CN101342941A - Disposal solidifying and molding method for fuselage ring and outer panel skin - Google Patents
Disposal solidifying and molding method for fuselage ring and outer panel skin Download PDFInfo
- Publication number
- CN101342941A CN101342941A CNA2008101369766A CN200810136976A CN101342941A CN 101342941 A CN101342941 A CN 101342941A CN A2008101369766 A CNA2008101369766 A CN A2008101369766A CN 200810136976 A CN200810136976 A CN 200810136976A CN 101342941 A CN101342941 A CN 101342941A
- Authority
- CN
- China
- Prior art keywords
- mould
- bulkhead
- fiber
- fuselage
- lay
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Images
Abstract
The invention provides a method for curing and molding an airframe and an aircraft skin at one time. The airframe adopts detachable grouped moulds which are vertically divided by bulkheads; the area between two bulkheads is one group and the moulds are connected through a bracket and a central track and can slide along the track; concave grooves matching with the airframe in shape are arranged on a carling/ stringer funicle and on the bulkheads; at the time of molding, fibre is laid in the concave grooves. The method includes the following steps: combining the mould groups, positioning and clamping the alloy ventral shields of the bulkheads, laying alar plate fibre inside and outside the bulkheads, and finally laying the aircraft skin fibre through winding. The airframe is disposed outside of a vacuum mould and is molded through vacuum pressure molding by injecting grease. Through the invention, the accessories of the airframe and the aircraft skin are reduced by more than 90% in number and by 50% in weight, the production period is reduced by 70% and the production cost is reduced by 30%; anti-corrosion capability and anti-fatigue capability are improved by more than two times and the safety factor is improved.
Description
(1) technical field
The present invention relates to the aircraft manufacturing technology, be specifically related to a kind of fuselage ring and covering curing molding technology.
(2) background technology
The conventional airframe forming process mainly is first assembling metal framework, i.e. bulkhead, longeron, stringers, then riveting dress metal skin on metal framework, fuselage will be made up of thousands of even up to ten thousand parts like this, the production process complexity, the production cycle is long and weight is big, and corrosion resistance is poor.
(3) summary of the invention
The object of the present invention is to provide a kind of number of parts to reduce more than 90%, weight reduces 50%, and the production cycle shortens 70%, and cost reduces by 30%, anticorrosive, anti-fatigue ability improves more than the twice, fuselage ring that safety factor is high and covering disposal solidifying forming method.
The object of the present invention is achieved like this: it all adopts built-in detachable grouping mould, and mould is vertically being separated by bulkhead, is one group between two bulkheads, and each set of molds links to each other with central orbit by support and can move along slide rail; At all longerons/length purlin fiber, the bulkhead place is provided with and the corresponding to groove of outling of truss, in groove, lay fiber during shaping, lay longeron/length purlin fiber earlier, the fleece-laying direction vertically becomes 0 degree angle with fuselage, lay bulkhead web fiber after longeron/length purlin fleece-laying is finished, the fleece-laying direction becomes miter angle with level in the horizontal; Merge set of molds, bulkhead alloy web location is clamped, utilize the winding method to lay outer wing plate fiber in the bulkhead, lay the covering fiber with the winding method at last; The external vacuum diaphragm of framework adopts the greasing of vacuum diaphragm platen press to be shaped.
The present invention also has some technical characterictics like this:
1, described fuselage bulkhead is " worker " font structure of " I " shape metal sheet and fiber composite materials fabricate block web;
2, described long purlin or stringers are the pure composite structure of " T " shaped sections;
Utilize the electron ray heating or heating or normal temperature cure technology implementation framework and the shaping of covering disposal solidifying in high temperature furnace according to fiber resin matrix bluk recombination technological requirement when 3, described vacuum diaphragm platen press greasing is shaped
4, on the described fuselage all opening frames such as window and covering simultaneously flange lay the fiber disposal solidifying and be shaped.
