WO2007099297A1 - Voilure d'aéronef et son procédé de montage - Google Patents

Voilure d'aéronef et son procédé de montage Download PDF

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Publication number
WO2007099297A1
WO2007099297A1 PCT/GB2007/000653 GB2007000653W WO2007099297A1 WO 2007099297 A1 WO2007099297 A1 WO 2007099297A1 GB 2007000653 W GB2007000653 W GB 2007000653W WO 2007099297 A1 WO2007099297 A1 WO 2007099297A1
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WO
WIPO (PCT)
Prior art keywords
wing
aircraft
section
subassembly
fuselage
Prior art date
Application number
PCT/GB2007/000653
Other languages
English (en)
Inventor
Hal Errikos Calamvokis
Original Assignee
Hal Errikos Calamvokis
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Filing date
Publication date
Application filed by Hal Errikos Calamvokis filed Critical Hal Errikos Calamvokis
Publication of WO2007099297A1 publication Critical patent/WO2007099297A1/fr

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/26—Attaching the wing or tail units or stabilising surfaces
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C3/00—Wings
    • B64C3/18—Spars; Ribs; Stringers
    • B64C3/185—Spars
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C3/00—Wings
    • B64C3/10—Shape of wings

Definitions

  • the present invention relates to aircraft wings.
  • the invention relates to wing subassemblies, and methods of wing and aircraft assembly.
  • a modern aircraft wing especially that of wide-bodied transport aircraft, has a complex three-dimensional shape which arises through mostly aerodynamic considerations.
  • the wing typically has a complex aerofoil section with a high lift-to-drag ratio, and makes an angle of dihedral or anhedral with the aircraft fuselage, when viewed from a front end of the aircraft.
  • Wing dihedral where the wingtips are above the wing-root
  • anhedral where the wingtips are below the wing-root
  • the wing is also typically swept backwards in plan-form to delay the formation of Shockwaves around the wing at transonic speeds (Mach 0.7 - 1.2). This permits the aircraft to cruise at a speed closer to the speed of sound than would otherwise be the case, reducing trip time and drag.
  • the wing has three main portions: a first exposed portion for generating lift on one side of the fuselage; a second exposed portion for generating lift on the other side of the fuselage; and a centre portion.
  • the simplest method of joining the wing and the fuselage is to have the wing centre portion pass either over or under the fuselage.
  • the drag penalty from having the wing centre portion pass fully above or below the fuselage has too great an impact on the economics of the aircraft (in terms of fuel burn and maximum economic speed) for it to be a viable solution.
  • this method of attachment is only normally used in military cargo aircraft, where the wing centre portion passes over the fuselage to give a large, uninterrupted cabin volume positioned relatively close to the runway when the aircraft is on the ground so that vehicles or cargo can be quickly loaded via a ramp on the aircraft; or in some business jets, where the thickness of the wing centre portion is small enough to pass under the fuselage so that the attributed drag penalty is more than offset by the production savings. For this reason the centre portion of the wing on most wide-bodied transport aircraft is embedded in the fuselage.
  • the wing centre portion of most modern wide-bodied transport aircraft normally passes below the passenger cabin floor / main cargo deck of the aircraft.
  • This 'low wing' design has dihedral to improve the natural stability of the aircraft.
  • This embedded wing centre portion is known as a "wing box”.
  • Modern transport aircraft normally have a fairing around the wing-fuselage join, again to reduce transonic drag.
  • the simplest wing box to manufacture has a substantially constant section passing straight through the fuselage, perpendicular to the aircraft centre-line. There are no aerodynamic considerations for the wing box, as opposed to the exposed wing portions. Due to the combination of dihedral and sweep on each exposed wing portion relative to the wing centre box, the most practical construction method has been to manufacture the three wing portions as subassemblies, as shown in Figures 1 to 3. The three wing subassemblies can then be joined together with rivets or fasteners, such as bolts.
