CN103144759A - Shock-resistant composite fuselage panel - Google Patents
Shock-resistant composite fuselage panel Download PDFInfo
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- CN103144759A CN103144759A CN2013100652477A CN201310065247A CN103144759A CN 103144759 A CN103144759 A CN 103144759A CN 2013100652477 A CN2013100652477 A CN 2013100652477A CN 201310065247 A CN201310065247 A CN 201310065247A CN 103144759 A CN103144759 A CN 103144759A
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Abstract
The invention provides a fuselage cover for an aircraft, which comprises a sandwich structure, wherein the sandwich structure is prepared from a metal/fiber/ceramic laminar composite material; and the fuselage cover is provided with at least one metal layer/fiber layer/ceramic layer sandwich structure. The invention is characterized in that the metal layer adopts aluminum, magnesium, titanium or a corresponding alloy material; the fiber layer adopts glass fibers, Kevlar fibers, carbon fibers, silicon nitride, silicon carbide or zirconium dioxide fibers; and the ceramic layer comprises zirconium oxide, yttrium oxide, aluminum oxide and mullite. The fuselage cover provided by the invention has the advantages of high hardness, favorable toughness, light weight and favorable shock resistance.
Description
Technical field
The present invention relates to a kind of shock resistance composite fuselage panel, relate in particular to a kind of fuselage cover with sandwich structure.
Background technology
Current, the large scale business jet airplane has all adopted high thrust turbofan aero-engine usually, the turbofan aero-engine of the type has all used large-sized fan blade, maximum fan blade diameter can reach 3m, during work, the tangential speed at fan blade tip place surpasses 450m/s, the development of following turbofan aero-engine, the tangential speed of turbofan blade tip can be higher.The blade of high-speed operation is subjected to the impact of foreign object strike damage or high-frequency vibration fatigue etc., inevitably the leaf destruction fault can occur.Broken blade has very high energy, if blade punctures engine nacelle, may produce infringement near the fuselage cover the driving engine installation site, and then jeopardize birdman's safety.Current airframe shell adopts light-weight metal magnalium titanium or their alloy to make usually, also some aircraft adopts composite material, but current fuselage cover still is difficult to keep out the broken blade that has as the aforementioned heavy impulse very or is other shock.
Summary of the invention
in order to overcome above-mentioned shortcoming and drawback, the invention provides a kind of fuselage cover for aircraft, comprise sandwich structure, it uses metal, fiber and ceramic laminar composite material are made, has the sandwich structure that at least one metal level/fibrage/ceramic layer consists of, it is characterized in that metal level adopts aluminium, magnesium, titanium or corresponding alloy material, fibrage adopts glass fibre, the kevlar fiber, carbon fiber, silicon nitride, carborundum or zirconia fiber, described ceramic layer comprises the zirconia of weight ratio 100:8:3:2 or 100:5:5:2, yttria, aluminium oxide and mullite.
Preferably, described zirconia ceramics material adopts zirconia, yttria, aluminium oxide and the mullite of weight ratio 100:5:5:2, add the distilled water heavy with aforementioned three's gross weight etc., ball milling is 5.5 hours in ball grinding mill, then drying, granulation, moulding, sintering is 2.0 hours at the temperature of 1650 degrees centigrade, and total temperature rise time is 9 hours; Be cooled to 1250 degrees centigrade of heat treatments 4.0 hours with 230 degrees centigrade of speed hourly again, then naturally cool to room temperature, then sample is warming up to 1480 degrees centigrade of heat treatments 1.5 hours, then again naturally cool to room temperature and obtain.
Preferably, described zirconia ceramics material adopts zirconia, yttria, aluminium oxide and the mullite of weight ratio 100:8:3:2, add the distilled water heavy with aforementioned three's gross weight etc., ball milling is 5 hours in ball grinding mill, then drying, granulation, moulding, sintering is 1.8 hours at the temperature of 1700 degrees centigrade, and total temperature rise time is 8 hours; Be cooled to 1200 degrees centigrade of heat treatments 3.5 hours with 220 comfort level speed hourly again, then naturally cool to room temperature, then sample is warming up to 1460 degrees centigrade of heat treatments 1.6 hours, then again naturally cool to room temperature and obtain.
Preferably, described fuselage cover is followed successively by metal level, fibrage, ceramic layer from inside to outside, lamination coating can be that winding of single layer can be also multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, thickness 1.5mm～the 3mm of metal level, the thickness 5mm～8mm of ceramic layer.
Preferably, fuselage cover is followed successively by metal level, fibrage, ceramic layer, fibrage, metal level from inside to outside, lamination coating can be that winding of single layer can be also multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, thickness 1.5mm～the 2mm of inner layer metal layer, the thickness 3mm～6mm of ceramic layer, the thickness 1.5mm～3.0mm of outer layer metal layer.
