阅读说明：本技术 用于飞行器的防冰系统 (Anti-icing system for aircraft ) 是由 维斯瓦纳达·古普塔·萨卡拉 拉文德拉·山卡尔·甘尼格 尼古拉斯·约瑟夫·克赖 于 2019-06-21 设计创作，主要内容包括：一种飞行器可以包括防冰系统。防冰系统可包括热联接到飞行器的至少第一暴露表面的碳纳米管阵列。防冰系统还可包括由飞行器携带并电联接到碳纳米管阵列的太阳能电池阵列。(An aircraft may include an anti-icing system. The anti-icing system can include an array of carbon nanotubes thermally coupled to at least a first exposed surface of the aircraft. The anti-icing system can also include a solar cell array carried by the aircraft and electrically coupled to the carbon nanotube array.)

1. An anti-icing system for an aircraft, comprising:

a carbon nanotube array thermally coupled to at least a first exposed surface of the aircraft; and

a solar cell array carried by the aerial vehicle and electrically coupled to the carbon nanotube array.

2. The anti-icing system according to claim 1, wherein said solar cell array further comprises carbon nanotubes embedded within a conductive polymer.

3. The anti-icing system according to claim 1, wherein said solar array is disposed on a second exposed surface of said aircraft remote from said first exposed surface.

4. The anti-icing system according to claim 1, further comprising a transmissive layer covering said solar cell array.

5. The anti-icing system according to claim 1 further comprising an insulation layer having opposing first and second sides, said second side coupled to said first exposed surface of said aircraft.

6. The anti-icing system according to claim 5 further comprising a channel in said insulation layer, said channel having an inner surface spaced from said second side.

7. The anti-icing system according to claim 6, wherein a heat generating layer is located within said channel and coupled to said inner surface of said channel.

8. The anti-icing system according to claim 7, wherein said array of carbon nanotubes is disposed within said heat generating layer coupled to said inner surface of said channel.

9. The anti-icing system according to claim 5, wherein said array of carbon nanotubes is disposed within a heat generating layer coupled to said insulating layer.

10. The anti-icing system according to claim 9, further comprising an erosion protection layer coupled to the heat generating layer.

Technical Field

The present disclosure relates to solar anti-icing systems for aircraft.

Background

Ice formation on aircraft structures (e.g., engine air intakes, wings, control surfaces, propellers, booster air intake vanes, air intake frames, etc.) can be a problem for modern aircraft. Ice adds weight, increases drag, and compromises the aerodynamic profiles of the airfoils, control surfaces, and air intakes, all of which reduce performance and increase fuel consumption. Furthermore, ice formed on the aircraft structure may shed, increasing the risk to other aircraft parts and engine components. Contemporary aircraft may include de-icing or anti-icing detection systems that utilize heat sources or heat-generating elements to provide heat to the aircraft structure to melt or prevent the formation of ice.

Disclosure of Invention

In one aspect, the present disclosure is directed to an anti-icing system for an aircraft that includes a carbon nanotube array thermally coupled to at least a first exposed surface of the aircraft, and a solar cell array carried by the aircraft and electrically coupled to the carbon nanotube array.

In another aspect, the present disclosure is directed to an aircraft including at least one aircraft component and an anti-icing system for the aircraft. The anti-icing system includes a carbon nanotube array thermally coupled to at least a first exposed surface of the aircraft, and a solar cell array carried by the aircraft and electrically coupled to the carbon nanotube array.

In yet another aspect, the present disclosure is directed to a method of preventing ice formation on an aircraft surface. The method includes supplying electrical energy from a solar cell array to a carbon nanotube array coupled to a first exposed aircraft surface, and transferring heat from the carbon nanotube array to the first exposed aircraft surface.

Drawings

In the drawings:

FIG. 1 is a perspective view of an aircraft having an anti-icing system according to various aspects described herein.

FIG. 2 is a schematic cross-sectional view of a turbine engine of the aircraft of FIG. 1 having an anti-icing system.

