WO1994026590A1

WO1994026590A1 - Overwing anti-ice system

Info

Publication number: WO1994026590A1
Application number: PCT/US1994/005167
Authority: WO
Inventors: George H. Grauer; David C. Wensley; Robert Erik Bylin
Original assignee: Technology Dynamics Group Inc.
Priority date: 1993-05-11
Filing date: 1994-05-10
Publication date: 1994-11-24
Also published as: AU6829594A

Abstract

A low profile heater (11) coupled with a good insulator (36) is mounted directly on the aircraft skin (30), preferably on an aerodynamic surface. The heater (11) has a low aerodynamic profile, resulting in only a 0.004 % increase in drag. The heater (11) can be mounted over the aluminum wing panel (30), particularly over a wing tank which sometimes contains cold soaked fuel. A sealant (52) on the wing panel bonds the insulated heating assembly (70). The heating device itself consists of a lower metal skin (53) bonded to an insulator layer (55) which in turn is bonded to a central metal skin (57), a heater element assembly (59), and an upper metal skin (61). Each element of the assembly may be made from a variety of materials. The preferred metal skin is aluminum, the preferred insulator is a honeycomb. The layers are preferably bonded with a structural adhesive.

Description

OVERWING ANTI-ICE SYSTEM

Technical Field

The present invention is directed to an economical device for preventing formation of ice on aircraft wings and other aerodynamic surfaces.

Background Art

Ice is a constant threat to the safety of aircraft operations and passengers. Numerous aircraft accidents have been attributed to ice, and aircraft designers are constantly in search of improved protection from this threat.

Ice on aerodynamic surfaces can cause roughness, increasing drag and decreasing lift, possibly to the point that the aircraft stalls. Ice formation may occur in many locations, including the top and the underside of wing surfaces. This happened on March 12, 1992 when an F-28 took off from La Guardia airport in New York City with an excess accumulation of ice. The plane could not develop sufficient lift and plunged into the icy East River. Twenty-seven of the 51 persons on board the flight died in this accident. Accumulated ice may shed or fall off during takeoff or flight and strike parts of the aircraft. This foreign object damage (FOD) can damage aircraft. This is a particular problem in certain aircraft configurations. If a jet engine is in a slip stream downstream of accumulated ice, there is a significant chance of damage to the engine at some time.

Ice build up has led to many devastating accidents and even incidents which cause aircraft damage but no deaths. For example, an SAS MD-80 took off in February 1992 and ice shedding from the wing surfaces was ingested by both engines, causing a flame out and complete loss of engine power. The plane crashed during takeoff into some trees, then skidded and ultimately broke in half. Fortunately, no lives were lost in this incident. Icing has also caused loss of life in many instances. This happened in January 1992 when a Boeing 737 took off in a snowstorm from runway 36 of Washington National Airport, struck the 14th Street Bridge and then plunged into the Potomac River. Seventy-four of 769 persons on board plus four others on the bridge lost their lives. There are many causes of ice formation, both in flight and on the ground, the primary application of the present invention. Ice can form on an aircraft on the ground if the right matrix or combination of conditions is present. First, there must be a cold surface on which to freeze water to ice, and second, there must be a source of moisture. Moisture is often present in the form of rain, hail, snow, or freezing rain, fog or even humidity. Wind conditions may affect icing, particularly on the ground - moving air tends to increase the amount of ice formation.

Some aircraft designs result in certain areas of an aircraft potentially acting as a cold thermal reservoir. The aircraft's recent operating history may affect the temperature of such a cold reservoir, and current environmental factors may influence the tendency to form ice on the aircraft. In many aircraft, the fuel in a wing can be an important thermal reservoir. A long flight, e.g. more than two hours at altitude, tends to lower fuel temperatures. Jet fuel temperatures can drop below -35° F and still be useable but jet fuel starts to become sludge at -45° F. This freezing point can be lowered with additives if needed. The cold soaked fuel remaining in tanks may provide a significant cold thermal mass.

