting lug (1312) which are arranged in a staggered mode;
the plurality of first connecting lugs (1311), the first heat insulating gasket (133), the base (137), the second heat insulating gasket (134), the adjusting gasket (135) and the check gasket (136) are sequentially arranged in series through the connecting piece (132);
9. the thermal protection system of claim 8,
the first coupling lug (1311) is interconnected with the base (137), and the second coupling lug (1312) is interconnected with the connecting portion (1122);
the support (131) is an elastic alloy piece.
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| CN110834712B (en) * | 2019-10-17 | 2021-03-26 | 中国运载火箭技术研究院 | Be applicable to high temperature resistant thermal plug structure of dismantling |
| CN113978696B (en) * | 2021-11-08 | 2024-04-09 | 湖北航天技术研究院总体设计所 | Spacecraft and thermal resistance type end cap mounting structure thereof |
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| US5657795A (en) * | 1993-07-01 | 1997-08-19 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Method to produce flexible ceramic thermal protection system resistant to high aeroacoustic noise |
| DE102007054873A1 (en) * | 2006-11-16 | 2009-05-28 | Johann Valentin Krier | Deflection device for preventing the contamination of the boundary layer on wings of a missile in subsonic and trans-sonic flow through the fuselage or similar bodies |
| FR3049571A1 (en) * | 2016-03-30 | 2017-10-06 | Snecma | AIRCRAFT WING WITH IMPROVED ACOUSTIC TREATMENT |
| CN106586036A (en) * | 2016-12-09 | 2017-04-26 | 大连理工大学 | Launch vehicle composite storage tank with stiffening beams and processing method of storage tank |
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