FIG. 3 is a schematic cross-sectional view of a portion of the turbine engine of FIG. 2 including an anti-icing system having an anti-icing structure and a power supply structure, in accordance with various aspects described herein.

Fig. 4 is a schematic diagram of the power supply configuration of fig. 3.

FIG. 5 is a schematic cross-sectional view of the anti-icing construction of FIG. 3.

FIG. 6 is a schematic perspective view of another anti-icing structure that may be used in the anti-icing system of FIG. 1.

FIG. 7 is a schematic cross-sectional view of the anti-icing construction of FIG. 6 along line VII-VII.

Detailed Description

Embodiments of the present disclosure are described relating to an anti-icing structure for a turbine engine. Anti-icing structures may include carbon nanotubes, which are cylindrical structural arrangements of carbon atoms that may be formed in a variety of ways, including single-walled, double-walled, or multi-walled. Depending on the specific arrangement of carbon atoms in the nanotube, such carbon nanotubes can have very high tensile strength (up to 60GPa in one example), high thermal conductivity in the direction of the tube (up to 3500W/m · K in one example), and electrical conductivity similar to that of metals or semiconductors. When supplied with an electric current, the carbon nanotubes may dissipate heat to surrounding structures.

For purposes of illustration, the present disclosure will be described with respect to an aircraft having a turbine engine. Further, aspects of the present disclosure may be applicable to aircraft during flight or in non-flight operation. However, it should be understood that the present disclosure is not so limited and may have general applicability in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.

As used herein, the term "forward" or "upstream" refers to moving in a direction toward the engine intake, or a component being relatively closer to the engine intake than another component. The term "aft" or "downstream" as used in conjunction with "forward" or "upstream" refers to a direction toward the rear or outlet of the engine, or relatively closer to the engine outlet than another component.

As used herein, a "set" may include any number of the correspondingly described elements, including only one element. Additionally, as used herein, the terms "radial" or "radially" refer to a dimension extending between a central longitudinal axis of the engine and an outer engine periphery.

All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, forward, rearward, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, rearward, etc.) are only used for identification purposes to aid the reader's understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of the disclosure. Connection references (e.g., attached, coupled, connected, and engaged) are to be construed broadly and may include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. Thus, joinder references do not necessarily infer that two elements are directly connected and in fixed relation to each other. The exemplary drawings are for illustrative purposes only, and the dimensions, locations, order and relative sizes reflected in the drawings may vary.

Fig. 1 shows an aircraft 1 that includes a fuselage 2, a cockpit 4 positioned in the fuselage 2, and a wing assembly 6 extending outwardly from the fuselage 2. The aircraft 1 may also include a plurality of engines, including a turbine engine 10, the turbine engine 10 being, by way of non-limiting example, a turbojet engine, a turbofan engine or a turboprop engine. While a commercial aircraft 1 has been shown, it is contemplated that the aspects of the present disclosure described herein may be used with any type of aircraft. Further, while two turbine engines 10 have been shown on each wing assembly 6, it should be understood that any number of turbine engines 10 may be included, including a single turbine engine 10 on the wing assembly 6, or even a single turbine engine 10 mounted in the fuselage 2.

The aircraft 1 may include an anti-icing system 100 on at least one aircraft component, schematically illustrated as being positioned on the wing assembly 6 and the nacelle 11 of the turbine engine 10. As used herein, "anti-icing" refers to removing ice that has accumulated on aircraft components, or preventing ice accumulation when environmental conditions favor ice formation. Additionally, it should be understood that anti-icing system 100 may be positioned anywhere on aircraft 1, including any desired portion of fuselage 2.

FIG. 2 is a schematic cross-sectional view of an exemplary gas turbine engine 10 that may be used in aircraft 1. The engine 10 is housed within a nacelle 11 and has a generally longitudinally extending axis or centerline 12 that extends from a forward portion 14 to an aft portion 16. Engine 10 includes in downstream serial flow relationship: a fan section 18 including a fan 20; a compressor section 22 including a booster or Low Pressure (LP) compressor 24 and a High Pressure (HP) compressor 26; a combustion section 28 including a combustor 30; a turbine section 32 including a HP turbine 34 and a LP turbine 36; and an exhaust section 38.

