Ice protection system for aircraft

Abstract

The invention relates to an anti-icing system for an aircraft. A system and method includes a heating layer including a perforation extending from a first surface to a second surface opposite the first surface. The aperture provides an acoustic path. The heating layer is configured to provide resistive heating to the component. The acoustic layer may be coupled to the heating layer. The through hole extends to the acoustic layer.

Description

[0001] Citations of Related Applications

[0002] This application is related to and claims priority to U.S. Provisional Patent Application No. 63 / 589,679, filed on October 12, 2023, which is hereby incorporated by reference in its entirety. Technical Field

[0003] Examples of the present disclosure generally relate to an anti-icing system for an aircraft. Background Art

[0004] Composite materials have been used as structural elements in a variety of applications. Composite materials such as fiber reinforced polymers have relatively high strength to weight ratios. The high strength to weight ratio makes these materials desirable for vehicle construction such as aircraft. These materials typically include multiple layers or plies of reinforcement material combined in a matrix material to form a heterogeneous material with combined properties that differ from the reinforcement material and matrix material alone.

[0005] For some applications, it is desirable for composite materials to provide multiple functions. For example, it may be desirable to have a single laminate structure that provides structural support and conducts electrical current to power electrical loads, communication transmissions, etc. Power distribution to embedded conductive materials within a laminate composite structure requires a means to deliver current to known locations within the thickness of the laminate structure and in specific areas.

[0006] Some existing systems embed one or more insulated wires into the thickness of a laminate structure. The assembly of the laminate may include forming cuts or splices in multiple plies and feeding wires through the cuts as the plies are sequentially applied to a mandrel or other tool to form a laminate stack. This process of embedding wires into the structure at multiple locations during the structure formation process may be detrimental to the structural performance of the laminate structure and may inhibit proper interfaces between the structural plies of the laminate structure and adjacent components (e.g., a honeycomb core). For example, routing wires through the surface of a laminate structure at multiple locations may result in discontinuities within the laminate structure, such as bulges and stress concentrations. Discontinuities may degrade the structural performance of the laminate, requiring additional plies (and therefore additional weight) to meet material strength requirements. In addition, the insulation of the conductors (wires) may contaminate the matrix material (e.g., resin) of the laminate structure, degrading material properties. Wire stripping and other wire preparation and routing tasks may increase the risk of foreign object debris (FOD) contaminating the laminate structure. Embedded wires may also introduce a risk of failure due to the handling of the wires and interference between the wires and adjacent components. During the life of the system, the embedded wires may experience fatigue and wear over time, which may cause the wires to fail. In addition, in at least one application, perforations are formed throughout the laminated structure. The perforations can be used for noise attenuation, distributing fluids through the laminated structure, etc. In the case where the embedded wires are present within the thickness of the laminated structure, there is a risk that the wires may be damaged during the perforation formation process, which may make the electrical load powered by the wires inoperable. On the contrary, avoiding embedded wires may undesirably limit the area of ​​the perforations, thereby limiting the effectiveness of the perforations.

[0007] Various aircraft include engines having nacelles. Typical nacelles rely on ejected exhaust (i.e., bleed air) extracted from the engine core to prevent ice from forming and / or accumulating on the inlet leading edge. Bleed air extraction reduces the efficiency of the engine and limits material choices at the leading edge due to the high exhaust temperatures involved. Furthermore, such systems typically extend along the length of the nacelle, thereby adding weight and further reducing efficiency. Furthermore, extracting bleed air can reduce engine performance and efficiency. Summary of the invention

[0008] There is a need for an improved anti-icing system for an aircraft engine.There is a need for a lighter anti-icing system that improves the operating efficiency of an aircraft engine.

[0009] In view of these needs, some examples of the present invention provide a system including a heating layer including perforations extending from a first surface to a second surface opposite the first surface. The perforations attenuate aircraft noise. The perforations provide an acoustic path. The heating layer is configured to provide resistive heating to a component.

[0010] In at least one example, the system further includes an acoustic layer coupled to the heating layer. The perforations may extend into the acoustic layer. In at least one example, the heating layer overlays at least a portion of the acoustic layer. The heating layer may be a carbon shield or other material including carbon fibers or carbon nanotubes.

