US20120219818A1 - Panel member, aircraft main wing, and method for forming panel member - Google Patents

Panel member, aircraft main wing, and method for forming panel member Download PDF

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Publication number
US20120219818A1
US20120219818A1 US13/402,099 US201213402099A US2012219818A1 US 20120219818 A1 US20120219818 A1 US 20120219818A1 US 201213402099 A US201213402099 A US 201213402099A US 2012219818 A1 US2012219818 A1 US 2012219818A1
Authority
US
United States
Prior art keywords
thickness
panel member
areas
wing
panel
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.)
Abandoned
Application number
US13/402,099
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English (en)
Inventor
Hideto MOTOHASHI
Yoshifumi Hosokawa
Toshinari FUJII
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Mitsubishi Aircraft Corp
Original Assignee
Mitsubishi Aircraft Corp
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Mitsubishi Aircraft Corp filed Critical Mitsubishi Aircraft Corp
Assigned to Mitsubishi Aircraft Corporation reassignment Mitsubishi Aircraft Corporation ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: FUJII, TOSHINARI, HOSOKAWA, YOSHIFUMI, MOTOHASHI, HIDETO
Publication of US20120219818A1 publication Critical patent/US20120219818A1/en
Priority to US14/792,103 priority Critical patent/US10081418B2/en
Abandoned legal-status Critical Current

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C3/00—Wings
    • B64C3/26—Construction, shape, or attachment of separate skins, e.g. panels
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49616—Structural member making
    • Y10T29/49622—Vehicular structural member making
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/12—All metal or with adjacent metals
    • Y10T428/12389—All metal or with adjacent metals having variation in thickness
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/24—Structurally defined web or sheet [e.g., overall dimension, etc.]
    • Y10T428/24479—Structurally defined web or sheet [e.g., overall dimension, etc.] including variation in thickness

