CN110901885B - Thermal protection system of aircraft - Google Patents
Thermal protection system of aircraft Download PDFInfo
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- CN110901885B CN110901885B CN201911340061.1A CN201911340061A CN110901885B CN 110901885 B CN110901885 B CN 110901885B CN 201911340061 A CN201911340061 A CN 201911340061A CN 110901885 B CN110901885 B CN 110901885B
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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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- Aviation & Aerospace Engineering (AREA)
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Abstract
The invention discloses a thermal protection system of an aircraft, and relates to the technical field of thermal protection of aircrafts. The device comprises an aircraft body, wherein a heat insulation layer and a heat-proof layer are laid on the outer wall of the aircraft body from inside to outside, an air rudder is arranged on one side, away from the heat insulation layer, of the heat-proof layer, the heat-proof layer comprises a first heat-proof area and a second heat-proof area, the second heat-proof area is located below a rudder tip of the air rudder, the remaining areas of the heat-proof layer are the first heat-proof areas, multiple layers of prepreg cloth are laid on the first heat-proof area and the second heat-proof area, and the laying density of the second heat-proof area is larger than that of the first heat-proof area. According to the thermal protection system of the aircraft, the local ablation resistance is improved by improving the laying density of the second heat-proof area, and the problem that the aerodynamic parameters and the appearance of the aircraft are influenced by additionally adding the rectification block in the second heat-proof area is solved.
Description
Technical Field
The invention relates to the technical field of thermal protection of aircrafts, in particular to a thermal protection system of an aircraft.
Background
In order to adapt to the severe aerodynamic thermal environment of the flight condition of the aerospace craft with long endurance and high Mach number, the thermal protection design of the aerospace craft is more and more important, the requirement on the heat insulation performance of a thermal protection material is high, meanwhile, the thermal protection structure is more and more complex, and particularly, the heat protection performance of the heat protection layer near the rudder tip of the air rudder of the aerospace craft needs to be considered particularly.
Generally, a heat protection structure of a traditional air rudder interference area is formed by adding a rectifying block with a certain size right in front of a rudder tip, wherein the rectifying block is used for blocking high-speed hot air flow coming from the rudder tip on one hand, and guiding and evacuating the high-speed hot air flow to the far end of an aircraft on the other hand, so that the high-speed hot air flow and a steering engine are prevented from generating a shock wave interference effect. However, the increase of the rectification block firstly increases the self weight of the aircraft, and secondly, the existing rectification block has a certain volume, so that the aerodynamic parameters of the aircraft in high-speed flight can be influenced to a certain extent.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a thermal protection system of an aircraft, which improves the local ablation resistance by improving the laying density of a second heat-proof area and avoids the problem that the aerodynamic parameters and the appearance of the aircraft are influenced by additionally adding a rectification block in the second heat-proof area.
In order to achieve the above purposes, the technical scheme adopted by the invention is as follows:
the aircraft comprises an aircraft body, wherein a heat insulation layer and a heat-proof layer are laid on the outer wall of the aircraft body from inside to outside, an air rudder is arranged on one side, away from the heat insulation layer, of the heat-proof layer, the heat-proof layer comprises a first heat-proof area and a second heat-proof area, the second heat-proof area is located below a rudder tip of the air rudder, the remaining areas of the heat-proof layer are the first heat-proof areas, multiple layers of prepreg cloth are laid on the first heat-proof area and the second heat-proof area, and the laying density of the second heat-proof areas is larger than that of the first heat-proof areas.
On the basis of the technical scheme, the number of the prepreg of the second heat-proof area is at least one layer more than that of the prepreg of the first heat-proof area, so that the density of the second heat-proof area is improved.
On the basis of the technical scheme, the first heat-proof area and the second heat-proof area are the same in thickness and are both 9 mm.
On the basis of the technical scheme, the first heat-proof area comprises 4 layers of prepreg cloth with the thickness of 2mm and 1 layer of prepreg cloth with the thickness of 1 mm.
On the basis of the technical scheme, the prepreg cloth with the thickness of 1mm in the first heat-proof area is positioned between the heat-insulating layer and the 2 nd prepreg cloth with the thickness of 2mm from inside to outside.
On the basis of the technical scheme, the second heat-proof area comprises 4 layers of prepreg cloth with the thickness of 2mm and at least 2 layers of prepreg cloth with the thickness of 1 mm.
On the basis of the technical scheme, the thickness of each of the 4 layers of prepreg cloth with the thickness of 2mm in the second heat-proof area and the thickness of 1 layer of prepreg cloth with the thickness of 1mm in the second heat-proof area are integrally formed with the prepreg cloth corresponding to the first heat-proof area.