The fuselage forming process of the inventive method and Air Passenger A350, Boeing B787 also has bigger difference.A350 lays the inside panel fiber earlier on the interior forming mould of segmentation fuselage, glued joint the assembling metal framework in the inside panel outside after being shaped, and glueds joint or the riveted joint exterior skin in the framework outside at last.B787 lays the inside panel fiber earlier on the interior one-tenth mould of segmentation fuselage, on the inside panel fiber, lay vertical, horizontal framework fiber afterwards, behind the above curing molding, again at framework outside second bonding exterior skin, the A350 shortcoming is a number of parts or more relatively, and weight is bigger; Though B787 quantity significantly reduces, for improving the framework resistance to shear, fiber consumption must strengthen, thereby weight is higher relatively, and cost also improves relatively.
Fuselage forming process of the present invention has following characteristics: at first on the framework materials, stressed bulkhead is the fabricate block of " I " shape metal sheet (mainly being titanium alloy, aluminum alloy) with fiber composite materials (carbon fiber composite material, E---glass fibre or aramid fiber reinforced composite); Long purlin or longeron are pure fiber composite materials.Metal sheet in the stressed bulkhead is one of web material of " worker " font bulkhead, improved the resistance to shear of bulkhead, this midwing monoplane for similar land-effect plane, that wing body melts aircraft mutually is particularly important, wingbar can connect with the fuselage bulkhead like this, cancellation central wing box has guaranteed the fuselage usage space.This also has high anti-crash ability for other aircrafts.Secondly, a great difference is arranged also on the formation method, we adopt detachable grouping internal shaping mould, are one group of mould between bulkhead and the bulkhead, form to be convenient to several of inside detouchables, and each group mould can slide on the slide rail of center; On fuselage internal shaping mould, be provided with these frameworks required groove that is shaped in longeron, long purlin, bulkhead position, earlier separately each group mould, in groove, lay longeron, long purlin framework fiber, re-lay bulkhead web fiber after groove is paved, merge the interior outer wing plate fiber that mould is laid bulkhead at last; The unified covering fiber of laying after the framework fleece-laying is finished; External vacuum diaphragm is an anode membrane, adopts the greasing of vacuum bag compression moulding to be shaped, the electron ray that adopts people to use always be heating and curing or high temperature furnace in be heating and curing and then implementation framework and the shaping of covering disposal solidifying.Respectively organize mould after the shaping and can pull down taking-up respectively.
Technology of the present invention is compared with A350 technology, changes the metal longeron, stringers is a full composite material, and bulkhead is metal and composite material fabricate block, and each is brought into play, and it is long.Therefore in light weight, operation cost is low.Technology of the present invention is compared with B787, and the bulkhead web has adopted strong " I " shape metal sheet of shearing resistance, and therefore not only fiber consumption reduces, and also increases its anti-crash, anti-thunderbolt ability, so operation cost reduction, safer.Compare with the aircraft that adopts the traditional handicraft manufacturing by the aircraft that adopts above-mentioned technology manufacturing, number of parts reduces more than 90%, weight reduces 50%, production cycle shortens 70%, cost reduces by 30%, anticorrosive, anti-fatigue ability improves more than the twice, thereby has improved safety factor, has reduced operation cost.
(4) description of drawings
Fig. 1 is required an internal shaping composite die and " I " metal sheet scheme drawing of specific embodiment part fuselage shaping;
Fig. 2 respectively organizes mould Assembly ﹠Disassembly process scheme drawing;
Fig. 3 is a framework fleece-laying sequential schematic;
Fig. 4 is that mould merges outer wing plate fiber scheme drawing in the back laying bulkhead;
Fig. 5 lays covering fiber scheme drawing;
Fig. 6 is a shaping fuselage lateral plan;
Fig. 7 is the A-A section;
Fig. 8 is B-B section rotation scheme drawing.
Fig. 9 is the part fuselage scheme drawing behind the taking-up mould.