[0011] The component may include a nacelle of an aircraft engine.

[0012] In at least one example, the heating layer includes a dielectric layer, a conductive heater coupled to the dielectric layer, and one or more films coupled to the conductive heater. The dielectric layer may include glass fiber. The one or more films may include a nickel chemical vapor deposition film. In at least one example, the perforations extend through the dielectric layer, the conductive heater, and the one or more films. The dielectric layer, the conductive heater, and the one or more films may be laminated together.

[0013] Some examples of the invention provide an aircraft including a propulsion system, the propulsion system including a nacelle having an inlet cowl. The inlet cowl includes an acoustic layer coupled to an anti-icing system. The anti-icing system includes a heating layer, the heating layer including perforations extending from a first surface to a second surface opposite the first surface. The perforations may include an acoustic liner. The perforations provide an acoustic path. The heating layer is configured to provide resistive heating to the nacelle.

[0014] Some examples of the present disclosure provide a method comprising forming perforations in a heating layer, the perforations extending from a first surface to a second surface opposite the first surface, wherein the perforations provide an acoustic path, and wherein the heating layer is configured to provide resistive heating to a component; and coupling the heating layer to an acoustic layer, wherein the perforations extend to the acoustic layer. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] These and other features, aspects and advantages of the present disclosure will become better understood when the following detailed description is read with reference to the accompanying drawings, in which like reference numerals refer to like parts throughout, and in which:

[0016] Figure 1 A block diagram of an ice protection system according to an example of the present disclosure is shown.

[0017] Figure 2 A plan view of an ice protection system according to an example of the present disclosure is shown.

[0018] Figure 3 Shown by Figure 2 A cross-sectional view of the anti-icing system of line 3-3.

[0019] Figure 4 A block diagram of a laminated composite structure according to an example of the present disclosure is shown.

[0020] Figure 5 A perspective view of an aircraft is shown.

[0021] Figure 6 An isometric side view of a nacelle showing the propulsion system of an aircraft.

[0022] Figure 7 Shows Figure 6 A front view of the air intake hood of the engine compartment is shown.

[0023] Figure 8 Shows Figure 7 A portion of the air intake hood indicated by a dotted circle.

[0024] Fig. 9 A cross-sectional view of a portion of an air intake hood of an engine compartment according to an example of the present disclosure is shown. DETAILED DESCRIPTION

[0025] When read in conjunction with the accompanying drawings, the above overview and the following detailed description of some embodiments will be better understood. As used herein, elements or steps stated in singular form and preceded by the word "one" or "a kind of" should be understood to not necessarily exclude the plural number of elements or steps. In addition, reference to "an embodiment" is not intended to be interpreted as excluding the existence of other embodiments that also combine the features. In addition, unless explicitly stated to the contrary, embodiments that "include" or "have" an element or multiple elements with a specific attribute may include other elements that do not have the attribute.

[0026] Engine manufacturers continue to increase fan diameters and reduce engine core size to increase bypass ratios, which can improve fuel efficiency. To offset the increased weight and drag from a larger diameter nacelle, significant benefits can be achieved by reducing the axial length of the nacelle. This in turn reduces the amount of area available for acoustic treatment. Accordingly, some examples of the present disclosure provide an anti-icing system that covers at least a portion of the acoustic treatment, thereby allowing the acoustic treatment portion of the inlet to extend to the water capturing area of ​​the leading edge.

[0027] In at least one example, an anti-icing system replaces the high temperature bleed air system. The anti-icing system is an electrical anti-icing system machined into the structural laminate. The anti-icing system is configured to operate at sufficiently low temperatures to allow the use of composite materials conventionally used for acoustic attenuation. The anti-icing system enables a multifunctional structure in which the acoustic liner can be pulled further forward to the leading edge of the nacelle. The anti-icing system can cover one or more portions of the acoustic area, thereby allowing the nacelle to be shortened, which reduces weight and drag without compromising noise protection or ice protection.

[0028] Examples of the present disclosure facilitate extension of acoustic pads to previously hard-walled areas to increase acoustic attenuation while reducing the overall length of the nacelle. Lower operating temperatures and an electrical anti-icing system reduce the need for high temperature resistant materials at the leading edge of the nacelle. In at least one example, the anti-icing system integrates a continuous conductive grid heater that enables anti-icing to be co-located with a perforated acoustic attenuation zone.

