CN108423154B - 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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- CN108423154B CN108423154B CN201810205982.6A CN201810205982A CN108423154B CN 108423154 B CN108423154 B CN 108423154B CN 201810205982 A CN201810205982 A CN 201810205982A CN 108423154 B CN108423154 B CN 108423154B
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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 invention discloses a hypersonic aircraft front edge thermal protection method based on a gradient porous material, wherein a porous front edge with gradient porosity is prepared by adopting a high-temperature resistant material, the porosity of a stagnation point area of the porous front edge is maximum, and the backward porosity is reduced; the rear part of the porous front edge is fixedly connected with a cooling pipeline, and a coolant is injected into the cooling cavity through the cooling pipeline and sprayed out of the surface of the front edge; when the coolant flows through the porous front edge, the convective heat exchange is forced to be carried out, the temperature of the porous front edge is reduced, meanwhile, the coolant is injected into a high-temperature main flow through micropores of the porous front edge, a thicker air film covering layer is formed in a stagnation point area of the porous front edge, and the porous front edge is separated from the heat flow. The traditional sweating cooling mode is optimized by applying the gradient porous material, high-precision positioning and quantitative injection of the coolant are realized, and an ideal thermal protection effect is further achieved.
Description
Technical Field
The invention relates to an active thermal protection method for a high-temperature structure surface, in particular to a hypersonic aircraft leading edge thermal protection method based on a gradient porous material.
Background
The aerospace technology is a mark for measuring the national technology level and comprehensive national strength, and has a great pulling effect on the national technology, military, civil and commercial fields. The hypersonic flight vehicle is taken as the development direction of future flight technology, and has become a key strategic item of competitive development of various aerospace big countries. With the development of the technology of the hypersonic aircraft, the thermal protection problem of the hypersonic aircraft is increasingly prominent, particularly at some key parts of the hypersonic aircraft, such as a leading edge nose cone, a wing leading edge, an air inlet overflow port and the like. Designing and developing an efficient active thermal protection system becomes a key technology in the field.
The existing active thermal protection technology mainly comprises three types of convection heat transfer, air film cooling and sweating cooling. As shown in fig. 1, transpiration cooling utilizes dense and uniform micron-scale pores within a porous media to uniformly transport coolant to the high temperature outer surface of the structure. The porous medium has a larger specific surface area, so that the coolant and the structural material can exchange heat sufficiently, and the temperature of the structural material is reduced; meanwhile, the flowing coolant forms an air film covering layer on the surface of the structure, and the structure is effectively isolated from high heat flow.
However, as shown in FIG. 2, aerodynamic forces and aerodynamic heat are strongest in the stagnation region of the hypersonic aircraft leading edge, where the cooling effect tends to be worst and the temperature before and after cooling is always highest. Therefore, the method has very important significance for realizing high-precision positioning and quantitative injection of the coolant according to the space distribution characteristics of severe changes of aerodynamic heat and force and aiming at the high-efficiency active heat protection requirement of the front edge of the hypersonic aircraft for bearing extremely high heat flow load.
Disclosure of Invention
The invention aims to provide a hypersonic aircraft leading edge thermal protection method based on a gradient porous material.
The purpose of the invention is realized by the following technical scheme:
the hypersonic aircraft leading edge thermal protection method based on the gradient porous material adopts a high-temperature resistant material to prepare a porous leading edge with gradient porosity, wherein the porosity of a stagnation point area of the porous leading edge is the largest, and the backward porosity is reduced;
the rear part of the porous front edge is fixedly connected with a cooling pipeline, and a coolant is injected into the cooling cavity through the cooling pipeline and sprayed out of the surface of the front edge;
when the coolant flows through the porous front edge, the convective heat exchange is forced to be carried out, the temperature of the porous front edge is reduced, meanwhile, the coolant is injected into a high-temperature main flow through micropores of the porous front edge, a thicker air film covering layer is formed in a stagnation point area of the porous front edge, and the porous front edge is separated from the heat flow.
According to the technical scheme provided by the invention, the hypersonic aircraft front edge thermal protection method based on the gradient porous material provided by the embodiment of the invention has the advantages that the porous front edge with the gradient porosity is adopted, the traditional sweating cooling mode is optimized by using the gradient porous material, the high-precision positioning and quantitative injection of the coolant are realized, and the ideal thermal protection effect is further achieved.
Drawings
FIG. 1 is a schematic view of the cooling principle of sweating;
FIG. 2 is a schematic diagram of the heat flow distribution of the leading edge surface of a hypersonic aircraft;
FIG. 3 is a schematic structural view of a porous leading edge having a gradient porosity in an embodiment of the present invention.
