Leading edge thermal ice protection system and method

Abstract

A leading edge thermal anti-icing system and method are disclosed, the anti-icing system comprising an airfoil structure having a leading edge portion facing a wind direction. The airfoil structure comprises an outer skin and an inner skin forming a plurality of channels, each channel having an air inlet and an air outlet. An air delivery duct within the airfoil structure extends transversely to the wind direction and has a plurality of openings pointing toward the air inlet of the channel. A heat exchanger is configured to provide hot air to the air delivery duct.

Description

Technical Field

[0001] The present disclosure relates to systems and methods for preventing ice accumulation. More specifically, the disclosed examples relate to preventing ice from forming on the leading edge of an aircraft wing. Background Art

[0002] Aircraft that encounter cold and humid air are susceptible to ice formation on various surfaces. Under these conditions, ice accumulation on wings, engine nacelles, turbine components or other surfaces can impair the flight characteristics of the aircraft. Therefore, those components most susceptible to ice formation can be equipped with anti-icing devices and systems suitable for the specific aircraft. Various different mechanical, chemical and thermal systems have been developed for preventing or removing ice accumulation on aircraft surfaces.

[0003] Some thermal ice protection systems use heated air that is redirected from an aircraft's engines to areas where ice formation may occur. When used to prevent ice formation on the leading edge of an aircraft's wings, such systems currently take up a large portion of the available space within the wing, limiting the amount of space available for other systems in the wing, such as actuators for flight control surfaces.

[0004] It is with respect to these and other considerations that the present invention has been proposed. Summary of the invention

[0005] The present disclosure provides systems, devices and methods related to leading edge anti-icing systems. In some examples, the anti-icing system may include an airfoil structure including a leading edge portion facing the wind direction. The airfoil structure may also include an outer skin and an inner skin forming a plurality of channels, each channel having an air inlet and an air outlet. The air delivery duct within the airfoil structure may extend transversely to the wind direction and may have a plurality of openings pointing to the air inlet of the channel. The heat exchanger may be configured to provide hot air to the air delivery duct.

[0006] In some examples, an aircraft may include a fuselage having a first wing extending from the fuselage and a second wing extending from an opposite side of the fuselage. Each wing may include a leading edge portion facing the wind direction. Each leading edge portion may include an outer skin and an inner skin forming a plurality of inner channels, each channel having an internal air inlet and an external air outlet. Each wing may include an air delivery duct extending transverse to the wind direction. The air delivery duct may have a plurality of openings pointing to the air inlet of the channel. A heat exchanger may be configured to provide hot air to the channel.

[0007] In some examples, a method of preventing ice from forming on a leading edge portion of an airfoil may include directing hot air from a heat exchanger to an air delivery duct extending along an interior of a leading edge portion of the airfoil. The method may also include dispersing air from the air delivery duct toward channels formed between an inner skin and an outer skin of a leading edge portion of the airfoil structure. Each channel may extend between an inlet and an outlet between an upper section and a lower section of the leading edge portion. The method may also include exhausting air from the channel through the outlet.

[0008] The features, functions, and advantages can be achieved independently in various examples of the present disclosure or may be combined in yet other examples further details of which can be seen with reference to the following description and drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] Figure 1 is an isometric view of an exemplary aircraft.

[0010] Figure 2 yes Figure 1 Schematic diagram of the air delivery system of an aircraft

[0011] Figure 3 A partially transparent isometric view of the leading edge ice protection system.

[0012] Figure 4 yes Figure 3 A cross-sectional view of the anti-icing system.

[0013] Figure 5 is a schematic cross-sectional illustration of an exemplary leading edge ice protection system according to aspects of the present disclosure.

[0014] Figure 6 yes Figure 5 Isometric view of an anti-icing system.

[0015] Figure 7 yes Figure 5 Isometric view of the inner skin of the anti-icing system.

[0016] Figure 8 is a cross-sectional view of the wing leading edge, including Figure 5 Anti-ice system and extended Krueger flaps

[0017] Fig. 9 is a flow chart depicting the steps of an exemplary method for preventing ice formation on the leading edge of an airfoil according to the present teachings.

[0018] Fig.10 is a flow chart depicting the steps of an exemplary aircraft manufacturing and service method.

[0019] Fig.11 is a schematic diagram of an exemplary aircraft. DETAILED DESCRIPTION

[0020] Various aspects and examples of thermal ice protection systems and related methods are described below and shown in the relevant drawings. Unless otherwise specified, the ice protection system and / or its various components according to the present teachings may, but need not necessarily, include at least one of the structures, components, functions and / or variations described, shown and / or incorporated herein. In addition, unless expressly excluded, the process steps, structures, components, functions and / or variations described, shown and / or incorporated herein in conjunction with the present teachings may be included in other similar devices and methods, including being interchangeable between the disclosed examples. The following description of various examples is merely exemplary in nature and is in no way intended to limit the present disclosure, its application or use. In addition, the advantages provided by the examples described below are exemplary in nature, and not all examples provide the same advantages or the same degree of advantages.

[0021] The detailed description then includes the following sections: (1) Overview; (2) Examples, Compositions, and Alternatives; (3) Exemplary Combinations and Additional Examples; (4) Advantages, Features, and Benefits; and (5) Conclusion. Examples, Compositions, and Alternatives are further divided into subsections A through C, each of which is labeled accordingly.

