FOR TEMPERATURE CONTROL OF SURFACES

TECHNICAL FIELD

The invention relates to a device for preventing the formation of ice upon surfaces of a vehicle, and in particular to a bonded laminate connected to a source of electrical energy and including control means adapted to control the temperature of the surface.

BACKGROUND ART

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 of the aircraft falls below 0 degrees C. and there is a 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, air- craft 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 Admin¬ istration 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 air¬ craft 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 and 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 0 degrees C. 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.

In U.S. Patent 4,250,397, Gray described a paper impregnated with graphite fiber and saturated with a bonder 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. Patent 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. Patent 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. Patent 3,900,654, Stinger described an electric heating element comprising a layer of electrically conductive elastomer containing carbon black dispersed in a fluorocarbσn 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. Patent 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. Patent 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. Patent 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. Patent 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.

DISCLOSURE 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 thermal conductive ply and through the lattice work in that ply to bond compatibly 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 sub¬ strate 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, nickel, 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 nickel 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 an electrical conductivity of less than one ohm per square, in the fabric form, with a range of 500 to 0.1 ohms. This thermally conductive ply can be used to selectively control the temperature of a surface to which it is applied by connecting it to a source of 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 an 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 thermally 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 thermal 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 thermally conductive ply bonded to these surfaces. A pliable wiring and bus arrangement connects the thermally conductive plies of these removable/movable components to the main grid. Also, electrical connectors can 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 can be added on one or on both sides of the thermal 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. BRIEF DESCRIPTION OF DRAWINGS Figure 1 is a perspective view of the laminate 101, applied to an aircraft wing 100.

Figure 2 is a perspective partially broken cross section of Figure 1 along line 2-2.

Figure 3 is an illustration of the invention used in various portions of an aircraft.

Figure 4 is a logic diagram illustrating the control and power systems most desirably used to operate the invention.

Figure 5 is a close-up view of the conductive bus at the edge of the laminate to an aircraft part 100 showing a soldered wire junction.

BEST MODE FOR CARRYING OUT THE INVENTION

The laminate 101 is applied to surfaces 100 of the aircraft, especially the wings as shown in Figure l but also may be applied to the fuselage and tail surfaces as shown in Figure 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 Figure 1.

The laminate 101 can be applied to the upper and lower surfaces of wings as shown in Figure 2. The

laminate is comprised of a plurality of integrally bonded plies. 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 thermal 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 0.00508 + 0.00127 cm. 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 thermally 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 fibers or to a ply comprised of said substrate fibers. The substrate fibers may be made of 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 - e ib rs ma ran e in di meter

from 4 to 100 microns and in length from 0.5 to continuous form. The resistance of the ply can range from 500 to 0.1 ohm. The metallic coating on the substrate is a good conductor of electrical energy including copper, silver, aluminum, nickel, 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 about 2.54 cm. The metallic coating is nickel applied such that the average weight of the thermal conductive ply 103 is 50% graphite and 50% nickel and the resistance of this ply is less than one ohm per square. The thermally 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 thermally conductive ply 103 and has an electrical connector 107 or other suitable wire interface as described herein- below which joins to complementary connector 108 which is joined to an electrical conductor 109 and 110. In one embodiment, the opposite edges of the thermally 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 con¬ nected to the negative pole of the power source. The thermally 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 Figure 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 in

V.O.C. Admixed 641 gr/liter Recommended film thickness - .00203 cm to .00305 cm

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, S.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 thermally 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 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 Figure 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 degrees C. A Chromel-Alumel Foil thermocouple Style I, as manufactured by Omega Engineering, Inc., Stamford, Ct., 0.00127 cm. thick, with a 0.0254 cm. diameter 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.

Thermistors are manufactured from 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 stable 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 of time. In the event a fixed energy level with time variations is chosen, a wide range of target temperatures should be selected to eliminate oscillating 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 0 degrees C. and shut off temperature should be below temperatures at which compon- ents of the aircraft part or any associated fuel or lubricants would be damaged or ignited. In general, less than 93.33 degrees C. would be a safe stop point. An automatic reset may be added based upon known performance of the system to stop heating to avoid exceeding the max- imu temperature. The use of 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 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 exceed¬ ed. 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 0 degrees C. and high at approx¬ imately 93.33 degrees C. 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 heating 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 alter¬ natively 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 of the in-board unit.

A preferred embodiment of the control system and heat source is shown in Figure 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 1.27 cm. in width and 0.00127 cm. thick. It may be a continuous strip or contain per¬ forated 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 complementary connector 108 which is connected to an electrical conductor 109 which runs in a sheltered area of 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 tempera¬ ture 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 a most desirably a microprocessor programmed to display the temperature of the aircraft sur¬ face 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 0 degrees C. or exceeds 93.33 degrees C. 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 respectively, 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 connec- tion 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 Figure 5. The electrical conductor 109 has its end stripped of in insulation revealing a bare conductor 121. The bus 106 has a pre-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 of epoxy or other 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 thermally 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 105 is applied and while wet, the conductive ply 103 is placed into the uncured adhesive ply 104. In a preferred embodiment of the invention, when the aircraft part 100 has a aluminum exterior surface, it is cleaned, degreased and primed for the preferred epoxy

adhesive at room temperature. The thermally 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 0.00254 cm. The thermally 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 thermally conductive ply. The subassembly is allowed to cure for about 30 minutes. The overall thickness of the laminate is now 0.00254 cm. or less. During the bonding process, one or more temperature sensors are embedded into the laminate with sensor wiring. The next outer protective ply 102 is applied and allowed to cure.

The following is an example of the invention in practice. A 30.48 cm. by 30.48 cm. 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 0.000254 cm. An 27.94 cm. by 27.94 cm. piece of the preferred thermally conductive ply was applied to the uncured epoxy followed by two 27.94 cm. long, 1.27 cm. wide strips of 0.00127 cm. thick cooper 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 0.16 cm. to 0.64 cm. The surface temperature measured -40 degrees C. 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 4.44 degrees C.

On the leading edge of a wing, using the same tech- nique as mentioned hereinabove, the enveloping of the laminate may be accomplished by folding over the laminate and adhering it to the understructure.