After the aircraft is on the ground, this cold soaked fuel remains at a lowered temperature, that is, below ambient temperature . Icing over wings during ground turnarounds has become a common problem, particularly for certain aircraft designs. This ice formation can be a problem even when ambient temperatures are well above freezing. Commercial air transport operators have had problems even during California summers with air time temperatures up to 80° F.

Ice can build up to a thickness of more than one inch over large areas of a wing surface. This region is generally coincident with a large thermal mass, e.g. on wings over or under fuel tanks containing cold-soaked fuel. As discussed above, this is a particular problem with MD-80 aircraft.

Previous attempts to solve this problem have included painting wing surfaces with black paint or using film heaters or fiber heaters encapsulated in paint to raise the wing surface temperature and prevent ice formation. Eddy current systems seek to expel ice by charging and discharging one or more large capacitors, causing vibration and breaking off ice. Some have used ice-phobic coatings (Teflon, etc.) to prevent ice adhesion. Other techniques seek to reduce the cold thermal mass. Some use heat exchangers in fuel tanks. Boeing directs aircraft hydraulic fluid through tubes immersed in the fuel tanks to raise the average temperature in the tanks. Some airlines manage fuel during flight using alternate fuel burn to redistribute fuel in tanks and leave less fuel in contact with the upper wing surface, thereby reducing the cold reservoir. When refueling, inboard refueling through an inlet near a problem area, brings in warm fuel to push out cold fuel. Refueling with ambient (ground) temperature fuel may warm the bulk of the fuel somewhat. None of these techniques are entirely sufficient.

Still other techniques are designed to modify the properties of water. In icing conditions, propylene or ethylene glycols are sprayed onto wing surfaces to depress the freezing point of any surface water. In the presence of light to heavy precipitation, glycol can wash off in five minutes. In cold and busy airports, some aircraft leaving the gate with a fresh application of glycol must take off within this time frame or return to the gate for another glycol application.

Residual glycol may be shed as the aircraft takes off, resulting in distribution of glycol over the area under the flight path. This can result in a serious environmental contamination. Miami, for example, has a very high water table and does not allow use of any glycol, yet the humid environment can cause icing problems, even on warm days.

Disclosure of the Invention The present invention provides a low profile heater coupled with a good insulator mounted directly on the aircraft skin on an area of concern. It is especially helpful when mounted on an aerodynamic surface. The present invention has a low aerodynamic profile, resulting in only a 0.004% increase in drag.

The new device is mounted over the aluminum wing panel, over a wing tank which sometimes contains cold soaked fuel. Fasteners and/or a sealant on the wing panel attaches the insulated heating assembly to the aircraft wing. The heating device itself consists of a lower metal skin bonded to an insulator layer which in turn is bonded to a central metal skin, a heater element assembly, and an upper metal skin. Each element of the assembly may be made from a variety of materials. The preferred metal skin is aluminum, the preferred insulator is a honeycomb. The layers are preferably bonded with a structural adhesive.

Brief Description of the Drawings Figure 1 illustrates an aircraft and areas of an airframe which may require ice protection.

Figure 2 illustrates a jet aircraft and potentially sensitive ice-forming areas.

Figures 3A, 3B and 3C illustrate cross-sections of prior electric heaters.

Figure 4 illustrates the heating assembly of the present invention.

Best Mode for Carrying Out the Invention The new device prevents ice formation on cold aircraft components, and to some extent can melt existing ice on the components. Prior art devices require too much power and add too much aerodynamic drag to be commercially satisfactory.

Figure 1 illustrates an aircraft and areas of an airframe which may require ice protection. Referring to Figure 1 , aircraft 10 includes several sensitive areas: wing leading edges 11; propellers 12; windshield 13; radome 14; essential instruments 15 (which may be positioned in a variety of locations); auxiliary air inlets 16; engine air inlets 17; balance horns 18 on wing or tail tips, antenna 19; and empennage leading edges 20. Ice may form on any of these areas, particularly during flight, but also on the ground and can dangerously impair aircraft performance.