The fan section 18 includes a fan housing 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 disposed radially about the centerline 12. The HP compressor 26, combustor 30, and HP turbine 34 form a core 44 of the engine 10, which generates combustion gases. The core 44 is surrounded by a core housing 46, and the core housing 46 may be coupled with the fan housing 40.

A HP shaft or spool 48, disposed coaxially about the centerline 12 of the engine 10, drivingly connects the HP turbine 34 to the HP compressor 26. An LP shaft or spool 50, coaxially disposed about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48, drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20. The spools 48,50 may rotate about the engine centerline and are coupled to a plurality of rotatable elements, which may collectively define a rotor 51.

The LP and HP compressors 24, 26 each include a plurality of compressor stages 52,54 with a set of compressor blades 56,58 rotating relative to a corresponding set of stationary compressor vanes 60,62 to compress or pressurize a flow of fluid through the stages. In a single compressor stage 52,54, a plurality of compressor blades 56,58 may be arranged in a ring and may extend radially outward from the blade platform to the blade tip relative to the centerline 12, while corresponding static compressor vanes 60,62 are positioned upstream of and adjacent to the rotating blades 56, 58. It should be noted that the number of blades, vanes, and compressor stages shown in FIG. 1 is chosen for illustration purposes only, and other numbers are possible.

The vanes 56,58 for the compressor stages may be mounted to a disk 61 (or integrally formed with the disk 61), with the disk 61 being mounted to a respective one of the HP and LP spools 48, 50. The buckets 60,62 for the compressor stages may be mounted to the core casing 46 in a circumferential arrangement.

The HP and LP turbines 34, 36 each include a plurality of turbine stages 64,66, with a set of turbine blades 68,70 rotating relative to a corresponding set of stationary turbine vanes 72,74 (also referred to as nozzles) to extract energy from the flow of fluid through the stages. In a single turbine stage 64,66, a plurality of turbine blades 68,70 may be arranged in a ring and may extend radially outward relative to the centerline 12, while respective static turbine buckets 72,74 are positioned upstream of and adjacent to the rotating blades 68, 70. It should be noted that the number of blades, buckets, and turbine stages shown in FIG. 1 is chosen for illustration purposes only, and other numbers are possible.

The blades 68,70 for the turbine stages may be mounted to a disc 71, the disc 71 being mounted to a respective one of the HP and LP spools 48, 50. The buckets 72,74 for the compressor stages may be mounted to the core housing 46 in a circumferential arrangement.

Complementary to the rotor portions, the stationary portions of the engine 10, such as the static vanes 60,62,72,74 in the compressor and turbine sections 22,32, are also referred to individually or collectively as the stator 63. As such, the stator 63 may refer to a combination of non-rotating elements throughout the engine 10.

In operation, the airflow exiting fan section 18 is divided such that a portion of the airflow is channeled into LP compressor 24, and LP compressor 24 then supplies pressurized air 76 to HP compressor 26, which HP compressor 26 further pressurizes the air. Pressurized air 76 from the HP compressor 26 is mixed with fuel in the combustor 30 and ignited, thereby generating combustion gases. The HP turbine 34 extracts some work (work) from these gases, which drives the HP compressor 26. The combustion gases are discharged into the LP turbine 36, the LP turbine 36 extracts additional work to drive the LP compressor 24, and the exhaust gases are ultimately discharged from the engine 10 via an exhaust section 38. The driving of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24.

A portion of the pressurized airflow 76 may be withdrawn from the compressor section 22 as bleed air 77. Bleed air 77 may be withdrawn from pressurized airflow 76 and provided to engine components requiring cooling. The temperature of the pressurized air stream 76 entering the combustor 30 increases significantly. Thus, the cooling provided by the bleed air 77 is necessary to operate such engine components in an elevated temperature environment.