[0029] Figure 1 A block diagram of an anti-icing system 100 according to an example of the present disclosure is shown. The anti-icing system 100 includes a heating layer 102 having perforations 104. The heating layer 102 is an electrical layer configured to generate heat. The heating layer 102 is electrically coupled to a power source 106, such as through one or more leads 108. Although shown as an anti-icing system 100, other types of systems may be used. For example, the heating system may include a heating layer 102 coupled to an acoustic layer 110.

[0030] In at least one example, the heating layer 102 is a carbon shroud that includes randomly oriented carbon fibers. As another example, the heating layer 102 includes carbon nanotubes. The heating layer 102 can be any type of continuous conductive material configured to provide a resistive heater. For example, the current generated by the power supply 106 flows into the heating layer 102, thereby increasing the temperature of the heating layer 102.

[0031] The perforations 104 are holes, openings, channels, etc. formed through the heating layer 102. For example, the perforations 104 may be holes with a diameter of 0.05 inches or less. In at least one example, the perforations 104 may have a diameter of 0.2 inches or less. Optionally, the perforations 104 may have a diameter less than or greater than 0.2 inches (e.g., 0.1 inches or 0.5 inches). The perforations 104 may be spaced apart from each other by a distance of 1 inch or less. For example, adjacent (i.e., nearest) perforations 104 may be spaced apart from each other by less than 0.2 inches. Optionally, the perforations 104 may be spaced apart from each other by a distance greater than 1 inch (such as 2 inches or more).

[0032] The perforations 104 extend through the heating layer 102. For example, the perforations are located between a first surface (such as a top surface) and a second surface (such as a bottom surface) opposite the first surface and pass through the first surface (such as a top surface) and the second surface (such as a bottom surface). The perforations 104 provide an acoustic path configured to attenuate acoustic energy. Therefore, the heating layer 102 is acoustically perforated.

[0033] In at least one example, heating layer 102 is integrated into the laminated composite structure. The composite structure may further include an acoustic layer 110, such as an acoustic liner of a nacelle of a propulsion system (such as an engine) of an aircraft. In at least one example, anti-icing system 100 including heating layer 102 covers (e.g., covers) at least a portion of acoustic layer 110. For example, the entirety of anti-icing system 100 covers at least a portion of acoustic layer 110. As another example, anti-icing system 100 covers the entirety of acoustic layer 110. As another example, acoustic layer 110 includes one or more portions that are not covered by anti-icing system 100.

[0034] Acoustic layer 110 may be part of component 111. For example, component 111 may include acoustic layer 110 and anti-icing system 100, which may include heating layer 102. Heating layer 102 and acoustic layer 110 may be integrated into component 111. Optionally, heating layer 102 and acoustic layer 110 may be separately secured to component 111.

[0035] In at least one example, component 111 is a nacelle of an aircraft engine. Optionally, component 111 can be various other types of structures configured to be heated and also have attenuated acoustic energy. For example, component 111 can be part of a ventilation, heating, cooling, refrigeration system, part of a residential or commercial building, etc.

[0036] As described herein, the system includes a heating layer 102 including perforations 104 extending from a first surface to a second surface opposite the first surface. The heating layer 102 is configured to provide resistive heating to a component 111. An acoustic layer 110 is coupled to the heating layer 102. The perforations 104 extend to the acoustic layer 110. The heating layer 102 covers at least a portion of the acoustic layer 110. In at least one example, the heating layer 102 is laminated to the acoustic layer 110 to form a laminated composite structure. The heating layer 102 can be part of an anti-icing system 100 or other type of heating system.

[0037] Some examples of the present disclosure provide a method including forming a perforation 104 in a heating layer 102. The perforation 104 extends from a first surface to a second surface opposite the first surface. The perforation 104 provides an acoustic path. The heating layer 102 is configured to provide resistive heating to a component 111. The method may also include coupling the heating layer 102 to an acoustic layer 110. The perforation 104 extends to the acoustic layer 110.