(5) specific embodiment
The present invention is further illustrated below in conjunction with the drawings and specific embodiments:
In conjunction with the accompanying drawings, Fig. 1-9 is uniform cross section fuselage, variable section fuselage and fuselage end forming principle.Fig. 1 is the shaping of part fuselage required internal shaping composite die and " I " metal sheet scheme drawing, among the figure, 1 is preceding group of mould left and right sides bulkhead shaping dies, 2 is mould rack, 3 is mold center's slide rail, 4 bolts for bulkhead shaping dies about connecting and left and right sides shaping dies, 5 for connecting the bolt between mould rack and the mould, 6 is the shaping dies of bulkhead up and down of inboard band overlap joint bead, 7 is uniform cross section fuselage bulkhead web fiber groove, 8 is uniform cross section fuselage bulkhead " I " shape metal sheet web, 9 is the mould up and down of middle group of die strip overlap joint bead, 10 is middle group of mould left and right sides shaping dies, 11 is bulkhead shaping dies up and down behind the middle group of mould, 12 is variable section fuselage bulkhead " I " shape metal sheet web, 13 is the mould up and down of end group die strip overlap joint bead, 14 is end group mould left and right sides mould, 15 is that end group die tip shaping dies and mold center's slide rail 3 are one, 16 is the preceding bulkhead of middle group of mould shaping dies up and down, 17 is bulkhead left and right sides shaping dies before the middle group of mould, 18 is bulkhead left and right sides shaping dies behind the middle group of mould, 19 is the outer up and down shaping dies of bulkhead before the end group mould, 20 is bulkhead left and right sides shaping dies before the end group mould, and 21 is longeron or the required groove of long purlin shaping.
Fig. 2 respectively organizes mould Assembly ﹠Disassembly process, and wherein, 22 is middle group of mould shaping dies support up and down; 23 is middle group of mould left and right sides shaping dies support; 24 is tie bolt before and after the mould; 25 is the tie bolt between the intermediate mold; 26 is the preceding frame of end group mould internal shaping mould up and down; 27 is end group mould shaping dies support up and down; 28 is end group mould left and right sides shaping dies support.
Fig. 3 is a framework fleece-laying order, and wherein 29,30 is uniform cross section fuselage bulkhead web fiber; 31 are long purlin fiber; 32,33 is variable section fuselage bulkhead web fiber.
Fig. 4 is that mould merges outer wing plate fiber in the back laying bulkhead, and wherein, 34 is uniform cross section fuselage bulkhead wing plate fiber; 35 is variable section fuselage bulkhead wing plate fiber.
Fig. 5 lays the covering fiber, and wherein 36 is the covering fiber; 37 are frame left and right sides internal shaping mould before the end group mould (only hydroairplane has this groove, to reduce the resistance of water)
Fig. 6 is a shaping fuselage lateral plan.
Fig. 7 is the A-A section, group mould left and right sides shaping dies 10 during 9 backs of mould up and down of group die strip overlap joint bead just can be pulled down in can finding to pull down from figure.
Fig. 8 can find from figure that for the rotation of B-B section uniform cross section fuselage bulkhead shaping dies takes off back displacement in the horizontal direction earlier at intermediate mold, and radially moving up and down to the centre afterwards, mould just can take out; Variable section fuselage bulkhead shaping dies is after intermediate mold takes off, internal shaping mould 26 about frame left and right sides internal shaping mould 37 or the preceding frame of end group mould before elder generation's parallel motion end group mould, radially move taking-up afterwards to the centre, afterwards radially before the group mould of middle mobile terminal about bulkhead outside before shaping dies 20 or the end group mould bulkhead up and down outside shaping dies 19 and taking out.
Fig. 9 is the part fuselage behind the taking-up mould.
1, group mould, middle group of mould, end group mould (the joint intermodule adds high temperature resistant elastic rubber fitting, is convenient to mould and takes out) before the assembling simultaneously.
