Electrically conductive laminate for temperature control of aircraft surface

A laminate coating comprising a durable outer ply which is resistant to abrasion and impermeable to water, bonded to and through a conductive layer of fibers and an integrally enveloping adhesive layer which is disposed on a vehicle component surface, such that the laminate coating upon curing is bonded to the vehicle component surface such as the surface of an aircraft part especially the fuselage, wing parts or tail section. The conductive layer is connected to a source of electrical energy and control means adapted to control the temperature of the coated surface of the aircraft part.

BACKGROUND OF THE INVENTION 
Aircraft are exposed to a variety of temperature conditions during flight 
and while out of service on the ground. Fuel, when tanked in areas 
directly adjoining exterior surfaces of the aircraft may change the 
temperature of those surfaces without warning. In the event the surface 
temperature or the aircraft falls below 32.degree. F. and there is mist, 
rain or high humidity in the ambient air; ice can form in thin films on 
these surfaces. If undetected, this ice can dislodge during operation of 
the aircraft and fly into other portions of the aircraft especially into 
engine parts, causing damage that may affect the performance of the 
aircraft and if severe, can completely stop engine operation. Aircraft 
which have wing fuel tanks with jet engines mounted behind the wings are 
especially prone to what is termed "ICE FOD" or "ice foreign object engine 
damage". In addition, aircraft taxiing behind aircraft with ice build-up 
may be exposed to showers of ice as the former aircraft begins take-off. 
In addition, sensors and control surfaces may become fouled with ice 
resulting in inaccurate readings and unreliable control during flight. 
The magnitude of the problem was reported by the U.S. Department of 
Transportation Federal Aviation Administration in General Aviation 
Airworthiness Alert Special Issue AC No. 43-16. In this report, 516 known 
ice related accidents occurred from April 1976 to April 1987 with 567 
fatalities. In this report certain aircraft were found to be more prone to 
ice formation than others. The recommendation of the report is as follows: 
RECOMMENDATION 
It is recommended that all owners and operators of airplanes listed in the 
Applicability list exercise extreme caution when planning a flight in 
areas where icing is known or forecast to exist. A special review of each 
individual airplane's records should be conducted to clarify the icing 
flight approach status of any anti-ice/de-ice equipment that is installed 
on the airplane, and flight into any kind of known icing environment made 
only when all of the equipment required for flight in that particular 
environment is "installed and approved." 
Methods to detect ice build-up on aircraft have been confined to visible 
observations aided by decals and movable tufts of material which will 
become fixed by the layer of ice. The removal of ice has been accomplished 
by using expensive trucks and crews which spray de-icing agents on the 
aircraft. This is only a temporary solution to the problems since delays 
in take-off can lead to a reaccumulation of the ice film before departure. 
Some aircraft have been equipped with heat manifolded from engines out to 
the leading edge of the wings that do not heat the wings adequately to 
remove surface ice. Heating pads of various types have been bonded to the 
surface of wings but are prone to detachment under extreme conditions of 
temperature, vibration and wind shear Raid have not been found to be safe 
or effective. It would advance the safety of air travel if a device were 
developed to sense conditions of surface temperatures at which ice would 
develop and rapidly heat these surfaces to prevent and/or melt ice 
accumulations which is integrally bonded onto the surface of the aircraft. 
It is also contemplated that the same technology can be applied to other 
types of vehicles and vehicle components which require ice melting or 
removal for safe and efficient operations. For example, movable components 
on ships such as gun turrets and the hoods of cars and trucks could be 
similarly protected. 
Therefore, it is an object of this invention to provide a means to detect 
temperature conditions on aircraft surfaces during which ice can 
accumulate (i.e. approximately 32.degree. F. or less) and to provide a 
means to heat these surfaces to temperatures sufficient to prevent and/or 
melt said ice rapidly prior to take-off. It is a further object of this 
invention to integrally bond a means of warming aircraft parts to that 
aircraft part. 
Heating elements or systems developed for other applications are not 
adaptable to heat the surfaces of aircraft especially exterior surfaces. 
Invention Disclosure Statement per 37.CFR 1.97 et seq. 
In U.S. Pat. No. 4,250,397, Gray described a paper impregnated with 
graphite fiber and saturated with a binder to adhere the top and bottom 
sheets to the saturated fiber matrix. Two segments of the paper are 
connected in series and encapsulated between cover sheets to form the 
heating element of a drapable heating pad. This invention doesn't teach 
the integral bonding of a conductive ply to a surface or the use of metal 
coated fibers as a conductive ply. 
In U.S. Pat. No. 3,923,697, Ellis described an electrically conductive 
composite comprising graphite, magnesium dioxide and zinc oxide for use on 
a substrate in electrically conductive coatings. This coating is not 
integrated into a laminate on application to a surface. 
In U.S. Pat. No. 3,935,422, Barnes and Sharpe described an electrically 
conductive mixture comprising vinylidene chloride polymer and carbon 
forming a woven glass fabric with a vapor barrier attached thereto. The 
amount of conductive material is varied to provide variation in the watt 
density of heat provided. The device can be glued to plaster board. This 
invention teaches that the density of the material must be varied in order 
to obtain changes in heat energy on a surface area. 
In U.S. Pat. No. 3,900,654, Stinger described an electric heating element 
comprising a layer of electrically conductive elastomer containing carbon 
black dispersed in a fluorocarbon elastomer attached to an insulator. The 
elastomer must be heated under pressure until it bonds to the insulator. 
This invention teaches that a conductive ply can be bonded to an insulator 
with heat and pressure, conditions that cannot be used in the current 
invention. 
In U.S. Pat. No. 3,839,134, Fujihara described a non-metallic web of carbon 
particles and non-metallic fibers coated with plastic for use as a heat 
generating sheet. This plastic coated conductive sheet is not bondable in 
the method described for the current invention. 
In U.S. Pat. No. 3,749,886, Michaelsen described a flexible conductive 
sheet including conductive particles in a matrix and channel-shaped 
electrodes covered by a flexible insulating envelope. This device is not 
applicable for lamination to a surface. 
In U.S. Pat. No. 3,657,516, Fujihara described a panel heating unit 
comprising an electrically resistive paper or porous material sealed in 
paper or cloth with a resin. This panel is a relatively thick 
self-contained heating unit intended to fit into a wall or ceiling system 
and is not adaptable to heating a vehicle surface. 
In U.S. Pat. No. 3,859,504, Kureha and Toyo described a panel heater 
comprising a sheet of carbon fiber paper with electrodes at each end and a 
layer of synthetic resin covered by aluminum foil which is coated by a 
synthetic resin. This panel is also a relatively thin self-contained 
heating unit intended to fit into a panel system.

SUMMARY OF INVENTION 
This invention discloses a unique, integrally bonded laminate which is used 
to thermally control a surface or a portion of a surface of an aircraft to 
which the laminate is bonded. The composite is an integrally bonded 
laminate comprising an outer ply which seals the interior of the laminate 
against penetration and water damage. It is most desirably a two component 
polyurethane top coating with adhesive properties which can bond to the 
electrically conductive ply and through the lattice work in that ply to 
bond comparably with the underlying adhesive. 
The next ply is comprised of substrate fibers which may be woven or 
non-woven, chopped or non-chopped non-woven materials or woven in 
continuous plies. The substrate may be graphite, ceramic fiber, aramid, 
polyester and other such substrates. A metal coating is a good conductor 
of electrical energy and can include copper, silver, nickle, gold and 
other similar metals and alloys. The fibers may be metallized individually 
and formed into the ply or the metallization can occur after the ply has 
been formed. In a significantly preferred embodiment the fibers are nickle 
coated graphite which are non-woven fibers manufactured from chopped 
fibers with a diameter of 8 microns with a range of 4 to 100 microns and 
length of 2.5 centimeters with a range of 0.5 to 5 centimeters; percent 
composition 50% carbon and 50% nickel by weight with a range of 5 to 95% 
of each component respectively. 
The conductive ply has a surface resistivity of less than one ohm per 
square in the fabric form with a range of 500 to 0.1 ohms per unit square. 
This electrically conductive ply can be used to selectively control the 
temperature of a surface to which it is applied by connecting it to a 
source or electrical energy. The amount of current can be varied using a 
control system. The temperature of the surface can be measured by a sensor 
and can be varied as required for the application of the invention. 