  • the most efficient aircraft final assembly method including the three wing portions is to build a centre fuselage section around the wing box as a centre fuselage subassembly. Front and rear fuselage subassemblies, and the left and right wing subassemblies can then be joined to this centre fuselage subassembly to form the complete fuselage and wing, as shown in Figure 11.
  • an aircraft wing subassembly comprises an outer wing section for forming an exposed wing area, and an inner wing section for forming a portion of a wing box and formed as an extension to the outer wing section, wherein the inner wing section lies substantially in a first plane and the outer wing section lies substantially in a second plane inclined with respect to the first plane, such that the outer wing section forms an angle of dihedral or anhedral with an aircraft fuselage when the subassembly is joined in an aircraft final assembly.
  • an aircraft wing comprises a mating pair of aircraft wing subassemblies in accordance with the first aspect of the present invention.
  • the wing subassemblies of the 'two section' form of final wing assembly in accordance with the present invention are more complex than the three wing subassemblies of the 'three section' form.
  • this poses no problems using modern composite lay-up methods where the primary wing structure is fabricated out of composite materials.
  • the joint since the two wing subassemblies may be joined by a single joint near the centre of the wing structure, the joint may be positioned at the point of minimum bending loads. As the joint mainly transfers shear loads between the skins of the inner wing sections that form the wing box, the joint can be made considerably simpler with fewer fasteners than the two joins of the 'three section' form.
  • the reduced number of wing joints and consequential reduction in fasteners achieves saving of weight, cost and manufacturing time.
  • Another advantage of the invention is that the transition from the outer to the inner wing sections can be blended to reduce local stress concentrations compared to prior art (where the nature of the joins forces a "sharp" angle).
  • each subassembly has a joining face, the mating wing subassemblies being joined together by a joint at said opposing joining faces.
  • the outer wing sections are aerofoils adapted to generate lift.
  • the outer wing sections may form an angle of sweep with the aircraft fuselage when the subassembly is joined in the aircraft final assembly.
  • each wing subassembly has a fuel tank formed therein.
  • the fuel tank may be formed in both the inner and outer wing sections.
  • the first plane of the inner wing sections form an angle of substantially zero dihedral or anhedral with the aircraft fuselage when the subassemblies are joined in the aircraft final assembly.
  • the wing subassemblies are made, at least in part, of glass-fibre composite, or carbon-fibre composite.
  • glass-fibre composite or carbon-fibre composite.
  • the advantages of the present invention may be realised using alternative composite materials, or even metals such as aluminium or titanium.
  • the joint comprises rivets, bolts, or other suitable fasteners between said opposing joining faces.
  • an aircraft includes a fuselage, and a wing in accordance with the second aspect of the present invention.
  • the wing box formed by joining the two inner portions of the wing subassemblies preferably passes through the fuselage adjacent a circumference thereof.
  • a method of aircraft wing assembly comprises the steps of providing first and second mating aircraft wing subassemblies, each in accordance with the first aspect of the present invention, and joining the first and second aircraft wing subassemblies together by a joint.
  • a method of aircraft final assembly includes the steps of providing a first aircraft fuselage section subassembly having a cut-away portion for accommodating at least a portion of an aircraft wing box, providing, first and second mating aircraft wing subassemblies, each in accordance with the first aspect of the present invention, and joining the first and second aircraft wing subassemblies and the first aircraft fuselage section subassembly together by joints, such that the inner wing sections of the first and second wing subassemblies, which form a wing box, are accommodated by the fuselage cut-away portion.
  • the first aircraft fuselage section subassembly having the cut-away portion accommodates the entire wing box.
  • the method may then comprise the further steps of providing a second aircraft fuselage section subassembly, and joining the second aircraft fuselage section subassembly and the wing box together by a joint.