The remainder of airflow 78 bypasses LP compressor 24 and engine core 44 and exits engine assembly 10 through a row of stationary vanes at fan exhaust side 84 (more specifically, an outlet guide vane assembly 80), outlet guide vane assembly 80 including a plurality of airfoil guide vanes 82. More specifically, a circumferential row of radially extending airfoil vanes 82 is used adjacent to fan section 18 to impart some directional control over airflow 78.

A portion 90 of the compressor 22 is shown in FIG. 3, with an exemplary anti-icing system 100 (FIG. 1) shown in greater detail with the exposed surfaces of the aircraft 1. As used herein, "exposed surface" or "exposed aircraft surface" refers to a surface that is directly or indirectly exposed to ice accretion, or to an external airflow or water stream, such that ice may accumulate on the surface of the aircraft. One exemplary aircraft component having an exposed surface that may be prone to icing is shown as a diverter nose 85. A splitter nose 85 is formed to the leading edge of the air intake of the LP compressor 24 and splits the incoming air into a bypass airflow 78 flowing over the engine core and a pressurized airflow 76 flowing through the engine core. It is contemplated that the diverter nose 85 may be formed from a composite material, including a carbon-filled epoxy. Alternatively, the diverter nose 85 may be metallic.

In a non-limiting example, the anti-icing system 100 may include an anti-icing structure 110 coupled to the splitter nose 85, and a power supply structure 120 coupled to a second aircraft component (e.g., the nacelle 11). The anti-icing structure 110 and the power supply structure 120 may be electrically or thermally coupled by conductors 115. Such conductors 115 may include carbon nanotubes, such as carbon nanotube wires or fibers, or may also include wiring, such as aluminum or copper wiring.

An anti-icing structure 110 is schematically shown coupled to at least a first exposed surface 95, such as on a first aircraft component shown as diverter nose 85. It is also contemplated that anti-icing structure 110 may be at least partially embedded within first exposed surface 95 to optimize airflow streamlines on diverter nose surfaces and anti-icing structure 110.

The power supply structure 120 is schematically shown as being carried by an aircraft. More specifically, the power supply structure 120 may be coupled to the second exposed surface 96 of the aircraft 10, such as on a second aircraft component, shown as the nacelle 11. It should be understood that the relative size and location of the power supply structures 120 may vary. Additionally, it should be understood that the power supply structure 120 may be contiguous with the second exposed surface 96 or at least partially embedded within the second exposed surface 96 in order to optimize airflow streamlines on the power supply structure 120.

The example of fig. 3 shows that the power supply structure 120 may be located remotely from the anti-icing structure 110. It is also contemplated that the power supply structure 120 may be positioned adjacent to the anti-icing structure 110. Additionally, the anti-icing structure 110 and the power supply structure 120 may be positioned on the same exposed surface or on adjacent exposed surfaces of the same aircraft component. Other non-limiting examples of aircraft components or exposed surfaces 95,96 thereof include the engine air intake, the booster 24, the outer fan casing 40, or the wing 6 (FIG. 1).

Referring now to fig. 4, the power supply structure 120 may also include an array 122 of solar cells 124. The solar cell 124 may have photovoltaic properties such that incident electromagnetic radiation (e.g., visible light) may be converted into electrical energy within the solar cell 124. It is further contemplated that the solar cell 124 may be in the form of a conductive polymer 125 coupled with carbon nanotubes 126, as shown. In the example shown, carbon nanotubes 126 are embedded within a conductive polymer solar cell 124 to form a nanowire. In another example, the solar cell 124 may be embedded within the carbon nanotube 126. Such coupling may provide increased power conversion efficiency or increased wavelength range that can be converted to electrical energy compared to conventional solar cells. It is also contemplated that the solar cell may be an array of photovoltaic cells spaced apart from and/or interspersed with an array of carbon nanotubes.