[0022] Overview

[0023] Typically, an anti-icing system according to the present teachings includes an air delivery duct extending along a portion of an airfoil such as an aircraft wing. The air delivery duct may also be referred to as a piccolo tube, which may be supplied with hot air from a source such as a heat exchanger of a bleed air system and / or a compressor driven by an engine shaft. The duct is installed in a leading edge portion of the airfoil, such as in the nose of an aircraft wing. The duct may be installed close to the upper portion or upper surface of the leading edge portion, and / or above the camber of the airfoil. The duct may extend above the airfoil transversely to the direction of airflow and may be closed by a removable cover.

[0024] The leading edge portion of the airfoil includes an inner skin and an outer skin that together form a plurality of channels. Each channel may extend from an air inlet adjacent an upper surface of the leading edge portion to an air outlet adjacent a lower surface of the leading edge portion. Each air inlet may be located above the camber line and open to the interior of the airfoil, and each air outlet may be located below the camber line and open to the exterior of the airfoil.

[0025] The air delivery duct includes a plurality of openings. Each opening is positioned to direct air from the duct to an air inlet of a corresponding channel. The anti-icing system may thus be configured to direct heated air along the air delivery duct, through the plurality of openings, into the inlets of the plurality of channels, and out the outlets of the plurality of channels. As the air passes through the channels, an outer skin of a leading edge portion of the airfoil may be heated by the air and thereby prevent ice from forming on the leading edge portion of the airfoil.

[0026] The profile of the inner skin is configured to form a desired channel geometry, which can adjust the characteristics of the air flow through the channel. For example, the inner skin can be shaped to provide a desired pressure and flow rate through the channel. The channel can be uniform along the airfoil, and / or can vary in shape or size. Each channel can have a varying cross-sectional dimension, such as width and / or depth, between a corresponding air inlet and air outlet.

[0027] The inner skin may be configured to serve as a structural member of an airfoil. For example, the inner skin may include a material having sufficient strength and / or rigidity so that the inner skin may replace the nose of an aircraft wing.

[0028] An anti-icing system according to the present teachings may also be described as an anti-icing system or a de-icing system. The system may be configured to completely evaporate water on a selected surface, may be configured to limit ice formation to a selected size or degree, and / or may be configured to remove accumulated ice. An anti-icing system may operate continuously, such as to prevent ice formation, and / or may operate selectively, such as after ice formation reaches a predetermined level.

[0029] Examples, components and alternatives

[0030] The following sections describe selected aspects of exemplary ice protection systems and related systems and / or methods. The examples in these sections are intended to be illustrative and should not be construed as limiting the overall scope of the present disclosure. Each section may include one or more different examples, and / or contextual or related information, functions, and / or structures.

[0031] A. Exemplary Ice Protection Systems

[0032] like Figures 5 to 8 As shown, this section describes an exemplary ice protection system 300. System 300 is an example of an ice protection system as described above.

[0033] Figure 1 An exemplary aircraft 100 that may include an anti-icing system is depicted. The aircraft is an example of an aircraft 600 produced by the exemplary method 500, as described below. The aircraft 100 includes a fuselage 110, a first wing 112, and a second wing 114. The second wing 114 extends from the fuselage 110 on a side opposite the first wing 112, and each wing includes a leading edge 116. In the depicted example, the aircraft 100 includes two engines 118, each engine mounted under a wing. In some examples, the aircraft may include more or fewer engines, and the engines may be mounted elsewhere on the aircraft, such as on the fuselage 110.

[0034] The leading edge 116 and the trailing edge of each wing 112, 114 include flight control surfaces 120. The control surfaces may include, but are not limited to, flaps, slats, ailerons, and / or spoilers. The actuators for the flight control surfaces 120 may be housed inside the wings 112, 114. An anti-icing system that uses hot air to prevent ice from forming on the wings may also be housed inside the wings. Such a system may be configured to prevent ice from forming by heating selected surfaces of the wings 112, 114. For example, the anti-icing system may heat one or more fixed surfaces of the wing and / or may heat one or more flight control surfaces of the wing.

[0035] Figure 2 An exemplary air delivery device 130 for a leading edge anti-icing system is described. In the depicted example, the air delivery device 130 is installed in the first wing 112 of the aircraft 100 and is connected to the heat exchanger 122 of the engine 118. The heat exchanger can be configured to provide heated air and can be connected to the engine 118 in any effective manner. For example, the heat exchanger can provide bleed air at a desired temperature and / or pressure from a bleed air system of the engine 118. Typically, the heated air can have a temperature in a range between 200 degrees Fahrenheit and 250 degrees Fahrenheit. In some examples, the hot air provided can have a temperature in a range between 100 degrees Fahrenheit and 300 degrees Fahrenheit.

[0036] The inboard air delivery duct 132 and the outboard air delivery duct 134 are connected to the heat exchanger 122. Each duct can direct hot air from the heat exchanger 122 along the leading edge 116 of the wing 112. The ducts 132, 134 can be described as extending parallel to the leading edge 116, and / or extending transverse to the wind direction 124. The wind direction can also be described as the direction of airflow over the wing 112. The inboard air delivery duct 132 directs air from the engine 118 toward the fuselage 110. The outboard air delivery duct 134 directs air from the engine 118 toward the end of the wing 112.

[0037] Each duct 132, 134 may supply hot air to prevent ice formation at multiple locations along the leading edge 116 of the wing 112. The ducts may be configured to provide a consistent air pressure along the length of the leading edge 116. For example, the diameter of each duct may decrease with distance from the heat exchanger 122.