[0038] Figure 2 1 shows a plan view of an anti-icing system 100 according to an example of the present disclosure. The heating layer 102 may include a plurality of conductive strips 112, each conductive strip including a perforation 104 (in Figure 1 1). The heating layer 102 may include more conductive strips 112 than shown. Optionally, the heating layer 102 may include a single integral sheet covering the substrate (as opposed to multiple strips separated by gaps). The leads 108 are electrically coupled to the heating layer 102 to provide power thereto, so that the heating layer 102 can provide resistive heating to a structure, such as a portion of an aircraft engine.

[0039] Figure 3 Shown by Figure 23-3 of the cross-sectional view of the anti-icing system 100. The heating layer 102 can cover a portion of the acoustic layer 110. In at least one example, the heating layer 102 includes a dielectric layer 120 (such as one or more insulators) covering a film adhesive 122 located on the acoustic layer 110. The dielectric layer 120 can include a glass fiber prepreg, other insulator compatible laminates, etc. A conductive heater 124 is disposed above the dielectric layer 120. A membrane 126 is disposed on the conductive heater 124. In at least one example, the membrane 126 is a nickel chemical vapor deposition membrane. The membrane 126 can be any type of conductive membrane. Optionally, instead of a membrane, a metal mesh can be used. In addition, a surface film 128 formed of a material similar to the membrane 126 can be further disposed on the heating layer 102. In at least one example, the dielectric layer 120, the conductive heater 124, and the membrane 126 are laminated together to form the heating layer 102. In at least one example, the perforations 104 extend through the entire thickness 103 of the heating layer 102 from a first surface 130 (such as an exposed surface) to a second surface 132 (such as a covered surface on the acoustic layer 110) opposite the first surface 130. For example, the perforations 104 extend through the dielectric layer 120, the conductive heater 124, and the membrane 126. Thus, the perforations 104 provide an acoustic path to the acoustic layer 110. In this way, acoustic energy can be guided through the heating layer 102 into the acoustic layer 110, which is configured to attenuate the acoustic energy. Because the conductive heater 124 is resistively heated, the perforations do not inhibit the heating ability of the conductive heater 124. Instead, the conductive heater 124 generates heat around the perforations. In this way, wires and leads do not extend through the heating layer 102, which thereby reduces the size, weight, and complexity of the heating layer 102.

[0040] As shown, the acoustic layer 110 can include a composite core 140 disposed on a composite sheet or skin 142. The acoustic layer 110 can be a unitary laminate or a sandwich structure. The acoustic layer 110 can be an acoustic honeycomb structure. The anti-icing system 100 including the heating layer 102 can be integrated into a laminate structure that also includes the acoustic layer 110.

[0041] As depicted, the heating layer 102 includes perforations 104. The heating layer 102 can be disposed on at least a portion of a composite structure, such as an acoustic layer 110. The acoustic layer 110 can be acoustically active, perforated, or non-perforated. The heating layer 102 is configured to heat a component, such as a nacelle of an aircraft engine, when electrically coupled to a power source 106. The perforations 104 of the heating layer 102 allow acoustic energy (e.g., noise) to enter the acoustic layer 110, where such acoustic energy is attenuated. Because the heating layer 102 is a continuous network of perforations 104, electrical continuity with the power source 106 is maintained, thereby allowing the heating layer 102 to heat the component.

[0042] In at least one example, anti-icing system 100 is configured to heat a component such as a nacelle of an engine, while acoustic layer 110 is configured to attenuate acoustic energy. Perforations 104 of heating layer 102 allow acoustic energy to pass through heating layer 102 and into acoustic layer 110, which cancels or attenuates the acoustic energy. In this way, a component such as a nacelle of an engine can include heating layer 102 disposed over (e.g., covering) a portion of acoustic layer 110 (compared to nacelles known in the prior art having heaters separate from the acoustic layer). In this way, a shorter, lighter, and more streamlined nacelle can be formed.

[0043] Figure 4 A block diagram of a laminated composite structure 150 according to an example of the present disclosure is shown. The laminated composite structure 150 includes a heating layer 102 coupled to an acoustic layer 110, as described herein. The heating layer 102 and the acoustic layer 110 can be laminated together and processed to form the laminated composite structure 150. In at least one example, the heating layer 102 is a Figure 1 A portion of an ice protection system 100 is shown that may be used to prevent, minimize, or reduce ice formation on and / or in a nacelle of an aircraft engine.

[0044] The laminated composite structure 150 may have a rectangular shape, but may have other shapes in other examples. In addition, before use, the laminated composite structure 150 may be bent, molded or formed into a non-planar shape to have a specified contour. In an exemplary application, the laminated composite structure 150 may form a panel of an air intake hood of an engine compartment of an aircraft. The laminated composite structure 150 may include an integrated electrical load defining a power supply area. The power supply area may be positioned along a central area of ​​the laminated composite structure 150. The power supply area may occupy most of the area of ​​the laminated composite structure 150. In one example, the power supply area represents a heating area. The power supply area may be heated to prohibit ice accumulation, thereby maintaining the temperature of the laminated composite structure 150 within a specified range, etc. The electrical load may include a resistive layer that converts electrical energy into thermal energy.

[0045] Figure 5 It is a perspective view of the aircraft 200 . Figure 1 The anti-icing system 100 shown in Figure 4The laminated composite structure 150 shown in FIG. 1 may be mounted on an aircraft 200. The aircraft 200 may include a fuselage 202 extending from a nose 203 to an empennage 204. The empennage 204 may include one or more tail surfaces for directional control of the aircraft 200. The aircraft 200 includes a pair of wings 206 extending from the fuselage 202. One or more propulsion systems 208 propel the aircraft 200. The propulsion systems 208 are supported by the wings 206 of the aircraft 200, but may be mounted to the fuselage or the tail in other types of aircraft. Each propulsion system 208 includes a rotor assembly 219 having a rotor that rotates to direct air.

[0046] The rotor assembly 219 of each propulsion system 208 is surrounded by a nacelle 210. The nacelle 210 is an outer shell or housing that holds the rotor assembly 219. The nacelle 210 includes an inlet section, called an air intake cowl, at the front or forward end of the nacelle 210. The nacelle 210 may also include a fan cowl, a thrust reverser section, and a rear fairing section located behind the air intake cowl along the longitudinal length of the nacelle 210. The air intake cowl has an inner barrel that defines an air intake duct for directing air to the rotor assembly 219. The nacelle 210 may have an exhaust nozzle 212 (e.g., a main exhaust nozzle and a fan nozzle) located at the rear end of the propulsion system 208. In one example, each propulsion system 208 may include or represent a gas turbine engine. The rotor assembly 219 may be a portion of an engine. The engine burns fuel, such as gasoline, kerosene, biofuel, or other fuel source, to generate thrust for propelling the aircraft 200.

[0047] Figure 6 An embodiment of a nacelle 210 of one of the propulsion systems 208 of an aircraft 200 according to an example is shown. The nacelle 210 extends a length from a front end 220 to a rear end 222 (opposite to the front end 220). The nacelle 210 may include an air intake cover 224, a fan cover 226 disposed at a rear portion of the air intake cover 224, and at least one tail section 228 disposed at a rear portion of the fan cover 226. The air intake cover 224 defines a leading edge 230 of the nacelle 210 at the front end 220 to guide air into a core 232 of the nacelle 210.

[0048] Figure 7 for Figure 6 The air intake hood 224 has an annular barrel shape defining a central opening 234, which is fluidly connected to the core 232 of the engine compartment 210 (at Figure 6230). The term "annular barrel" means that the air intake hood 224 defines a closed annular shape when viewed from the front. The annular barrel is oriented about a central longitudinal axis 236 extending through the central opening 234. The air intake hood 224 can have a generally cylindrical shape. For example, the leading edge 230 can be circular. The air intake hood 224 guides air through the central opening 234 into the core 232.

[0049] The inlet shroud 224 has a leading edge 230, an outer side 238, and an inner side 240. The outer side 238 extends from the leading edge 230 to an outer trailing edge 242. The inner side 240 extends from the leading edge 230 to an inner trailing edge 244. The outer side 238 is located radially outward of the inner side 240 and surrounds the inner side 240. The inner side 240 may define a central opening 234 that operates as an inlet duct to supply air into the core 232 for the rotor assembly 219. The inlet shroud 224 may define a cavity (not shown) located aft of the leading edge 230 and radially disposed between the outer side 238 and the inner side 240.

[0050] In one example, the laminated composite structure 150 is used to define the leading edge 230 of the inlet cowl 224. For example, the laminated composite structure 150 can be assembled and then formed into an annular barrel shape. Figure 1 ) and / or the heating layer 102 may define a surface layer of the air intake hood 224 that is exposed to the surrounding environment. Optionally, Figure 4 The laminated composite structure 150 shown in the figure is one panel of a plurality of different panels that are assembled side by side along the circumference of the air inlet hood 224. For example, a first laminated composite structure 150 can be assembled adjacent to a second laminated composite structure 150, and the second laminated composite structure 150 has the same or similar composition, shape, and size as the first laminated composite structure 150. In another example, the laminated composite structure 150 can be an integral single-piece structure that forms the entire annular leading edge 230 without any seams. In this single-piece embodiment, the laminated composite structure 150 optionally can include multiple electrical loads within the thickness, and the electrical loads can be separated from each other at the seams.

[0051] The laminated composite structure 150 may be part of an electrical ice protection system (EIPS), such as Figure 1 1 and 10, which is used to meet engine ice accumulation and ingestion requirements. The electrical load may be a heater element (e.g., a resistive layer). The heater element converts electrical energy into thermal energy to heat the surface of the inlet shroud 224 along the leading edge 230. In one example, the laminated composite structure 150 includes a plurality of heater elements spaced apart from each other by seams. The heater elements are arranged along the circumference of the inlet shroud 224 so that each heater element defines a corresponding heating area 252 representing a circumferential segment of the leading edge 230.

[0052] Continue to refer Figure 7 , then Figure 8 Shown by Figure 7 251 in the dotted circle 250. In one example, the seam 250 is not oriented parallel to the central longitudinal axis 236 and / or the flow direction of the airflow through the central opening 234. For example, the heater elements may be arranged so that the seam 250 defined between adjacent heater elements can define an inclination angle 256 relative to the central longitudinal axis 236 and the flow direction of the airflow 258 passing through the air intake hood 224. The inclination angle 256 is not parallel or perpendicular. For example, the inclination angle 256 is 1 degree to 89 degrees. In one example, the inclination angle 256 can be 20 degrees to 70 degrees. Optionally, the inclination angle 256 can be 30 degrees to 60 degrees. In one example, the seam 250 is oriented so that the seam 250 defines an inclination angle relative to the entire range 260 of the airflow flow direction that the aircraft 200 can experience in flight. For example, the range 260 can be defined between a maximum rise angle 262 (when the aircraft is ascending) and a maximum descent angle 264 (when the aircraft is descending). Range 260 may be maximum rise angle 262 plus maximum descent angle 264. Thus, regardless of the angle of attack of the aircraft at a given time during typical flight conditions, seam 250 is oriented obliquely relative to the current flow direction of airflow 258.

[0053] Generally speaking, the seams 250 represent areas along the air intake hood 224 where there is a risk of ice accumulation because the seams 250 may have a lower heat flux than the adjacent heating areas 252. Ice accumulation may adversely affect the aerodynamic characteristics of the aircraft 200 and may also potentially damage the rotor assembly 219 or increase wear on the rotor assembly 219. The air intake hood 224 avoids ice accumulation by orienting the heater elements so that the seams 250 between the heating areas 252 are inclined relative to the flow direction of the airflow 258. For example, cold air moving in the direction of the airflow 258 intersects the seams 250 only in a small area before flowing through the heating areas 252. Most of the cold air is heated along the first heating area 252 and then flows through the seams 250 before being heated along the second heating area 252.

[0054] Fig. 9 A transverse cross-sectional view of a portion of an air intake hood 224 of a nacelle 210 according to an example of the present disclosure is shown. As shown, an anti-icing system 100 can be disposed at a leading edge 230 within an inner side 240. A heating layer 102 including perforations 104 covers (e.g., covers) at least a portion of an acoustic layer 110. Thus, the portion of the acoustic layer 110 is sandwiched between an inner structure 270 of the air intake hood 224 and the heating layer 102. The overlapping portion of the heating layer 102 and the acoustic layer 110 reduces the weight, size, and drag of the nacelle 210.

[0055] In addition, the present disclosure includes examples according to the following items:

[0056] Item 1. A system comprising:

[0057] A heating layer comprising perforations extending from a first surface to a second surface opposite the first surface, wherein the perforations provide an acoustic path, wherein the heating layer is configured to provide resistive heating to the component.

[0058] Item 2. The system of Item 1, further comprising an acoustic layer coupled to the heating layer, wherein the perforations extend to the acoustic layer.

[0059] Item 3. The system of Item 2, wherein the heating layer covers at least a portion of the acoustic layer.

[0060] Item 4. The system of any one of items 1-3, wherein the heating layer is a carbon shield comprising carbon fibers or carbon nanotubes.

[0061] Clause 5. The system of any of clauses 1-4, wherein the component comprises a nacelle of an aircraft engine.

[0062] Item 6. The system of any one of items 1-5, wherein the heating layer comprises:

[0063] Dielectric layer;

[0064] a conductive heater coupled to the dielectric layer; and

[0065] One or more membranes coupled to the conductive heater.

[0066] Item 7. The system of Item 6, wherein the dielectric layer comprises glass fibers.

[0067] Item 8. The system of Item 6 or 7, wherein the one or more films comprise a nickel chemical vapor deposition film.

[0068] Item 9. The system of any of Items 6-8, wherein the perforations extend through the dielectric layer, the conductive heater, and the one or more membranes.

[0069] Item 10. The system of any of items 6-9, wherein the dielectric layer, the conductive heater, and the one or more membranes are laminated together.

[0070] Item 11. An aircraft comprising:

[0071] A propulsion system comprising a nacelle having an air inlet cowl, wherein the air inlet cowl comprises an acoustic layer coupled to an anti-icing system, wherein the anti-icing system comprises:

[0072] A heating layer includes perforations extending from a first surface to a second surface opposite the first surface, wherein the perforations provide an acoustic path, wherein the heating layer is configured to provide resistive heating to the engine compartment.

[0073] Item 12. An aircraft according to Item 11, wherein the perforations extend to the acoustic layer.

[0074] Item 13. The aircraft of Item 12, wherein the heating layer covers at least a portion of the acoustic layer.

[0075] Item 14. The aircraft of any one of Items 11-13, wherein the heating layer is a carbon shroud comprising carbon fibers or carbon nanotubes.

[0076] Item 15. An aircraft according to any one of Items 11 to 14, wherein the heating layer comprises:

[0077] Dielectric layer;

[0078] a conductive heater coupled to the dielectric layer; and

[0079] One or more membranes coupled to the conductive heater.

[0080] Item 16. An aircraft according to Item 15, wherein the dielectric layer comprises glass fibers.

[0081] Item 17. An aircraft according to Item 15 or 16, wherein the one or more films include a nickel chemical vapor deposition film.

[0082] Clause 18. The aircraft of any of Clauses 15-17, wherein the perforations extend through the dielectric layer, the conductive heater, and the one or more membranes.

[0083] Clause 19. The aircraft of any one of clauses 15-18, wherein the dielectric layer, the conductive heater and the one or more films are laminated together.

[0084] Item 20. A method comprising:

[0085] forming a perforation in the heating layer, wherein the perforation extends from a first surface to a second surface opposite the first surface, wherein the perforation provides an acoustic path, and wherein the heating layer is configured to provide resistive heating to the component; and

[0086] The heating layer is coupled to the acoustic layer, wherein the perforations extend to the acoustic layer.

[0087] As described herein, examples of the present disclosure provide an improved anti-icing system for an aircraft engine. Additionally, examples of the present disclosure provide a lighter anti-icing system that improves the operating efficiency of the aircraft engine.

[0088] Although various spatial and directional terms such as top, bottom, lower, middle, side, horizontal, vertical, front, etc. may be used to describe examples of the present disclosure, it should be understood that these terms are used only with respect to the orientation shown in the drawings. The orientation may be reversed, rotated, or otherwise changed so that top is lower and vice versa, horizontal becomes vertical, etc.

[0089] As used herein, a structure, constraint, or element "configured to" perform a task or operation is specifically structurally formed, constructed, or adapted in a manner corresponding to the task or operation. For clarity and avoidance of doubt, an object that can merely be modified to perform a task or operation is not "configured to" perform a task or operation as used herein.