  • the method of aircraft final assembly further comprises the steps of providing a second aircraft fuselage section subassembly having a cut-away portion for accommodating at least a portion of the aircraft wing box, and joining the second aircraft fuselage section and the wing box together by a joint, such that the wing box is also accommodate by the cut-away portion of the second aircraft fuselage section.
  • the method of aircraft final assembly further comprises the step of providing and joining a fairing to an outside of the fuselage section(s) adjacent the wing box.
  • Figure 1 shows a schematic plan view of a conventional "three-section" aircraft wing prior to final assembly
  • Figure 2 shows a schematic section view of the conventional "three-section" aircraft wing of Figure 1
  • Figure 3 shows a schematic section view of the conventional "three-section” aircraft wing of Figures 1 and 2 after final assembly;
  • Figure 4 shows a schematic plan view of a "two-section" aircraft wing prior to final assembly in accordance with the present invention
  • Figure 5 shows a schematic section view of the "two-section" aircraft wing of Figure 4;
  • Figure 6 shows a schematic section view of the "two-section" aircraft wing of Figures
  • Figure 7 shows an alternative "two-section" aircraft wing similar to that shown in Figure
  • Figure 8 shows a schematic section view of the conventional "three-section" aircraft wing of Figure 1 with structural detail;
  • Figure 9 shows a schematic section view of the "two-section" aircraft wing of Figure 4 with structural detail
  • Figure 10 shows an alternative "two-section" aircraft wing similar to that shown in Figure 9;
  • Figure 11 shows a schematic partially cut-away side view of conventional aircraft fuselage subassemblies prior to final assembly
  • Figure 12 shows a schematic partially cut-away side view of a first example of aircraft fuselage subassemblies and one of the wing subassemblies in accordance with the present invention prior to final assembly;
  • Figure 13 shows a schematic partially cut-away side view of a second example of aircraft fuselage subassemblies and one of the wing subassemblies in accordance with the present invention prior to final assembly.
  • FIG. 1 A schematic plan view of a conventional "three-section" aircraft wing positioned just prior to final assembly is shown in Figure 1.
  • the aircraft wing comprises a starboard wing subassembly 1 , a port wing subassembly 2 and a central wing box 3.
  • the position of a fuselage 4 is indicated in dotted lines.
  • the starboard wing subassembly 1 is positioned adjacent wing box 3 for a joint to be made at a joint position 5.
  • the port wing subassembly 2 is positioned adjacent wing box 3 for a joint to be made at a joint position 6.
  • Figure 2 is a section view taken along line A-A of Figure 1.
  • the position of a belly fairing 7 is shown in dotted lines on the underside of the fuselage 4.
  • FIG. 3 illustrates a similar view to that of Figure 2, after the final assembly of the wing once the starboard and port wing subassemblies 1 , 2 have been joined to the wing box 3.
  • Figure 4 illustrates an embodiment of the present invention in which the aircraft wing is in the form of a "two-section" aircraft wing prior to final assembly.
  • a starboard wing subassembly comprises an outer wing section 11 for forming an exposed wing area. That is, the portion of the wing which is substantially not to be embedded within the aircraft fuselage and which is exposed to moving fluid in order to generate lift.
  • the outer wing section 11 illustrated schematically in Figure 4 may be either a primary frame structure comprising, for example, ribs and spars; or may be a completed outer wing section comprising, for example, ribs, spars, and at least some of upper and lower aerodynamic surfaces, a leading edge, a trailing edge and flaps.
  • the starboard wing subassembly shown schematically in Figure 4 further comprises an inner wing section 18 for forming a portion of a wing box which is to be substantially embedded within the aircraft fuselage.
  • the inner wing section 18 is formed as an extension to the outer wing section 11.
  • the aircraft fuselage is generally indicated by dotted lines 14.
  • a mating port wing subassembly comprises an outer wing section 12 and an inner wing section 19.
  • the port and starboard wing subassemblies of Figure 4 are shown positioned just prior to wing final assembly.