[0038] In this example, the air delivery device 130 includes both an outer air delivery duct and an inner air delivery duct. In some examples, the device includes either an outer air delivery duct or an inner air delivery duct. The ducts 132, 134 can extend together the entire span of the wing 112, or can extend along only a portion of the wing. Each duct can be integral, or can include a system of connected ducts, tubes, and / or other components.

[0039] Figure 3 to Figure 4 A typical leading edge ice protection system 200 is depicted, which may be installed in an aircraft such as aircraft 100 and supplied with hot air via a device such as air delivery device 130. Figure 3 As shown, the fixed leading edge portion or nose 210 of an aircraft wing includes a front cavity 212. An air delivery duct 214 extends through the front cavity 212 parallel to the leading edge 216 of the wing. The front cavity 212 is defined between an outer skin 218 of the nose 210 and a nose bridge 220 of the wing.

[0040] like Figure 4 As shown, the air delivery duct 214 includes a plurality of openings 222 from which hot air flows into the front cavity 212. This air flow is depicted by dashed lines 224. Each opening 222 of the duct is directed toward one or more impingement points of the outer skin 218. Openings at the upper edge of the nose bridge 220 and behind the nose bridge in the lower portion of the outer skin 218 allow hot air from the front cavity 212 to be discharged to the outside of the wing. The heating of the rear cavity extends to the closed boundary in front of the fixed leading edge nose 210.

[0041] In order to achieve sufficient heating of the outer skin 218, a large amount of hot air may need to be supplied from the air delivery duct 214 to the front cavity 212. The air may tend to flow as shown by arrows 224, but heating of the entire outer skin 218 may require a flow rate and pressure of hot air sufficient to fill the front cavity. The temperature may also vary along the outer skin. For example, a peak temperature may exist near each opening 222. For another example, a cross flow may occur, where air moves through the front cavity 212 in a direction parallel to the air delivery duct.

[0042] Accessing the air delivery duct 214 for repair or routine maintenance may be challenging and require removal of the outer skin 218 . Figure 4 Also shown is a flight control surface actuator 226 . The location and size of the actuator 226 may be limited by the location and extent of the forward cavity 212 .

[0043] Figure 5 3 is a schematic diagram of an exemplary ice protection system 300, depicting a cross-section of a fixed leading edge portion or nose 310 of an aircraft wing 302. Nose 310 is formed by an upper wing panel 312 and a lower wing panel 314. An aerodynamic outer skin 316 extends between these wing panels, defining a leading edge 318 of the wing. Wing panels 312, 314 may include one or more materials having any thermal properties, including but not limited to aluminum alloys and composite materials. Outer skin 316 may include a thermally conductive material, such as a metal alloy, and may be configured to conduct heat from an inner surface to an outer surface.

[0044] In this example, ice protection system 300 is configured to heat the fixed leading edge of the inboard portion of aircraft wing 302. In some examples, the ice protection system may be configured to heat other fixed surfaces of the wing and / or control surfaces of the inboard, outboard, leading edge and / or trailing edge portions of the wing and / or tail or canard surfaces. The ice protection system may be positioned according to the ice protection needs of a particular aircraft. For example, the system may be used to heat surfaces that are identified in a simulation analysis as being most susceptible to ice formation.

[0045] Ice protection system 300 includes air delivery duct 320 and inner skin 322. Air delivery duct 320 is enclosed in cavity 326 between outer skin 316 and cover 324. Air delivery duct may carry hot air along wing nose 310 from a source such as a heat exchanger or bleed air system. Air delivery duct 320 may be an example of air delivery device 130. Duct 320 is mounted proximate to an upper portion of outer skin 316, but spaced therefrom. Duct may be described as being above leading edge 318 and / or above camber 350 of wing 302.

[0046] The cover 324 can be described as having an open C-shape and / or a parabolic shape. A first end of the cover is fastened to a stringer 329, which is located at the junction between the outer skin 316 and the upper wing panel 312. A second end of the cover is fastened to a flange 330 of the inner skin 322. The stringer 329 and the flange 330 both extend generally perpendicular to the outer skin 316 toward the interior of the wing nose 310. The cover 324 is positioned so that the first end of the cover overlaps the stringer 329 and the second end of the cover overlaps the flange 330. The first and second ends of the cover can also be described as having outer surfaces that contact a majority of the stringer and the flange.

[0047] The cover 324 is removably connected to the stringer and the flange by a fastener assembly 332. In the example shown, each of the first and second ends of the cover 324 includes a fixing nut element. A bolt extends through a hole in each of the stringer 329 and the flange 330 to threadably engage the fixing nut element. The fixing nut element and the bolt together form the fastener assembly 332.

[0048] Any effective number and / or combination of fasteners or fastener assemblies may be used to attach cover 324. Cover may be removably and / or permanently fixed in place and may be attached and / or mounted to any suitable structure or component of wing nose 310. Cover 324 may be shaped to define a desired geometry and / or volume of cavity 326. Cover may be positioned and / or shaped based on the location of air delivery duct 320 and / or nearby structural components.

[0049] The size of the air delivery duct 320 may vary along the wing nose 310. For example, the diameter of the duct may decrease toward the fuselage or wing tip to maintain a consistent air pressure throughout the duct. The cover 324 may similarly vary in size so that the volume of the cavity 326 remains constant along the wing nose 310. In the depicted example, the air delivery duct 320 has approximately 2 3 / 4 inch diameter. The duct 320 and cover 324 may be of any size suitable for the desired air flow characteristics, such as velocity, pressure, and / or volume.

