CN101659137A - Heat protective coating structure for reducing temperature of aircraft kernel - Google Patents
Heat protective coating structure for reducing temperature of aircraft kernel Download PDFInfo
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- CN101659137A CN101659137A CN200910112588A CN200910112588A CN101659137A CN 101659137 A CN101659137 A CN 101659137A CN 200910112588 A CN200910112588 A CN 200910112588A CN 200910112588 A CN200910112588 A CN 200910112588A CN 101659137 A CN101659137 A CN 101659137A
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- overlay
- aircraft
- coating structure
- dielectric constant
- electromagnetic wave
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Abstract
A heat protective coating structure for reducing the temperature of an aircraft kernel relates to a heat protective structure of an aircraft. The invention provides a heat protective coating structurefor reducing the temperature of the aircraft kernel, which can prevent the temperature of the aircraft kernel from being affected by electromagnetic wave radiation, and maintain the stability of thetemperature of the aircraft kernel. The heat protective coating structure is provided with an outer coating layer, a middle coating layer and an inner coating layer, wherein the outer coating layer adopts a high-temperature resistant material with the dielectric constant of n1; the middle coating layer adopts a wave-transmitting material with the dielectric constant of n2; the inner coating layeradopts a high-temperature resistant material with the dielectric constant of n3; n2 is more than n1 and n3, and n3R23/(n1R12) is less than or equal to 1, wherein R12 is the curvature radius of an incident point of incident electromagnetic waves on the separating surface between the outer coating layer and the middle coating layer, and R23 is the curvature radius of the incident point of the incident electromagnetic waves on the separating surface between the middle coating layer and the inner coating layer.
Description
Technical field
The present invention relates to the thermal protection structure of aircraft, especially relate to a kind of thermal protection coating structure that is used for the interior nuclear temperature of reduction aircraft of aircraft outer surface coating.
Background technology
The thermal protection of aircraft is to guarantee that aircraft is under the high-speed flight situation, (be not the air force heating because of pneumatic heating, be meant when object and air or other gas are done the high speed relative motion, the aircraft ambient air produces high-temperature gas because of being subjected to violent compression, diabatic process to object) and the synergy of other thermal force and mechanical load etc. and influence a key technology of aircraft interior components, be to prevent that aircraft is damaged and the safeguard procedures that must take.At present, that can consult relates to thermo-lag pertinent literature of aircraft or patent, mainly all is to study to adopt reduce aircraft someway pneumatic to be heated to be purpose.But the thermal protection of aircraft not only comprises the protection to pneumatic heating, also should comprise the electromagnetic protection of laser, various cosmic rays or the like being carried big energy.Because when these electromagnetic wave irradiations during at aircraft surface, according to the Fresnel formula of revising under the moving interface situation (Wen Shengle as can be known, Yi Huixian. the Fresnel formula during the dielectric surface motion, College Physics, 2001 (11): 4-7), variation along with incident angle, to there be a certain proportion of electromagnetic wave energy to incide in the middle of aircraft coating or the kernel, and along with reducing of incidence angle, be refracted to coating and can become big gradually (according to the change in dielectric constant of different media with energy in the kernel, rate of change between 60～90 degree increases suddenly), no matter but the energy in the coating of being refracted to and the kernel what, all can have more or less influence to the aircraft surface temperature controlling.
In addition, in today that the Mach number of the develop rapidly of modern missilery and space technology, aircraft improves constantly, a kind of special material---electromagnetic wave transparent material in the application in aerospace craft field also more and more widely requires more and more stricter.Along with a large amount of appearances of the high-performance electromagnetic wave transparent material of the saturating ripple of various wave bands, its stable high-frequency dielectric performance, good hot property, good mechanical performance and environmental resistance provide the space of development for the thermal protection of aircraft.
Can see in sum, except factors such as pneumatic heating to the aircraft kernel Temperature Influence, the electromagnetic wave energy radiation influences also obviously, can not ignore it, utilization has the novel electromagnetic wave transparent material of premium properties and adopts special structure to be improved this influence, the technology that adopts new structure and method to solve this influence yet there are no relevant report but at present.
