Heat blanket buffer assembly

An apparatus and method is disclosed using a graphite fiber blanket (20) of high thermal conductivity to repair an aircraft structure (10) made of composite fiber materials. The graphite fiber blanket (20) is placed in thermal contact with a silicone rubber heat blanket (22), with an internal heat source, and the aircraft structure, such as composite laminates (12, 14). Use of the graphite fiber blanket (20) increases the rate at which the temperature of the laminates can be elevated to the proper range for repair and transfers heat from hot spots to cold spots to ensure a uniform temperature throughout the laminates.

TECHNICAL FIELD OF THE INVENTION 
This invention relates to the repair of aircraft structures made of 
composite laminates. 
BACKGROUND OF THE INVENTION 
Aircraft structure formed of carbon fiber composite laminates are becoming 
more common, particularly in high performance military aircraft. These 
materials replace the traditional materials, such as aluminum, from which 
aircraft structures have traditionally been made. 
One of the advantages of the use of carbon fiber materials is the ease of 
repair should the aircraft structure become damaged, as by an accident 
such as a tool or the like damaging the surface or even combat related 
damage. The repair can be performed by providing sufficient bonding resins 
and carbon fiber material to the place of repair and bonding the carbon 
fibers with heat to cure the bonding materials. 
In performing such repairs, it is desirable to maintain the temperature of 
the aircraft structure being repaired at a precise, and uniform 
temperature consistent with the curing requirements of the bonding 
material. In the past, it has been common to use a silicone rubber heat 
blanket with heater wires distributed therethrough to heat the aircraft 
structure to a desired temperature. However, because of integral heat 
sinks and nonuniform structure in the aircraft assemblies, it has been 
difficult to achieve a uniform temperature distribution. A need exists for 
an apparatus and method to provide a more uniform heat distribution to 
provide a uniform temperature for the repair. 
SUMMARY OF THE INVENTION 
In accordance with one aspect of the present invention, an apparatus is 
provided for repairing an aircraft structure which is repairable with the 
use of heat. The apparatus includes a heat source and a graphite fiber 
element in thermal contact with the heat source and the aircraft 
structure. The graphite fiber element ensures a uniform temperature at the 
aircraft structure for repair. In accordance with another aspect of the 
present invention, a silicone rubber heat blanket is provided in thermal 
contact with the graphite fiber element. 
In accordance with another aspect of the present invention, a method is 
provided for repairing an aircraft structure which is repairable with the 
use of heat, including the step of placing a graphite fiber element in 
thermal contact with the aircraft structure to be repaired and placing a 
heat source in thermal contact with the graphite fiber element. The method 
further includes a step of heating the aircraft structure with heat from 
the heat source distributed uniformly by the graphite fiber element to 
provide a uniform temperature in the aircraft structure for repair. A 
further aspect of the method is the placement of a silicone rubber heat 
blanket in thermal contact with the graphite fiber element.

DETAILED DESCRIPTION 
With reference now to the accompanying figures and the following detailed 
description, an apparatus and method for uniformly heating an aircraft 
structure 10 to provide a uniform temperature in this structure for 
repairing the structure is disclosed. 
In FIGS. 1 and 2, the aircraft structure 10 includes a left hand composite 
laminate 12 and a right hand composite laminate 14 separated by an 
aluminum heat sink 16. The laminates 12 and 14 and heat sink 16 rest on a 
composite base panel 18. However, it will be understood that the present 
invention need not use a heat sink 16 and the laminate or laminates being 
repaired can be of any configuration. 
Lying atop the laminates 12 and 14 and heat sink 16 is a graphite fiber 
blanket 20 having highly conductive pitch graphite fibers. Lying atop the 
blanket 20 is a silicone rubber heat blanket 22. Blanket 22 is of the type 
commonly known in the industry that contains heater wires therein through 
which current can pass to provide heat to perform a repair on the 
laminates 12 and 14. 
In the absence of the graphite fiber blanket 20, as the laminates 12 and 14 
are heated by the heat blanket 22, the temperature distribution in the 
laminates and aluminum heat sink will vary considerably with time, as seen 
in FIG. 3. However, by use of the graphite fiber blanket 20, the 
temperature distribution between the aluminum bar heat sink 16 and the 
laminates 12 and 14 and within the laminates is much more uniform, as seen 
in FIG. 4, providing a much more uniform temperature throughout the 
laminates. In addition, the temperature rise of the entire aircraft 
structure is much quicker, providing for a faster and more uniform bonding 
of the repair material to the aircraft structure. The graphite fiber 
blanket transfers heat from hot spots to cold spots and assists in 
maintaining uniform temperature with resulting improved repair patch 
integrity. 
The graphite fiber blanket 20 utilized is formed of the type of material 
discussed in U.S. Pat. No. 5,316,080, which patent is incorporated herein 
by reference in its entirety, which has a very high thermal conductivity, 
exceeding even that of copper. The thermal conductivity of the fibers can 
be, for example, about three times that of copper. 
One graphite fiber blanket 20 found usable in the process of the present 
invention is manufactured by AMOCO Performance Products, Inc. as 
Thermalgraph.TM.fabric under the trademark Thornel.RTM.. Both the 
Thermalgraph.TM.fabric EWC-300X and the Thermalgraph fabric EWC-500X have 
been found suitable. The test illustrated in FIG. 4 was with EWC-300X 
fabric. These materials are pitch fiber based high thermal conductivity 
woven fabrics. Due to the orthotropic nature of the weave and the high 
longitudinal thermal conductivity of the fibers, biaxial thermal 
conductivity is achieved. The EWC-300X is a plain weave fabric constructed 
from four-thousand filament continuous pitch tows. EWC-500X is available 
as an eight harness satin weave fabric constructed from two-thousand 
filament continuous pitch tows. The EWC-300X material has a count (warp 
and fill) of 11.times.10 tows per inch and a weight of 599 g per square 
meter. The fabric electrical resistivity (warp and fill) is 0.05 
.OMEGA./sq. The density is 2.1 g per cubic centimeter and the yarn 
electrical resistivity is 4.0-5.0 micro ohm meters. The estimated thermal 
conductivity is 200-300 W/m.degree.K. The EWC-500X material has a count 
(warp and fill) of 20.times.20 tows per inch and a weight of 485 g per 
square meter. The thickness is 0.61 mm and the resistivity is 0.03 
.OMEGA./sq. The density is 2.15 g per cubic meter and the yarn electrical 
resistivity is 2.3 to 2.8 micro ohm meters. The estimated thermal 
conductivity is 400-500 W/m.degree.K. 
In the preferred embodiment, a three-ply graphite fiber blanket 20 is 
formed composed of three plies of the EWC-300X material stitched together. 
In the assembly illustrated in FIGS. 1 and 2, the silicone rubber heat 
blanket 22 is a rectangle of 12 inches by 18 inches. The aluminum heat 
sink is a bar 5/16 inch thick by 21/2 inches wide by 18 inches long. 
Although a single embodiment of the invention has been illustrated and 
described with numerous specific details in the forgoing description and 
accompanying drawings, it will be understood that the invention is not 
limited to the embodiment disclosed, but is capable of numerous 
rearrangements, modifications and substitutions of parts and elements 
without department from the spirit and scope of the invention.