[0050] The cover 324 may include a flexible, resilient material to allow deformation during installation or removal of the cover. An installer or maintenance worker may compress the first end of the cover toward the second end of the cover so that the two ends may be inserted between the stringer 329 and the flange 330. The cover may be similarly compressed to allow the two ends of the cover to be withdrawn from between the stringer and the flange. When released from this compression, the resilient material of the cover 324 may bias the cover to return to the open C-shape.

[0051] When the cover 324 is properly positioned between the stringers 329 and the flange 330, the material of the cover can bias the ends of the cover against the stringers and the flange. That is, the ends of the cover can be pushed into contact with the stringers and the flange. The cover 324 can thus form an airtight seal with the stringers 329 and the flange 330. In some examples, a seal can be provided between the ends of the cover and the stringers and the flange to ensure that the seal is sufficiently airtight.

[0052] like Figure 6 As shown, the air delivery duct 320 is supported by a series of brackets 348. The brackets may be fixed and / or secured to any suitable structure, including the outer skin 316, the inner skin 322, and / or the stringers 329.

[0053] Refer again Figure 5 , the inner skin 322 is positioned adjacent to the outer skin 316 to form a plurality of channels 328. Each channel includes an inlet 334 in fluid communication with the cavity 326. The air delivery duct 320 includes a plurality of openings 336, each corresponding to an inlet 334. Each opening 336 is positioned on the duct 320 so that air flowing through the opening is directed toward the corresponding inlet 334. In other words, the duct 320 is configured to exhaust hot air from the opening 336 toward the inlet 334. Directing the openings 336 toward the inlet 334 can improve the uniformity of heating of the outer skin 316 and reduce the peak temperature on the outer skin near the opening.

[0054] Each passage 328 extends from an inlet 334 to an outlet 338 in fluid communication with the exterior of the wing 302. Each inlet 334 may be described as being in a gap between the outer skin 316 and an upper end 335 of the inner skin 322 leading to the cavity 326. The inlet 334 may be described as being above the leading edge 318, near the upper portion of the outer skin 316, and / or above the camber line 350. Each outlet 338 may be described as being in a gap between the outer skin 316 and the lower wing panel 314 that is open to the exterior of the wing 302. The outlet 338 may be described as being below the leading edge 318, near the lower portion of the outer skin 316, and / or below the camber line 350.

[0055] The anti-icing system 300 is configured to direct hot air from the air delivery duct 320 as shown by the dashed arrow 360. The air may flow from the opening 336 through the cavity 326 and into the inlet 334. Some of the hot air may also fill the cavity 326. As the air of the airflow 360 passes downward from the inlet 334 along the channel 328, heat may be transferred from the air to the outer skin 316. The outer skin 316 and the inner skin 322 may be referred to as being in heat exchange. The air may then exit the outlet 338 and continue under the wing 302. Due to the location of the duct 320, the inlet 334, and the outlet 338, the air may travel downward past the leading edge 318. Therefore, the outer skin 316 and the inner skin 322 may be referred to as being in counter-flow heat exchange.

[0056] The passage 328 is defined between the inner surface of the outer skin 316 and the outer surface 323 of the inner skin 322. The geometry of the passage 328 is determined by the outer contour of the inner skin 322. In other words, the curvature and features of the inner skin 322 facing the outer skin 316 determine the properties of the passage 328, such as the cross-sectional area and cross-sectional shape. The inner skin 322 can be contoured relative to the outer skin 316 so that the passage 328 is adjusted to achieve desired airflow characteristics, such as pressure and flow rate through the passage.

[0057] The depth 340 of each channel can be defined as the distance between the outer skin and the inner skin. The depth 340 can vary along each channel 328. Each channel 328 can be described as having an upper portion 342 near the inlet 334, a middle portion 344 near the leading edge 318, and a lower portion 346 near the outlet 338. In this example, the depth 340 has a first value for the upper portion 342 and the lower portion 346 and a second value for the middle portion 344, the first value being greater than the second value. For example, the upper portion 342 and the lower portion 346 can each have a depth of 0.2 inches, while the middle portion 344 has a depth of 0.1 inches. In some examples, the depth 340 can be different between the upper portion 342 and the lower portion 346. In some examples, the depth 340 can be constant throughout the channel 328 and / or can be appropriately varied to achieve the desired airflow characteristics through the channel. The depth 340 can be in the range of approximately one hundredth of an inch and one inch.

[0058] Increasing the depth of the channel 328 near the inlet 334 and outlet 338 can increase the flow velocity at the inlet and outlet, thereby facilitating a uniform flow through the channel with a higher overall velocity. The uniform flow can improve temperature uniformity and reduce hot spots on the outer skin 316. The inlet and outlet can also be configured to avoid blocking or restricting the air flow through the channel. For example, the size of the gap between the outer skin 316 and the lower wing panel 314 forming each outlet 338 can be determined to avoid restriction of airflow.

[0059] For another example, Figure 5 As shown, the inner skin includes a chamfer 352 adjacent to the outlet 338. The chamfer may form an angle of approximately one hundred and thirty-five degrees with the adjacent surface of the inner skin 322, or any angle greater than approximately ninety degrees. The chamfer 352 shown may also be described as a forty-five degree chamfer. The chamfer may reduce the restriction to the flow through the outlet 338 and increase the flow velocity.

[0060] Increasing the depth of the passage 328 near the inlet 334 and the outlet 338 can reduce the pressure of the airflow through the passage 328. The inlet, air delivery duct 320, and cavity 326 can also be configured to reduce the air pressure in the anti-icing system 300. For example, the air delivery duct 320 and the opening 336 can have a diameter selected to reduce the pressure required to deliver a desired air flow mass. For another example, the volume of the cavity 326 and the location of the opening 336 on the air delivery duct 320 can be selected to reduce the air pressure required to achieve a desired flow rate in the passage 328. For another example, the size of the gap between the inner skin 322 and the outer skin 316 forming each inlet 334 can be designed to reduce the pressure in the passage at a desired flow rate.

[0061] The reduced pressure used by anti-icing system 300 can free up a limited supply of heated bleed air from the aircraft engine for use by other systems. For example, another anti-icing system according to the present disclosure can be installed on the outboard leading edge of wing 302 and use the available bleed air. The reduced pressure in passage 328 can also improve the loads in wing 302 caused by airflow and reduce the weight driven into the wing.

[0062] Figure 7 35 is a view of an outer surface 323 of the inner skin 322. The skin includes a plurality of ridges 354 and a plurality of protrusions 356. Each ridge 354 extends from an upper end 335 of the inner skin 322 to a lower end 337 of the skin and is configured to contact the outer skin. The ridges 354 may be described as sealing against the outer skin and / or forming an airtight seal with the outer skin. Thus, each channel is defined between two adjacent ridges 354. The distance between adjacent ridges 354 may be referred to as a channel width 358. In the depicted example, the channel has a constant width. In some examples, the channel width 358 may vary from the upper end 335 to the lower end 337 of the inner skin 322 and / or may vary between channels.

[0063] Dividing the space between the inner skin 322 and the outer skin of the wing nose into distinct channels with ridges 354 can improve uniformity of outer skin heating by restricting lateral flow along the leading edge. Channel width 358 can be selected in conjunction with other characteristics of channel 328 to reduce pressure through the channel and increase temperature and / or heating uniformity across the wing nose.

[0064] Protrusions 356 facilitate connection of inner skin 322 to outer skin while spacing the inner skin from the outer skin. Each protrusion is configured to contact outer skin 316 and may include features suitable for a selected fastener or connection type. For example, each protrusion 356 may include a threaded hole sized to receive a bolt and corresponding to a hole in the outer skin. Figure 6 , a projection 356 including a fastener is shown. For another example, each projection 356 can have an outer surface suitable for bonding to an outer skin.

[0065] Each protrusion 356 has a thickness corresponding to the depth of the adjacent channel 328 so that the protrusion is positioned along the channel. The bosses can be positioned at any point on the outer surface of the inner skin 322 in any pattern suitable for connecting the inner skin to the outer skin. In this example, the protrusions 356 are arranged in two rows at the upper end 335 and the lower end 337 of the inner skin 322. The protrusions 356 along the upper end of the inner skin 322 have a thickness corresponding to the inlet 334, and the protrusions along the lower end of the inner skin have a thickness corresponding to the outlet 338.

[0066] In this example, inner skin 322 is also configured to function as a structural member of the wing. Inner skin may be described as a nose bridge, or as replacing the nose bridge of a wing. Inner skin 322 may include a heat resistant material having structural properties suitable for a structural member. For example, inner skin may be machined from an aluminum alloy. In some examples, replacing the nose bridge with inner skin 322 may provide additional space inside the nose of the wing. In some examples, the structural properties of inner skin 322 may allow the outer skin of the nose of the wing to be thinner and / or lighter.

[0067] Figure 8 302, including anti-icing system 300 and high lift system 370. In this example, the system controls Krueger flap 372. High lift system includes actuator, electrical system and Krueger hinge 374. Air delivery duct 320 is positioned above the camber of the wing, on the upper surface of the nose of the wing. Inner skin 322 fits outer skin 316 closely, and serves as the nose of the wing at the same time. Krueger hinge 374 can therefore be positioned in the nose of the wing. This positioning can allow hinge 374 and / or the actuator for the hinge to be reduced in size, and can improve the trajectory of Krueger flap 372. This positioning can also make room for other systems in wing 302.

[0068] B. Exemplary Methods of Preventing Ice Formation

[0069] This section describes the steps of an exemplary method 400 for preventing ice from forming on the leading edge of an airfoil; see Fig. 9 Aspects of the aircraft and / or ice protection system described above may be used in the method steps described below. Where appropriate, reference may be made to components and systems that may be used to perform each step. These references are for illustrative purposes only and are not intended to limit the possible ways in which any particular step of the method may be performed.

[0070] Fig. 9 is a flowchart showing steps performed in an exemplary method and may not describe the complete process or all steps of the method. Fig. 9 4. Various steps of method 400 are depicted in FIG. 4, but these steps do not have to be performed in full, and in some cases can be performed simultaneously or in an order different from that shown.

[0071] At step 410, the method includes directing hot air from the heat exchanger to a duct. For example, the heat exchanger may be part of a bleed air system of an aircraft engine, and the duct may be a Piccolo tube connected to the heat exchanger. The duct may extend along a portion of a wing, such as an aircraft wing. For example, the duct may extend parallel to the leading edge of the wing between a wing-mounted engine and a fuselage of the aircraft. The duct may be mounted inside the airfoil, above the camber of the airfoil. For example, the duct may be supported by a bracket from an upper portion of an outer aerodynamic skin of the airfoil.

[0072] Step 412 of the method includes dispersing air from the duct toward an inlet of a channel in a leading edge portion of the airfoil. The duct may include a plurality of openings along the length of the duct configured to dissipate hot air. Each opening may be positioned on the duct so that the dispersed air is directed toward the inlet of the channel. The duct may be enclosed by a cover, and dispersing air from the duct may include filling the enclosed volume with hot air.

[0073] The leading edge portion of the airfoil may include an inner skin and an outer skin, with a channel defined between the inner skin and the outer skin by a plurality of ridges on the inner skin. The inner skin may be fixed to the outer skin and may serve as a structural member of the airfoil. The distance between the inner skin and the outer skin may define a depth of the channel, which may vary along the channel.

[0074] Step 414 of the method includes exhausting air from the outlet of the channel. Each channel may include an inlet at a first end and an outlet at a second end. Expelling air may include conveying the air dispersed from the duct in step 412 along the channel from the inlet to the outlet. Each inlet may be disposed above the leading edge of the airfoil, and each outlet may be disposed below the leading edge. Conveying air along the channel may include directing air downwardly in a reverse flow direction along the inner surface of the leading edge of the outer skin of the airfoil. Conveying air along the channel may also include transferring heat from the air to the outer skin, thereby uniformly heating the outer skin.

[0075] Each channel may have a varying cross-sectional size between the inlet and the outlet. For example, the channel may have a greater depth in the upper portion near the inlet and in the lower portion near the outlet than in the middle portion near the leading edge of the airfoil. Such varying sizes may allow air to be transported through the channel at a consistent flow rate, high velocity, and low pressure.

[0076] C. Exemplary Aircraft and Related Methods

[0077] Examples disclosed herein may be used in exemplary aircraft manufacturing and service method 500 (see Fig.10 ) and exemplary aircraft 600 (see Fig.11). The method 500 includes a number of processes, periods, or stages. During pre-production, the method 500 may include a specification and design phase 504 of the aircraft 600 and a material procurement phase 506. During production, a component and subassembly manufacturing phase 508 and a system integration phase 510 of the aircraft 600 may occur. Thereafter, the aircraft 600 may undergo a certification and delivery phase 512 to be placed in an in-service phase 514. While in service (e.g., by an operator), the aircraft 600 may be scheduled for routine maintenance and repair 516 (which may also include modification, reconfiguration, refurbishment, etc. of one or more systems of the aircraft 600). Although the examples described herein generally relate to operational use of the aircraft 600 during the in-service phase 514, they may be implemented at other stages of the method 500.

[0078] Each process of method 500 may be performed or carried out by a system integrator, a third party, and / or an operator (e.g., a customer). For purposes of this specification, a system integrator may include, but is not limited to, any number of aircraft manufacturers and primary system subcontractors; a third party may include, but is not limited to, any number of suppliers, subcontractors, and vendors; and an operator may be an airline, a leasing company, a military entity, a service organization, etc.

[0079] like Fig.11 As shown, an aircraft 600 produced by exemplary method 500 may include a fuselage 602 having a plurality of systems 604 and an interior 606. Examples of the plurality of systems 604 include one or more of a propulsion system 608, an anti-icing system 610, a pneumatic system 612, and a flight control system 614. Each system may include various subsystems, such as controllers, processors, actuators, effectors, motors, generators, etc., depending on the functions involved. Any number of other systems may be included. Although an aerospace example is shown, the principles disclosed herein may be applied to other industries, such as the automotive industry, the rail transportation industry, and the marine engineering industry. Therefore, in addition to the aircraft 600, the principles disclosed herein may be applied to other vehicles, such as land vehicles, marine vehicles, etc.

[0080] During any one or more stages of the manufacturing and service method 500, the apparatus and methods shown or described herein may be employed. For example, components or subassemblies corresponding to the component and subassembly manufacturing stage 508 may be manufactured or produced in a manner similar to components or subassemblies produced when the aircraft 600 is operated during the service stage 514. Also, one or more examples of apparatus, methods, or a combination thereof may be utilized during the manufacturing stage 508 and the system integration stage 510, for example, by substantially expediting or reducing the cost of assembling the aircraft 600. Similarly, for example, but not limited to, one or more examples of apparatus or method implementations, or a combination thereof, may be utilized while the aircraft 600 is in the service stage 514 and / or during the maintenance and repair stage 516.

[0081] Exemplary Combinations and Further Examples

[0082] This section describes additional aspects and features of ice protection systems and related methods, which are presented, without limitation, as a series of paragraphs, some or all of which may be alphanumeric for clarity and effectiveness. Each of these paragraphs may be combined with one or more other paragraphs, and / or combined with disclosure from elsewhere in this application in any suitable manner. Some of the following paragraphs explicitly refer to and further limit other paragraphs, providing, without limitation, examples of some of the suitable combinations.

[0083] A0. An anti-icing system, the anti-icing system comprising:

[0084] An airfoil structure comprising a leading edge portion facing the wind direction and comprising an outer skin and an inner skin forming a plurality of channels, each channel having an air inlet and an air outlet,

[0085] an air delivery duct located within the airfoil structure and extending substantially transversely to the wind direction, the air delivery duct having a plurality of openings directed toward the air inlets of the plurality of channels, and

[0086] An air supply device is configured to provide hot air to the air delivery duct.

[0087] A1. An anti-icing system according to A0, wherein each channel has a cross-sectional dimension, and each cross-sectional dimension varies between the air inlet and the air outlet.

[0088] A2. The ice protection system of A1, wherein each channel has an upper portion, a middle portion, and a lower portion, at least one of the upper portion and the lower portion having an increased cross-sectional area compared to the middle portion.

