Aircraft Fuselage and Method of Forming Same

This invention relates to the construction of aircraft fuselages, and more particularly to an aircraft fuselage and method of forming same through the use of a plurality of independently formed splice longerons and independent fastening members which allow a fuselage portion having a non-constant diameter to be formed from a plurality of extruded skin panels.

Most aircraft employ a fuselage having a constant-section (i.e., center) portion and doubly curved forward and aft sections. While such structures could potentially be produced using a variety of processes, large parts such as stiffened aircraft fuselage skins presently can be made from large axially stiffened near-net extrusions or, alternatively, directly machined out of plate. The use of extrusions are favored from a cost-effectiveness standpoint because extruding involves less labor and utilizes a much greater fraction of the raw material used for producing the fuselage skins. Patents involving the formation of aircraft fuselage skins or other curving structures are U.S. Pat. Nos. 2,230,393; 2,427,065; 2,458,686; 3,096,958; 3,920,206; 3,940,891; 4,310,132; 4,471,609 and 5,154,373, the disclosures of which are hereby incorporated by reference.

For the constant-section portion of the fuselage, the application of extrusions if fairly straightforward because the fuselage skins can be formed with integrally formed stiffeners which are parallel to one another. Wing planks often employ parallel stiffeners in both integral and multi-piece construction because they are substantially flat or singly curved. However, the application of extrusions in the doubly curved forward and aft sections of the fuselage is complicated by the fact that the spacing of the stiffeners typically varies as the fuselage diameter increases or decreases. As will be appreciated, with present day extrusion processes it is not possible to achieve variable spacing of the stiffeners when extruding large stiffened fuselage panels. Thus, doubly curved portions of the fuselage, following present day design and fabrication concepts, can only be made with integrally formed stiffening portions by machining out of non-near-net stock, such as plate, with the associated inefficient material usage.

It would therefore be highly desirable to provide an aircraft fuselage and a method for manufacturing same which allows a plurality of extruded fuselage panels, each having a plurality of integrally formed stiffening elements, to be secured to one another by a plurality of independent members such that doubly curved sections of the fuselage can be formed in a cost effective manner.

It would also be highly desirable to provide a method of forming a complexly shaped aircraft fuselage through conventional extrusion techniques, wherein independent skin panels of the fuselage each have integrally formed stiffening members, and where the integrally formed stiffening members of adjacent skin panels can be secured together by independent fastening members to provide added structural strength and rigidity to the fuselage.

Still further, it would be very advantageous to be able to form an aircraft fuselage having one or more doubly curved portions in which a plurality of fuselage panels are extruded, and where each of the fuselage panels include integrally formed stiffening portions which could be coupled together via a plurality of splice longerons and also via independent fastening members to fasten the stiffening portions of each of the fuselage panels to the splice longerons at the areas where the stiffening portions run out into contact with the splice longerons.

The above and other objects are provided by an aircraft fuselage and method of forming same in accordance with the present invention. The method of the present invention involves extruding a plurality of fuselage panels, wherein each fuselage panel includes at least one integrally formed stiffening portion extending longitudinally therealong. Each of the fuselage panels is then formed into a fuselage skin panel having at least one doubly curved portion. Each of the fuselage skin portions are then trimmed such that when the fuselage skin panels are secured together they will form a portion of the aircraft fuselage having the desired complex contour. The fuselage skin panels are then secured to one another along longitudinal edges thereof by a plurality of independently formed splice longerons. The splice longerons couple adjacent longitudinal edges of adjacently positioned pairs of fuselage skin panels and may be riveted, welded or adhesively bonded to the longitudinal edges. Alternatively, adjacent skin panels may be secured to one another by welding, such as friction stir welding, or other suitable welding techniques, or even adhesives.

In one embodiment independent fastening members are used to couple those integrally formed stiffening portions of adjacent fuselage skin panels which run out towards and into contact with an intermediately disposed splice longeron to the splice longeron. The independent fastening members may be riveted, adhesively bonded or welded to the stiffening portions and the splice longeron to provide structural strength and rigidity to the assembled fuselage.

With the apparatus and method of the present invention, the aircraft fuselage panels and the splice longerons may be extruded, thus saving significant material costs over other manufacturing processes for constructing aircraft fuselages. More specifically, the method of the present invention enables complexly shaped fuselages to be formed through efficient extrusion processes, therefore resulting in significantly less wasted material and reduced labor costs.