Definitions

  • the present invention relates to a panel member, an aircraft main wing, and a method for forming the panel member.
  • a wing panel For an aircraft main wing, the outer surface thereof is formed by a panel-form member (hereinafter, referred to as a wing panel).
  • This wing panel is configured so that the thickness thereof is set according to the portion of the wing. That is, in the base end portion on the airframe side of main wing, and in the portion around the engine-mounting position, the thickness of wing panel is set so as to be large because high strength is required in these portions. In contrast, in the tip end portion of wing or the like portions, the thickness of wing panel is set so as to be small. Thus, the thickness of wing panel is prevented from being increased than necessary while a necessary strength is ensured depending on the respective portions, whereby the weight of wing panel is restrained.
  • Some wing panels are formed by being machined out of a metallic base material (for example, refer to Takeshi Yamada et al. “Development of Shot Peen Forming Technology of Main Wing Integral Skin of Continental Business Jet” Mitsubishi Heavy Industries Technical Review, Vol.39, No.1 (2002), p.36).
  • the position of the cutting tool of a cutting machine with respect to the metallic base material is changed stepwise.
  • the change amount of the thickness must be set to a dimension equal to or larger than the machining tolerance.
  • the machining tolerance is, for example, ⁇ 0.1 mm
  • the thickness is preferably increased or decreased stepwise for each dimension equal to or larger than the machining tolerance, for example, for each dimension equal to or larger than 0.2 mm (this increasing/decreasing amount is referred to as a step amount as appropriate). That is, in the case where the design value of thickness in a certain portion P 1 of a wing panel 1 is 6.0 mm as shown in FIGS.
  • the thickness increases from the portion P 1 to the portion P 2 in design, actually, in some cases, the thickness does not change, or inversely the thickness decreases from the portion P 1 to the portion P 2 . Even in such a case, no problem occurs in terms of strength as long as the thicknesses of the portions P 1 and P 2 are within the machining tolerance. However, a hindrance may occur in the manufacturing process.
  • the wing panel machined out of a metallic base material by cutting machine is inspected at least visually. If a flaw or the like induced while machining is present on the wing panel, the surroundings of the flaw must be sanded to correct the flaw to the predefined standard or higher grade.
  • the step amount of thickness is small, the step existing in the portion in which the thickness changes becomes a minor step, and therefore, at the inspection time, it is sometimes impossible to distinguish between a normal step and a flaw. For this reason, correction is sometimes made as the result of mistaking the normal step for a flaw, which leads to an increase in the manufacturing cost.
  • the thickness design value of 6.4 mm is an excess value by 0.3 mm.
  • the present invention has been made to solve the above-described technical problems, and accordingly an object thereof is to provide a panel member capable of restraining an increase in weight of a wing panel while restraining wasteful work in the manufacturing process, an aircraft main wing formed by the said panel member, and a method for forming a panel member.
  • the present invention provides a panel member forming the outer surface of an aircraft, wherein a portion of the panel member in which the thickness thereof changes gradually is divided into a plurality of areas each having a certain thickness; the thicknesses of the adjacent areas are different from each other; and in the boundary portion between the adjacent areas, there is formed a ridge having a height larger than the dimensional difference in thickness between the areas on both sides of the boundary portion.
  • Such a ridge is preferably formed so as to have a height larger than the machining tolerance of the panel member.
  • the panel member is formed by being machined out of a metallic base material.
  • Such a panel member can also be used in forming a portion other than the outer surface of main wing.
  • the present invention provides an aircraft main wing wherein the outer surface of the wing is formed by the panel member described above.
  • the present invention provides a method for forming a panel member, in which the panel member is machined out of a metallic base material, wherein a portion in which the thickness changes gradually is divided into a plurality of areas; each of the areas is formed so as to have a certain thickness; and in the boundary portion between the adjacent areas, there is formed a ridge having a height larger than the dimensional difference in thickness between the areas on both sides of the boundary portion.
  • the present invention is not limited to the case where the panel member and stringers (reinforcing ribs) are integrally machined out of a thick plate, and can be applied to a so-called built-up panel in which the stringers are fastened to the panel member with fasteners as far as the panel member is machined out of a thick plate.
  • the boundary portion between the adjacent areas there is formed the ridge having a height larger than the dimensional difference in thickness between the areas on both sides of the boundary portion.
  • the boundary portion between the adjacent areas each having a different thickness can be visually confirmed easily. Thereby, only a portion in which a flaw occurs in each area can be corrected, so that wasteful work in the manufacturing process can be restrained.
  • the boundary portion between the adjacent areas can also be formed so as to have a thickness equal to or smaller than the machining tolerance, and thereby the increase in weight of the wing panel can be restrained.
  • FIG. 1 is a perspective view showing a part of a wing panel in accordance with an embodiment of the present invention
  • FIG. 2 is a sectional view of the wing panel shown in FIG. 1 ;
  • FIG. 3 is sectional views showing other examples in which the cross-sectional shape of a ridge formed on a wing panel is different;
  • FIG. 4 is a perspective view showing a part of a conventional wing panel.
  • FIG. 5 is a sectional view of a conventional wing panel.
  • FIG. 1 is a perspective view for explaining a wing panel 10 in accordance with this embodiment.
  • the wing panel 10 forms the outer surface of an aircraft main wing.
  • the wing panel 10 is formed so as to have the necessary minimum thickness or a larger thickness depending on a portion of a main wing.
  • the wing panel 10 is formed by being machined out of a metallic base material.
  • the thickness of the wing panel 10 is changed by dividing the wing panel 10 into a plurality of areas A 1 , A 2 , . . . and by gradually increasing or decreasing the thicknesses t 1 , t 2 , . . . of the areas A 1 , A 2 , . . . , respectively.
  • the change amounts of thicknesses of the adjacent areas are set so as to be a specified value s or larger.
  • the specified value s can be set optionally regardless of the machining tolerance set at the design time. That is, the specified value s can be made a dimension equal to or larger than the machining tolerance.
  • ridges 20 are formed so as to be continuous.
  • This ridge 20 has a height h equal to or larger than the machining tolerance and a width w.
  • both of the height h and the width w can be set at about 1 mm.
  • the worker who handles the wing panel 10 can visually confirm the step portions of the areas A 1 , A 2 , . . . easily. Thereby, if a flaw or the like is found in the areas A 1 , A 2 , . . . , corrective actions can be properly taken on the actually existing flaw without mistaking the flaw for the step portion of the areas A 1 , A 2 , . . . .
  • a configuration can be provided in which no hindrance occurs even if the change amounts of thicknesses of the areas A 1 , A 2 , . . . (the difference between thickness t 1 and thickness t 2 , the difference between thickness t 2 and thickness t 3 ) are equal to or smaller than the machining tolerance.
  • each of the thicknesses of the areas A 1 , A 2 , . . . can be made the necessary minimum thickness, and the wing panel 10 can be formed without having a wasteful thickness.
  • the weight of the wing panel 10 is not increased, so that the weight thereof can be reduced as compared with the conventional configuration.
  • the ridges 20 can be caused to function as a reinforcing element for the wing panel 10 .
  • the thicknesses of the areas A 1 , A 2 , . . . are decreased further, so that the weight of the wing panel 10 can also be reduced.
  • the ridge is formed into a rectangular shape in cross section.
  • the ridge shape is not limited to this shape, and as shown in FIGS. 3A and 3B , the ridge can be formed into a triangular shape or an inverse U shape in cross section.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Shaping Metal By Deep-Drawing, Or The Like (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US13/402,099 2011-02-24 2012-02-22 Panel member, aircraft main wing, and method for forming panel member Abandoned US20120219818A1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
US14/792,103 US10081418B2 (en) 2011-02-24 2015-07-06 Panel member, aircraft main wing, and method for forming panel member

