CN102825428B - Aircraft fairing and method for manufacturing same - Google Patents
Aircraft fairing and method for manufacturing same Download PDFInfo
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- CN102825428B CN102825428B CN201210295425.0A CN201210295425A CN102825428B CN 102825428 B CN102825428 B CN 102825428B CN 201210295425 A CN201210295425 A CN 201210295425A CN 102825428 B CN102825428 B CN 102825428B
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Abstract
The invention relates to an aircraft fairing and a method for manufacturing the same, comprising the following steps: (1) blanking a blank; (2) hot-press molding the middle section of the cover body; (3) machining the hood head and the hood tail of the fairing; (4) and (6) welding. The invention solves the problems of poor performance of the head part of the fairing and the processing of the thermal forming die, has low cost, is suitable for titanium alloy with the plate thickness of 0.5-2mm, and can meet the design requirement of products.
Description
Technical field
The present invention relates to aerospace industry technical field, particularly a kind of aircraft cowling and manufacture method thereof.
Background technology
Along with countries in the world are to the attention of aerospace vehicle correlative study, require more and more higher to the service behaviour of the entirety of aircraft.During aerospace vehicle flight, wave resistance and frictional resistance sharply increase, and cause aerodynamic drag to increase, and produce serious to push away resistance matching problem.Radome fairing part, as the outside appurtenances of aircraft, has a great impact the raising of the aeroperformance of aircraft.Described radome fairing is that a kind of button is located at external side, aircraft cabin, for walking the long narrow shape outer cover (as shown in Figure 1 and Figure 2) of cable (control line), also cabling radome fairing is claimed, it comprises capouch 1, cover body 2 and cover tail 3, and being provided with multiple installing hole 4, material therefor is titanium alloy.Traditional radome fairing to be fabricated to titanium alloy plate integral heat pressure shaping, then boring forms in designed installation position.Because the die size of thermoforming is long, Mould Machining is difficult and long in time limit, expensive; During use, mobile and assembling difficulty, press tonnage used is large; Owing to adopting sheet material hot pressing, the thickness of the capouch 1 of the radome fairing after shaping is identical with the thickness of other parts, when flying, owing to generating heat with windage, easily producing thermal deformation, making head degradation.
As can be seen here, above-mentioned existing aircraft cowling, in structure and manufacture method, obviously still has deficiency, and is urgently further improved.
Summary of the invention
Main purpose of the present invention is, the defect that the manufacture method overcoming existing aircraft cowling exists, and a kind of aircraft cowling and manufacture method thereof of new structure are provided, technical problem to be solved be adopt segmentation hot-forming-welding, in conjunction with machining, to overcome, overlength Mould Machining is expensive, awkward problem, improves surface roughness, reduces air drag, increase capouch thickness, reduce thermal deformation.
The object of the invention to solve the technical problems realizes by the following technical solutions.
The manufacture method of a kind of aircraft cowling proposed by the invention, described aircraft cowling comprises capouch, cover body and cover tail, and wherein said cover body is made up of multistage, and described multistage cover body and described capouch and cover tail are welded;
Described manufacture method comprises the following steps:
(1) blank blanking
Capouch, cover tail adopt machining, and select thickness to be greater than the titanium alloy plate blanking of cover body height, surrounding leaves certain allowance; Cover status section adopt hot-forming, by divide each segment length, selects thickness to be greater than the titanium alloy plate of cover body wall thickness, carry out blanking by expanded view, and finish-milling outer surface is to cover body thickness, roughness Ra≤1.6 μm of finished surface;
(2) in the middle of, each section of cover body is hot-forming
By even for the die joint of hot-die brushing alcohol graphite solution, air-dry rear fine sandpaper is slightly polished, and puts into workpiece, aligns punch and matrix, carries out hot pressing, temperature 500-750 DEG C, pressurize 5-30 second, is cooled to less than 150 DEG C, and workpiece is taken out in die sinking;
(3) capouch, the cover tail of machined radome fairing
Finish-milling outer surface roughness reaches Ra≤1.6 μm; Machined parameters: the speed of mainshaft: 1500-1800r/min, chip speed: 150-200mm/min, tool feeding amount: 0.2-0.5mm.
(4) weld
The each section of cover body that sweating heat is shaping and the capouch that machine adds, cover tail become aircraft cowling.
The object of the invention to solve the technical problems also can be applied to the following technical measures to achieve further.
The manufacture method of aforesaid aircraft cowling, the welding wherein described in step (4) is Laser Welding or electron beam welding or argon arc welding.
