CN103144759B - Shock-resistant composite fuselage panel - Google Patents
Shock-resistant composite fuselage panel Download PDFInfo
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- CN103144759B CN103144759B CN201310065247.7A CN201310065247A CN103144759B CN 103144759 B CN103144759 B CN 103144759B CN 201310065247 A CN201310065247 A CN 201310065247A CN 103144759 B CN103144759 B CN 103144759B
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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-resistant composite fuselage panel, particularly relate to a kind of fuselage cover with sandwich structure.
Background technology
Current, large scale business jet airplane all have employed high thrust turbofan aero-engine usually, the turbofan aero-engine of the type all employ large-sized fan blade, maximum fan blade diameter can reach 3m, during work, the tangential speed at fan blade tip place is more than 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 subject to the impact of foreign object strike damage or high-frequency vibration fatigue etc., inevitably leaf destruction fault.Broken blade has very high energy, if blade punctures engine nacelle, then may produce infringement to the fuselage cover near engine mounting positions, and then jeopardize birdman's safety.Current aircraft fuselage cover adopts light-weight metal magnalium titanium or their alloy to manufacture usually, also some aircraft then adopts composite material, but current fuselage cover is still difficult to keep out the broken blade as the aforementioned with very heavy impulse or the shock being other.
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, there is the sandwich structure that at least one metal level/fibrage/ceramic layer is formed, 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, 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 the zirconia of weight ratio 100:5:5:2, yttria, aluminium oxide and mullite, add the distilled water weighed with aforementioned four gross weights etc., ball milling 5.5 hours in ball grinding mill, then drying, granulation, shaping, at the temperature of 1650 degrees Celsius, sinter 2.0 hours, total temperature rise time is 9 hours; Be cooled to 1250 degrees Celsius of heat treatments 4.0 hours with 230 degrees Celsius of speed hourly again, then naturally cool to room temperature, then sample is warming up to 1480 degrees Celsius of heat treatments 1.5 hours, then again naturally cool to room temperature and obtain.
Preferably, described zirconia ceramics material adopts the zirconia of weight ratio 100:8:3:2, yttria, aluminium oxide and mullite, add the distilled water weighed with aforementioned four gross weights etc., ball milling 5 hours in ball grinding mill, then drying, granulation, shaping, at the temperature of 1700 degrees Celsius, sinter 1.8 hours, total temperature rise time is 8 hours; Be cooled to 1200 degrees Celsius 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 Celsius 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 winding of single layer also can be 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 winding of single layer also can be multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, thickness 1.5mm ~ the 2mm of interior 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 winding of single layer also can be 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, described sandwich structure adopts the macromolecule resin material such as epoxy resin or polyimide metal level, fibrage and ceramic layer bonding to be got up by solidification process as adhesive agent.
Owing to have employed high tenacity, porous zirconia stupalith in the present invention, fuselage cover according to the present invention has very excellent shock resistance, has lower density simultaneously.
Detailed description of the invention
Aircraft fuselage cover 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 is formed.Metal layer is wherein as adopted 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 the zirconia of weight ratio 100:8:3:2, yttria, aluminium oxide and mullite, add the distilled water weighed with aforementioned four gross weights etc., ball milling 5 hours in ball grinding mill, then drying, granulation, shaping, at the temperature of 1700 degrees Celsius, 32.5MPa fired under pressure 1.8 hours, total temperature rise time is 8 hours; Be cooled to 1200 degrees Celsius 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 Celsius of heat treatments 1.6 hours, then again naturally cool to room temperature.
In another embodiment, described zirconia ceramics material adopts the zirconia of weight ratio 100:5:5:2, yttria, aluminium oxide and mullite, add the distilled water weighed with aforementioned four gross weights etc., ball milling 5.5 hours in ball grinding mill, then drying, granulation, shaping, at the temperature of 1650 degrees Celsius, 32.5MPa fired under pressure 2.0 hours, total temperature rise time is 9 hours; Be cooled to 1250 degrees Celsius of heat treatments 4.0 hours with 230 degrees Celsius of speed hourly again, then naturally cool to room temperature, then sample is warming up to 1480 degrees Celsius of heat treatments 1.5 hours, then again naturally cool to room temperature.
Described sandwich structure adopts the macromolecule resin material such as epoxy resin or polyimide metal level, fibrage and ceramic layer bonding to be got up by solidification process as adhesive agent.