  • the inner wing sections 18, 19 of the starboard and port wing subassemblies, respectively, are to be joined by a single joint at a joint position 15.
  • Figure 5 shows a schematic section view of the "two-section" aircraft wing of Figure 4 taken from line B-B.
  • a belly fairing 17 is indicated by dashed lines.
  • Figure 6 illustrates the section of Figure 5 after final assembly wherein the port and starboard wing subassemblies have been joined by a joint at joint position 15.
  • the joint as in the prior art, may be implemented by use of rivets, fasteners, or other known suitable joining means such as glueing. A combination of joining means may be employed.
  • the joint between the inner wing sections 18, 19 in the preferred embodiment of the present invention shown with reference to Figures 4 to 6 is located substantially along a centre line of the aircraft fuselage in a substantially vertical plane when considered in a zero pitch and zero roll attitude of the aircraft.
  • the inner wing sections 18, 19 have opposing faces which are joined together by a joint at joint position 15.
  • the inner wing sections 18, 19 lie substantially in a first, horizontal plane.
  • the outer wing sections 11, 12 lie substantially in respective planes inclined with respect to the first plane, such that the outer wing sections 11 , 12 form an angle of dihedral, ⁇ , as indicated in Figure 3. Since the outer wing sections 11 , 12 lie in an inclined plane with respect to a plane of the inner wing sections 18, 19, the wing subassemblies have a more complex spar structure than the "three-section" conventional aircraft wing of Figures 1 to 3. However, the use of composite material for the wing subassemblies makes such a complex wing structure relatively easy to manufacture.
  • an aircraft constructed in accordance with the present invention having a "two-section" aircraft wing may reduce the number of fasteners for joining the wing subassemblies together and to the fuselage by up to around 40%. Not only does this reduction in the number of fasteners required reduce the overall weight of the aircraft giving rise to significant reductions in fuel bum during aircraft operation, but also significant reductions in manufacturing cost and time.
  • Figure 5 (as well as Figures 6, 7, 9 & 10) is shown with a discrete angle, ⁇ , at the transition between the outer 11 ,12 and inner 18,19 wing sections in the vertical plane.
  • ⁇ a discrete angle
  • the joint at position 15 between the inner wing sections 18, 19 is shown in a substantially vertical plane along a centre line of the aircraft
  • the opposing faces of the inner wing sections 18, 19 to be joined may take many forms.
  • the opposing faces may have a stepped configuration, may be formed in an inclined plane to the vertical plane described above, or may be offset from the centre line of the aircraft. All such alternatives are envisaged within the scope of the present invention.
  • FIG. 7 there is shown an alternative "two-section" aircraft wing similar to that shown in Figure 6.
  • the inner wing sections 18, 19 have extensions 18a, 19a extending downwardly from the inner sections 18, 19, respectively.
  • the extensions 18a, 19a do not increase the volume of the overall aircraft but would allow the volume of any fuel tanks in the wing to be increased in volume.
  • the conventional "three-section" aircraft wing such as illustrated schematically in Figure 3
  • three fuel tanks are typically provided, one in each of the starboard 2, port 1 and centre 3 wing sections.
  • the barriers between the tanks are usually formed by sides of the centre wing box 3, near the boundary of the fuselage.
  • the "two-section" wing in accordance with the present invention permits the fuel system to be simplified such that there are only two tanks, one formed in each of the starboard 11 and port 12 wing subassemblies.
  • the inner wing sections 18, 19 may have opposing side walls to be joined when the joint is formed at position 15 during aircraft final assembly.
  • the starboard and port wing subassemblies may then be formed having an open ended structure adjacent one another.
  • integral fuel tanks of such wing subassemblies may be sealed by joining each of the starboard and port wing subassemblies to a generally vertical barrier wall disposed between them, again at positions 15.
  • Conventional fuel systems may connect the integral fuel tanks of the "two-section" wing in accordance with the present invention.