Summary of the invention
The purpose of this invention is to provide a kind of interior nuclear temperature of aircraft that prevents and be subjected to the influence of electromagenetic wave radiation, keep the thermal protection coating structure of the interior nuclear temperature of reduction aircraft of aircraft kernel temperature stabilization.
The present invention is provided with outer overlay, middle overlay and interior overlay, and it is n that outer overlay adopts dielectric constant
1Exotic material, it is n that middle overlay adopts dielectric constant
2Electromagnetic wave transparent material, it is n that interior overlay adopts dielectric constant
3Exotic material, n
2＞n
1, n
2＞n
3, n
3R
23/ (n
1R
12)≤1, wherein, R
12Be the incident electromagnetic wave radius of curvature of the incidence point on the interface of overlay and middle overlay outside, R
23Radius of curvature for the incidence point of incident electromagnetic wave on the interface of middle overlay and interior overlay.
Described dielectric constant is n
1Exotic material preferably ceramic based composites.As select α-Dan Huagui (α-Si for use
3N
4) ceramic matric composite.
Described dielectric constant is n
2The preferred saturating ripple pottery of electromagnetic wave transparent material, as magnesium silicate (Mg
2SiO
4) matrix, this ripple pottery can be to the saturating ripple of the electromagnetic wave of microwave band, saturating Bob is more than 70%.
Described dielectric constant is n
3Exotic material preferably ceramic based composites.As select silica (SiO for use
2) ceramic matric composite.
Described n
3R
23/ (n
1R
12) revolve anyway and preferably be less than or equal to the electromagnetic wave incidence angle during overlay, i.e. arcsin (n from outer overlay is incident to
3R
23/ (n
1R
12))≤i
1, i
1Be electromagnetic wave incidence angle during overlay from outer overlay is incident to.
When the present invention uses, be coated in interior overlay, middle overlay, outer overlay on the aircraft casing successively, because the present invention adopts said structure and requirement, in entering with certain incidence angle (incidence angle outside being incident to by outer overlay on the interface of overlay and middle overlay), electromagnetic wave behind the overlay, can on the interface of middle overlay and interior overlay total reflection take place.Full emission back electromagnetic wave has following two paths and propagates:
1) reflects on the interface of overlay and middle overlay outside, the outside overlay of most of energy is radiate.
2) in middle overlay, transmit, and can transfer to the aircraft end, be radiated outside the space from terminal (empennage) with the waveguide form.
This shows, will significantly reduce the incident electromagnetic wave energy to aircraft kernel Temperature Influence through the present invention.During practical application, middle overlay can be adopted the electromagnetic wave transparent material of the saturating ripple of different frequency, then will reduce various frequency irradiation of electromagnetic waves energy to a greater degree, thereby more effectively the aircraft kernel be played the thermal protection effect.
Description of drawings
Fig. 1 is the structure and the simulation application principle schematic of embodiments of the invention.
The specific embodiment
The invention will be further described below in conjunction with embodiment and accompanying drawing.
Referring to Fig. 1, the present invention constitutes (promptly outer overlay 1, middle overlay 2 and interior overlay 3) by three layers of overlay.n
1, n
2, n
3Be respectively the dielectric constant of outer overlay 1, middle overlay 2 and interior overlay 3.Outer overlay 1 is n
1Refractory ceramics based composites (α-Dan Huagui (α-Si of=1.5
3N
4)), middle overlay 2 is n
2=3.64 saturating ripple pottery (magnesium silicate (Mg
2SiO
4) matrix), this magnesium silicate (Mg
2SiO
4) the saturating ripple pottery of matrix can be to the saturating ripple of the electromagnetic wave of microwave band, saturating Bob is more than 70%.Interior overlay 3 is n
3=1.00 ceramic matric composite (SiO
2).