Ice may also accumulate over cold thermal masses, such as fuel stored in wing fuel tanks. Referring to Figure 2, aircraft 25 has wings 26 with fuel tanks 28 (inside wing, not shown) near root 27 of wing 26. If ice accumulates over a fuel tank 28, it may be loosened or shed at any time, including during takeoff, and may be sucked directly into the intake of an engine 24.

Many surface heaters have been designed and used in the past. Referring to Figure 3A, a standard NACA heater construction requires 22 watts per square inch and can be bonded to an aerodynamic surface. Aluminum skin 30 is covered by a

0.125" (0.318 cm) synthetic rubber inner layer 31, then 3 mil (0.076 mm) glass cloth 32, 1 mil (0.025 mm) heating elements 33, another 3 mil glass cloth 34, and topped with a 12 mil (0.305 mm) Neoprene outer layer 35. Figure 3B illustrates a more efficient system, a typical heater construction with a metal overshoe, which requires 18-22 watts per square inch. This construction typically is laminated to a basic aircraft structure. In this heater, 25 mil (0.635 mm) aluminum skin 30 is covered by 12 mil insulation 36, 3 mil ribbon heater 37, 12 mil insulation 38, and 5 mil (0.127 mm) stainless steel layer 39. Another typical surface heater molded as a structural member requires still less energy (16-20 watts per square inch). Referring to Figure 3C, one or more 10 mil (0.254 mm) layers 40 of epoxy-impregnated glass cloth are covered by 2 mil (0.051 mm) expanded metal heating element 41, then 8 mil (0.203 mm) outer layer 42 (which may be two 4 mil (0.102 mm) layers) of epoxy- impregnated glass cloth, covered by 5 mil stainless steel layer 43.

The new device offers significant improvements over old devices. It is mounted over an aluminum wing panel, preferably over a wing tank which may contain cold soaked fuel. One particularly useful location is on each wing over the bulkhead between the center fuel tank and the outboard fuel tank on an MD-80. Heating an area of about forty sq. feet (3.7 m²) on each wing provides satisfactory de-icing. The heating assembly should not be too thick (high) or it will adversely impact aerodynamic performance. A preferred embodiment has overall thickness (height) of 0.128" (3.18 mm). A particularly useful range is 0.010"-.250" (0.025 cm - 0.635 cm).

Mechanical fasteners and/or a sealant on the wing panel attaches the multi- layer heating assembly to the wing panel. The heating assembly consists of a lower metal skin bonded to an insulator which in turn is bonded to a central metal skin, a heater element, and an upper metal skin. Each element of the assembly may be made from a variety of materials. The preferred metal skin is aluminum, the preferred insulator is a honeycomb. The layers are preferably bonded with a structural adhesive.

The heating assembly is finished on all edges with an aerodynamic extrusion to provide an aerodynamic profile and smooth transition from the wing surface to the heating assembly. This minimizes drag so air flowing over the wing surface is essentially undisturbed when passing from the wing panel to the mounted heating assembly. The entire assembly is designed so that wing loading is unchanged when the heating assembly is secured to the wing.

Referring to Figure 4, sealant 52 between wing skin or panel 51 and heating assembly 70 may be any of a number of sealants used commercially in the aircraft industry. One preferred sealant is PR-1422, formerly known as PRC-1422, available from Products Research & Chemical Corporation, 5454 San Fernando Rd., PO Box 1800, Glendale, CA 91209. This material is ductile down to about -65° F and up to 250° F, and widely used on aircraft, e.g. sealing fuel tanks. In prior art heating devices, composites tend to absorb water, which can lead to potential corrosion problems in the skin. Sealant 52 provides corrosion inhibition, useful in case any moisture is condensed or trapped between heating assembly 70 and the surface of wing panel 51.

The structural adhesive used for adhesive layers 54, 56, 58 and 60 should inhibit corrosion and provide good mechanical properties with low water absorption properties. The structural adhesive transfers loads from one layer to the next, and provides shear strength between different materials. This is particularly important when materials with different coefficients of thermal expansion are held at different temperatures. A preferred implementation uses AF-163-2K, 2U and 2M, available from 3M, which are often used for bonding structural layers of composite materials.