Temperature sensors can also be bonded in the laminate. The amount of 
current can be varied in response to the outer surface temperature using a 
logic system such as a microprocessor. The edges of the conductive ply are 
connected to the source of power using all edge connector or bus bar and 
wiring system. The conductive ply is then bonded to the surface of the 
panel to be protected using an adhesive which will maintain its bonding 
capability over a wide range of temperatures. In a preferred embodiment of 
the invention, the adhesive is a one part epoxy coating. Due to the porous 
nature of the conductive layer the outer ply is intermittently bonded to 
the adhesive ply thereby integrally enveloping the electrically conductive 
ply. 
An important application of this invention is the application of the 
laminate to the wings and fuselage of aircraft. The exterior surface of 
aircraft can form layers of ice on the wings and fuselages. This layer of 
ice can affect the operation of controls and sensors as well as shear off 
and damage other parts of the aircraft especially the engines as described 
hereinabove. 
This invention can be used on aircraft parts to permit controlled heating 
of certain surfaces, especially the wings. The rate of heating can be 
controlled to rapidly heat and efficiently prevent and/or melt ice from 
these surfaces. The electrically conductive layer can be die cut to 
facilitate coating the irregular shapes and to surround inspection plates 
and control surfaces. The latter removable and movable surfaces can also 
be treated with the laminate, using precisely cut pieces of the 
electrically conductive ply bonded to these surfaces. A pliable wiring and 
bus arrangement connects the electrically conductive plies of these 
removable/movable components to the main grid. Also, electrical connectors 
call be employed to facilitate rapid removal of these elements. 
In order to add further strength and in some embodiments to insulate the 
laminate from the surface of the part, a resilient/insulating ply of 
material can be added if desired, to enhance the strength of the laminate. 
The resilient/insulating ply call be added on one or on both sides of the 
electrically conductive ply if desired. The material should be porous or a 
lattice of fibers to permit integral bonding of the components of the 
laminate. (Examples of materials that could be employed include aramid and 
similar inert, structurally strong materials. The adhesive is added on 
either side of the resilient/insulating ply to facilitate integral bonding 
of the plies. 
DESCRIPTION OF THE INVENTION 
The laminate 101 is applied to surfaces 100 of the aircraft, especially the 
wings as shown in FIG. 1 but also may be applied to the fuselage and tail 
surfaces as shown in FIG. 3. It can be applied over flat or curved 
surfaces and also can be used on movable and removable structures as shown 
for the inspection plate 100A and flap 100B in FIG. 1. 
The laminate 101 can be applied to the upper and lower surfaces of wings as 
shown in FIG. 2. The laminate is comprised of a plurality of integrally 
bonded piles. The outer ply of the laminate is waterproof, resistant to 
abrasion and penetration and hard when cured. In a preferred embodiment 
the outer ply 102 is a polyurethane top coating with adhesive properties 
that is applied to the electrically conductive ply 103. The preferred 
polyurethane coating is a two component coating comprising a polyurethane 
base mixed in equal volumes with a polyurethane curing solution, such as 
that manufactured by Crown Metro Aerospace Coatings, Inc. of Greenville, 
S.C. 24-72 Series which is compatible to bond with the preferred epoxy 
coating for the adhesive ply 104 as described herebelow. Said outer ply 
can be spray applied and a dry film with thickness of 0.002.+-.0.0005 
inches thick results after 9 hours at room temperature. Use of different 
thinners can affect drying time and electrostatic properties. The cured 
coating 102 in the preferred embodiment is a two component chemically 
cured polyurethane topcoat designed to provide outstanding resistance to 
weathering with maximum gloss and color retention. This quality coating is 
a carefully balanced formulation that will give maximum chemical 
resistance coupled with sufficient flexibility, is water impermeable and 
is capable of approximately 30% elongation, to minimize chipping, flaking 
and erosion. This topcoat is available in all color and gloss ranges 
including clears and metallics. 
The electrically conductive ply 103 is comprised of a lattice work of 
conductive fibers which are metallically coated substrate fibers. The 
metallic coating can be applied individually to the substrate members or 
to a ply comprised of said substrate fibers. The substrate fibers may be 
made or graphite ceramic fiber, polyaramid polyester and the like. These 
substrate fibers may be chopped or not chopped to an average length. The 
chopped or unchopped fibers may be used either in a woven or non-woven ply 
form. The fibers may range in diameter from 4 to 100 microns and in length 
from 0.5 to continuous form. The surface resistivity of the ply can range 
from 500 to 0.1 ohm per unit square. The metallic coating on the substrate 
is a good conductor or electrical energy including copper, silver, 
aluminum, nickle, gold or other similar metals and alloys. 