In Fig. 1, P1 represents the aircraft head, P2 represents the external freedom space, the hollow arrow direction is represented the direct of travel of plane nose P1, solid arrow is represented the direction of propagation of electromagnetic wave in the thermal protection coating structure, and S1 is the interface of outer overlay 1 and middle overlay 2, and S2 is the interface of middle overlay 2 and interior overlay 3, letter O, E, A, B, C represent the incidence point of electromagnetic wave when three interfaces, i respectively
1Be the incident angle value on the S1.R
12Be the residing camber line radius of curvature of incidence point E, R
23Be the residing camber line radius of curvature of incidence point A, R
23/ R
12=0.98.
The frequency f of the incident electromagnetic wave that adopts in the experiment is that 9GHZ, projectile energy are 49.2MW.When electromagnetic wave from the point of O foremost of aircraft head P1 of simulation with certain incidence angle i
1During incident, test result sees Table 1.
Table 1
| Incidence angle i 1(degree) | The transmission potential (mw) of overlay in being transmitted through | Percent transmission=transmission potential/projectile energy (%) |
| ??0 | ??41.6 | ?84.5 |
| ??15 | ??40.8 | ?82.9 |
| ??30 | ??40.0 | ?81.3 |
| ??35 | ??39.2 | ?79.7 |
| ??38 | ??38.5 | ?78.2 |
| ??39 | ??38.1 | ?77.4 |
| ??40 | ??37.6 | ?76.4 |
| ??40.2 | ??36.7 | ?74.6 |
| ??40.4 | ??36.1 | ?73.4 |
| ??40.6 | ??34.6 | ?70.3 |
| ??40.7 | ??34.4 | ?69.9 |
| ??40.8 | ??0.62 | ?1.26 |
| ??40.9 | ??0.62 | ?1.26 |
| ??41 | ??0.62 | ?1.26 |
| ??41.2 | ??0.61 | ?1.24 |
| ??41.4 | ??0.62 | ?1.26 |
| ??41.6 | ??0.60 | ?1.22 |
| ??42 | ??0.60 | ?1.22 |
| ??45 | ??0.58 | ?1.18 |
| ??50 | ??0.57 | ?1.16 |
| ??60 | ??0.53 | ?1.08 |
| ??75 | ??0.51 | ?1.04 |
| ??90 | ??0.50 | ?1.02 |
From table 1, can find out, as incidence angle i
1When 0 °～40.8 ° scope, through equipment such as light power meter and silicon photoelectromagnetic can record be transmitted in Electromgnetically-transparent energy 〉=34.4mw in the overlay, and incidence angle i
1When being 40.8 °, the Electromgnetically-transparent energy is undergone mutation, and only is 0.62mw, and when firing angle 〉=40.8 °, the Electromgnetically-transparent energy of actual measurement is below 0.62mw (theory should be 0mw, and measured data is due to the electromagnetic interference wave energy in the experimental situation) all.This illustrates i
1Being 40.8 ° is critical angle, works as i
1In the time of 〉=40.8 °, electromagnetic wave is sent out the phenomenon that will penetrate at A point place, that is to say that electromagnetic wave energy can be transmitted through interior overlay hardly, therefore, aircraft head P1 (aircraft kernel) has been played the thermal protection effect.
Satisfying n
2＞n
1, n
2＞n
3, n
3R
23/ (n
1R
12Under the prerequisite of)≤1, if change parameter n1, n2, n3, R
23/ R
12, f (frequency of incident electromagnetic wave), critical angle i
1Can change.Table 2 has provided change parameter n1, n2, n3, R
23/ R
12, f the time, critical angle i
1Test result.
Table 2
| Change parameter | ??n 1 | ??n 2 | ??n 3 | ??R 23/R 12 | ??f(GHz) | ??i 1 |
| Only change n 1Value | ??1.74 | ??3.64 | ??1 | ??0.98 | ??9 | ??34.3° |
| Only change n 2Value | ??1.5 | ??3.96 | ??1 | ??0.98 | ??9 | ??40.8° |
| Only change n 3Value | ??1.5 | ??3.64 | ??1.3 | ??0.98 | ??9 | ??58.1° |
| Only change R 23/R 12Value | ??1.5 | ??3.64 | ??1 | ??0.94 | ??9 | ??38.8° |
| Only change the frequency f value | ??1.5 | ??3.64 | ??1 | ??0.98 | ??8.5 | ??40.8° |
From table 2, can find out the influence that the parameter that changed changes critical angle.n
1With n
3Between difference big more, the critical angle that satisfies condition is more little; R
23/ R
12More little, the critical angle that satisfies condition is also more little; The variation of n2 can not have influence on incidence angle i
1Size; Changing wave frequency f can be to critical angle i
1Exert an influence.
