Spliced laminate for aircraft fuselage

A laminated body panel for aircraft applications comprises first and second metal layers and an adhesive layer. First and second sections in each metal layer are generally coplanar and separated by a splice line. A first splice line in a first metal layer is generally parallel to a second splice line in a second metal layer, and spaced laterally from the second splice line. An adhesive layer between the metal layers preferably contains reinforcing fibers bridging across the splice lines. An aircraft fuselage or wing or empennage made with spliced body panels of the invention has fewer joints and weighs less than one made with narrower body panels containing unspliced metal layers.

FIELD OF THE INVENTION 
The present invention relates to laminated metal-polymer-metal panels 
suitable for aircraft applications. More particularly, the invention 
relates to splicing such laminated panels in order to reduce the number of 
joints between adjacent panels in an aircraft fuselage or wing or 
empennage. 
BACKGROUND OF THE INVENTION 
Metal-polymer laminates for aircraft applications are known in the prior 
art. For example, Schijve et al U.S. Pat. Nos. 4,489,123 and 4,500,589 
disclose laminates reinforced with aromatic polyamide fibers, and 
Vogelesang et al U.S. Pat. No. 5,039,571 discloses a laminate reinforced 
with S2-glass fibers. These laminates are being accepted for some aircraft 
applications because they generally have better mechanical properties and 
lower specific gravity than aluminum alloys. In addition, they are 
available at lower cost than fiber-reinforced polymer matrix composites. 
Metal-polymer laminates for aircraft applications are presently limited to 
a maximum width of approximately 152 centimeters (60 inches). This maximum 
width results from the fact that metal sheet material in the specific 
alloys, tempers and thicknesses needed for aircraft laminates is 
restricted by present manufacturing technology to a maximum width of only 
about 152-165 centimeters (60-65 inches). 
The weight of an aircraft fuselage increases in direct relationship with 
the number of joints between adjacent body panels in the fuselage. The 
number of joints is inversely related to the width of the body panels. 
Accordingly, increasing the width of aircraft body panels by splicing or 
otherwise will directly reduce the fuselage weight. 
Although the desirability of splicing adjacent laminated panels to increase 
their width is readily apparent, until the present invention, there has 
not yet been developed a satisfactory technique for increasing the width 
of such panels without detrimentally affecting their mechanical 
properties. 
A principal objective of the invention is to provide a laminated 
metal-polymer body panel for aircraft applications made with spliced metal 
layers and having satisfactory mechanical properties. Surprisingly, 
spliced laminates of the invention have increased residual strength 
compared with unspliced laminates. 
It is a related objective of the present invention to provide an aircraft 
structure such as a fuselage or wing or empennage, made with laminated 
metal-polymer body panels and having reduced weight. 
Another objective of the invention is to provide an aircraft structure made 
with laminated metal-polymer body panels and having satisfactory 
mechanical properties. 
Additional objectives and advantages of the present invention will become 
apparent to persons skilled in the art from the following detailed 
description. 
SUMMARY OF THE INVENTION 
In accordance with the present invention, there is provided an aircraft 
fuselage comprising a plurality of frame members supporting a plurality of 
generally longitudinally extending stringers. A plurality of 
longitudinally and circumferentially extending body panels are attached to 
the frame, preferably by through fasteners which are rivets in a preferred 
embodiment. Body panels of the invention may also be attached to the wings 
or empennage by utilizing analogous techniques. 
The body panels are metal-polymer laminates made with two or more metal 
layers and an adhesive layer between each of the metal layers. Adjacent 
metal layers in the laminate are generally laterally coextensive. In other 
words, the metal layers are generally coextensive in both the longitudinal 
and circumferential directions when the panels are attached to an aircraft 
frame. The laminates are preferably generally rectangular and contain 2 to 
about 20 metal layers and about 1 to about 19 adhesive layers. The metal 
layers are less than about 1.5 mm thick, preferably less than 1 mm, more 
preferably about 0.1-0.9 mm, and most preferably about 0.2-0.5 mm. In a 
preferred embodiment, the metal layers have a thickness of about 0.3 mm 
(0.012 inch). 
The metal layers may be made from a metal having a tensile strength of more 
than 0.20 GPa. Some suitable metals are aluminum alloys, steel alloys, 
titanium alloys, copper alloys, magnesium alloys and aluminum matrix 
composites. Aluminum-copper alloys of the AA2000 series, aluminum-zinc 
alloys of the AA7000 series and aluminum-magnesium-silicon alloys of the 
AA6000 series are preferred. Some particularly preferred alloys are AA2024 
aluminum-copper; AA7075 aluminum-zinc and AA6013 
aluminum-magnesium-silicon. 
The adhesive layers are made from synthetic polymers such as epoxy resins, 
unsaturated polyesters, vinyl esters, phenolic resins and thermoplastic 
resins. Epoxy resins are particularly preferred. 
The adhesive layers are similar in thickness to the metal layers. The 
adhesive layers are less than about 1.5 mm thick, preferably less than 1 
mm, more preferably about 0.1-0.9 mm and most preferably about 0.2-0.5 mm. 
Adhesive layers of about 0.3 mm (0.012 inch) thickness are utilized in a 
preferred embodiment. 
The adhesive layers are preferably reinforced with continuous fibers of a 
material such as glass, aromatic polyamides ("aramids") and carbon. The 
preferred glass fibers are S-2 glass or R-glass fibers each containing 
about 58-69 wt. % SiO.sub.2, 18-29 wt. % Al.sub.2 O.sub.3 and 7-19 wt. % 
MgO. Also suitable are less expensive E-glass fibers which contain about 
55 wt. % SiO.sub.2, 15 wt. % Al.sub.2 O.sub.3, 19 wt. % CaO, 7 wt. % 
B.sub.2 O.sub.3 and 3 wt. % MgO. One suitable aramid fiber is made from 
poly-paraphenyleneterephthalamide. The fibers may have a modulus of 
elasticity of about 60-650 GPa and an elongation at break of about 0.2-8%. 
The fibers are preferably continuous filaments each having a diameter of 
about 3-30 microns. 
A preferred laminate is reinforced with S-2-glass fibers in the adhesive 
layers. The S-2 glass fibers preferably have a diameter of about 8-12 
microns, and they make up about 35-75% of the total volume of adhesive and 
fiber in the adhesive layers. 
Unspliced metal-polymer laminates containing materials similar to those 
utilized in the present invention are disclosed in Schijve et al U.S. Pat. 
Nos. 4,489,123 and 4,500,589 and Vogelesang et al U.S. Pat. Nos. 4,992,323 
and 5,039,571. The Schijve et al patents cover laminates made with 
aluminum alloy sheets, a thermohardening adhesive layer and aromatic 
polyamide fibers. Vogelesang et al U.S. Pat. No. 4,992,323 covers a 
laminate provided with a synthetic thermoplastic layer which may contain a 
polyarylate, polysulphone, polyethersulphone, polyetherimide, 
polyphenylene ether, polyphenylene sulphide, polyether-ether ketone, 
polyether ketone, polyether ketone-ketone or a liquid crystal polymer. 
Vogelesang et al U.S. Pat. No. 5,039,571 covers metal-polymer laminates 
reinforced with S-2-glass fibers. The disclosures of these Schijve et al 
and Vogelesang et al patents are incorporated herein by reference, to the 
extent consistent with the present invention. 
At least one of the fuselage body panels has a width greater than about 165 
centimeters (65 inches). More preferably, the panel width is at least 
about 300 centimeters (118 inches). In a particularly preferred 
embodiment, the panel width is about 305 centimeters (120 inches). As used 
herein, the term "width" may refer either to the longitudinal dimension of 
the panels or to their circumferential dimension when they are attached to 
an aircraft fuselage. 
Aluminum alloy metal layers in the laminates usually have a maximum width 
of about 60 inches (152.4 centimeters). In order to provide greater width 
in the body panels, the metal layers are manufactured in sections. At 
least one of the metal layers includes a first section, a second section 
generally coplanar with the first section, and a splice line between the 
sections. 
The body panels of the invention preferably include a first metal layer 
comprising first and second sections having a first splice line between 
them and a second metal layer comprising first and second sections having 
a second splice line between them. The first and second splice lines are 
preferably generally parallel and spaced apart by a distance of at least 
several centimeters. More preferably, the parallel splice lines are 
spliced apart by at least 10 centimeters or 20 centimeters. 
There may also be a third metal layer between the first and second metal 
layers and comprising first and second sections separated by a third 
splice line. When the third splice line is generally parallel to the first 
and second splice lines, it should be spaced from both the first and 
second splice lines by at least several centimeters and preferably by at 
least 10 centimeters or 20 centimeters. 
The adhesive layers preferably contain glass fibers bridging the splice 
lines. The glass fibers are generally uninterrupted adjacent the splice 
lines. The fibers may be oriented in one direction or in groups 
corresponding to several different directions, depending upon the loading 
condition of the structure. Preferably, at least about one-half of the 
fibers extend perpendicular to splice lines in adjacent metal layers. In a 
particularly preferred embodiment, about one-half of the fibers are 
oriented in a 0.degree. (longitudinal) direction and about one-half are 
oriented in a 90.degree. (circumferential) direction. Alternatively, about 
one-third of the fibers may be oriented at 0.degree. and about two-thirds 
at 90.degree. or about two-thirds may be oriented at 0.degree. and about 
one-third at 90.degree.. 
In an alternative embodiment, the third splice line in the third metal 
layer may extend generally parallel to the fuselage longitudinal axis and 
transverse or perpendicular to the other splice lines. When the third 
splice line extends longitudinally, the body panel should be joined to a 
stringer in the frame by fasteners on opposite sides of the third splice 
line. For example, a stringer having a base straddling the third splice 
line will be joined to the panel with a first through fastener joining the 
first section in the third layer to the base and with a second through 
fastener joining the second section in the third layer to the base. The 
through fasteners are preferably rivets.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT 
There is shown in FIG. 1 an aircraft 15 having a fuselage 16 made with 
several body panels or skin panels 20 of the present invention. The 
aircraft 15 has a frame which includes circumferentially extending frame 
members 22 and longitudinally extending stringers 24 supported by the 
frame members 22. The body panels 20 are spaced outwardly from a principal 
or longitudinal axis 26 of the fuselage 16. 
The first body panel 20a shown in FIG. 1 adjoins a second body panel 20b 
along a first lateral joint 30. Similarly, a second lateral joint 31 joins 
the first body panel 20a to a third body panel 20c. In addition, a fourth 
body panel 20d joins the first panel 20a along a first longitudinal joint 
32, and a fifth body panel 20e joins the first panel 20a along a second 
longitudinal joint 33. 
The spliced laminates of the invention are made up of alternating layers of 
metal and a fiber-reinforced adhesive. FIGS. 2A, 2B and 2C show 
configurations of some metal layers 40, 41, 42. The top metal layer 40 in 
FIG. 2A is made up of a first section 40a and a second section 40b 
separated by a first splice line 45. The bottom metal layer 41 in FIG. 2C 
is made up of a first section 41a and a second section 41b separated by a 
second splice line 46. A middle metal layer 42 in FIG. 2B includes a first 
section 42a and a second section 42b separated by a third splice line 47. 
Some alternative embodiments of spliced laminates 20 made in accordance 
with the invention are shown in FIGS. 3A, 3B, 4A, 4B, 5A and 5B. The 
laminates shown therein are all made up of alternating metal layers 50 and 
adhesive layers 51. The metal layers 50 are made of sections separated by 
splice lines 55, 56, 57, 58. The metal layers are sheets of a 2024-T3 
aluminum-copper alloy having a thickness of 0.3 mm (0.012 inch). The 
adhesive layers contain an AF-163 epoxy resin reinforced with S-2-glass 
fibers. About half of the fibers extend in a 0.degree. (longitudinal) 
direction and about half extend in a 90.degree. (circumferential) 
direction. FIG. 6 shows another laminate 20 made up of alternating spliced 
metal layers 50 and adhesive layers 51. 
There is shown in FIG. 7 a stringer 24 attached to a spliced body panel 20 
of the invention. The panel 20 includes a first metal layer 61, a second 
metal layer 62 and a third metal layer 63 between the other layers 61, 62. 