Additionally, transmissive layer 128 may optionally be coupled to solar array 122 and cover solar array 122. As used herein, "transmissive" refers, in non-limiting examples, to a property of a material that is transparent to a wide range of electromagnetic wavelengths, including radio, microwave, infrared, visible, and ultraviolet. The transmissive layer 128 may be formed of a protective material such as glass or polymer. In a non-limiting example, the transmissive layer 128 can include multiple layers of transmissive materials, including combinations of transmissive materials. In this manner, incident electromagnetic radiation (indicated by arrow 129) may be transmitted through the transmissive layer 128 to reach the array 122 of solar cells, and the transmissive layer 128 may provide protection for the power supply structure 120 during operation of the aircraft 1.

Turning to FIG. 5, anti-icing arrangement 110 is shown in greater detail. It is contemplated that anti-icing structure 110 may include an insulation layer 130, a heat generating layer 140, and an erosion protection layer 150.

The insulating layer 130 may have a first side 131 and an opposing second side 132, the second side 132 coupled to the shunt nose 85 (fig. 3). As shown, an insulating layer thickness 135 can be defined between the first and second sides 131, 132. In a non-limiting example, the insulating layer thickness may be between 200 and 1400 μm. Further, in non-limiting examples, the insulating layer can be made of various electrical or thermal insulating materials, such as fiberglass, polymer/plastic, or composite materials.

The heat generating layer 140 may have a first side 141 and an opposing second side 142, wherein the second side 142 of the heat generating layer 140 is coupled to the first side 131 of the insulating layer 130, as shown. As shown, a heating layer thickness 145 can be defined between the first and second sides 141,142, such as between 200 and 1400 μm in a non-limiting example.

It is contemplated that the carbon nanotube array 146 may be disposed within the heat generating layer 140. The carbon nanotubes 148 within the array 146 may have various orientations; although shown as randomly oriented within the array 146, it is contemplated that the carbon nanotubes 148 may be aligned in any desired direction. In this manner, the carbon nanotube array 146 may be thermally coupled to the shunt nose 85. Additionally, the conductor 115 may be electrically coupled to the carbon nanotube array 146, thereby providing an electrical coupling between the array 146 and the power supply structure 120.

Although not shown, it is also contemplated that the solar array 122 may also be located within the heat generating layer. In this case, the solar cells 124 and the carbon nanotubes 148 may be dispersed throughout the heat generating layer. Alternatively, the solar cells 124 may be grouped to form the array 122 in a first portion of the heat generating layer and the carbon nanotubes 148 may be grouped to form the array 146 in a second portion of the heat generating layer. It can thus be appreciated that anti-icing system 100 can have solar cell array 122 and carbon nanotube array 146 coupled to the same exposed surface of the aircraft component.

Erosion protection layer 150 may be coupled to heat generating layer 140 within anti-icing structure 110. In a non-limiting example, erosion protection layer 150 may be metallic and have a protection layer thickness 155, for example, between 200 and 1400 μm. It is further contemplated that protective layer thickness 155 can be the same size as insulating layer thickness 135. In another example, the heating layer thickness 145 can be twice one of the insulating layer thickness 135 or the protective layer thickness 155.

In operation, electromagnetic radiation (arrow 129 in fig. 4) incident on the power supply structure 120 may pass through the transmissive layer 128 and encounter the solar cell array 122. The solar cell array 122 may convert such incident electromagnetic radiation into electrical energy and may provide the converted electrical energy to a carbon nanotube array 146 (fig. 5) within the heat generating layer 140. Electrical energy may be supplied via a conductor 115 (e.g., copper wiring or carbon nanotube fibers). The carbon nanotube array 146 may generate heat due to the supplied electrical energy, and the generated heat may be conducted through the erosion protection layer 150 and melt any accumulated ice on the erosion protection layer 150 or the exposed surface 95. It will also be appreciated that the conducted heat may also prevent ice formation on the erosion protection layer 150 or the exposed surface 95.