(1) group mould before the assembling is overlapped on preceding group of mould left and right sides bulkhead shaping dies 1 that bolt together becomes complete ring body (in conjunction with Fig. 1) on the interior side flanges of the shaping dies of bulkhead up and down 6 that inboard band overlaps bead;
(2) organize mould in the assembling, in group mould left and right sides shaping dies 10 be overlapped on the interior side flanges of mould up and down 9 of group die strip overlap joint bead bolt together become complete ring body, in bulkhead left and right sides shaping dies 17 before the group mould, in during bulkhead left and right sides shaping dies 18 is overlapped on behind the group mould on the bead of group mould left and right sides shaping dies 10 and bolt together, in bulkhead shaping dies 11 up and down behind the group mould, in before the group mould on the interior side flanges of bulkhead mould up and down 9 of group die strip overlap joint bead during up and down shaping dies 16 is overlapped on and bolt together (in conjunction with Fig. 1-2), like this, intermediate mold is formed complete ring body by 12 moulds;
(3) assembling end group mould, end group mould left and right sides mould 14 is overlapped on bolt together becomes complete ring body on the interior side flanges of mould up and down 13 of end group die strip overlap joint bead, the outer up and down shaping dies 19 of bulkhead before the end group mould, before the end group mould frame up and down internal shaping mould 26 be overlapped on the interior side flanges of mould up and down 13 of end group die strip overlap joint bead and bolt together, before the end group mould about bulkhead outer shaping dies 20 and cooresponding before bulkhead left and right sides internal shaping mould 37 be connected on the interior side flanges of end group mould left and right sides mould 14 and bolt together, this group mould by end group mould shaping dies support 27 up and down, end group mould left and right sides shaping dies support 28 is fixed on the mould heart slide rail 3 and with end group die tip shaping dies 15 and combines formation end group mould (combining Fig. 1-2);
(4) uniform cross section fuselage bulkhead " I " shape metal sheet web 8, variable section fuselage bulkhead " I " shape metal sheet web 12 and preceding group mould, middle group of mould are utilized carrying sleeve to be contained on mold center's slide rail 3 by the order of Fig. 1, and leave certain spacing with convenient-laying bulkhead web fiber.
2, lay the framework fiber
(1) in the groove of mold head purlin, lays long purlin fiber respectively, lay direction and vertically become 0 degree angle with fuselage;
(2) after the fleece-laying in the above-mentioned groove is finished, lay bulkhead web fiber, the fleece-laying direction becomes miter angle with transverse horizontal;
(3) merge mould, clamping and positioning 8,12;
(4) lay outer wing plate plate fiber in the bulkhead, adopt the winding method to lay.
3, lay the covering fiber, adopt the winding method to lay.
4, vacuum is inhaled in external vacuum diaphragm greasing.
5, utilize inspect by instrument deep-seated blowhole defective.
6, defective is got rid of in vibration, for hydroairplane, suppresses horizontal drag reduction groove at the fuselage downside.
7, be heating and curing.
8, dismounting mould is dismantled mould: bulkhead front the bulkhead of group mould frame outer shaping dies 20--end group die tip shaping dies 15 (mold center's slide rail 3 and its one) about the front bulkhead of front bulkhead left and right sides internal shaping mould 37--end group mould of outer shaping dies 20 correspondences about the front bulkhead of lower outside shaping dies 19--end group mould on the front bulkhead of internal shaping mould 26--end group mould up and down before the upper/lower die 13--end group mould left and right sides mould 14--end group mould of bulkhead left and right sides shaping dies 18--end group die strip overlap joint bead behind the group mould among front the bulkhead of the group mould left and right sides shaping dies 17--among the shaping dies 16--up and down among the shaping dies 11--up and down behind the group mould among the group mould left and right sides shaping dies 10--among the upper/lower die 9--of group die strip overlap joint bead among front group mould left and right sides of the up and down bulkhead shaping dies 6--bulkhead shaping dies 1--of inboard band overlap joint bead in the following order.
9, repair surface defects is finished integral body.
Claims (5)
1, a kind of fuselage ring and covering disposal solidifying forming method, it is characterized in that its all built-in detachable grouping mould of employing, mould is vertically being separated by bulkhead, is one group between two bulkheads, and each set of molds links to each other with central orbit by support and can move along slide rail; At all longerons/length purlin fiber, the bulkhead place is provided with and the corresponding to groove of outling of truss, in groove, lay fiber during shaping, lay longeron/length purlin fiber earlier, the fleece-laying direction vertically becomes 0 degree angle with fuselage, lay bulkhead web fiber after longeron/length purlin fleece-laying is finished, the fleece-laying direction becomes miter angle with level in the horizontal; Merge set of molds, bulkhead alloy web location is clamped, utilize the winding method to lay outer wing plate fiber in the bulkhead, lay the covering fiber with the winding method at last; The external vacuum diaphragm of framework adopts the greasing of vacuum diaphragm platen press to be shaped.