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
JP2011-038161 2011-02-24
JP2011038161A JP5361009B2 (ja) 2011-02-24 2011-02-24 パネル材、航空機の主翼、パネル材の形成方法

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US14/792,103 Division US10081418B2 (en) 2011-02-24 2015-07-06 Panel member, aircraft main wing, and method for forming panel member

Publications (1)

Publication Number Publication Date
US20120219818A1 true US20120219818A1 (en) 2012-08-30

Family

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Family Applications (2)

Application Number Title Priority Date Filing Date
US13/402,099 Abandoned US20120219818A1 (en) 2011-02-24 2012-02-22 Panel member, aircraft main wing, and method for forming panel member
US14/792,103 Active 2032-08-24 US10081418B2 (en) 2011-02-24 2015-07-06 Panel member, aircraft main wing, and method for forming panel member

Family Applications After (1)

Application Number Title Priority Date Filing Date
US14/792,103 Active 2032-08-24 US10081418B2 (en) 2011-02-24 2015-07-06 Panel member, aircraft main wing, and method for forming panel member

Country Status (4)

Country Link
US (2) US20120219818A1 (ja)
EP (1) EP2492191B1 (ja)
JP (1) JP5361009B2 (ja)
CA (1) CA2763175C (ja)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9657489B2 (en) * 2015-06-15 2017-05-23 The Boeing Company Fractal stiffening

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20080296433A1 (en) * 2003-01-16 2008-12-04 Fraunhofer-Gesellschaft Zuer Foerderung Der Angewandten Forschung E.V. Lightweight structural component in particular for aircraft and method for its production
US20110268926A1 (en) * 2010-04-30 2011-11-03 Airbus Operations S.L. Internal structure for aircraft in composite material

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE10007995C2 (de) * 2000-02-22 2002-03-07 Airbus Gmbh Strukturbauteil, insbesondere für ein Flugzeug und Verfahren zur Herstellung eines Strukturbauteils
EP1343609A4 (en) * 2000-12-12 2007-10-03 Remmele Engineering Inc MONOLITHIC PART AND METHOD FOR THE PRODUCTION THEREOF
JP2003191028A (ja) * 2001-12-26 2003-07-08 Mitsubishi Heavy Ind Ltd ピーン成形方法
GB0207239D0 (en) * 2002-03-27 2002-05-08 Airbus Uk Ltd Wing skin and method of manufacture thereof
US20040035979A1 (en) * 2002-08-23 2004-02-26 Mccoskey William Robert Integrally stiffened axial load carrying skin panels for primary aircraft structure and closed loop manufacturing methods for making the same
US8074694B2 (en) * 2009-05-28 2011-12-13 The Boeing Company Stringer transition method

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20080296433A1 (en) * 2003-01-16 2008-12-04 Fraunhofer-Gesellschaft Zuer Foerderung Der Angewandten Forschung E.V. Lightweight structural component in particular for aircraft and method for its production
US20110268926A1 (en) * 2010-04-30 2011-11-03 Airbus Operations S.L. Internal structure for aircraft in composite material

Also Published As

Publication number Publication date
CA2763175C (en) 2014-04-15
EP2492191B1 (en) 2017-01-04
JP2012171576A (ja) 2012-09-10
US10081418B2 (en) 2018-09-25
CA2763175A1 (en) 2012-08-24
JP5361009B2 (ja) 2013-12-04
US20150307181A1 (en) 2015-10-29
EP2492191A1 (en) 2012-08-29

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AS Assignment

Owner name: MITSUBISHI AIRCRAFT CORPORATION, JAPAN

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:MOTOHASHI, HIDETO;HOSOKAWA, YOSHIFUMI;FUJII, TOSHINARI;SIGNING DATES FROM 20120111 TO 20120112;REEL/FRAME:027749/0294

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION