Adhesive bonding and miscellaneous chemical manufacture – Methods – Surface bonding and/or assembly therefor
Reexamination Certificate
1999-11-26
2002-04-16
Aftergut, Jeff H. (Department: 1733)
Adhesive bonding and miscellaneous chemical manufacture
Methods
Surface bonding and/or assembly therefor
C156S166000, C156S299000, C442S204000
Reexamination Certificate
active
06372072
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates to the design and manufacture of composite laminates, particularly fibre reinforced composite laminates, and more particularly to such laminates which are constructed from a series of plies of fabric, each ply being comprised of reinforcing fibres layed substantially in a single orientation. Such laminates find utility in the manufacture of highly stressed structures, for example in the manufacture of aircraft wings, where the ratio of strength and stiffness to weight is critical.
2. Description of Prior Art
An optimised aircraft wing, especially one made of carbon fibre based composite, is a very complex structure due to the variety of functions required of different types of wing component, eg. skins, stringers, ribs and spars. A potential advantage of composites over metals is that the strength and stiffness of the structure can be tailored by varying the percentage content of reinforcing fibre, the fibre orientation and sometimes the fibre modulus. In this way, for example, different spanwise and chordwise moduli in a skin can be achieved, varying from wing root to tip.
A unidirectional “prepreg” tape, typically 0.125 mm-0.25 mm thick and 75 mm to 300 mm wide, is traditionally used within the aerospace industry for primary structure applications to achieve these qualities.
Prepreg material is fibrous reinforcing material pre-impregnated with a controlled viscosity plastics matrix material, usually epoxy resin. The above method of laying up the laminate allows a very high degree of design tailorability and thus the obtaining of an optimum strength and stiffness laminate. However the layup process for a large thick structure is very slow for a number of reasons.
Firstly, as 4 to 8 layers of tape are required for every 1 mm of laminate thickness, a very large number of “passes” of the tape head can be required. Secondly, in order to lay up the tape in the required fibre orientations (the fibre orientation of the tape itself is always lengthwise), the tape laying head will need to move across the layup in all those directions of orientation, thus requiring an expensive head movement apparatus and complex computer software. Thirdly, for a 3-dimensional laminate like portions of an aircraft wing, a further degree of freedom of the tape laying head will be required to accommodate the 3-D form.
Hence, to lay up a laminate of the size and shape of a modern passenger aircraft wing requires both very expensive tape laying apparatus and a long layup time. Also the amount and cost of scrap consumables, eg. release paper, can be high for tape laying methods.
An alternative design approach for laying up laminates, in particular thick laminates, is to use thicker multi-ply fabrics which may also be much wider than tapes, eg 1.26 m or more, and therefore much faster to lay up.
Non crimp fabrics, “NCFs”, are a type of fabric in which the fibre tows in a ply of the fabric lie in their particular direction of orientation substantially without crimping. Any number of plies of fabric may be held together, prior to the application of matrix material, by stitching or by a low level of binder to form a fabric of a given “style”. In the context of unidirectional fabric plies the “style” is taken to mean the areal mass, for the purpose of this invention. Also included within the definition of NCF, for the purposes of this invention, are fabrics known as “uniweaves” which are comprised of unidirectional fibres lightly woven about thin thread.
Multi-ply NCFs can currently be in styles having up to seven plies of fabric. Each ply will normally comprise unidirectionally orientated fibres and each ply will normally have fibres orientated in a direction different from its neighbour. In the manufacture of aerospace vehicle structures, and in other high strength applications, commonly used fibre orientations are as follows: 0 degrees (“0 deg”)/warp (along a datum line, eg spanwise of an aircraft wing); plus or minus 45 deg; 90 deg/weft. It is possible to use other directions of fibre orientation, eg 30 deg and 60 deg, and any such other directions should be taken as included within the scope of this invention. The use of the following terms: 0 deg; plus or minus 45 deg; 90 deg in this specification is not intended to be limiting and other fibre orientations not differing substantially from the quoted figures are included within the meanings of these terms.
Such fabrics which include several plies of unidirectional fibres at different orientations are known as “multiaxial NCFs”. A key advantage of using multi-axial fabrics is that they can be laid up in a single direction whilst providing fibre orientations in all the required directions for optimum stress-bearing characteristics. Huge time savings can result for the layup process when these fabrics are used.
An example of such a multiaxial fabric comprising a style of seven plies, as mentioned above, could have through-thickness fibre orientations of 0 deg (20%); +45 deg (12.5%); −45 deg (12.5%); 90 deg (10%); −45 deg (12.5%); +45 deg (12.5%); 0 deg (20%). Such a style would be both balanced and symmetrical (ie non handed) about its neutral axis and could be made up by the combination of a triaxial fabric (preferably NCF) with a second identical but reversed, ie oppositely handed, triaxial, both sandwiching a 90 deg uniweave fabric between them. Alternatively, a triaxial fabric could be combined with a quadraxial fabric.
The thickness of NCF fabrics, for aerospace applications, can be from 0.1 mm (for say a single ply NCF) to 1.5 mm for a multi-ply NCF. Based on unidirectional tape experience, ply thicknesses may be suitably 0.5 mm possibly up to 1.0 mm, for an angular change in fibre orientation between plies of 45 deg, and say 0.25 mm thick for an angular change of 90 deg.
Historically, a number of design rules have arisen for this type of laminate, based on analysis and experience with unidirectional tape. Such rules may differ depending on the design/certification philosophy of the aircraft manufacturer concerned but are generally stated here to be as follows:
45 deg plies are required on the exterior of a composite structure for damage tolerance and softening of stress inputs into adhesive bondlines;
the thickness of a unidirectional ply (ply of fabric having unidirectional fibre orientation) should not normally exceed 0.5 mm if the angle change between it and the adjacent ply is 45 deg;
layups should ideally be balanced and symmetrical about the neutral axis of the laminate;
90 deg plies are generally recommended at a level of at least 10% largely to control Poisson ratio effect, or higher if required for bolt-bearing applications;
the distribution of the ply orientations should be as evenly spaced throughout the laminate as possible;
the distance between ply dropoffs should not be less than 20×the ply thickness, spanwise and 10×the ply thickness chordwise - also the lowest possible dropoff height, i.e. thickness of plies at dropoff, is always desired.
The layup design for an aircraft composite wing will depend on the design approach and the method of assembly, ie. bonded or bolted. If a soft skin - hard stringer approach is used the multiaxial fabrics may vary in style from 30-60% for the 0 deg plies; 40-60% for the 45 deg plies and approximately 10% for the 90 deg plies. A higher requirement for 90 deg plies will apply for higher load input areas. For stringers, a layup of 67% for 0 deg plies; 23% for 45 deg plies and 10% for 90 deg plies would be suitable. Layups for spar and rib areas could suitably be 20% for 0 deg plies; 60% for 45 deg plies and 20% for 90 deg plies. Other design approaches may have layups intermediate these mentioned here, although layups outside these ranges are quite possible.
A complex laminate such as an aircraft wingskin formed from a single multiaxial fabric style would be likely to suffer an unacceptable weight penalty caused by the inability of that style to provide the optimum ratio of fibre orientations in a
Aftergut Jeff H.
BAE Systems plc
Nixon & Vanderhye P.C.
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