On the basis of the technical scheme, the size of the second heat-proof area is 100mmX100 mm.
On the basis of the technical scheme, the prepreg cloth is quartz fiber cloth.
On the basis of the technical scheme, the heat-proof layer is formed by the multiple layers of prepreg cloth through high-temperature die pressing.
Compared with the prior art, the invention has the advantages that:
the thermal protection system of the aircraft comprises a first thermal protection area and a second thermal protection area, wherein the number of the prepreg cloth of the second thermal protection area is at least one layer more than that of the prepreg cloth of the first thermal protection area so as to improve the laying density of the second thermal protection area, improve the ablation resistance of the second thermal protection area and meet the requirement of local thermal protection close to a rudder tip. Compared with the prior design, the system can also avoid the complex design that the rectifying block needs to be additionally added on the outer wall of the aircraft on the basis of solving the problem that the heat-proof layer of the existing aircraft is insufficient in heat-proof near the rudder tip, does not need to change the aerodynamic appearance of the aircraft, effectively improves the heat-proof performance of the heat-proof layer of the aircraft near the rudder tip, and ensures that the whole heat-proof layer has good heat-proof performance and reliability.
Drawings
Fig. 1 is a schematic view of a thermal protection system for an aircraft according to an embodiment of the present invention.
In the figure: 1-an aircraft body, 2-a thermal insulation layer, 3-a thermal protection layer, 30-prepreg cloth and 4-an air rudder.
Detailed Description
Embodiments of the present invention will be described in further detail below with reference to the accompanying drawings.
Referring to fig. 1, an embodiment of the present invention provides an aircraft thermal protection system, which includes an aircraft body 1, a thermal insulation layer 2 and a thermal protection layer 3 are laid on an outer wall of the aircraft body 1 from inside to outside, an air rudder 4 is disposed on a side of the thermal protection layer 3 away from the thermal insulation layer 2, the thermal protection layer 3 includes a first thermal protection region and a second thermal protection region, the second thermal protection region is located below a rudder tip of the air rudder 4, remaining regions of the thermal protection layer 3 are both the first thermal protection region, a plurality of layers of prepreg 30 are laid on the first thermal protection region and the second thermal protection region, and a laying density of the second thermal protection region is greater than a laying density of the first thermal protection region.
Specifically, referring to fig. 1, the number of the prepreg 30 in the second heat-shielding region is at least one more layer than the number of the prepreg 30 in the first heat-shielding region, thereby increasing the density of the second heat-shielding region. In order to not change the appearance design of the aircraft as much as possible, the thicknesses of the first heat-proof area and the second heat-proof area are the same and are both 9 mm. Because the number of the layers of the prepreg 30 at the second heat-proof area is more than that of the prepreg 30 at the first heat-proof area, and the two areas are equal in thickness, compared with the first heat-proof area with a larger area, the laying density at the second heat-proof area is improved, so that the ablation resistance of the second heat-proof area is improved, and the requirement of local heat protection of the heat-proof layer 3 close to the rudder tip is met.
Specifically, referring to fig. 1, the first heat-proof area includes 4 layers of prepreg 30 each having a thickness of 2mm and 1 layer of prepreg 30 having a thickness of 1mm, wherein the prepreg 30 having a thickness of 1mm in the first heat-proof area is located between the heat-insulating layer 2 and the 2 nd prepreg 30 having a thickness of 2mm from inside to outside, that is, is attached to the heat-insulating layer 2. The second heat-proof area comprises 4 layers of prepreg 30 with the thickness of 2mm and at least 2 layers of prepreg 30 with the thickness of 1mm, and from the angle of structural design, the 4 layers of prepreg 30 with the thickness of 2mm and the 1 layer of prepreg 30 with the thickness of 1mm are integrally formed with the prepreg 30 corresponding to the first heat-proof area.
Specifically, the size of the second heat-proof area is 100mmX100mm, and the size of the second heat-proof area can also be adjusted according to the actual requirement of the aircraft. In addition, the prepreg 30 is unified into a quartz fiber cloth, and in the laying process, one layer is laid in sequence from inside to outside, wherein the prepreg 30 which is not integrally formed with the first heat-proof area and has a smaller area in the second heat-proof area is positioned between the 3 rd layer and the 4 th layer from inside to outside, and after all the prepreg 30 are laid, the final heat-proof layer 3 is formed through high-temperature die pressing. The structure improves the heat-proof performance of the part near the rudder tip of the heat-proof layer 3 in a targeted manner by locally increasing the density of the heat-proof layer 3 while hardly increasing the total weight of the aircraft, and realizes the heat-proof adaptive design of different parts of the aircraft.