In a significantly preferred embodiment, the substrate is graphite fibers 
which are chopped, non-woven fibers with an average diameter of 12 microns 
and an average length of one inch. The metallic coating is nickle applied 
such that the average weight of the electrically conductive ply 103 is 50% 
graphic and 50% nickle and the surface resistivity of this ply is less 
than one ohm per square. The electrical conductive ply 103 is a very thin 
sheet of a gauze-like web of fibers, having mechanical properties similar 
to tissue paper and transparent to light. A nominal thickness of such a 
ply of nickel coated fibers, used in the preferred embodiment is on the 
order of 0.003 inches. The conductive ply of the preferred embodiment 
would have a nominal weight of 0.3 ounces per square yard. A typical sheet 
of such material is available from VERATEC, a division of International 
Paper Co., located in Walpole, Mass. Their non-woven mat consisting of 
nickel coated carbon fibers is marketed as Grade Number 8000855. Its 
surface resistivity is nominally 1.3 ohms per unit square. The 
electrically conductive ply 103 is connected to a source of electrical 
energy, using a connecting edge bus and wiring system. The electrical 
energy may be provided as direct current or alternating current. 
The edge bus 106 is bonded directly to the electrically conductive ply 103 
and has an electrical connector 107 other suitable wire interface as 
described hereinbelow which joins to complimentary connector 108 which is 
joined to an electrical conducutor 109 and 110. In one embodiment, the 
opposite edges of the electrically conductive ply 103 are connected to 
opposing poles of direct or alternating electrical source 115 with 
conductor 109 being connected to the positive pole and 110 being connected 
to the negative pole of the power source. The electrically conductive ply 
103 is then bonded to the surface 105 of the aircraft component using an 
adhesive ply 104. The conductive ply can be readily die cut prior to 
lamination in the composite to facilitate covering a variety of shapes and 
sizes of component parts. The adhesive ply 104 should maintain a strong 
bond between the laminate 101 and the surface of the component 105 as 
shown in FIG. 2 over a wide range of temperature and humidity conditions. 
In a preferred embodiment of the invention in which the aircraft part is 
metallic, the adhesive is a one part epoxy coating which is aluminized and 
chemically cured. It has the following characteristics: 
Admixed Viscosity--(#2 Zahn) 19 Sec.+-.2 
Grind Fine--5 min. 
V.O.C. Admixed 641 gr/liter 
Recommended film thickness--0.8 to 1.2 mil 
A typical epoxy, as mentioned above but not limited to, such as 10-P1-3 
Aluminized Epoxy Coating, Chemical Resistant as manufactured by Crown 
Metro Aerospace Coatings, Inc. Greenville, N.C. 
All aluminum surfaces to be treated by the epoxy primer coating must be 
anodized, chromate conversion coated or primed before application thereto. 
In the event that the aircraft part to be coated is fiberglass or other 
non-metallic surfaces, then a compatible adhesive with similar performance 
properties to the above referenced metallic adhesive can be used in the 
practice of the invention. 
In some embodiments of the invention one or more resilient/insulating plies 
125 may be incorporated into the laminate. In most embodiments, the ply 
will be added between the electrically conductive ply and the surface of 
the panel. In this event, the adhesive ply will be duplicated on either 
side of the resilient/insulating ply 125. The resilient/insulating ply is 
desirably a network of non-conductive fibers which may be either woven or 
non-woven; chopped or non-chopped fibers which may be non-woven or woven 
in continuous plies. The thickness of the ply may range from 10 to 100 
microns with an average of 30 microns in a preferred embodiment. The 
fibers may be aramid polyester, ceramic fiber and other similar inert 
materials with good electrical and heat insulating properties. This ply is 
porous to permit integral bonding of the adhesive 104 from the thermal 
conductive ply 103 to the surface 105 of the part being treated with the 
laminate 101. 
In order to provide a means to assess the amount of heat generated on the 
aircraft part 100 treated by the laminate 101, a means of measuring 
surface temperature at one or more areas in the laminate 101 is desirable. 