In one embodiment, the dull and stereotyped layered composite structure fuselage cover of manufactured one, be followed successively by metal level, fibrage, ceramic layer from inside to outside, lamination coating can be winding of single layer also can be 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, the dull and stereotyped layered composite structure fuselage cover of manufactured one, be followed successively by metal level, fibrage, ceramic layer, fibrage, metal level from inside to outside, lamination coating can be winding of single layer also can be multiple wraps, if multiple wraps, preferably between multilayer, machine direction is crisscross arranged, the thickness 1.5mm ~ 2mm of interior metal layer, thickness 3mm ~ the 6mm of ceramic layer, the thickness 1.5mm ~ 3.0mm of outer layer metal layer.
In another embodiment, the dull and stereotyped layered composite structure fuselage cover of manufactured one, be followed successively by metal level, fibrage, ceramic layer, fibrage, ceramic layer from inside to outside, lamination coating can be winding of single layer also can be 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.
The aforementioned different embodiment about zirconia ceramics and above-mentioned three specific embodiments about fuselage cover can combine.And those skilled in the art can make replacement or modification according to content disclosed by the invention and the art technology grasped to content of the present invention; but these replacements or modification should not be considered as disengaging the present invention design, and these replacements or modification are all in the interest field of application claims protection.
Claims (1)
1. the fuselage cover for aircraft, comprise sandwich structure, it uses metal, fiber and ceramic laminar composite material to make, there is the sandwich structure that at least one metal level and fibrage and ceramic layer are formed, metal level adopts aluminium, magnesium or titanium, fibrage adopts glass fibre, kevlar fiber, carbon fiber, silicon nitride, carborundum or zirconia fiber, and described ceramic layer comprises the zirconia of weight ratio 100:8:3:2, yttria, aluminium oxide and mullite; It is characterized in that the zirconia of zirconia ceramics material employing weight ratio 100:8:3:2, yttria, aluminium oxide and mullite, add the distilled water weighed with aforementioned four gross weights etc., ball milling 5 hours in ball grinding mill, then drying, granulation, shaping, at the temperature of 1700 degrees Celsius, 32.5MPa fired under pressure 1.8 hours, total temperature rise time is 8 hours; Be cooled to 1200 degrees Celsius of heat treatments 3.5 hours with 220 degrees Celsius of speed hourly again, then naturally cool to room temperature, then sample is warming up to 1460 degrees Celsius of heat treatments 1.6 hours, then again naturally cool to room temperature and obtain.
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| Application Number | Priority Date | Filing Date | Title |
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| 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 CN103144759A (en) | 2013-06-12 |
| CN103144759B true CN103144759B (en) | 2015-06-10 |
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Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
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| DE102016201838A1 (en) * | 2016-02-08 | 2017-08-10 | Siemens Aktiengesellschaft | Method for producing a component and device |
| CN105537594B (en) * | 2016-03-08 | 2017-11-14 | 许晓丽 | A kind of resin aluminum-based layered composite material fan blade |
| 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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| CN101186499A (en) * | 2007-12-14 | 2008-05-28 | 天津大学 | Zirconium oxide quaternary system composite ceramic material containing mullite component |
| CN100497089C (en) * | 2006-09-27 | 2009-06-10 | 北京航空航天大学 | Fibre-reinforced metal/ceramic sheet-like composite container casing and its manufacture method |
| 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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| Publication number | Priority date | Publication date | Assignee | Title |
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| DE10310945A1 (en) * | 2003-03-13 | 2004-10-07 | Sgl Carbon Ag | Fiber-reinforced ceramic material |
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| Publication number | Priority date | Publication date | Assignee | Title |
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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 |
| CN1288794A (en) * | 1999-08-12 | 2001-03-28 | 印杰克斯有限公司 | Method for producing screw |
| CN1418848A (en) * | 2002-12-25 | 2003-05-21 | 天津大学 | Heterogeneous ceramic material containing silicon phase quatermary system zicronium oxide |
| CN1724465A (en) * | 2005-06-03 | 2006-01-25 | 中国科学院上海硅酸盐研究所 | The yttrium aluminum garnet transparent ceramic material and the preparation method of codope |
| CN100497089C (en) * | 2006-09-27 | 2009-06-10 | 北京航空航天大学 | Fibre-reinforced metal/ceramic sheet-like composite container casing and its manufacture method |
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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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