Preferably, fuselage cover is followed successively by metal level, fibrage, ceramic layer, fibrage, ceramic layer from inside to outside, lamination coating can be that winding of single layer can be also multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, thickness 1.5mm～the 3mm of metal level, the thickness 3mm～5mm of internal layer ceramic layer, the thickness 4mm～6mm of outer pottery.
Preferably, the Polymer resin materials such as described sandwich structure employing epoxy resin or polyimide are processed by curing as adhesive agent metal level, fibrage and ceramic layer are bondd.
Owing to having adopted high tenacity, porous zirconia stupalith in the present invention, fuselage cover according to the present invention has very excellent shock resistance, has simultaneously lower density.
The specific embodiment
Airframe shell in the present invention has sandwich structure, and it uses metal, fiber and ceramic laminar composite material to make, and has the sandwich structure that at least one metal level/fibrage/ceramic layer consists of.Metal level wherein for example adopts aluminium, magnesium, titanium or corresponding alloy material, fibrage adopts glass fibre, kevlar fiber, carbon fiber, silicon nitride, carborundum or zirconia fiber, and described stupalith is a kind of zirconia ceramics of high tenacity porous.
Described zirconia ceramics material adopts zirconia, yttria, aluminium oxide and the mullite of weight ratio 100:8:3:2, add the distilled water heavy with aforementioned three's gross weight etc., ball milling is 5 hours in ball grinding mill, then drying, granulation, moulding, in sintering at the temperature of 1700 degrees centigrade, under 32.5MPa pressure 1.8 hours, total temperature rise time was 8 hours; Be cooled to 1200 degrees centigrade of heat treatments 3.5 hours with 220 comfort level speed hourly again, then naturally cool to room temperature, then sample is warming up to 1460 degrees centigrade of heat treatments 1.6 hours, then again naturally cool to room temperature and get final product.
In another embodiment, described zirconia ceramics material adopts zirconia, yttria, aluminium oxide and the mullite of weight ratio 100:5:5:2, add the distilled water heavy with aforementioned three's gross weight etc., ball milling is 5.5 hours in ball grinding mill, then drying, granulation, moulding, in sintering at the temperature of 1650 degrees centigrade, under 32.5MPa pressure 2.0 hours, total temperature rise time was 9 hours; Be cooled to 1250 degrees centigrade of heat treatments 4.0 hours with 230 degrees centigrade of speed hourly again, then naturally cool to room temperature, then sample is warming up to 1480 degrees centigrade of heat treatments 1.5 hours, then again naturally cool to room temperature and get final product.
Described sandwich structure adopts the Polymer resin materials such as epoxy resin or polyimide to process by curing as adhesive agent metal level, fibrage and ceramic layer is bondd.
In one embodiment, a kind of dull and stereotyped layered composite structure fuselage cover of manufacturing, be followed successively by from inside to outside metal level, fibrage, ceramic layer, lamination coating can be that winding of single layer can be also multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, the thickness 1.5mm～3mm of metal level, the thickness 5mm～8mm of ceramic layer.
In another embodiment, a kind of dull and stereotyped layered composite structure fuselage cover of manufacturing, be followed successively by from inside to outside metal level, fibrage, ceramic layer, fibrage, metal level, lamination coating can be that winding of single layer can be also multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, the thickness 1.5mm～2mm of inner layer metal layer, thickness 3mm～the 6mm of ceramic layer, the thickness 1.5mm～3.0mm of outer layer metal layer.
In another embodiment, a kind of dull and stereotyped layered composite structure fuselage cover of manufacturing, be followed successively by from inside to outside metal level, fibrage, ceramic layer, fibrage, ceramic layer, lamination coating can be that winding of single layer can be also multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, thickness 1.5mm～the 3mm of metal level, the thickness 3mm～5mm of internal layer ceramic layer, the thickness 4mm～6mm of outer pottery.
Certainly, the fuselage cover in the present invention also can only be applied to partly near aero-engine installation site near zone.
Aforementioned different embodiment and above-mentioned three specific embodiments about fuselage cover about zirconia ceramics can make up.And those skilled in the art can make replacement or modification to content of the present invention according to content disclosed by the invention and the art technology of grasping; but these replacements or modification should not be considered as breaking away from the present invention's design, and these replacements or modification are all in the claimed interest field of the present invention.
Claims (7)
1. fuselage cover that is used for aircraft, comprise sandwich structure, it uses metal, fiber and ceramic laminar composite material to make, has the sandwich structure that at least one metal level/fibrage/ceramic layer consists of, it is characterized in that metal level adopts aluminium, magnesium, titanium or corresponding alloy material, fibrage adopts glass fibre, kevlar fiber, carbon fiber, silicon nitride, carborundum or zirconia fiber, and described ceramic layer comprises zirconia, yttria, aluminium oxide and the mullite of weight ratio 100:8:3:2 or 100:5:5:2.