Detailed Description
The embodiments of the present invention will be described in further detail below. Details which are not described in detail in the embodiments of the invention belong to the prior art which is known to the person skilled in the art.
The invention discloses a hypersonic aircraft leading edge thermal protection method based on a gradient porous material, which comprises the following preferred specific embodiments:
preparing a porous leading edge with gradient porosity by adopting a high-temperature resistant material, wherein the porosity of a stagnation point region (front end) of the porous leading edge is the maximum, and the backward porosity is reduced;
the rear part of the porous front edge is fixedly connected with a cooling pipeline, and a coolant is injected into the cooling cavity through the cooling pipeline and sprayed out of the surface of the front edge;
when the coolant flows through the porous front edge, the convective heat exchange is forced to be carried out, the temperature of the porous front edge is reduced, meanwhile, the coolant is injected into a high-temperature main flow through micropores of the porous front edge, a thicker air film covering layer is formed in a stagnation point area of the porous front edge, and the porous front edge is separated from the heat flow.
The gradient porosity is continuously or stepwise.
The coolant is gaseous or liquid coolant.
The high temperature resistant material for preparing the porous leading edge is steel material, high temperature alloy or ceramic material.
The cooling pipeline is a single pipeline or a plurality of pipelines.
The fundamental solution of the method for protecting the leading edge of a hypersonic aircraft from heat in high-temperature and high-speed airflow is how to reasonably arrange the distribution amount of the coolant at the stagnation point of the leading edge so that more coolant flows to the stagnation point area.
The hypersonic aircraft leading edge thermal protection method based on the gradient porous material has the following advantages:
1. the method of the invention realizes thermal protection of the front edge of the hypersonic aircraft in a sweating cooling mode by means of the microporous structure of the porous medium, namely, coolant can be injected into a high-temperature main flow through micropores, and forced convection heat exchange is carried out when the coolant flows through the porous structure material, so that the temperature of a nose cone is reduced, and meanwhile, an air film covering layer is formed on the surface of the front edge, so that the front edge is effectively separated from heat flow.
2. The porous media of the present invention have a gradient porosity profile with a locally increased porosity in the high heat flow region at the leading edge stagnation point and a reduced porosity in the low heat flow region downstream. Under the same coolant injection pressure, the reasonable distribution of the coolant is realized, the coolant dosage of the stagnation point area can be locally increased, the cooling efficiency of the stagnation point area is effectively improved, the uniformity of the temperature distribution of the surface of the nose cone can be improved, the temperature gradient is reduced, the continuous increase of the thermal stress of the material is avoided, and meanwhile, the carrying capacity of the cooling medium can be obviously reduced.
3. The invention aims at the space distribution characteristics of severe aerodynamic heat and force changes of the front edge of the hypersonic aircraft, optimizes the traditional sweating cooling mode by using the gradient porous material, realizes high-precision positioning and quantitative injection of the coolant, and further achieves the ideal heat protection effect.
The specific embodiment is as follows:
as shown in FIG. 3, the method of the present invention utilizes the concept of transpiration cooling to achieve thermal protection of the outer surface of the hypersonic aircraft leading edge based on the permeability characteristics of the porous media. The method comprises a porous front edge 1 made of high-temperature resistant material, a cooling pipeline 2 is fixedly connected to the rear part of the porous front edge, and coolant is injected into a cooling cavity 3 through the cooling pipeline 2. The coolant penetrates into the high-temperature main flow from the micropores under the action of the pressure in the cavity.
As shown in FIG. 3, the material of the porous media used for the porous leading edge 1 and the shape of the leading edge 1 in the method of the present invention are substantially the same as those of the prior art, except that the present invention improves the characteristics of the porous media used for the leading edge 1. The method of the invention is mainly characterized by comprising the following steps: the first porous leading edge 1 has a gradient porosity distribution, and the porosity is locally increased in a stagnation point area of the porous leading edge 1; the porosity of the second porous leading edge 1 may vary continuously or in steps.
The gradient porosity material of the porous leading edge 1 of the present invention can be obtained by using various existing high temperature resistant materials and various different processing techniques, such as the conventional powder sintering method and the die pressing method, or the novel centrifugal deposition technique, the wet spraying and brushing, the three-dimensional printing forming technique, the selective laser sintering and the powder injection forming technique, etc.
In the above embodiment, the coolant is preferably a liquid coolant.