  • Figure 8 illustrates the conventional wing spar structure of a "three-section" wing in accordance with Figures 2 and 3. Such a wing spar structure is typically utilised to form the front and rear elements of the primary wing structure.
  • Figure 9 illustrates the wing spar structure of a "two-section" wing in accordance with the present invention as illustrated in Figures 5 and 6.
  • Figure 10 illustrates the internal structure of a "two- section" wing in accordance with the embodiment of Figure 7 of the present invention.
  • exemplary layouts of the wing ribs can be clearly seen.
  • the aircraft wing subassembly in accordance with the present invention finds particular application when constructed of composite materials such as glass fibre or carbon fibre.
  • composite materials such as glass fibre or carbon fibre.
  • traditional aircraft materials such as aluminium or titanium may be similarly employed and their use is envisaged within the scope of the present invention.
  • the present invention also provides a method of assembling the aircraft wing subassemblies in a wing, or final, aircraft assembly. It will be appreciated by those skilled in the art that the wing subassemblies may be constructed at a different location from that where aircraft final assembly occurs, and that aircraft wing assembly from the two wing subassemblies may also occur at a different location to that of aircraft final assembly.
  • FIG. 11 illustrates the relationship between various aircraft subassemblies just prior to aircraft final assembly.
  • the wing box 3 is typically integrally formed with a centre section 42 of an aircraft fuselage.
  • Fore and aft aircraft fuselage sections 40, 41 are joined to the aircraft centre fuselage section 42 at positions 44, 45, respectively during aircraft final assembly.
  • the starboard 1 and port 2 wing subassemblies are joined to the centre wing box 3 at positions 5, 6, respectively.
  • the belly fairing 7 may then be joined to the aircraft fuselage beneath the centre wing box 3.
  • Figure 12 illustrates the positions of various aircraft subassemblies just prior to aircraft final assembly in accordance with the present invention. Since the wing box is formed as extensions to the starboard and port wings, the wing box is not formed as part of a centre fuselage section. This makes it possible to join the fore and aft fuselage sections 50, 51 by a single join at position 53. To make this possible, the fore and aft fuselage sections 50, 51 include cut-away portions 52 for accommodating the inner wing portions 18, 19 which constitute the wing box.
  • Various joining possibilities of the aircraft subassemblies will be apparent to those skilled in the art in considering the layout of Figure 12.
  • the fore and aft fuselage sections 50, 51 may be joined at position 53 and separately the starboard and port wing subassemblies may be joined to form the wing prior to joining of the wing to the fuselage.
  • one of the two wing subassemblies may be joined to either the fore or the aft fuselage section 50, 51 , followed by joining the other wing subassembly and the other fuselage section, or the four subassemblies may be joined concurrently.
  • the method of aircraft final assembly illustrated by Figure 12 reduces the number of joints between aircraft fuselage subassemblies and also makes it possible to position the joint 53 at a point of minimum bending load.
  • the joint mainly transfers shear loads between the skins of the fuselage sections 50, 51, the joint can be made considerably simpler with fewer rivets or fasteners than the method of assembly of Figure 11.
  • FIG 13 An alternative method for the final assembly is shown in Figure 13. Again, a cut-away portion 57 is formed in the fore or the aft fuselage sections, 54, 55 (cut-away portion 57 is formed in the fore fuselage section 54 in Figure 13) whereby a single joint between the fore and aft fuselage sections 54, 55 may be made at position 56.
  • the aircraft wing may be joined to the fuselage either prior to, or after, joining of the fore and aft fuselage sections 54, 55. Whilst the aircraft produced by the method of final assembly illustrated by Figure 13 may not place the top part of the joint at 56 in an optimal portion in the fuselage with respect to bending loads, the benefit of a single joint between fore and aft fuselage sections remains as for Figure 12. In either method of construction of Figure 12 or Figure 13, the belly fairing 17 may be subsequently joined to the aircraft fuselage.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Connection Of Plates (AREA)