2. fuselage cover according to claim 1, it is characterized in that described zirconia ceramics material adopts zirconia, yttria, aluminium oxide and the mullite of weight ratio 100:5:5:2, add the distilled water heavy with aforementioned three's gross weight etc., ball milling is 5.5 hours in ball grinding mill, then drying, granulation, moulding, sintering is 2.0 hours at the temperature of 1650 degrees centigrade, and total temperature rise time is 9 hours; Be cooled to 1250 degrees centigrade of heat treatments 4.0 hours with 230 degrees centigrade of speed hourly again, then naturally cool to room temperature, then sample is warming up to 1480 degrees centigrade of heat treatments 1.5 hours, then again naturally cool to room temperature and obtain.
3. fuselage cover according to claim 1, it is characterized in that described zirconia ceramics material adopts zirconia, yttria, aluminium oxide and the mullite of weight ratio 100:8:3:2, add the distilled water heavy with aforementioned three's gross weight etc., ball milling is 5 hours in ball grinding mill, then drying, granulation, moulding, sintering is 1.8 hours at the temperature of 1700 degrees centigrade, and total temperature rise time is 8 hours; Be cooled to 1200 degrees centigrade of heat treatments 3.5 hours with 220 comfort level speed hourly again, then naturally cool to room temperature, then sample is warming up to 1460 degrees centigrade of heat treatments 1.6 hours, then again naturally cool to room temperature and obtain.
4. according to claim 1-3 described fuselage covers, it is characterized in that described fuselage cover is followed successively by metal level, fibrage, ceramic layer from inside to outside, lamination coating can be that winding of single layer can be also multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, thickness 1.5mm～the 3mm of metal level, the thickness 5mm～8mm of ceramic layer.
5. according to claim 1-3 described fuselage covers, it is characterized in that fuselage cover is followed successively by metal level, fibrage, ceramic layer, fibrage, metal level from inside to outside, lamination coating can be that winding of single layer can be also multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, thickness 1.5mm～the 2mm of inner layer metal layer, the thickness 3mm～6mm of ceramic layer, the thickness 1.5mm～3.0mm of outer layer metal layer.
6. according to claim 1-3 described fuselage covers, it is characterized in that fuselage cover is followed successively by metal level, fibrage, ceramic layer, fibrage, ceramic layer from inside to outside, lamination coating can be that winding of single layer can be also multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, thickness 1.5mm～the 3mm of metal level, the thickness 3mm～5mm of internal layer ceramic layer, the thickness 4mm～6mm of outer pottery.
7. fuselage cover according to claim 1, is characterized in that described sandwich structure adopts the Polymer resin materials such as epoxy resin or polyimide to process by curing as adhesive agent metal level, fibrage and ceramic layer are bondd.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201310065247.7A CN103144759B (en) | 2013-03-01 | 2013-03-01 | Shock-resistant composite fuselage panel |
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| CN201310065247.7A CN103144759B (en) | 2013-03-01 | 2013-03-01 | Shock-resistant composite fuselage panel |
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| CN103144759A true CN103144759A (en) | 2013-06-12 |
| CN103144759B CN103144759B (en) | 2015-06-10 |
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| CN201310065247.7A Active CN103144759B (en) | 2013-03-01 | 2013-03-01 | Shock-resistant composite fuselage panel |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105537594A (en) * | 2016-03-08 | 2016-05-04 | 苏州珍展科技材料有限公司 | Resin-aluminum-based layered composite fan blade |
| CN107226192A (en) * | 2017-05-28 | 2017-10-03 | 珠海磐磊智能科技有限公司 | A kind of composite board and aircraft |
| CN108602126A (en) * | 2016-02-08 | 2018-09-28 | 西门子股份公司 | Method and apparatus for manufacturing component |
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| CN102701735A (en) * | 2012-06-08 | 2012-10-03 | 武汉工程大学 | Method for preparing stable zirconia/mullite ceramic material |
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2013
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| US4875616A (en) * | 1988-08-10 | 1989-10-24 | America Matrix, Inc. | Method of producing a high temperature, high strength bond between a ceramic shape and metal shape |
| DE19628105A1 (en) * | 1996-07-12 | 1997-11-06 | Daimler Benz Ag | Multilayered light armour element |
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| CN108602126A (en) * | 2016-02-08 | 2018-09-28 | 西门子股份公司 | Method and apparatus for manufacturing component |
| CN105537594A (en) * | 2016-03-08 | 2016-05-04 | 苏州珍展科技材料有限公司 | Resin-aluminum-based layered composite fan blade |
| CN107226192A (en) * | 2017-05-28 | 2017-10-03 | 珠海磐磊智能科技有限公司 | A kind of composite board and aircraft |
| CN107226192B (en) * | 2017-05-28 | 2020-10-23 | 珠海磐磊智能科技有限公司 | Composite board and aircraft |
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| Publication number | Publication date |
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| CN103144759B (en) | 2015-06-10 |
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