[0089] A3. An ice protection system according to A1 or A2, wherein each channel has an upper portion, a middle portion and a lower portion, and the upper portion and the lower portion each have an increased cross-sectional area compared to the middle portion.

[0090] A4. An anti-icing system according to any one of A0 to A3, wherein the air inlet receives air from within the airfoil structure and the air outlet discharges air to the outside of the airfoil structure.

[0091] A5. An anti-icing system according to any one of A0 to A4, wherein the inner skin has a front surface facing the outer skin, the front surface includes an upper edge portion, a lower edge portion, and a plurality of ridges extending from the upper edge portion to the lower edge portion, the plurality of ridges forming walls of the plurality of channels.

[0092] A6. The ice protection system of A5, wherein the front surface has a plurality of protrusions along the upper edge portion and the lower edge portion configured to connect the inner skin and the outer skin.

[0093] A7. The anti-icing system according to any one of A0 to A6, further comprising:

[0094] A cover is used to close the air delivery duct.

[0095] A8. An ice protection system according to A7, wherein the cover is connected to a flange structure within the airfoil structure.

[0096] A9. An ice protection system according to A8, wherein the inner skin has a front surface facing the outer skin, and the front surface includes an upper edge portion, and the flange structure includes a first flange structure extending inwardly substantially orthogonally from the upper edge portion of the inner skin.

[0097] A10. The ice protection system of A8 or A9, wherein the cover is removably secured to the flange structure.

[0098] A11. An anti-icing system according to any one of A7 to A10, wherein the cover has an open C-shape.

[0099] A12. An anti-icing system according to any one of A0 to A11, wherein the heat exchanger is in fluid communication with the air delivery duct.

[0100] A13. An anti-icing system according to any one of A0 to A12, wherein the airfoil structure includes an upper surface and a lower surface, each air inlet is adjacent to the upper surface, and each air outlet is adjacent to the lower surface.

[0101] A14. An ice protection system according to any one of A0 to A13, wherein the airfoil structure includes an arc line, each air inlet is above the arc line, and each air outlet is below the arc line.

[0102] A15. An anti-icing system according to any one of A0 to A14, wherein the air delivery duct is arranged above the camber line of the airfoil structure.

[0103] A16. An anti-icing system according to any one of A0 to A15, wherein the air delivery duct is arranged close to an upper surface of the airfoil structure.

[0104] A17. An anti-icing system according to any one of A0 to A15, wherein the air supply device includes a heat exchanger.

[0105] B0. An aircraft, comprising:

[0106] body,

[0107] a first wing extending from the fuselage, a second wing extending from an opposite side of the fuselage, each wing including a leading edge portion facing the wind direction, each leading edge portion including an outer skin and an inner skin forming a plurality of internal passages, each internal passage having an internal air inlet and an external air outlet,

[0108] Each wing includes an air delivery duct extending substantially transversely to the wind direction, each air delivery duct having a plurality of openings directed toward the internal air inlet of the corresponding plurality of internal passages, and

[0109] An air supply device is configured to provide heated air to the plurality of internal channels.

[0110] B1. An aircraft according to B0, wherein each internal passage has a cross-sectional dimension, and each cross-sectional dimension varies between the air inlet and the air outlet.

[0111] B2. An aircraft according to B0 or ​​B1, wherein each internal channel has an upper portion, a middle portion and a lower portion, at least one of the upper portion and the lower portion having an increased cross-sectional area compared to the middle portion.

[0112] B3. An aircraft according to any one of B0 to B2, wherein each channel has an upper portion, a middle portion and a lower portion, and the upper portion and the lower portion each have an increased cross-sectional area compared to the middle portion.

[0113] B4. An aircraft according to any one of B0 to B3, wherein the inner skin has a front surface facing the outer skin, the front surface includes an upper edge portion, a lower edge portion and a plurality of ridges extending from the upper edge portion to the lower edge portion, the plurality of ridges forming walls of the plurality of internal channels.

[0114] B5. The aircraft according to B4, wherein the front surface has a plurality of protrusions along the upper edge portion and the lower edge portion configured to connect the inner skin and the outer skin.

[0115] B6. An aircraft according to any one of B0 to B5, the aircraft further comprising:

[0116] A cover is provided for closing the air delivery duct in each wing, each cover being connected to a flange structure in the corresponding wing, wherein one of the flange structures extends generally orthogonally from an upper edge portion of the inner skin.

[0117] B7. An aircraft according to any one of B0 to B6, wherein the air supply device includes a heat exchanger of a bleed air system.

[0118] B8. An aircraft according to any one of B0 to B7, wherein the air supply device includes an engine shaft-driven compressor.

[0119] C0. A method for preventing ice formation on a leading edge portion of an airfoil, the method comprising:

[0120] directing the hot air from the heat exchanger to an air delivery duct extending along the interior of the leading edge portion of the airfoil structure,

[0121] dispersing air from the air delivery duct towards channels formed between the inner skin and the outer skin of the leading edge portion of the airfoil-profile structure, wherein each channel extends between an inlet and an outlet between an upper section and a lower section of the leading edge portion, and

[0122] Air is exhausted from the passage through the outlet.

[0123] C1. A method according to C0, wherein each channel has a varying cross-sectional dimension between the inlet and the outlet.

[0124] C2. The method of C0 or C1, wherein discharging air from the passage comprises discharging air outside the airfoil.

[0125] C3. A method according to any one of C0 to C2, wherein the inlet of each channel is located at the first end of the channel, and the outlet of each channel is located at the second end of the channel.