The thermal protection system is mainly used for local heat protection of the heat protection layer 3 near the rudder tip in the high-Mach number and long-endurance aerodynamic thermal environment of the aircraft, the laying density of the second heat protection area is improved by increasing the number of the prepreg cloth 30 of the second heat protection area, the ablation resistance of the second heat protection area is improved, and the requirement of local heat protection close to the rudder tip is met. Compared with the prior design, the system can solve the problem that the heat-proof layer 3 of the existing aircraft is insufficient in heat-proof performance near the rudder tip, further avoid the complex design that a rectifying block needs to be additionally arranged on the outer wall of the aircraft, avoid changing the aerodynamic appearance of the aircraft, effectively improve the heat-proof performance of the heat-proof layer 3 of the aircraft near the rudder tip, and ensure that the whole heat-proof layer 3 has good heat-proof performance and reliability.
The present invention is not limited to the above-mentioned preferred embodiments, and any other products in various forms can be obtained by anyone with the teaching of the present invention, but any changes in the shape or structure thereof, which have the same or similar technical solutions as the present invention, are within the protection scope.
Claims (7)
1. A thermal protection system for an aircraft, characterized in that it comprises:
the aircraft comprises an aircraft body (1), wherein a heat insulation layer (2) and a heat-proof layer (3) are laid on the outer wall of the aircraft body from inside to outside, an air rudder (4) is arranged on one side, away from the heat insulation layer (2), of the heat-proof layer (3), the heat-proof layer (3) comprises a first heat-proof area and a second heat-proof area, the second heat-proof area is located below a rudder tip of the air rudder (4), the rest area of the heat-proof layer (3) is the first heat-proof area, a plurality of layers of prepreg cloth (30) are laid on the first heat-proof area and the second heat-proof area, and the laying density of the second heat-proof area is greater than that of the first heat-proof area; wherein the content of the first and second substances,
the thickness of the first heat-proof area is the same as that of the second heat-proof area, the number of the prepreg cloth (30) of the second heat-proof area is at least one layer more than that of the prepreg cloth (30) of the first heat-proof area so as to improve the density of the second heat-proof area, the prepreg cloth (30) is quartz fiber cloth, and the heat-proof layers (3) are formed by high-temperature mould pressing of the plurality of layers of prepreg cloth (30).
2. A thermal protection system for an aircraft according to claim 1, characterized in that: the thickness of the first heat-proof area and the second heat-proof area is 9 mm.
3. A thermal protection system for an aircraft according to claim 2, characterized in that: the first heat-proof area comprises 4 layers of prepreg (30) with the thickness of 2mm and 1 layer of prepreg (30) with the thickness of 1 mm.
4. A thermal protection system for an aircraft according to claim 3, characterized in that: the prepreg (30) with the thickness of 1mm in the first heat-proof area is positioned between the heat-insulating layer (2) and the 2 nd prepreg (30) with the thickness of 2mm from inside to outside.
5. A thermal protection system for an aircraft according to claim 2, characterized in that: the second heat-proof area comprises 4 layers of prepreg (30) with the thickness of 2mm and at least 2 layers of prepreg (30) with the thickness of 1 mm.
6. A thermal protection system for an aircraft according to claim 5, characterized in that: and 4 layers of prepreg (30) with the thickness of 2mm in each layer and 1 layer of prepreg (30) with the thickness of 1mm in the second heat-proof area are integrally formed with the prepreg (30) corresponding to the first heat-proof area.
7. A thermal protection system for an aircraft according to claim 1, characterized in that: the size of the second heat-proof area is 100mmX100 mm.
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| CN201911340061.1A CN110901885B (en) | 2019-12-23 | 2019-12-23 | Thermal protection system of aircraft |
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| CN201911340061.1A CN110901885B (en) | 2019-12-23 | 2019-12-23 | Thermal protection system of aircraft |
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| CN110901885A CN110901885A (en) | 2020-03-24 |
| CN110901885B true CN110901885B (en) | 2021-11-30 |
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Families Citing this family (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111924089B (en) * | 2020-06-28 | 2021-09-07 | 北京临近空间飞行器***工程研究所 | Rudder shaft heat-proof structure with separated heat-proof and force-bearing functions |
| CN113022842B (en) * | 2021-03-26 | 2023-03-17 | 宁波中科祥龙轻量化科技有限公司 | High-temperature-resistant high-bearing foldable air rudder |
| CN113665850B (en) * | 2021-08-02 | 2023-06-13 | 湖北航天技术研究院总体设计所 | Phase-change heat-proof structure of rudder shaft and aircraft |
| 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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