Lower metal layer or skin 53, center metal skin 57 and upper metal skin 61 may be made of a number of materials. Standard aircraft aluminum is a well known and proven material that provides durability and good impact strength at a reasonable cost. In a preferred embodiment, the thickness of each of lower skin 53, center skin 57, and upper skin 61 is 12 mils (0.305 mm). Upper skin 61 should be thick enough to resist impact damage, for example if a tool is dropped on the surface, but it is desireable to make the material as thin as possible. The skin is an effective conductor of thermal energy, acting as a dead thermal short for thermal conductivity, so the thickness of each skin should be minimized. The metal skins, particularly upper skin 61, also provide lightning protection by passing electrical current to the wing structure. The entire heating assembly 70 is electrically bonded to the wing, that is, it doesn't change the electrical characteristics of the wing. Many other metals were studied for use as a skin but were more costly. In general, the material should have a low coefficient of thermal expansion. Other useful materials include titanium, stainless and other steels, graphite-epoxy composites, other epoxies, thermoplastic composites and other materials known to one skilled in the art.

Insulating layer 55 should be very light weight with good insulating properties, yet provide some structural rigidity. In a preferred embodiment, insulating layer 55 is a honeycomb material. Other insulators also useful as structural materials include syntactic type foams of which closed cell polyurethane foams are candidates. A vacuum might be used, but no useful systems are currently known. Glass or plastic bubbles might also be incorporated in honeycomb to provide insulation.

One useful honeycomb is Nomex, from DuPont, with a 1/8" (0.318 cm) cell size and "9 lb" (9 lb/ft³) density (weight per cubic foot, 144.17 kg/m³). This cell size provides a good balance of insulating and structural properties. A larger cell size provides better insulation but lower shear strength, while a small cell size provides the opposite properties. Cell size refers to the dimension of each honeycomb cell, with 1/8" the smallest currently available cell size from DuPont. The cells are formed from paper or aramid fiber, then impregnated phenolic resin. The cells in the finished honeycomb are hollow, which provides insulating properties. A density of 6 lb (96.11 kg/m³) is useable but higher densities are preferable.

Blocks of honeycomb can be cut to varying thicknesses. A thickness of 60 mil (0.15 cm) to which a 0.6 W/in² (930 W/m³) heat flux is applied to one surface provides a 50° F (27.8° C) thermal gradient, with the honeycomb yielding a thermal conductivity of 0.104 BTU/hr ft °F (0.06 W/m °K). One skilled in the art may choose to use 20-200 mil (0.051 - 0.51 cm) thick honeycomb. Thicker slabs would be useful but the overall dimension of heating assembly 70 is limited by other constraints, such as aerodynamic drag.

Heater element 59 can be made from a variety of materials well known in the art. In a preferred embodiment, heater element 59 is etched foil (30% copper, 70% nickel) surrounded by epoxy glass dielectric skins. Such an element is available from Tayco Engineering, Inc., 10874 Hope St., Cypress, CA 90630. The epoxy skins provide electric insulation and some incidental thermal insulation. Other suitable metals can be used for the heater, and other dielectric materials can be used in place of the epoxy material. The epoxy glass provides good impact resistance and structural strength. A preferred embodiment of the heater has the capacity to deliver 10 W/in² (15.5 kW/m²) although only 0.6 W/in² (930 W/m²) is required. The heater element is approximately the size of the heater assembly, 40 ft² (3.7 m²).