Claims (8)
1, reduce the thermal protection coating structure of nuclear temperature in the aircraft, it is characterized in that being provided with outer overlay, middle overlay and interior overlay, it is n that outer overlay adopts dielectric constant
1Exotic material, it is n that middle overlay adopts dielectric constant
2Electromagnetic wave transparent material, it is n that interior overlay adopts dielectric constant
3Exotic material, n
2＞n
1, n
2＞n
3, n
3R
23/ (n
1R
12)≤1, wherein, R
12Be the incident electromagnetic wave radius of curvature of the incidence point on the interface of overlay and middle overlay outside, R
23Radius of curvature for the incidence point of incident electromagnetic wave on the interface of middle overlay and interior overlay.
2, the thermal protection coating structure of nuclear temperature in the reduction aircraft as claimed in claim 1 is characterized in that described dielectric constant is n
1Exotic material, be ceramic matric composite.
3, the thermal protection coating structure of nuclear temperature in the reduction aircraft as claimed in claim 2 is characterized in that described ceramic matric composite is the ceramic matric composite of α-Dan Huagui.
4, the thermal protection coating structure of nuclear temperature in the reduction aircraft as claimed in claim 1 is characterized in that described dielectric constant is n
2Electromagnetic wave transparent material, be saturating ripple pottery.
5, the thermal protection coating structure of nuclear temperature in the reduction aircraft as claimed in claim 4 is characterized in that described ripple pottery is the saturating ripple pottery of magnesium silicate.
6, the thermal protection coating structure of nuclear temperature in the reduction aircraft as claimed in claim 1 is characterized in that described dielectric constant is n
3Exotic material, be ceramic matric composite.
7, the thermal protection coating structure of nuclear temperature in the reduction aircraft as claimed in claim 6 is characterized in that described ceramic matric composite is the ceramic matric composite of silica.
8, the thermal protection coating structure of nuclear temperature in the reduction aircraft as claimed in claim 1 is characterized in that described n
3R
23/ (n
1R
12) revolve anyway and be less than or equal to electromagnetic wave incidence angle i during overlay from outer overlay is incident to
1, i.e. arcsin (n
3R
23/ (n
1R
12))≤i
1
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN200910112588A CN101659137A (en) | 2009-09-25 | 2009-09-25 | Heat protective coating structure for reducing temperature of aircraft kernel |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN200910112588A CN101659137A (en) | 2009-09-25 | 2009-09-25 | Heat protective coating structure for reducing temperature of aircraft kernel |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN101659137A true CN101659137A (en) | 2010-03-03 |
Family
ID=41787483
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN200910112588A Pending CN101659137A (en) | 2009-09-25 | 2009-09-25 | Heat protective coating structure for reducing temperature of aircraft kernel |
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| Country | Link |
|---|---|
| CN (1) | CN101659137A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103723269A (en) * | 2013-09-11 | 2014-04-16 | 太仓派欧技术咨询服务有限公司 | Thermal protection structure |
-
2009
- 2009-09-25 CN CN200910112588A patent/CN101659137A/en active Pending
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103723269A (en) * | 2013-09-11 | 2014-04-16 | 太仓派欧技术咨询服务有限公司 | Thermal protection structure |
| CN103723269B (en) * | 2013-09-11 | 2018-01-30 | 太仓派欧技术咨询服务有限公司 | A kind of thermal protection structure |
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|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
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| RJ01 | Rejection of invention patent application after publication | ||
| RJ01 | Rejection of invention patent application after publication |
Application publication date: 20100303 |