[0126] Advantages, Features and Benefits

[0127] The various examples of anti-icing systems described herein provide several advantages over known solutions for preventing ice from forming on the leading edge of an airfoil. For example, the exemplary embodiments described herein allow for greater space and clearance within the wing for other systems such as flight control surface actuators.

[0128] Additionally, the illustrative examples described herein allow for the inner skin of an ice protection system to be used as a structural component of an airfoil, among other benefits.

[0129] Additionally, the exemplary embodiments described herein reduce the volume of heated air required to effectively prevent ice formation, among other benefits.

[0130] Additionally, the illustrative examples described herein reduce the pressure and weight of air driven into the airfoil, among other benefits.

[0131] Additionally, the illustrative examples described herein provide, among other benefits, consistent hot air flow over the aerodynamic outer skin of an airfoil.

[0132] Additionally, the illustrative examples described herein allow for maintenance access to air delivery ducts, among other benefits.

[0133] No known system or device can perform these functions, particularly while achieving effective uniform heating of the aerodynamic outer skin of the airfoil. Therefore, the exemplary examples described herein are particularly useful for preventing ice from forming on aircraft wings. However, not all examples described herein provide the same advantages or the same degree of advantages.

[0134] in conclusion

[0135] The disclosure set forth above may include multiple different examples with independent utility. Although each of these is disclosed in its preferred form, its specific examples as disclosed and shown herein should not be considered restrictive, because many variations are possible. With regard to the section titles used in this disclosure, such titles are only used for organizational purposes. The subject matter of the present disclosure includes all novel and non-obvious combinations and sub-combinations of various elements, features, functions and / or properties disclosed herein. The attached claims specifically point out certain combinations and sub-combinations that are considered to be novel and non-obvious. Other combinations and sub-combinations of features, functions, elements and / or characteristics may be claimed in applications claiming priority to this application or a related application. Such claims, whether their scope is wider, narrower, the same or different than the original claims, are considered to be included in the subject matter of the present disclosure.

Claims

1. An anti-icing system (300), the anti-icing system comprising: An airfoil structure (112, 302) having a camber line (350), the airfoil structure comprising a leading edge portion (310) facing a wind direction (124) and comprising an outer skin (316) and an inner skin (322) forming a plurality of channels (328), each channel having an air inlet (334) above the camber line (350) and an air outlet (338) below the camber line (350), an air delivery duct (320) located within the airfoil structure above the camber line (350) and extending substantially transversely to the wind direction, the air delivery duct having a plurality of openings (336) directed toward the air inlets of the plurality of channels, and An air supply device (122) is configured to supply hot air to the air delivery duct.

2. The anti-icing system (300) according to claim 1, wherein: Each passage (328) has a cross-sectional dimension (340), and each cross-sectional dimension varies between the air inlet (334) and the air outlet (338).

3. The anti-icing system (300) according to claim 2, wherein: Each channel (328) has an upper portion (342), a middle portion (344), and a lower portion (346), at least one of the upper portion and the lower portion having an increased cross-sectional area compared to the middle portion.

4. The anti-icing system (300) according to claim 2, wherein: Each channel (328) has an upper portion (342), a middle portion (344), and a lower portion (346), the upper portion and the lower portion each having an increased cross-sectional area compared to the middle portion.

5. The anti-icing system (300) according to any one of claims 1 to 4, wherein: The air inlet (334) receives air from within the airfoil structure (112, 302), and the air outlet (338) discharges air outside the airfoil structure.

6. The anti-icing system (300) according to any one of claims 1 to 4, wherein: The inner skin (322) has a front surface (323) facing the outer skin (316), the front surface including an upper edge portion (335), a lower edge portion (337) and a plurality of ridges (354) extending from the upper edge portion to the lower edge portion, the plurality of ridges forming walls of the plurality of channels (328).

7. The anti-icing system (300) according to claim 6, wherein: The front surface (323) has a plurality of protrusions (356) along the upper edge portion (335) and the lower edge portion (337), the plurality of protrusions being configured to connect the inner skin (322) and the outer skin (316).

8. The anti-icing system (300) according to claim 1, further comprising: A cover (324) is used to close the air delivery duct (320).

9. The anti-icing system (300) according to claim 8, wherein: The cover (324) is connected to a flange structure (330, 329) within the airfoil structure (112, 302).

10. The anti-icing system (300) according to claim 9, wherein: The inner skin (322) has a front surface (323) facing the outer skin, and the front surface includes an upper edge portion (335), and the flange structure (330, 329) includes a first flange structure (330) extending inwardly substantially orthogonally from the upper edge portion of the inner skin.

11. The anti-icing system (300) according to any one of claims 8 to 10, wherein: The cover (324) has an open C-shape.

12. A method (400) of preventing ice from forming on a leading edge portion of an airfoil, the method comprising: directing (410) the hot air to an air delivery duct (320) extending along an interior (326) of a leading edge portion (310) of an airfoil structure (112, 302), the air delivery duct (320) being located above a camber line (350) of the airfoil structure (112, 302), dispersing (412) air from the air delivery duct toward channels (328) formed between an inner skin (322) and an outer skin (316) of the leading edge portion of the airfoil profile structure, wherein each channel extends between an upper portion (342) and a lower portion (346) of the leading edge portion between an air inlet (334) located above the camber (350) and an air outlet (338) located below the camber (350), and Air is exhausted (414) from the passage through the outlet.

13. The method (400) of claim 12, wherein: Each passage (328) has a varying cross-sectional dimension (340) between the air inlet (334) and the air outlet (338).

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