CN108423154A - Hypersonic aircraft leading edge thermal protection method based on gradient porous material - Google Patents
Hypersonic aircraft leading edge thermal protection method based on gradient porous material Download PDFInfo
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- CN108423154A CN108423154A CN201810205982.6A CN201810205982A CN108423154A CN 108423154 A CN108423154 A CN 108423154A CN 201810205982 A CN201810205982 A CN 201810205982A CN 108423154 A CN108423154 A CN 108423154A
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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
- B64C1/38—Constructions adapted to reduce effects of aerodynamic or other external heating
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Abstract
The hypersonic aircraft leading edge thermal protection method based on gradient porous material that the invention discloses a kind of preparing the porous leading edge with gradient porosity using heat-resisting material, and porous up-front stationary point region porosity is maximum, and porosity reduces backward；It is fixedly connected with cooling pipe at porous up-front rear portion, injected coolant in cooling chamber by cooling pipe and sprays leading edge surface；Coolant is when flowing through porous leading edge, it forces to carry out heat convection, reduces porous up-front temperature, while coolant is injected by porous up-front micropore in high temperature mainstream, one layer thicker of film overcast layer is formed in porous up-front stationary point region, porous leading edge is separated with hot-fluid.By being optimized with to traditional Sweat coolling mode for gradient porous material, the high accuracy positioning and calibrated shot of coolant are realized, and then reach ideal thermal protection effect.
Description
Technical field
The present invention relates to a kind of thermal structure surface active thermal means of defences, more particularly to one kind being based on gradient porous material
Hypersonic aircraft leading edge thermal protection method.
Background technology
Aeronautical and space technology be weigh a national science and technology level and overall national strength mark, to country science and technology, military affairs,
Civilian and commercial field generates huge pulling function.And development side of the hypersonic aircraft as the following airmanship
To having become the key strategy project that each aerospace big country competitively develops.With the development of hypersonic aircraft technology,
Its thermal protection problem becomes increasingly conspicuous, especially some key positions of hypersonic aircraft for example leading edge nose cone, the leading edge of a wing,
Spillover mouth etc..Efficient active thermal-protection system is designed and developed, the key technology in the field is become.
Existing active thermal protection technology includes mainly heat convection, gaseous film control and Sweat coolling three types.Such as Fig. 1
It is shown, Sweat coolling utilize porous media interior solid and uniform micrometer grade hole, equably transport coolant to structure height
Warm outer surface.It since porous media has larger specific surface area, is fully exchanged heat, is reduced between coolant and structural material
The temperature of structural material；Meanwhile the coolant of outflow forms one layer of film overcast layer in body structure surface, effectively by structure and height
Hot-fluid completely cuts off.
However, as shown in Fig. 2, in the up-front stationary point region aerodynamic force of hypersonic aircraft and Aerodynamic Heating be it is strongest,
The cooling effect in the region is often worst, cools down front and back temperature highest always.Therefore, before for hypersonic aircraft
Edge bears the efficient active thermal protection needs of high hot-fluid load, and according to Aerodynamic Heating, the spatial distribution characteristic of power acute variation,
The high accuracy positioning and calibrated shot for realizing coolant have very important significance.
Invention content
The hypersonic aircraft leading edge thermal protection method based on gradient porous material that the object of the present invention is to provide a kind of.
The purpose of the present invention is what is be achieved through the following technical solutions：
The hypersonic aircraft leading edge thermal protection method based on gradient porous material of the present invention, using heat-resisting material
The porous leading edge with gradient porosity is prepared, the porous up-front stationary point region porosity is maximum, and porosity subtracts backward
It is small；
It is fixedly connected with cooling pipe at porous up-front rear portion, is injected coolant in cooling chamber simultaneously by cooling pipe
Spray leading edge surface；
Coolant is forced to carry out heat convection when flowing through porous leading edge, reduces porous up-front temperature, while coolant
It is injected in high temperature mainstream by porous up-front micropore, one layer of thicker film overcast is formed in porous up-front stationary point region
Layer, porous leading edge is separated with hot-fluid.
As seen from the above technical solution provided by the invention, provided in an embodiment of the present invention to be based on gradient porous material
Hypersonic aircraft leading edge thermal protection method, due to using gradient porosity porous leading edge, pass through gradient porous material
Optimized with to traditional Sweat coolling mode, realize the high accuracy positioning and calibrated shot of coolant, and then reach
Ideal thermal protection effect.
Description of the drawings
Fig. 1 is Sweat coolling principle schematic；
Fig. 2 be hypersonic aircraft leading edge surface heat flux distribution schematic diagram；
Fig. 3 is the porous up-front structural schematic diagram with gradient porosity in the embodiment of the present invention.
