INTERDIGITATING PORTIONS

FIELD OF THE INVENTION The present invention relates to a fastenerless joint, and more particularly to such a joint for composite aircraft structures.

BRIEF DESCRIPTION OF THE PRIOR ART In many fabrication applications laminated structures are attached to support structures by common fastening techniques. In one particular application, composite laminated aircraft wing skins are attached to a support substructure by means of rivets. Within the aircraft environment this creates several problems. In the fiifst place, the rivets passing through the various plies of the laminated structures create stress forces which could cause rupture and ultimate destruction of the aircraft.

In modern fabrication techniques ^' for composite aircraft skins, fuel or other liquids are contained within the skin so that the aircraft structure itself creates a reservoir for the fluid. A serious problem with current construction techniques utilizing fasteners is the presence of leaks where rivets penetrate the composite laminate.

For these and other reasons, it would be highly desirable to create a mechanical joint between aircraft composite laminated skin structures and support subasse blies without the utilization of discrete fasteners .

BRIEF DESCRIPTION OF THE INVENTION The present invention utilizes mechanically interleaved plies of a composite laminate wing structure

which enforces the normal bonding between plies of the laminate.

Alternating projections are formed from the wing laminate and these receive mating recesses in a composite wing substructure so that interdigitation of the wing skin structure and support substructure results. The interdigitating members are hollow so that an elongated bar, which may be fabricated from the composite laminate, may be slid therethrough thereby locking the members together.

The result is securement between the wing skin structure and support substructure without the need for discrete metallic fasteners. By avoiding metal fasteners such as rivets, the previously discussed problems of high shear stress and leakage can be eliminated.

Further advantages of the present invention relate to automated precut material use which lends itself to rapid repeatable production and reduction of layup time. The tooling required is relatively uncomplicated, which serves as a distinctive economic advantage of the invention.

BRIEF DESCRIPTION OF THE FIGURES The above-mentioned objects and advantages of the present invention will be more clearly understood when considered in conjunction with the accompanying drawings, in which:

FIG. 1 is a perspective view illustrating interlocking engagement between a wing skin structure and support structure in accordance with the present invention; FIG. 2A is a perspective view of a top ply as employed in the present invention;

FIG. 2B is a perspective view of a bottom ply as employed in the present invention;

FIG. 2C is an illustration of juxtaposed top and bottom plies;

FIG. 3 is an illustration of a finished construction in accordance with the present invention wherein multiple top and bottom plies are employed.

DETAILED DESCRIPTION OF THE INVENTION Reference is made to FIG. 1 wherein interlocking engagement between a wing skin structure and support substructure is illustrated. The interlocking members are generally indicated by reference numeral 10 and are seen to include a wing support substructure 12 secured to a wing skin structure 14. In order to achieve the interlocking engagement between members 12 and 14, trapezoidal projections 16 are formed in the wing skin structure 14 at longitudinally spaced intervals. The wing support substructure 12 has, along a triangular base portion, slots 13 formed therein - along similarly spaced intervals. The.slots 13 receive corresponding trapezoidal projections 16 so that the engaging portions form interdigitations. The triangular base portion of the substructure 12 and the trapezoidal projections 16 are hollowed so that a bar 20 can slide therethrough to secure interlocking engagement between the substructure 12 and the projections 16.

It is to be noted that the wing support substructure 12 is shown mounted in longitudinal offset relationship with the wing skin substructure 14 so that details of the invention are better illustrated. In actual utilization of the invention, the substructure 12 would be shifted to more completely overlie the skin structure 14.

In order to better appreciate the construction of the present invention, reference is made to FIGS. 2A-2C.

Referring to FIG. 2A, a top ply for the wing skin structure is generally indicated by reference numeral 21. The ply may be fabricated from a sheet 22 made of Fiberglass or other suitable composite material. Three

spaced openings 24 , 2β and 28 are illustrated, by way of example . A bottom ply, which will be positioned in underlying juxtaposition with the top ply is shown in FIG. 2B and is generally indicated by reference numeral 30. The bottom ply 30 is also fabricated from a sheet 32 o f appropriate materia l .- The bottom ply has a longitudinal medial opening 33 interrupted by integrally formed connecting trapezoidal projections 34, 36 and 38. By lowering the top ply over the bottom ply so that proj ections 34 , 36 and 38 respectively project upwardly through openings 24 , 26 and 28 , the two plies become geometrically locked, as shown in FIG. 2C. The top and bottom plies together form an elementary laminate. As is the case in existing composite aircraft laminated structures, the individual plies must be impregnated with resin and cured. Conventional pre- impregnated resins may be successfully used in accordance with known practices.