FIG. 6 illustrates another anti-icing structure 210 that may be used in anti-icing system 100 of FIG. 1. Anti-icing structure 210 is similar to anti-icing structure 110; accordingly, similar components will be described with the same reference numerals incremented by 100, with the understanding that the description of the same components of anti-icing structure 110 applies to anti-icing structure 210 unless otherwise specified.

The anti-icing structure 210 may include an insulating layer 230 and a heat generating layer 240. One difference is that: at least one channel 236 may be formed in the insulating layer 230, and a heat generating layer 240 may be positioned within the channel 236. For example, at least a portion of the channel 236 may have a serpentine profile as shown. In other non-limiting examples, the channel 236 may have a rectangular, circular, linear, or irregular geometric profile.

Turning to fig. 7, the channel 236 can also include an inner surface 238 spaced apart from the second side 232 of the insulating layer 230. The carbon nanotube array 246 may be disposed within the heat generating layer 240, and the layer 240 may be coupled to the inner surface 238 of the channel 236, as shown. Additionally, as shown, an erosion protection layer 250 may cover the insulation layer 230 and the heat generating layer 240 within the channel 236.

In one example, the heat generating layer 240 may include a first side 241 that is flush or coplanar with the first side 231 of the insulating layer 230, as shown. It is also contemplated that the first side 241 of the heat generating layer 240 may extend beyond the first side 231 of the insulating layer 230. In this case, the erosion protection layer 250 may have a sufficient thickness to account for any extension of the heat generating layer 240 beyond the insulating layer 230. Thus, erosion protection layer 250 may have a smooth outer surface 251.

A method of preventing ice formation on a surface of an aircraft 1 includes supplying electrical energy from a solar cell array (such as array 124) to an array 146 of carbon nanotubes 148 coupled to an exposed aircraft surface (e.g., exposed surface 95). Heat may be transferred from the carbon nanotube array 146 to the exposed surface 95. Optionally, the method may include providing power via carbon nanotube conductors 115 electrically coupled to the solar cell array 122 and the carbon nanotube array 146. Optionally, the method may include generating heat within a heat generating layer (e.g., heat generating layer 140,240) having an array of carbon nanotubes. Alternatively, supplying electrical energy may include supplying electrical energy to a heat generating layer 140,240, the heat generating layer 140,240 being coupled to the second exposed aircraft surface 96 remote from the first exposed aircraft surface 95.

Aspects of the present disclosure provide a number of benefits. Conventional anti-icing solutions include the use of heated bleed air from the engine to remove or prevent ice build-up on aircraft components such as superchargers and engine air intake structures, which limits engine performance. Furthermore, such heated bleed air has been led to aircraft components for anti-icing with complex air duct structures, which also increases the weight of the engine. The use of carbon nanotubes in the heat generating layer can reduce engine weight and complexity by eliminating the need for dedicated piping. In one example, weight savings in excess of 50 pounds are achieved by using the anti-icing system of the present disclosure. The heat generating carbon nanotubes may also improve engine performance because more air may be retained within the engine for combustion rather than being pulled as bleed air. Furthermore, the use of carbon nanotube-coupled solar cells to power the heat generating layer may also improve efficiency and reduce operating costs, as the power system on the aircraft may be used for other operations than anti-icing or de-icing. It is understood that the coupling of the carbon nanotubes to the solar cell may increase the power conversion efficiency of the solar cell, thereby providing more power to the heat generating layer.

Various features, aspects, and advantages of the present disclosure may also be embodied in any permutation of aspects of the present disclosure, including, but not limited to, the following technical solutions defined in the enumerated aspects:

1. an anti-icing system for an aircraft, comprising:

a carbon nanotube array thermally coupled to at least a first exposed surface of the aircraft; and

a solar cell array carried by the aircraft and electrically coupled to the carbon nanotube array.

2. The anti-icing system of aspect 1, wherein the solar cell array further comprises carbon nanotubes embedded in a conductive polymer.

3. The anti-icing system according to any of aspects 1-2, wherein the solar cell array is disposed on a second exposed surface of the aircraft remote from the first exposed surface.

4. The anti-icing system according to any of aspects 1-3, wherein the first exposed surface is located on an engine air intake, a supercharger, a splitter nose, an outer fan casing, a nacelle, or a wing.

5. The anti-icing system according to any of aspects 1-4, further comprising a transmissive layer covering the solar cell array.

6. The anti-icing system according to any of aspects 1-5, further comprising an insulation layer having opposing first and second sides, the second side coupled to the first exposed surface of the aircraft.

7. The anti-icing system of aspect 6 further comprising a channel in the insulation layer, the channel having an inner surface spaced apart from the second side.

8. The anti-icing system according to aspect 7, wherein the heat generating layer is located within the channel and coupled to an inner surface of the channel.

9. The anti-icing system of aspect 8, wherein the array of carbon nanotubes is disposed within a heat generating layer coupled to an inner surface of the channel.

10. The anti-icing system of aspect 7 wherein at least a portion of the channel has a serpentine profile.

11. The anti-icing system of aspect 6, wherein the array of carbon nanotubes is disposed within a heat generating layer coupled to the insulating layer.

12. The anti-icing system according to aspect 11 further comprising an erosion protection layer coupled to the heat generating layer.

13. The anti-icing system according to aspect 12, wherein the insulating layer comprises an insulating layer thickness, the heat generating layer comprises a heating layer thickness, and the erosion protection layer comprises a protective layer thickness.

14. The anti-icing system according to aspect 13, wherein at least one of the thickness of the insulation layer, the thickness of the heating layer and the thickness of the protective layer is between 200 and 1400 μm.

15. The anti-icing system of aspect 13, wherein the heating layer thickness is at least twice as thick as one of the insulating layer thickness or the protective layer thickness.

16. An aircraft, comprising:

at least one aircraft component; and

an anti-icing system for an aircraft, comprising:

a carbon nanotube array thermally coupled to at least a first exposed surface of the aircraft; and

a solar cell array carried by the aircraft and electrically coupled to the carbon nanotube array.

17. The aircraft of aspect 16, wherein the at least one aircraft component comprises an engine air intake, a supercharger, a separator nose, an outer fan casing, a nacelle, or a wing.

18. The aircraft of any of aspects 16-17, wherein the array of carbon nanotubes is thermally coupled to a first aircraft component and the array of solar cells is disposed on a second aircraft component.

19. The aircraft of any of aspects 16-18, wherein at least one aircraft component comprises a carbon-filled epoxy composite.

20. The aircraft of aspects 16-19, wherein the solar cell array further comprises carbon nanotubes embedded within the conductive polymer.

21. A method of preventing ice formation on an aircraft surface, the method comprising:

supplying electrical energy from a solar cell array to a carbon nanotube array coupled to a first exposed aircraft surface; and

heat is transferred from the carbon nanotube array to the first exposed aircraft surface.

22. The method of aspect 21, wherein supplying electrical energy further comprises supplying electrical energy from a solar cell array coupled to the carbon nanotubes.

23. The method of any of aspects 21-22, wherein supplying electrical energy further comprises supplying electrical energy via carbon nanotubes electrically coupled to one of a solar cell array or a carbon nanotube array coupled to a first exposed aircraft surface.

24. The method of any of aspects 21-23, further comprising generating heat within the heat-generating layer having the array of carbon nanotubes.

25. The method of aspect 24, wherein supplying electrical energy further comprises supplying electrical energy to the heat generating layer from a solar cell array coupled to a second exposed aircraft surface remote from the first exposed aircraft surface.

To the extent not already described, different features, structures or aspects of the various versions, embodiments or examples may be used in combination or substituted for one another as desired. Not all embodiments showing a feature, structure or aspect are not meant to be construed as so illustrative, but rather as so made for the sake of brevity of description. Thus, various features, structures, or aspects of different embodiments may be mixed and matched as desired to form new versions, implementations, or embodiments, whether or not such new versions, implementations, or embodiments are explicitly described. The present disclosure encompasses all combinations or permutations of features described herein.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.