Specific implementation mode
The embodiment of the present invention will be described in further detail below.What is be not described in detail in the embodiment of the present invention is interior
Appearance belongs to the prior art well known to professional and technical personnel in the field.
The hypersonic aircraft leading edge thermal protection method based on gradient porous material of the present invention, it is preferably specific real
The mode of applying is：
Porous leading edge with gradient porosity, the porous up-front stationary point region are prepared using heat-resisting material
(front end) porosity is maximum, and porosity reduces backward；
It is fixedly connected with cooling pipe at porous up-front rear portion, is injected coolant in cooling chamber simultaneously by cooling pipe
Spray leading edge surface；
Coolant is forced to carry out heat convection when flowing through porous leading edge, reduces porous up-front temperature, while coolant
It is injected in high temperature mainstream by porous up-front micropore, one layer of thicker film overcast is formed in porous up-front stationary point region
Layer, porous leading edge is separated with hot-fluid.
The gradient porosity is in consecutive variations or stepped change.
The coolant is gaseous state or liquid coolant.
It is Steel material, high temperature alloy or ceramic material to prepare porous up-front heat-resisting material.
The cooling pipe is single pipeline either multi-pipeline.
The up-front thermal protection method of hypersonic aircraft is directed in high-temperature high-speed airflow, how basic solution is
It is rational to arrange coolant in the sendout in leading edge stationary point, make more coolant flows to stationary point region.
The hypersonic aircraft leading edge thermal protection method based on gradient porous material of the present invention, with following excellent
Point：
1, the method for the present invention makes hypersonic aircraft leading edge with Sweat coolling by the multi-cellular structure of porous media
Mode realizes thermal protection, i.e., coolant can be injected by micropore in high temperature mainstream, when flowing through porous structure material, force into
Row heat convection, reduces the temperature of nose cone, while forming film overcast layer on up-front surface, effectively by leading edge and hot-fluid every
It opens.
2, porous media of the invention is distributed with gradient porosity, and hole is locally increased in the high heat flux regions in leading edge stationary point
Gap rate, the low heat flux regions in downstream reduce porosity.Under same coolant injection pressure, the reasonable distribution of coolant is realized, no
The coolant dosage that only can locally increase stationary point region, effectively improves the cooling efficiency in stationary point region, and can improve nose cone
The uniformity of surface temperature distribution reduces temperature gradient, avoids continuing to increase for material thermal stress, while also can significantly reduce cold
But the carrying amount of medium.
3, the present invention passes through ladder for hypersonic aircraft leading edge Aerodynamic Heating, the spatial distribution characteristic of power acute variation
Being optimized with to traditional Sweat coolling mode for degree porous material, realizes the high accuracy positioning of coolant and quantitative note
It penetrates, and then reaches ideal thermal protection effect.
Specific embodiment：
As shown in figure 3, Penetration Signature of the method for the present invention according to porous media, utilizes the concept of Sweat coolling, completion pair
The thermal protection of hypersonic aircraft leading edge outer surface.The method of the present invention include one be prepared by heat-resisting material it is porous
Leading edge 1 is fixedly connected there are one cooling pipe 2 at porous up-front rear portion, coolant is injected cooling chamber by cooling pipe 2
Interior 3.Coolant penetrates into high temperature mainstream under cavity pressure effect from micropore.
As shown in figure 3, the material for the porous media that the porous leading edge 1 in the method for the present invention uses and the shape of leading edge 1
It is essentially identical with the prior art, the difference is that the present invention improves the characteristic of the porous media of the use of leading edge 1.The present invention
The main feature of method includes：First porous leading edge 1 is distributed with gradient porosity, in the stationary point region of porous leading edge 1 part
Increase porosity；The porosity of second porous leading edge 1 can be with consecutive variations, can also stepped change.
In the present invention gradient porosity material of porous leading edge 1 may be used existing various heat-resisting materials and it is various not
With processing technology obtain, such as traditional powder sintering and die pressing or novel centrifugal deposition technology, wet spray and
Brushing, 3 D-printing forming technique, precinct laser sintering and powder injection molding technology etc..
In above-described embodiment, coolant preferentially selects liquid coolant.
In above-described embodiment, coolant flow in pipes 2 can be changed to multi-pipeline by single pipeline, and then increase cooling chamber
Quantity.
In above-described embodiment, the porosity gradient difference of porous leading edge 1 can in real time be adjusted according to the needs of actual condition
It is whole, preferentially select larger difference.
In above-described embodiment, the structure of porous leading edge 1 is preferentially designed using non-uniform thickness, cuts down stationary point region as needed
Thickness, the superposition of the two can be readily apparent that mitigate problems of the prior art.