The various advantages of the present invention will become apparent to one skilled in the art by reading the following specification and subjoined claims and by referencing the following drawings in which:

FIG. 1 is an elevational view of an aircraft having a fuselage having doubly curved forward and aft sections;

FIG. 2 is a plan view of a plurality of independent fuselage panel portions, wherein the dashed lines indicate portions of each fuselage panel which have been trimmed away prior to each of the sections being secured together, and which also indicate a plurality of longitudinally extending, integrally formed stiffening portions formed on each fuselage panel;

FIG. 3 is a cross sectional view in accordance with section line 3--3 in FIG. 2 illustrating the plurality of integrally formed stiffening portions of one fuselage panel;

FIG. 4 is a fragmentary elevational view of an inside surface portion of a portion of an aircraft fuselage with the integrally formed stiffening portions indicated in solid parallel lines;

FIG. 5 is a perspective view of a pair of fuselage panels which have been formed into fuselage skin panels and a portion of one splice longeron which is used to secure the two fuselage skin panels together adjacent one another; and

FIG. 6 is a perspective view of two independent fastening members securing portions of integrally formed stiffening members from two adjacently positioned fuselage skin panels to an intermediately disposed splice longeron.

Referring to FIG. 1, there is shown an aircraft 10 having a fuselage 12. The fuselage 12 has a generally constant diameter center section 14, a doubly curved forward section 16 and a doubly curved aft section 18. By "doubly curved", it is meant a cross sectional portion of the fuselage 12 which has a non-constant or changing radius. As explained previously herein, the manufacture of such a fuselage 12 by present day conventional methods can be costly and inefficient. The method and apparatus of the present invention, however, allows for the manufacture of complexly shaped fuselages in a more efficient and cost effective manner, as will now be described.

Referring to FIG. 2, the method of the present invention involves forming the fuselage 12 by extruding a plurality of fuselage panels 12a-12d. It will be appreciated that the number of fuselage panels 12a-12d will depend upon the overall diameter of the fuselage 12. It will also be appreciated that for the purpose of illustration, panel 12a has been illustrated split into two sections 12a1 and 12a2. Thus, the illustration of four fuselage panels 12a-12d is meant to be representative only of the fact that several independent fuselage panels are typically required to form the fuselage 12.

With further reference to FIG. 2, each of the fuselage panels 12a-12d includes a plurality of integrally formed and longitudinally extending stiffening portions 20a-20d, respectively. Again, for the purpose of illustration, portion 20a has been split into portions 20a1 and 20a2. With brief reference to FIG. 3, the stiffening portions 20c can be seen in cross section. Stiffening portions 20c each form generally inverted L-shaped members protruding from an inner surface 22c of fuselage panel 12c. The stiffening portions 20c further are formed to extend, in one embodiment, generally parallel to one another along preferably the entire length or a major portion of the length of the fuselage panel 12c, as can be seen in FIG. 2. It will also be appreciated that in some instances it may be desirable to orient the stiffeners such that they do not extend perfectly parallel to each other or to adjacent panels 12a-12d, but rather diverge or converge as may be required or desired.

The fuselage panels 12a-12d are extruded using conventional extrusion techniques. The extrusion process permits the stiffening portions 20a-20d of each panel 12a-12d to be integrally formed with its fuselage panel, thereby reducing the labor required to form each fuselage panel 12a-12d. The extrusion process also utilizes a much greater portion of the raw material required to make each fuselage panel 12a-12d.

Referring to FIG. 4, prior to being assembled together, each of the fuselage panels 12a-12d are formed into contoured fuselage skin panels such as by stretch forming, creep forming, peen forming or other known processes. The skin panels are then trimmed into gores by removing the material indicated in dashed lines in FIG. 2 from each panel 12a-12d. When spliced together, the panels 12a-12d form a surface having a desired contour. The trimming step could, however, just as easily precede the forming step if desired. Because the stiffening portions 20 of each of the panels 12a-12d have a constant spacing relative to one another prior to each of the fuselage panels 12a-12d being formed into fuselage skin panels, the stiffening portions 20 of each fuselage skin panel 12a-12d will "run-out" at various places along the longitudinal edges of each of the panels, as indicated by reference numerals 24 in FIG. 2.

With further reference specifically to FIG. 4, when any two of the fuselage skin panels 12a-12d are assembled together, the curvature or contour of each of the fuselage skin panels 12a-12d causes the stiffening portions 20c and 20d to run out to the longitudinal edges 26 at approximately the same points, as indicated by reference numerals 28 in FIG. 4. It will be appreciated immediately that in some instances there will usually be several areas where stiffening portions 20 from adjacently positioned fuselage skin panels 12a-12d run out and meet. The number of areas at which the stiffening portions run out and meet will vary depending upon the curvature of the doubly curved portion of the fuselage 12.