2, fuselage ring according to claim 1 and covering disposal solidifying forming method is characterized in that " worker " font structure of described fuselage bulkhead for " I " shape metal sheet and fiber composite materials fabricate block web.
3, fuselage ring according to claim 2 and covering disposal solidifying forming method is characterized in that described long purlin or stringers are the pure composite structure of " T " shaped sections.
4, fuselage ring according to claim 3 and covering disposal solidifying forming method utilize the electron ray heating or heating or normal temperature cure technology implementation framework and the shaping of covering disposal solidifying in high temperature furnace according to fiber resin matrix bluk recombination technological requirement when it is characterized in that described vacuum diaphragm platen press greasing is shaped.
5, fuselage ring according to claim 4 and covering disposal solidifying forming method, it is characterized in that on the described fuselage all opening frames such as window and covering simultaneously flange lay the fiber disposal solidifying and be shaped.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CNA2008101369766A CN101342941A (en) | 2008-08-21 | 2008-08-21 | Disposal solidifying and molding method for fuselage ring and outer panel skin |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CNA2008101369766A CN101342941A (en) | 2008-08-21 | 2008-08-21 | Disposal solidifying and molding method for fuselage ring and outer panel skin |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN101342941A true CN101342941A (en) | 2009-01-14 |
Family
ID=40245092
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CNA2008101369766A Pending CN101342941A (en) | 2008-08-21 | 2008-08-21 | Disposal solidifying and molding method for fuselage ring and outer panel skin |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN101342941A (en) |
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103341987A (en) * | 2013-06-17 | 2013-10-09 | 沈阳飞机工业（集团）有限公司 | Co-curing process for omega-stringer vertical and horizontal reinforcement composite integral wallboard |
| CN105035354A (en) * | 2015-08-04 | 2015-11-11 | 航天材料及工艺研究所 | Composite material skin net edge coordination tools and assembling method for skins |
| CN105619834A (en) * | 2014-10-28 | 2016-06-01 | 中航通飞研究院有限公司 | Application of curing furnace forming technology to airplane composite material pressurized cabin |
| CN105966596A (en) * | 2016-06-13 | 2016-09-28 | 江西洪都航空工业集团有限责任公司 | Thickness-variable thin-wall skin structure |
| CN106892083A (en) * | 2017-04-12 | 2017-06-27 | 北京建中数字科技有限公司 | A kind of bionical frame for intersecting twin-rotor helicopter |
| CN108248890A (en) * | 2017-12-15 | 2018-07-06 | 石家庄飞机工业有限责任公司 | A kind of baby plane covering combined type molding die |
| CN111037204A (en) * | 2019-12-31 | 2020-04-21 | 湖北三江航天红阳机电有限公司 | Welding tool and welding method |
| CN112405953A (en) * | 2020-09-18 | 2021-02-26 | 航天特种材料及工艺技术研究所 | Composite material machine body integrated forming assembly die and manufacturing method |
| CN115071163A (en) * | 2022-06-22 | 2022-09-20 | 沈阳飞机工业（集团）有限公司 | Integral co-curing forming process for multi-partition-frame carbon fiber composite S-shaped air inlet channel |
-
2008
- 2008-08-21 CN CNA2008101369766A patent/CN101342941A/en active Pending
Cited By (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103341987B (en) * | 2013-06-17 | 2015-09-09 | 沈阳飞机工业（集团）有限公司 | Ω long purlin orthogonal stiffeners composites wallboard co-curing reaction |
| CN103341987A (en) * | 2013-06-17 | 2013-10-09 | 沈阳飞机工业（集团）有限公司 | Co-curing process for omega-stringer vertical and horizontal reinforcement composite integral wallboard |
| CN105619834A (en) * | 2014-10-28 | 2016-06-01 | 中航通飞研究院有限公司 | Application of curing furnace forming technology to airplane composite material pressurized cabin |
| CN105035354A (en) * | 2015-08-04 | 2015-11-11 | 航天材料及工艺研究所 | Composite material skin net edge coordination tools and assembling method for skins |
| CN105966596A (en) * | 2016-06-13 | 2016-09-28 | 江西洪都航空工业集团有限责任公司 | Thickness-variable thin-wall skin structure |