The manufacture method of aforesaid aircraft cowling; welding wherein described in step (4) is Laser Welding; and adopt six axis robot to weld complex-curved curved welding seam; welding parameter: laser power: 1.5-2.0KW; focal length 100-120mm, defocusing amount :-0.2mm, speed of welding: 900-1000mm/min; inert protective gas flow: He:11-15ml/min, Ar:14-20ml/min.
The manufacture method of aforesaid aircraft cowling, wherein the trade mark of step (1) titanium alloy plate used is TC1 or TC3 or TC4 or TA15.
The manufacture method of aforesaid aircraft cowling is wherein 400# fine sandpaper with the fine sandpaper that fine sandpaper is slightly polished described in step (2).
The manufacture method of aforesaid aircraft cowling, wherein said multistage cover body is 3 to 6 sections.
The manufacture method of aforesaid aircraft cowling, wherein said multistage cover body is 5 sections.
The present invention compared with prior art has obvious advantage and beneficial effect.By technique scheme, aircraft cowling of the present invention and manufacture method thereof can reach suitable technological progress and practicality, and have the extensive value in industry, and it at least has following advantages:
1, aircraft cowling of the present invention adopts segmental structure, adopts slab machine-shaping, improves the roughness of outer surface, reduce windage, suitably increase thickness, reduce aloft thermal deformation for capouch, cover tail.
2, the cover body of aircraft cowling of the present invention adopts segmentation hot-forming, and without the need to overlength hot-die, the manufacturing cost of mould reduces relatively, increase economic efficiency, in the practicality used and cost benefit, really meet completely needed for industry development, there is industrial utilization.
3, the present invention specifically provides a kind of aircraft cowling and manufacture method thereof, and solve head degradation and the hot forming tool processing problems of radome fairing, cost is low, is applicable to the titanium alloy of thickness of slab 0.5-2mm, can meet product design requirement.
In sum, the aircraft cowling of novelty of the present invention and manufacture method thereof, there is above-mentioned plurality of advantages and practical value, no matter it is in manufacture method, structure or functionally all have larger improvement, there is significant progress technically, and create handy and practical effect, and more existing aircraft cowling and manufacture method thereof have outstanding effect of enhancement, thus be more suitable for practicality, and there is the extensive value of industry, be really a new and innovative, progressive, practical new design.
Above-mentioned explanation is only the general introduction of technical solution of the present invention, in order to technological means of the present invention can be better understood, and can be implemented according to the content of description, and can become apparent to allow above and other object of the present invention, feature and advantage, below especially exemplified by preferred embodiment, and coordinate accompanying drawing, be described in detail as follows.
Accompanying drawing explanation
Fig. 1 is the structure chart of aircraft cowling of the present invention.
Fig. 2 is the side view of Fig. 1.
Fig. 3 is the A-A sectional view of Fig. 1.
Fig. 4 is the B-B sectional view of Fig. 1.
Fig. 5 is the C-C sectional view of Fig. 1.
Fig. 6 is the flow process chart of aircraft cowling of the present invention.
Detailed description of the invention
Refer to shown in Fig. 1, Fig. 2, Fig. 3, Fig. 4, Fig. 5, the aircraft cowling of present pre-ferred embodiments, comprise capouch 1, cover body 2 and cover tail 3, it is characterized in that described cover body 2 is made up of multistage, described multistage cover body and described capouch and cover tail are welded.
Wherein said multistage cover body is 3 to 6 sections.
In diagram specific embodiment, described multistage cover body is 5 sections, 2-1,2-2,2-3,2-4,2-5 namely in Fig. 1,2.Be in Fig. 2 centered by installing hole 4, point five sections of cover bodies thermoforming machines.
Due to the cover status of long narrow shape is become multistage, use mould heat pressure shaping respectively, thus solve the overlength pressing mold manufacturing cycle long, expensive; During use, mobile and assembling difficulty; The problem that press tonnage used is large.
Fig. 6 is the flow process chart of aircraft cowling of the present invention, comprises the following steps:
(1) blank blanking S01
Capouch, cover tail adopt machining, and select thickness to be greater than the titanium alloy plate blanking of cover body height, surrounding leaves certain allowance;
Cover status section adopt hot-forming, by divide each segment length, selects thickness to be greater than the titanium alloy plate of cover body wall thickness, carry out blanking by expanded view, and finish-milling outer surface is to cover body thickness, roughness Ra≤1.6 μm of finished surface; Expanded view is known technology, does not show in the present invention.