FIG. 3 illustrates a more customary utilization of multiple top and bottom plies to form a wing skin structure 14. In the illustration shown in FIG. 3 , two top plies 22a and 22b are indicated as being alternately positioned between two bottom plies 32a and 32b. Since two bottom plies are employed, each of the trapezoidal proj ections 16 is seen to include dual overlying plies (34a, 34b; 36a, 36b; and 38a, 38b) .

The final construction of the invention will now be explained in connection with FIG. 1. The four-ply wing skin structure 14 is seen to resemble that of FIG. 3, just discussed. The wing support substructure 12 is shown as fabricated from two plies of resin-impregnated fabric, which are folded over one another to form the flat portion 42. The illustrated layers 40a-40d make up the flat portion 42 and it should be understood that layers 40a and 40d are opposite ends of the same single ply while layers 40b and 40c are opposite ends of the same second ply. A lower portion of the wing support structure is

formed in a closed hollowed triangular portion generally indicated by reference numeral 44.

By positioning the triangular cross section slots 13 in interdigital relation with the trapezoidal projection 16, the wing support substructure 12 may be longitudinally fixed. However, in order to fasten the wing skin structure 14 and wing support substructure 12 together, the slidable bar 20 is longitudinally passed through the opening in the triangular portion 44 and the opening in the spaced trapezoidal projections 16 rising above the surface of the wing skin structure 14. The bar 20 is preferably made from a resin-impregnated composite fabric laminate (20a-20d) , such as Fiberglass, as is the case with the wing skin structure 14 and support substructure 12. After the bar is slid completely through the support substructure 12, mechanical secured interlocking therebetween is achieved.

In order to fabricate the composite wing skin structure and composite wing support substructure as rigid load-supportable members, the resin impregnation of the composite fabrics must be cured. In this regard the wing skin structure 14 may be cured separately from the support substructure 12 and then assembled. Alternatively, the wing skin structure 14 may be connected to the support substructure 12 while both are still wet. Then, they may be co-cured while a mandrel (not shown) is located within the recess that bar 20 occupies in FIG. 1. After co-curing the mandrel may be removed; and bar 20, separately cured, is passed through the substructure 12 and trapezoidal projection 16.

In order to increase the substructure's resistance to shear forces, the individual layers 40a-40d may be laced together. This is simply achieved by forming rectangular spaced slots 46 longitudinally along the flat portion 42 of the substructure 12. Then, a strip of fabric laminate 48 may be threaded through the slots 46 thereby passing

through all of the layers 40a-40d in interleaved securement thereof. The strip 48 is preferably threaded through slots 46 and then cured with the substructure 12. In a finally assembled aircraft, the wing support substructure 12 and trapezoidal projections 16 are positioned internally of the wing. The underside of the wing skin structure 14, opposite the surface 18 illustrated, will form the exterior surface of the wing skin. Although the invention thus far described is resistive to longitudinal loads, it will be appreciated that the disposition of wing support substructure 12 forms somewhat of a hinge connection with the wing skin structure 14. In order to resist rotational displacement of the wing skin structure 14 relative to substructure 12 , additional reinforcement means is preferred. A web reinforcement means is generally indicated in FIG. 1 by strip 50. As will be seen, a laminated fabrie composite is employed. For illustrative purposes, the reinforcement strip includes four plies 50a-50d. The strip is positioned longitudinally along the wing support substructure 12 so that it intimately contacts the surface of the flat portion 42 and the triangular portion 44. Also, the bottom of the reinforcing strip 50 is attached to the upper surface of wing skin structure 14. The strip, being impregnated with resin or the like, is attached to structures 12 and 14 to resist rotational displacement of the wing support substructure 12. Although a single strip 50 is shown on the near side of substructure 12 and wing skin structure 14, this has been done to simplify the view; and in fact, a second identical strip is located on the far side of structures 12 and 14.

Although the present invention has been explained in connection with trapezoidal projection 16 and a triangular-based wing support substructure 12, this has

been done to simplify the explanation and is only illustrative. Other mating shapes may be similarly employed.

As will be appreciated, by virtue of the connection of wing skin structure 14 to wing support substructure 12, without the aid of rivets, high stress points and leakage points between the two structures are eliminated.

Although the present invention has been described in terms of a fastenerless joint for securing an aircraft wing skin to a support structure, it should be appreciated that the invention contemplates other applications. Thus, the following claims envision general utilization of projections from a first laminated structure to form interdigitations with a support substructure to achieve a secure joint therebetween. Specifically, the present invention is envisioned for applications to securing other aircraft skin sections to support substructures, in addition to a wing skin as discussed herein.

It should be understood that the invention is not limited to the exact details of construction shown and described herein, for obvious modifications will occur to persons skilled in the art.