In the above embodiment, the coolant injection pipe 2 may be changed from a single pipe to a multi-pipe, thereby increasing the number of cooling chambers.
In the above embodiment, the porosity gradient difference of the porous leading edge 1 may be adjusted in real time according to the requirement of the actual working condition, and a larger difference is preferably selected.
In the above embodiment, the porous leading edge 1 is preferably designed to have a non-uniform thickness, and the thickness of the stagnation point region is reduced as required, so that the problem in the prior art can be remarkably alleviated by the superposition effect of the two.
The above embodiments are only used for illustrating the present invention, wherein the kind, the manufacturing process and the gradient porosity distribution parameters of the gradient porous material can be changed, and all equivalent changes and modifications based on the technical solution of the present invention should not be excluded from the protection scope of the present invention.
The above description is only for the preferred embodiment of the present invention, but the scope of the present invention is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present invention are included in the scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.
Claims (5)
1. A hypersonic aircraft leading edge thermal protection method based on gradient porous materials is characterized in that a porous leading edge with gradient porosity is prepared by adopting high-temperature resistant materials, the porosity of a stagnation point area of the porous leading edge is the largest, and the backward porosity is reduced;
the rear part of the porous front edge is fixedly connected with a cooling pipeline, and a coolant is injected into the cooling cavity through the cooling pipeline and sprayed out of the surface of the front edge;
when the coolant flows through the porous front edge, the convective heat exchange is forced to be carried out, the temperature of the porous front edge is reduced, meanwhile, the coolant is injected into a high-temperature main flow through micropores of the porous front edge, a thicker air film covering layer is formed in a stagnation point area of the porous front edge, and the porous front edge is separated from the heat flow.
2. The hypersonic aircraft leading edge thermal protection method based on gradient porous material of claim 1, wherein the gradient porosity is continuously or stepwise varied.
3. The method for thermal protection of a leading edge of a hypersonic aircraft based on gradient porous material of claim 1, wherein the coolant is gaseous or liquid.
4. The hypersonic aircraft leading edge thermal protection method based on gradient porous material as claimed in claim 1, wherein the high temperature resistant material for preparing the porous leading edge is steel material, high temperature alloy or ceramic material.
5. The method for thermal protection of a leading edge of a hypersonic aircraft based on gradient porous material of claim 1, wherein the cooling ducts are single ducts or multi-ducts.
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN109580694B (en) * | 2018-11-30 | 2021-07-09 | 中国航空工业集团公司沈阳飞机设计研究所 | Hot protective structure test fixture |
| CN109835466A (en) * | 2019-03-14 | 2019-06-04 | 中国科学技术大学 | Aircraft and its housing assembly |
| US11260953B2 (en) | 2019-11-15 | 2022-03-01 | General Electric Company | System and method for cooling a leading edge of a high speed vehicle |
| US11352120B2 (en) * | 2019-11-15 | 2022-06-07 | 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 |
| US11260976B2 (en) * | 2019-11-15 | 2022-03-01 | General Electric Company | System for reducing thermal stresses in a leading edge of a high speed vehicle |
| CN111824391B (en) * | 2020-07-27 | 2021-11-23 | 清华大学 | Phase-change sweating cooling heat protection structure and construction method thereof |
| WO2022051912A1 (en) * | 2020-09-08 | 2022-03-17 | 西门子股份公司 | Laval nozzle and manufacturing method therefor |
| 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 |
| CN115556948B (en) * | 2022-11-21 | 2023-03-21 | 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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| US4949920A (en) * | 1989-12-14 | 1990-08-21 | The United States Of America As Represented By The Secretary Of The Navy | Ablative cooling of aerodynamically heated radomes |
| CN101418391B (en) * | 2008-12-15 | 2010-08-25 | 哈尔滨理工大学 | Method for preparing gradient porous material |
| US8844877B1 (en) * | 2010-09-02 | 2014-09-30 | The Boeing Company | Stay sharp, fail safe leading edge configuration for hypersonic and space access vehicles |
| CN102145747A (en) * | 2011-03-22 | 2011-08-10 | 北京航空航天大学 | Impact and micro straight channel cooling structure for front edge of hypersonic vehicle |
| CN104859835B (en) * | 2015-04-27 | 2017-06-16 | 清华大学 | A kind of hypersonic aircraft nose cone based on Compound cooling mode |
| CN106516072B (en) * | 2016-11-10 | 2018-06-29 | 清华大学 | A kind of thermal protection structure at the leading edge position of hypersonic vehicle |
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