(2) each section of hot-forming S02 of cover body in the middle of
By even for the die joint of hot-die brushing alcohol graphite solution, air-dry rear fine sandpaper is slightly polished, and puts into workpiece, aligns punch and matrix, carries out hot pressing, temperature 500-750 DEG C, pressurize 5-30 second, is cooled to less than 150 DEG C, and workpiece is taken out in die sinking;
(3) capouch, the cover tail S03 of machined radome fairing
Carry out machining with end mill or machining center, finish-milling outer surface roughness reaches Ra≤1.6 μm; Machined parameters: the speed of mainshaft: 1500-1800r/min, chip speed: 150-200mm/min, tool feeding amount: 0.2-0.5mm;
The titanium alloy head section intensity that machining process goes out is higher, effectively can solve when flying the degradation problem caused with windage heat-dissipating.
(4) S04 is welded
The each section of cover body that sweating heat is shaping and the capouch that machine adds, cover tail become aircraft cowling.
Welding described in step (4) is Laser Welding; and adopt six axis robot to weld complex-curved curved welding seam; welding parameter: laser power: 1.5-2.0kw; focal length 100-120mm; defocusing amount :-0.2mm; speed of welding: 900-1000mm/min, inert protective gas flow: He:11-15ml/min, Ar:14-20ml/min.
Described welding also can be electron beam welding or argon arc welding.
The trade mark of step (1) titanium alloy plate used is TC1 or TC3 or TC4 or TA15.
400# fine sandpaper with the fine sandpaper that fine sandpaper is slightly polished described in step (2).
The above, it is only preferred embodiment of the present invention, not any pro forma restriction is done to the present invention, although the present invention discloses as above with preferred embodiment, but and be not used to limit the present invention, any those skilled in the art, do not departing within the scope of technical solution of the present invention, make a little change when the technology contents of above-mentioned announcement can be utilized or be modified to the Equivalent embodiments of equivalent variations, in every case be the content not departing from technical solution of the present invention, according to any simple modification that technical spirit of the present invention is done above embodiment, equivalent variations and modification, all still belong in the scope of technical solution of the present invention.
Claims (7)
1. a manufacture method for aircraft cowling, described aircraft cowling comprises capouch, multistage cover body and cover tail, and described multistage cover body and described capouch and cover tail are welded, and described manufacture method comprises the following steps:
(1) blank blanking
Capouch, cover tail adopt machining, and select thickness to be greater than the titanium alloy plate blanking of cover body height, surrounding leaves certain allowance; Cover status section adopt hot-forming, by divide each segment length, selects thickness to be greater than the titanium alloy plate of cover body wall thickness, carry out blanking by expanded view, and finish-milling outer surface is to cover body thickness, roughness Ra≤1.6 μm of finished surface;
(2) in the middle of, each section of cover body is hot-forming
By even for the die joint of hot-die brushing alcohol graphite solution, air-dry rear fine sandpaper is slightly polished, and puts into workpiece, aligns punch and matrix, carries out hot pressing, temperature 500-750 DEG C, pressurize 5-30 second, is cooled to less than 150 DEG C, and workpiece is taken out in die sinking;
(3) capouch, the cover tail of machined radome fairing
Finish-milling outer surface roughness reaches Ra≤1.6 μm; Machined parameters: the speed of mainshaft: 1500-1800r/min, chip speed: 150-200mm/min, tool feeding amount: 0.2-0.5mm;
(4) weld
The each section of cover body that sweating heat is shaping and the capouch that machine adds, cover tail become aircraft cowling.
2. the manufacture method of aircraft cowling according to claim 1; it is characterized in that the welding described in step (4) is Laser Welding; and adopt six axis robot to weld complex-curved curved welding seam; welding parameter: laser power: 1.5-2.0KW; focal length 100-120mm, defocusing amount :-0.2mm, speed of welding: 900-1000mm/min; inert protective gas flow: He:11-15ml/min, Ar:14-20ml/min.
3. the manufacture method of aircraft cowling according to claim 1, is characterized in that the welding described in step (4) is electron beam welding or argon arc welding.
4. the manufacture method of aircraft cowling according to claim 1, is characterized in that the trade mark of step (1) titanium alloy plate used is TC1 or TC3 or TC4 or TA15.
5. the manufacture method of aircraft cowling according to claim 1, to is characterized in that described in step (2) with the fine sandpaper that fine sandpaper is slightly polished it being 400# fine sandpaper.
6. the manufacture method of aircraft cowling according to claim 1, is characterized in that described multistage cover body is 3 to 6 sections.
7. the manufacture method of aircraft cowling according to claim 1, is characterized in that described multistage cover body is 5 sections.