| CN106892083B (en) * | 2017-04-12 | 2023-11-21 | 北京清航紫荆装备科技有限公司 | Bionic rack of crossed double-rotor helicopter |
| CN106892083A (en) * | 2017-04-12 | 2017-06-27 | 北京建中数字科技有限公司 | A kind of bionical frame for intersecting twin-rotor helicopter |
| CN108248890A (en) * | 2017-12-15 | 2018-07-06 | 石家庄飞机工业有限责任公司 | A kind of baby plane covering combined type molding die |
| CN108248890B (en) * | 2017-12-15 | 2023-08-01 | 石家庄飞机工业有限责任公司 | Combined forming die for small aircraft skin |
| CN111037204B (en) * | 2019-12-31 | 2021-12-21 | 湖北三江航天红阳机电有限公司 | Welding tool and welding method |
| CN111037204A (en) * | 2019-12-31 | 2020-04-21 | 湖北三江航天红阳机电有限公司 | Welding tool and welding method |
| CN112405953A (en) * | 2020-09-18 | 2021-02-26 | 航天特种材料及工艺技术研究所 | Composite material machine body integrated forming assembly die and manufacturing method |
| CN112405953B (en) * | 2020-09-18 | 2022-05-03 | 航天特种材料及工艺技术研究所 | Composite material machine body integrated forming assembly die and manufacturing method |
| CN115071163A (en) * | 2022-06-22 | 2022-09-20 | 沈阳飞机工业（集团）有限公司 | Integral co-curing forming process for multi-partition-frame carbon fiber composite S-shaped air inlet channel |
| CN115071163B (en) * | 2022-06-22 | 2024-03-08 | 沈阳飞机工业（集团）有限公司 | Integral co-curing forming process for S-shaped air inlet channel of multi-bulkhead carbon fiber composite material |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN101342941A (en) | Disposal solidifying and molding method for fuselage ring and outer panel skin | |
| CN108622123B (en) | Low-cost carbon fiber composite air-iron vehicle body and manufacturing process | |
| CN101500787B (en) | Improved composite resin window frame constructions for airplanes | |
| CN103407175B (en) | A kind of integral forming method of fiber-reinforced resin matrix compound material wing box | |
| CN105035359B (en) | A kind of degressive-rate composite material integral formed rudder face structure and machining method | |
| CN103832576B (en) | Include the high integrity structure of leading edge rib and trailing edge rib for ROV lifting surface | |
| Hinrichsen et al. | The challenge of reducing both airframe weight and manufacturing cost | |
| JP2019502591A (en) | Process for manufacturing apron plates for equipment cabins in high-speed railway vehicles with composite materials | |
| CN101342942A (en) | Disposal solidifying and forming technique for frame and outer panel skin of wing profile | |
| CN104743095A (en) | Highly integrated infused box made of composite material and method of manufacturing | |
| CN103979016B (en) | Lightweight electric vehicle body skeleton | |
| CN103832574B (en) | Torsion box for the optimization of ROV | |
| CN109435272B (en) | Variable-thickness composite material integral framework forming method and forming die thereof | |
| CN205417779U (en) | Combined material passenger train roof | |
| CN107651163A (en) | Outer wing structure of a kind of fixed-wing unmanned plane and preparation method thereof | |
| CN206383959U (en) | A kind of track vehicle body structure | |
| CN108481762B (en) | Composite material core-level cover plate for rocket and preparation method thereof | |
| CN201261545Y (en) | Once solidified and molded fuselage ring and covering | |
| CN201276209Y (en) | Once solidified and molded wing type frame and covering | |
| CN106347684A (en) | Detachable composite tail fairing and machining method thereof | |
| CN106335644A (en) | Integrally-formed jet nozzle structure | |
| CN110524910A (en) | A kind of VARTM technique composite material mould and its manufacturing method | |
| CN112405953B (en) | Composite material machine body integrated forming assembly die and manufacturing method | |
| CN211543866U (en) | Composite material partition frame structure for aircraft wing trailing edge | |
| CN206765397U (en) | A kind of fiber-reinforced composite deck board |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| AD01 | Patent right deemed abandoned |
Effective date of abandoning: 20090114 |
|
| C20 | Patent right or utility model deemed to be abandoned or is abandoned |