The various embodiments described above are merely to illustrate the present invention, the wherein type, manufacture craft and gradient pore of gradient porous material
Gap rate distributed constant may be changed, and every equivalents carried out based on the technical solution of the present invention and change
Into should not exclude except protection scope of the present invention.
The foregoing is only a preferred embodiment of the present invention, but scope of protection of the present invention is not limited thereto,
Any one skilled in the art is in the technical scope of present disclosure, the change or replacement that can be readily occurred in,
It should be covered by the protection scope of the present invention.Therefore, protection scope of the present invention should be with the protection model of claims
Subject to enclosing.
Claims (5)
1. a kind of hypersonic aircraft leading edge thermal protection method based on gradient porous material, which is characterized in that use resistance to height
Adiabator prepares the porous leading edge with gradient porosity, and the porous up-front stationary point region porosity is maximum, to metapore
Gap rate reduces；
It is fixedly connected with cooling pipe at porous up-front rear portion, coolant is injected in cooling chamber and sprayed by cooling pipe
Leading edge surface；
Coolant is forced to carry out heat convection, reduces porous up-front temperature, while coolant passes through when flowing through porous leading edge
In porous up-front micropore injection high temperature mainstream, one layer thicker of film overcast layer is formed in porous up-front stationary point region, it will
Porous leading edge is separated with hot-fluid.
2. the hypersonic aircraft leading edge thermal protection method according to claim 1 based on gradient porous material, special
Sign is that the gradient porosity is in consecutive variations or stepped change.
3. the hypersonic aircraft leading edge thermal protection method according to claim 1 based on gradient porous material, special
Sign is that the coolant is gaseous state or liquid coolant.
4. the hypersonic aircraft leading edge thermal protection method according to claim 1 based on gradient porous material, special
Sign is that it is Steel material, high temperature alloy or ceramic material to prepare porous up-front heat-resisting material.
5. the hypersonic aircraft leading edge thermal protection method according to claim 1 based on gradient porous material, special
Sign is that the cooling pipe is single pipeline either multi-pipeline.
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Cited By (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109580694A (en) * | 2018-11-30 | 2019-04-05 | 中国航空工业集团公司沈阳飞机设计研究所 | A kind of thermal protection structure test fixture |
| CN109835466A (en) * | 2019-03-14 | 2019-06-04 | 中国科学技术大学 | Aircraft and its housing assembly |
| CN111824391A (en) * | 2020-07-27 | 2020-10-27 | 清华大学 | Phase-change sweating cooling heat protection structure and construction method thereof |
| CN112765913A (en) * | 2021-04-08 | 2021-05-07 | 中国空气动力研究与发展中心计算空气动力研究所 | Layered gradient porous material sweating cooling structure and aircraft |
| CN112758304A (en) * | 2021-04-07 | 2021-05-07 | 中国空气动力研究与发展中心计算空气动力研究所 | Self-adaptive porous material sweating cooling front edge structure based on pyrolysis |
| CN112810799A (en) * | 2019-11-15 | 2021-05-18 | 通用电气公司 | System and method for cooling leading edge of high speed vehicle |
| CN112810798A (en) * | 2019-11-15 | 2021-05-18 | 通用电气公司 | System for reducing thermal stress in the leading edge of a high speed vehicle |
| US11260953B2 (en) | 2019-11-15 | 2022-03-01 | General Electric Company | System and method for cooling a leading edge of a high speed vehicle |
| US11267551B2 (en) * | 2019-11-15 | 2022-03-08 | General Electric Company | System and method for cooling a leading edge of a high speed vehicle |
| WO2022051912A1 (en) * | 2020-09-08 | 2022-03-17 | 西门子股份公司 | Laval nozzle and manufacturing method therefor |
| CN115556948A (en) * | 2022-11-21 | 2023-01-03 | 中国科学院力学研究所 | Thermal protection method and system for sharp front edge of hypersonic vehicle |
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| CN111824391A (en) * | 2020-07-27 | 2020-10-27 | 清华大学 | Phase-change sweating cooling heat protection structure and construction method thereof |
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| CN112758304A (en) * | 2021-04-07 | 2021-05-07 | 中国空气动力研究与发展中心计算空气动力研究所 | Self-adaptive porous material sweating cooling front edge structure based on pyrolysis |
| CN112765913B (en) * | 2021-04-08 | 2021-06-29 | 中国空气动力研究与发展中心计算空气动力研究所 | Layered gradient porous material sweating cooling structure and aircraft |
| CN112765913A (en) * | 2021-04-08 | 2021-05-07 | 中国空气动力研究与发展中心计算空气动力研究所 | Layered gradient porous material sweating cooling structure and aircraft |
| CN115556948A (en) * | 2022-11-21 | 2023-01-03 | 中国科学院力学研究所 | Thermal protection method and system for sharp front edge of hypersonic vehicle |
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