Heater assembly 70 is secured over a fuel tank (not shown) containing cold soaked fuel 50 to wing panel 51 by sealant 52 and mechanical fasteners. The assembly can be installed using conventional techniques well known in the industry. Electrical power leads and resistive temperature device (RTD) sensors are positioned throughout the assembly and routed through slots cut in the honeycomb, then pulled down through an access panel, preferably an existing access panel already designed into the wing. The wires are routed along the trailing edge side of the rear spar of the wing and into the fuselage where the heater control units are mounted in the aft cargo pit, adjacent to the access door. A heater control unit provides closed loop heat control, turning the heater on and off as needed to maintain a preset temperature. In a preferred embodiment the preset temperature is set to 80° F (26.7° C). Cockpit controls include only an indicator light and an on-off switch to activate the unit. Alternative switches or electrical lockouts detect whether there is weight on the wheels so that the heating assembly cannot be activated when the plane is not on the ground. Other sensors detect throttle position to turn off the heating assembly when the throttles are advanced past a certain minimum position, as in takeoff.

One significant advantage of the current design is that very little power is required to achieve deicing or anti-icing. A preferred embodiment requires watt densities of only 0.6 Watts per square inch (930 W/m²), the lowest watt density that has been used in previous devices for anti-icing or de-icing.

One key element of the new construction is the insulating layer. In general, the better the insulator, the less power is required. Other insulation configurations should lead to still lower watt densities. The improved efficiency of the new device does not prevent using higher power, and higher watt densities let the heating assembly operate under harsher conditions. Under extreme conditions, with low temperatures or large temperature gradients, die preferred embodiment can handle 5 W/sq. in.

When de-icing or anti-icing, it is important to raise the exposed surface temperature above freezing, or 32° F (0° C). The temperature gradient is a function of many factors, including air temperature, fuel temperature and wind velocity. This system can maintain a surface temperature under a wider range of conditions than previously available. For example, the current device can prevent ice formation in 10° F (-12.2° C) ambient temperatures with fuel temperature as low as -24° C (-11° F) for winds from 0 to 30 mph (0 to 13.4 m/s).

A general description of the device and method of using the present invention as well as a preferred embodiment of the present invention has been set forth above. One skilled in the art will recognize and be able to practice many changes in many aspects of the device and method described above, including variations which fall within the teachings of this invention. The spirit and scope of the invention should be limited only as set forth in the claims which follow.

Claims

ClaimsWhat is claimed is:

1. A heating device for aeronautical applications, comprising: a lower layer, an insulating layer above the lower layer, a heating element above the insulating layer, and a top layer.

2. The heating device of claim 1 further comprising: a first adhesive layer between said lower layer and said insulating layer, a second adhesive layer between said insulating layer and said heating element, and a third adhesive layer between said heating element and said top layer.

3. The heating device of claim 1 wherein said lower layer is made from a material selected from the group consisting of titanium, stainless steel, other steels, graphite-epoxy composite, epoxy, and thermoplastic composites.

4. The heating device of claim 1 wherein said lower layer is aluminum.

5. The heating device of claim 1 wherein said lower layer is between about 5 and about 50 mils.

6. The heating device of claim 1 wherein said insulating layer is made of a syntactic-type foam.

7. The heating device of claim 1 wherein said insulating layer is made of a honeycomb material.

8. The heating device of claim 1 wherein said honeycomb has a density greater than about 3 pounds per cubic foot.

9. The heating device of claim 1 wherein said honeycomb has a thickness less than about 90 mils.

10. The heating device of claim 1 wherein said heating element is etched foil in epoxy glass.

11. The heating device of claim 10 wherein said etched foil is further comprised of about 70% nickel and about 30% copper.

12. The heating device of claim 1 wherein said upper layer is made from a material selected from the group consisting of titanium, stainless steel, other steels, graphite-epoxy composite, epoxy, and thermoplastic pre-preg.

13. The heating device of claim 1 wherein said upper layer is aluminum.

14. The heating device of claim 1 wherein said upper layer is between about 5 and about 50 mils.

15. The heating device of claim 1 wherein said heating device has an overall thickness of less than about 0.25 inches.

16. The heating device of claim 1 wherein said heating device draws less than about five Watts per square inch of heating device surface area.

17. A method of deicing an aircraft component using a surface-mounted heating device using less than about five Watts per square inch of heating device surface area.

18. The method of claim 17 further comprising less than about one Watt per square inch when ambient air temperatures are above about freezing.