The panel 20 also includes fiber reinforced adhesive layers 64, 65. 
The third metal layer 63 includes sections 63a and 63b separated by a 
splice line 70. 
The stringer 24 is attached to the panel 20 by means of rivets or through 
fasteners 71, 72 extending through opposed lateral sides of the base 54a, 
54b positioned on opposed sides of the third splice line 70. 
FIG. 8 shows a specimen of a spliced laminate 80 which was tested for 
fatigue behavior. The laminate 80 includes a spliced front metal layer 81, 
a spliced back metal layer 82 and an unspliced middle metal layer 83. The 
metal layers 81, 82, 83 each comprise 2024-T3 aluminum-copper alloy sheets 
having a thickness of 0.3 mm. The metal layers 81, 82, 83 are sized at 300 
mm in a longitudinal direction L and 100 mm in a transverse direction T. 
Adhesive layers 91, 92 between the metal layers 81, 82, 83 comprise an 
epoxy resin reinforced with S-2-glass fibers. About one-half of the fibers 
extend in the longitudinal (300 mm) direction and one-half of the fibers 
extend in the 100 mm (transverse) direction. The adhesive layers 91, 92 
also have a thickness of about 0.3 mm. 
FIG. 9 shows a test laminate 90 similar to the one in FIG. 8, except that 
there are five metal layers 81, 82, 83, 84, 85 and four adhesive layers 
91, 92, 93, 94. 
The spliced laminates 80, 90 each include a front side or front wall 81 
divided into a center section 81a and lateral sections 81b, 81c. 
Similarly, the back wall or back side 82 is divided into a center section 
82a flanked by lateral sections 82b, 82c. The spliced center sections 81a, 
82a each have a width W of 20 mm. The lateral sections 81b, 81c, 82b, 82c 
are each 40 mm wide. 
Fatigue behavior of these laminates 80, 90 was measured by providing the 
center sections 81a, 82a with transversely extending saw cuts to simulate 
cracks. The specimens were then subjected to fatigue loads of 6-120 MPa, 
and crack length was measured as a function of the number of fatigue 
cycles. Measurements were performed on both spliced and unspliced 
laminates. 
Fatigue test results for the laminate 80 of FIG. 8 are shown in FIG. 10. 
Fatigue behavior of the FIG. 9 laminate 90 is plotted in FIG. 11. 
The fatigue behavior of several different test specimens is plotted in 
FIGS. 10 and 11. Plot A represents an average of the unspliced laminates. 
Specimens B, C, D and E were all spliced. Crack locations on the center 
sections of the spliced laminates were as follows; B-front side, left; 
C-front side, right; D-back side, left; and E-back side, right. 
Fatigue behavior shown in FIGS. 10 and 11 is similar for both spliced and 
unspliced laminates up to about 20,000 fatigue cycles. Surprisingly, 
however, after 20,000 cycles, the spliced laminates all demonstrated 
reduced crack growth compared with the unspliced specimens. 
Residual strength of the spliced laminates 80, 90 of FIGS. 8 and 9 was also 
tested by measuring gross stress to failure in the longitudinal (300 mm) 
direction for both spliced and unspliced laminates. Test results are shown 
in FIG. 12. Surprisingly, the 3/2 laminate 80 had a residual strength of 
395.7 MPa compared with 329.3 MPa for an unspliced laminate. Similarly, 
the 5/4 laminate 90 had a residual strength of 378.4 MPa compared with 
345.6 MPa for an unspliced 5/4 laminate. 
Although the present invention is principally directed at production of 
wide (greater than about 165 cm) panels, it may be desirable to produce 
narrower spliced panels in order to take advantage of their enhanced 
residual strength. Spliced panels of the invention will generally have a 
width of at least about 100 cm (39 in), preferably in the range of about 
250 cm to 400 cm (98 in to 157 in). The maximum width presently 
contemplated is approximately 800 cm (315 in). 
The foregoing detailed description of a particularly preferred embodiment 
of the invention has been made for illustrative purposes only. Persons 
skilled in the art will understand that numerous changes and adaptations 
can be made therein without departing from the spirit and scope of the 
following claims.