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| CN201210295425.0A CN102825428B (en) | 2012-08-19 | 2012-08-19 | Aircraft fairing and method for manufacturing same |
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| CN104591736B (en) * | 2015-01-09 | 2016-09-21 | 中国科学院上海光学精密机械研究所 | The manufacture method of saturating infrared ZnS trousers pottery |
| CN104959775B (en) * | 2015-05-14 | 2017-06-16 | 芜湖市爱德运输机械有限公司 | Screw machine outer cover processing method |
| CN106425084B (en) * | 2015-08-07 | 2019-05-31 | 昆山汉鼎精密金属有限公司 | Automatic processing system and method |
| CN107717342B (en) * | 2017-09-27 | 2019-02-26 | 湖北三江航天江北机械工程有限公司 | Special-shaped depth blind hole antenna house processing method |
| CN108972159B (en) * | 2018-07-30 | 2020-04-07 | 大连理工大学 | Double-side grinding method for sapphire spherical cover |
| CN111992996B (en) * | 2020-08-25 | 2022-06-21 | 哈尔滨汽轮机厂有限责任公司 | Method for manufacturing nozzle fairing of 400 MW-grade G/H-grade heavy gas turbine |
| CN113953771A (en) * | 2021-11-12 | 2022-01-21 | 中国航发贵州黎阳航空动力有限公司 | Machining method of wide-edge metal plate hollow support plate-shaped fairing |
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| GB9224823D0 (en) * | 1992-11-26 | 1993-01-13 | Howden Group Plc | Ljungstrom heat exchanger |
| RU2220882C2 (en) * | 2001-10-10 | 2004-01-10 | Открытое акционерное общество "Казанский вертолётный завод" | Helicopter stabilizer |
| US7875132B2 (en) * | 2005-05-31 | 2011-01-25 | United Technologies Corporation | High temperature aluminum alloys |
| CN100442032C (en) * | 2006-10-17 | 2008-12-10 | 北京航空航天大学 | Carbon-fiber composite material high-speed air-craft rectifying cover surface transient temperature measuring apparatus |
| FR2918919B1 (en) * | 2007-07-17 | 2013-03-29 | Eurocopter France | METHOD AND DEVICE FOR STITCHING THE METAL CHAIN OF AN ATTACHMENT EDGE OF A CAR |
| DE102008007278B4 (en) * | 2008-02-01 | 2010-04-08 | Airbus Deutschland Gmbh | Bleedairduct segment, Bleedairduct arrangement with such Bleedairduct segments and Bleedairduct system with regulation device |
| EP2279943A4 (en) * | 2008-04-21 | 2012-05-09 | Boris Andreevich Polovinkin | Vertical take-off and vertical landing gyroplane |
| CN101497372B (en) * | 2009-02-18 | 2013-01-02 | 中国科学院力学研究所 | External cowling of scramjet engine and design method thereof |
| FR2960522B1 (en) * | 2010-05-27 | 2012-06-29 | Airbus Operations Sas | METHOD FOR MANUFACTURING BY SUPERPLASTIC FORMING AND BY LAUNDRYING A RIB FOR AERODYNAMIC FITTING OF AN AIRCRAFT ENGINE HITCHING MAT |
| CN201959567U (en) * | 2011-01-20 | 2011-09-07 | 杭州中天模型有限公司 | Three-way fixed-wing airplane |
| CN202130570U (en) * | 2011-05-16 | 2012-02-01 | 湖北航天技术研究院总体设计所 | Aircraft cowling shell |
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Address after: 618400 Shifang City, Sichuan Province Economic Development Zone (Ling Jie Park) Patentee after: SICHUAN FUTURE AEROSPACE INDUSTRIAL Co.,Ltd. Address before: 618400 Shifang City, Sichuan Province Economic Development Zone (Ling Jie Park) Patentee before: Shifang City Mingri Space Navigation Industry Co.,Ltd. |
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Effective date of registration: 20220510 Address after: Room 101, office building 42, No. 4, East Street, Nankou Town, Changping District, Beijing 102202 Patentee after: Beijing tomorrow Aerospace Engineering Technology Research Co.,Ltd. Address before: 618400 Shifang Economic Development Zone, Deyang City, Sichuan Province (Lingjie Park) Patentee before: SICHUAN FUTURE AEROSPACE INDUSTRIAL Co.,Ltd. |
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Address after: Room 101, office building 42, No. 4, East Street, Nankou Town, Changping District, Beijing 102202 Patentee after: Beijing SIA Titanium Aerospace Technology Co.,Ltd. Address before: Room 101, office building 42, No. 4, East Street, Nankou Town, Changping District, Beijing 102202 Patentee before: Beijing tomorrow Aerospace Engineering Technology Research Co.,Ltd. |