In a preferred embodiment of the invention, a bondable foil thermocouple 
111 as shown in FIG. 1 may be incorporated in the laminate and a 
thermocouple control wire 112 can be routed with the electrical wires to 
the control system. In the event the laminate 101 is applied to a 
removable 100A or movable 100B aircraft part, an electrical connector 107 
may be used on the thermocouple lead wire compatible with a second 
connector 108 applied to the control wire. The thermocouple sensor is 
ideally thin and flat and can sense temperatures up to 150.degree. C. A 
Chromel-Alumel Foil thermocouple Style I as manufactured by Omega 
Engineering, Inc., Stamford, Conn., 0.0005 inches thick with an 0.010 inch 
diamter control wire has a rapid response to temperature change and is 
readily employed in the laminate. It may be installed, exterior to or 
interior to, the conductive ply 103. Alternatively, other temperature 
sensors such as a 3 wire RTD can be used but these systems are sensitive 
to vibration and shock. Thermocouples can develop age related errors and 
thermocouples can exhibit non-linear temperature responses over wide 
temperature ranges. 
Thermisters are manufactured front a mixture of metal oxides sealed in 
glass or epoxy. These sensors require precise individual calibrations and 
are error prone at high temperatures. However, thermistors tend to become 
more s table with time and are highly accurate at the design temperature 
range. 
Control of temperature may be accomplished by having a varied amount of 
current delivered over either a fixed or varied amount of time or by 
providing a constant amount of current for a series of fixed intervals or 
time. In the event a fixed energy level with time variations chosen, a 
wide range of target temperatures should be selected to eliminate 
oscilating on-off current or "chatter" in the power circuit. In the 
current invention, if a fixed current with variable time application is 
chosen for control, the minimum activation temperature should be several 
degrees above 32.degree. F. and shut off temperature should be below 
temperatures at which components or the aircraft part or any associated 
fuel or lubricants would be damaged or ignited. In general, less than 
200.degree. F. would be a safe stop point. An automatic reset may be added 
based upon known performance of the system to slop heating to avoid 
exceeding the maximum temperature. The use or the system requires prior 
testing of the size grid being heated to establish the automatic reset 
point. 
In a current proportional system, the voltage can be varied in response to 
the temperature measured versus the temperature desired on the grid. 
Direct current voltage can be applied in a range from 8 volts to up to 32 
volts to heat the grid. If alternating current is used, 8 to 220 volts 
could be applied to the grid at 50 to 400 cycles per second. A 
programmable microprocessor based controller could be used to establish 
voltages to be applied depending upon the surface temperature of the grid 
prior to heating. 
Alternatively, a control system could be provided to simply heat the grid 
on a pre-programmed basis without sensor feedback. A manual over-ride 
switch could be used to terminate the current if safe temperatures are 
exceeded. The latter system is significantly less preferred due to the 
potential danger of over heating the part. 
In any event, an alarm logic loop incorporated into the temperature 
measurement circuit is present in a preferred embodiment of the invention 
with alarm limits for low temperature at 32.degree. F. and high at 
approximately 200.degree. F. with a temperature read-out in the aircraft 
cockpit as well as provision for an alarm and temperature signal access to 
the ground crew. A manual over-ride to disable or enable the healing 
circuit is also provided in a significantly preferred embodiment of the 
invention. The power and temperature control system may be contained in 
the aircraft or alternatively be housed in ground support equipment and 
"plugged in" to the aircraft during preparation for departure. The latter 
alternative may be more cost effective but is less preferred because the 
ground crew must be available if a second de-icing is needed prior to 
take-off. In a preferred embodiment, the aircraft is equipped with a 
control and heating system with a ground back-up unit available on an 
emergency basis in the event of failure or the in-board unit. 
