CN103158858B - A kind of aircraft wing shell - Google Patents
A kind of aircraft wing shell Download PDFInfo
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- CN103158858B CN103158858B CN201310066804.7A CN201310066804A CN103158858B CN 103158858 B CN103158858 B CN 103158858B CN 201310066804 A CN201310066804 A CN 201310066804A CN 103158858 B CN103158858 B CN 103158858B
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Abstract
The invention provides a kind of wing shell 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/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, and described ceramic layer comprises zirconia, magnesium oxide, calcium oxide, yttria and cerium oxide.Wing shell hardness of the present invention is high, good toughness, quality light, has good shock resistance.
Description
Technical field
The present invention relates to a kind of aircraft wing shell, particularly relate to a kind of wing shell 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 wing shell near engine mounting positions, and then jeopardize birdman's safety.Current aircraft wing shell adopts light-weight metal magnalium titanium or their alloy to manufacture usually, also some aircraft then adopts composite material, but current wing shell 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 wing shell 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 100:6:4:2:1 or 100:8:3:4:1, magnesium oxide, calcium oxide, yttria and cerium oxide.
Preferably, described zirconia ceramics material adopts the zirconia of weight ratio 100:6:4:2:1, magnesium oxide, calcium oxide, yttria and cerium oxide, add the distilled water weighed with aforementioned Multiple components gross weight etc., ball milling 5 hours in ball grinding mill, then drying, granulation, shaping, at the temperature of 1600-1750 degree 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.
Preferably, described zirconia ceramics material adopts the zirconia of weight ratio 100:8:3:4:1, magnesium oxide, calcium oxide, yttria and cerium oxide, add the distilled water weighed with aforementioned Multiple components gross weight etc., ball milling 5.5 hours in ball grinding mill, then drying, granulation, shaping, at the temperature of 1600-1700 degree 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.
Preferably, described wing shell 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, wing shell 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, wing shell 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, wing shell according to the present invention has very excellent shock resistance, has lower density simultaneously.
Detailed description of the invention
Aircraft wing 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 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 100:6:4:2:1, magnesium oxide, calcium oxide, yttria and cerium oxide, add the distilled water weighed with aforementioned Multiple components gross weight etc., ball milling 5 hours in ball grinding mill, then drying, granulation, shaping, at the temperature of 1600-1750 degree 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.
In another embodiment, described zirconia ceramics material adopts the zirconia of weight ratio 100:8:3:4:1, magnesium oxide, calcium oxide, yttria and cerium oxide, add the distilled water weighed with aforementioned Multiple components gross weight etc., ball milling 5.5 hours in ball grinding mill, then drying, granulation, shaping, at the temperature of 1600-1700 degree 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.
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, manufactured one dull and stereotyped layered composite structure wing shell, 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, manufactured one dull and stereotyped layered composite structure wing shell, 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, manufactured one dull and stereotyped layered composite structure wing shell, 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 wing shell 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 wing shell 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 wing shell for aircraft, comprise sandwich structure, it uses metal, fiber and ceramic laminar composite material are made, be followed successively by metal level from inside to outside, fibrage, ceramic layer, fibrage, metal level, metal level adopts aluminium, magnesium, titanium or their alloy material, fibrage adopts glass fibre, kevlar fiber, carbon fiber, silicon nitride, carborundum or zirconia fiber, it is characterized in that, ceramic layer adopts the zirconia ceramics of high tenacity porous, this zirconia ceramics material adopts the zirconia of weight ratio 100:6:4:2:1, magnesium oxide, calcium oxide, yttria and cerium oxide, add the distilled water weighed with aforementioned Multiple components gross weight etc., ball milling 5 hours in ball grinding mill, then dry, granulation, shaping, 1.8 hours are sintered at the temperature of 1600-1750 degree Celsius, 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 can obtain zirconia ceramics material.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201310066804.7A CN103158858B (en) | 2013-03-01 | 2013-03-01 | A kind of aircraft wing shell |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201310066804.7A CN103158858B (en) | 2013-03-01 | 2013-03-01 | A kind of aircraft wing shell |
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| Publication Number | Publication Date |
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| CN103158858A CN103158858A (en) | 2013-06-19 |
| CN103158858B true CN103158858B (en) | 2015-11-25 |
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| CN201310066804.7A Active CN103158858B (en) | 2013-03-01 | 2013-03-01 | A kind of aircraft wing shell |
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4759971A (en) * | 1985-10-07 | 1988-07-26 | Occidental Research Corporation | Multilayered structure |
| CN1052733A (en) * | 1989-12-21 | 1991-07-03 | 施宁 | The light compound material board shell of anti-springing and anti-striking |
| CN2406216Y (en) * | 1999-09-03 | 2000-11-15 | 北京航空航天大学 | Advanced composite-material anti-bullet board |
| CN1477081A (en) * | 2003-07-11 | 2004-02-25 | 清华大学 | High-toughness Al2O3/Ti3SiC2 laminate ceramic composite material and its hot-pressing preparation method |
| US6838162B1 (en) * | 1999-09-16 | 2005-01-04 | Sgl Technik Gmbh | Fiber-bundle-reinforced composite material having a ceramic matrix, method for manufacturing a composite material and method for manufacturing elements formed of a composite material |
| CN1708396A (en) * | 2002-10-30 | 2005-12-14 | 戈尔企业控股股份有限公司 | Porous polymeric membrane toughened composites |
-
2013
- 2013-03-01 CN CN201310066804.7A patent/CN103158858B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4759971A (en) * | 1985-10-07 | 1988-07-26 | Occidental Research Corporation | Multilayered structure |
| CN1052733A (en) * | 1989-12-21 | 1991-07-03 | 施宁 | The light compound material board shell of anti-springing and anti-striking |
| CN2406216Y (en) * | 1999-09-03 | 2000-11-15 | 北京航空航天大学 | Advanced composite-material anti-bullet board |
| US6838162B1 (en) * | 1999-09-16 | 2005-01-04 | Sgl Technik Gmbh | Fiber-bundle-reinforced composite material having a ceramic matrix, method for manufacturing a composite material and method for manufacturing elements formed of a composite material |
| CN1708396A (en) * | 2002-10-30 | 2005-12-14 | 戈尔企业控股股份有限公司 | Porous polymeric membrane toughened composites |
| CN1477081A (en) * | 2003-07-11 | 2004-02-25 | 清华大学 | High-toughness Al2O3/Ti3SiC2 laminate ceramic composite material and its hot-pressing preparation method |
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| CN103158858A (en) | 2013-06-19 |
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