Abstract

Sous-ensemble voilure d'aéronef comportant une section de voilure externe (11, 12) destinée à constituer une zone de voilure exposée, et une section de voilure interne (18, 19) destinée à constituer une partie de caisson de voilure et formée comme une extension de la section de voilure externe. La section de voilure interne se situe sensiblement dans un premier plan, la section de voilure externe se situant sensiblement dans un second plan, incliné par rapport au premier. La section de voilure externe forme un angle dièdre ou un dièdre négatif par rapport au fuselage (14) de l'aéronef lorsque ce sous-ensemble est accouplé lors de l'assemblage final de l'aéronef. L'invention concerne également un procédé d'assemblage final d'aéronef comprenant les étapes suivantes consistant à : réaliser un premier sous-ensemble de section de fuselage d'aéronef disposant d'une partie découpée destinée à recevoir au moins une partie de caisson de voilure d'aéronef; réaliser deux sous-ensembles d'accouplement de voilure d'aéronef de ce type; accoupler, par l'intermédiaire de jonctions, ces deux sous-ensembles de voilure d'aéronef et le premier sous-ensemble de section de fuselage d'aéronef, de façon à ce que les sections de voilure interne qui constituent un caisson de voilure, puissent se loger dans la partie découpée du fuselage.
PCT/GB2007/000653 2006-02-28 2007-02-26 Voilure d'aéronef et son procédé de montage WO2007099297A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GB0604017A GB2435457B (en) 2006-02-28 2006-02-28 Aircraft wings and their assembly
GB0604017.4 2006-02-28

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WO2007099297A1 true WO2007099297A1 (fr) 2007-09-07

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Cited By (4)

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Publication number Priority date Publication date Assignee Title
JP2008114707A (ja) * 2006-11-02 2008-05-22 Honda Motor Co Ltd 飛行機の主翼構造
CN101909989A (zh) * 2007-12-30 2010-12-08 空中客车营运有限公司 用于连接飞机上的两个主机翼和机身段的机翼-机身-结构件
CN109229373A (zh) * 2018-09-29 2019-01-18 中船重工（海南）飞船发展有限公司 一种全复合材料地效翼船的连接结构
US11174042B2 (en) 2018-10-09 2021-11-16 The Boeing Company Wing join system and method for a wing assembly

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FR2948099B1 (fr) * 2009-07-16 2012-05-11 Airbus Operations Sas Procede de fabrication d'un aeronef a voilure fixe
US10118686B2 (en) * 2012-09-27 2018-11-06 The Boeing Company Wing root insert system for an aircraft family
FR3091259A1 (fr) * 2018-12-26 2020-07-03 Airbus Operations Voilure d’aeronef comprenant deux ailes fixees l’une sur l’autre
EP3998192B1 (fr) * 2021-02-19 2024-03-20 Lilium eAircraft GmbH Structure de longeron d'aile

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US6114050A (en) * 1996-01-11 2000-09-05 The Boeing Company Titanium-polymer hybrid laminates
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EP1413782A1 (fr) * 2002-10-21 2004-04-28 The Boeing Company Procédé et dispositif de support pour éléments rotatifs
US20050230538A1 (en) * 2004-03-04 2005-10-20 Bruno Sarpy Fuselage spar for aircraft and central sparbox provided with such a spar

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JPS5529629A (en) * 1978-08-18 1980-03-03 Mitsubishi Heavy Ind Ltd Aircraft having monolithic wing and body
US6114050A (en) * 1996-01-11 2000-09-05 The Boeing Company Titanium-polymer hybrid laminates
US6190484B1 (en) * 1999-02-19 2001-02-20 Kari Appa Monolithic composite wing manufacturing process
EP1413782A1 (fr) * 2002-10-21 2004-04-28 The Boeing Company Procédé et dispositif de support pour éléments rotatifs
US20050230538A1 (en) * 2004-03-04 2005-10-20 Bruno Sarpy Fuselage spar for aircraft and central sparbox provided with such a spar

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2008114707A (ja) * 2006-11-02 2008-05-22 Honda Motor Co Ltd 飛行機の主翼構造
CN101909989A (zh) * 2007-12-30 2010-12-08 空中客车营运有限公司 用于连接飞机上的两个主机翼和机身段的机翼-机身-结构件
CN109229373A (zh) * 2018-09-29 2019-01-18 中船重工（海南）飞船发展有限公司 一种全复合材料地效翼船的连接结构
CN109229373B (zh) * 2018-09-29 2020-07-03 中船重工（海南）飞船发展有限公司 一种全复合材料地效翼船的连接结构
US11174042B2 (en) 2018-10-09 2021-11-16 The Boeing Company Wing join system and method for a wing assembly

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GB0604017D0 (en) 2006-04-05
GB2435457B (en) 2011-03-09

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