Referring now to FIG. 5, to secure adjacent fuselage skin panels 12 together, a plurality of preferably extruded splice longerons 30 are formed and cut to desired lengths. The splice longerons 30 are shown as having an inverted T-shape, but could just as readily take other shapes as desired. The desired length will depend on the overall length of each of the fuselage skin panels 12a-12d which are to be secured together. Preferably, the splice longeron 30 will be about equal to the overall length of the fuselage skin panels 12 such that the splice longeron 30 will overlap the longitudinal edges 26c and 26d of each of the fuselage panels 12c and 12d, respectively, for the entire length of each fuselage skin panel 12c and 12d.

Referring now to FIG. 6, the coupling of converging stiffening portions 20c and 20d of fuselage skin panels 12c and 12d, respectively, will be described. In the doubly curved portion of the fuselage 12, the stiffening portions 20c and 20d converge towards longitudinal edges 26c and 26d of each of the respective fuselage skin panels 12c and 12d. The first step in securing the fuselage skin panels 12c and 12d together, therefore, is the trimming of end portions of the converging stiffening portions 20c and 20d to allow room for the splice longeron 30 to be laid over the longitudinal edge portions 26c and 26d of adjacent skin panels 12c and 12d, respectively. This is shown in phantom by cut edge 32c of stiffening portion 20c. Cut edge 32c allows splice longeron 30 to be laid over the longitudinal edges 26c and 26d without interference from stiffening portion 20c. If not for cut edge 32c, the portion of the stiffening portion 20c which runs out to the longitudinal edge 26c would interfere with placement of the splice longeron 30 along the longitudinal edge 26c.

With further reference to FIG. 6, once each of the stiffening portions 20c and 20d have been trimmed or cut to make room for the splice longeron 30, the splice longeron is placed over the longitudinal edges 26c and 26d along preferably the entire length of each panel 12c and 12d. The splice longeron 30 includes an upright portion 30a and connecting portions 30b. The connecting portions 30b are preferably each formed with a plurality of openings 31 which allow rivets or mechanical fasteners to be used to secure the connecting portions 30b to each of the longitudinal edges 26c and 26d along the full length of each of the fuselage skin panels 12c and 12d. It will be appreciated, however, that other means of attaching the longeron 30 could also be employed, such as welding or adhesive bonding, in which case it would not be necessary to form the openings 31.

With further reference to FIG. 6, once the splice longeron 30 is secured to the fuselage skin panels 12c and 12d, independent fastening members 34 are used to secure portions of each of the stiffening portions 20c and 20d to the splice longeron 30. Each of the independent fastening members 34 is a component which may be produced through conventional casting, forging or machining techniques to provide a first portion 34a and a second portion 34b. The second portion 34b is formed at an angle relative to the first portion 34a, which is the angle at which each of the stiffening portions 20 converges toward the longitudinal edge 26 of its fuselage skin panel 12. The first portion 34a of each fastening member 34 includes an upright portion 36 and a connecting portion 38. Each of the portions 36 and 38 include openings 39 which allow the fastening member 34 to be secured such as by rivets to the splice longeron 30. The second portion 34b of each fastening member 34 is generally L-shaped and also includes a plurality of openings 40 which permit the second portion 34b to be secured to the stiffening portion 20c such as by rivets. In this regard it will be appreciated that once the fastening members 34 have been positioned in place, if rivets or mechanical fasteners are to be used for securing, then holes will need to be drilled at the appropriate places in the upright portion 30a of splice longeron 30 and in the stiffening portions 20c to allow the rivets to pass therethrough. Once the fastening members 34 are secured to the splice longeron 30, an extremely rigid and structurally strong joint is formed. It will be appreciated that while FIG. 6 illustrates only one point at which the stiffening portions 20c and 20d run out to the longitudinal edges 26c and 26d, respectively, that in fact there will almost always be multiple points at which the stiffening portions 20 of adjacently positioned fuselage skin panels 12 run out to the longitudinal edges of the panels 12a-12d and will need to be coupled together. Thus, the coupling arrangement shown in FIG. 6 will be present, in most cases, at a large plurality of points in a doubly curved fuselage 12.

The apparatus and method of the present invention allows the fuselage panels to be extruded, which is a very cost-effective method of manufacture because it involves less labor and utilizes a much greater fraction of raw material than if the fuselage panels (including the integrally formed stiffening portions) were machined directly out of plate.

Those skilled in the art can now appreciate from the foregoing description that the broad teachings of the present invention can be implemented in a variety of forms. Therefore, while this invention has been described in connection with particular examples thereof, the true scope of the invention should not be so limited since other modifications will become apparent to the skilled practitioner upon a study of the drawings, specification and following claims.