A preferred embodiment of the control system and heat source is shown in 
FIG. 4. The laminate 101 is connected to the power source 115 housed in 
the control/power system 110. The edge bus 106 is a layer of copper foil 
in a preferred embodiment of the invention which is approximately 0.5 
inches in width and 0.0005 inches thick. It may be a continuous strip or 
contain perforated or fenstrations if desired. It is desirably fastened to 
opposing edges of the conductive ply 103 when it is embedded into the 
still wet adhesive ply 104. Thus, the bus is integrally bonded as a 
portion of the laminate. The electrical connector 107 joined to the bus 
106 is attached to a complimentary connector 108 which is connected to an 
electrical conductor 109 which runs in a sheltered area or the aircraft 
preferrably covered by a water-tight conduit 113. One edge of the laminate 
101 is joined to the positive pole 116 of a source of electrical energy 
115. The bus at the opposite edge of the laminate 101 is connected to a 
negative or ground pole 117 of the source of electrical energy 115. At 
least one temperature sensor 111 is laminated into the component 101 and 
connected to the control wire 112 which terminates in the control unit 
114, passing through a conduit 113. The control unit 114 is most desirably 
a microprocessor programmed to display the temperature of the aircraft 
surface to displays 117 and 117A. The control unit 114 is desirably 
programmed to activate alarms 119 and 119A in the event the temperature 
falls below 32.degree. F. or exceeds 200.degree. F. or other temperatures 
found to be critical to the safe operation of the aircraft. In a preferred 
embodiment temperature display and alarm 117 and 119 respectfully would be 
in the cockpit of the aircraft connected to the control unit 114 by 
control wires 116 and 118 respectively while the respective element 
designated "A" would be duplicated for the ground crew while the aircraft 
is not in service. A manual over-ride switch 120 can disable the power 
circuit to the laminate 101 to prevent over-heating in the event of 
control error. 
In an alternative variation of bus bar 106 connection to the electrical 
connector 109, the use of a plug type connector is eliminated and the 
conductor 109 is attached directly to the bus 106 as shown in FIG. 5. The 
electrical conductor 109 has its end stripped of insulation revealing a 
bare conductor 121. The bus 106 has a pro-punched hole 122 at the position 
on the bus where the electrical connection is to be made. The conductor 
end 121 is placed in hole 122 of the bus 106 and the junction is soldered. 
The soldered junction 123 is desirably covered with a small amount or 
epoxy or ocher adhesive ply 104 and then the outer ply 102 to water seal 
the area. A strain relief wire clamp 124 may be bonded or fastened to a 
bare area of the aircraft skin or the area for the fastener can be pre-cut 
from the electrically conductive ply 103 prior to installation. 
The laminate 101 is applied and bonded to an aircraft part as described 
hereinbelow. The part is primed for application of the adhesive ply 104 
and allowed to dry. The adhesive 104 is applied and while wet, the 
conductive ply 104 is placed into the uncured adhesive ply 105. In a 
preferred embodiment of the invention, when the aircraft part 100 has an 
aluminum exterior surface, it is cleaned, degreased and primed for the 
preferred epoxy adhesive at room temperature. The electrically conductive 
ply is placed in the uncured adhesive which may be brushed or sprayed on 
the surface of the aircraft part to a thickness of less than 0.001 inches. 
The electrically conductive ply is pre-cut to fit the surface to be coated 
and is quite pliable to fit over curved surfaces. The edge bus 106 is 
pressed into the edges of the electrically conductive ply. The subassembly 
is allowed to cure for about 30 minutes. The overall thickness of the 
laminate is now 0.001 or less. During the bonding process, one or more 
temperature sensors are imbedded into the laminate with sensor wiring. 
Next the outer protective ply 102 is applied and allowed to cure. 
The following is an example of the invention in practice. A 12 inch by 12 
inch square of 0.040 gauge 6061 T3 aircraft aluminum was degreased and 
primed. After the prime coat was dry, a one part aluminized epoxy was 
applied at a thickness of less than 0.001 inch. An 11 inch by 11 inch 
piece of the preferred electrically conductive ply was applied to the 
uncured epoxy followed by two 11 inch long 1/2 inch wide strips of 0.0005 
inches thick copper was applied to opposing edges of the laminate. After 
30 minutes a coat of polyurethane was applied and the sample was allowed 
to cure for four hours. In a test, a block of dry ice was placed on the 
opposite side of the sample and atomized water was sprayed on the 
laminated surface. A layer of clear ice formed rapidly varying in 
thickness from 1/16" to 1/4". The surface temperature measured -40.degree. 
F. Using a 110 VAC current source, a variable voltage transformer was used 
to apply 20 volts to the laminate. Within 3.5 minutes the ice was melted 
and the surface temperature measured 40.degree. F. 
On the leading edge of a wing, using the same technique as mentioned 
hereinabove, the enveloping of the laminate may be accomplished by folding 
over the laminate and adhering it to the understructure.