[0090] It should be understood that the above description is intended to be illustrative rather than restrictive. For example, the above embodiments (and / or aspects thereof) may be used in combination with each other. In addition, without departing from the scope of the present disclosure, many modifications may be made to adapt specific situations or materials to the teachings of the various embodiments of the present disclosure. Although the sizes and types of materials described herein are intended to define the parameters of the various embodiments of the present disclosure, these embodiments are by no means restrictive, but exemplary embodiments. After reviewing the above description, many other embodiments will be apparent to those skilled in the art. Therefore, the scope of the different embodiments of the present disclosure should be determined by reference to the appended claims and the full range of equivalents granted by these claims. In the appended claims, the terms "including" and "wherein" are used as the plain English equivalents of the corresponding terms "including" and "wherein". In addition, the terms "first", "second", and "third", etc. are used only as markers and are not intended to impose numerical requirements on their objects. In addition, the limitations of the appended claims are not written in the format of means plus function, unless the phrase "means for..." is explicitly used in such claim limitations, followed by the function of avoiding other structures.

[0091] The description herein uses examples to disclose various embodiments of the present disclosure, including the best mode, and also enables those skilled in the art to practice various embodiments of the present disclosure, including making and using any device or system and performing any incorporated method. The scope of protection of the various embodiments of the present disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. If some examples have structural elements that are not different from the literal language of the claims, or if the examples include equivalent structural elements that are not substantially different from the literal language of the claims, these other examples are intended to fall within the scope of the claims.

Claims

1. An anti-icing system for an aircraft, comprising: A heating layer comprising perforations extending from a first surface to a second surface opposite the first surface, wherein the perforations provide an acoustic path, wherein the heating layer is configured to provide resistive heating to a component.

2. The system of claim 1, further comprising an acoustic layer coupled to the heating layer, wherein the perforations extend to the acoustic layer.

3. The system of claim 2, wherein the heating layer covers at least a portion of the acoustic layer.

4. The system of claim 1, wherein the heating layer is a carbon shield comprising carbon fibers or carbon nanotubes.

5. The system of claim 1, wherein the component comprises a nacelle of an aircraft engine.

6. The system of claim 1, wherein the heating layer comprises: Dielectric layer; a conductive heater coupled to the dielectric layer; as well as One or more films coupled to the conductive heater. The system of claim 6 , wherein the dielectric layer comprises fiberglass.

8. The system of claim 6, wherein the one or more films include a nickel chemical vapor deposition film.

9. The system of claim 6, wherein the perforations extend through the dielectric layer, the conductive heater, and the one or more membranes.

10. The system of claim 6, wherein the dielectric layer, the conductive heater, and the one or more membranes are laminated together.

11. An aircraft comprising: A propulsion system comprising a nacelle having an air inlet cowl, wherein the air inlet cowl comprises an acoustic layer coupled to an ice protection system, wherein the ice protection system comprises: A heating layer includes perforations extending from a first surface to a second surface opposite the first surface, wherein the perforations provide an acoustic path, wherein the heating layer is configured to provide resistive heating to an engine compartment.

12. The aircraft of claim 11, wherein the perforations extend to the acoustic layer.

13. The aircraft of claim 12, wherein the heating layer covers at least a portion of the acoustic layer.

14. The aircraft of claim 11, wherein the heating layer is a carbon shroud comprising carbon fibers or carbon nanotubes.

15. The aircraft of claim 11, wherein the heating layer comprises: Dielectric layer; a conductive heater coupled to the dielectric layer; as well as One or more films coupled to the conductive heater.

16. The aircraft of claim 15, wherein the dielectric layer comprises fiberglass.

17. The aircraft of claim 15, wherein the one or more films include a nickel chemical vapor deposition film.

18. The aircraft of claim 15, wherein the perforations extend through the dielectric layer, the conductive heater, and the one or more membranes.

19. The aircraft of claim 15, wherein the dielectric layer, the conductive heater, and the one or more films are laminated together.

20. A method for manufacturing an ice protection system, comprising: forming a perforation in the heating layer, wherein the perforation extends from a first surface to a second surface opposite the first surface, wherein the perforation provides an acoustic path, and wherein the heating layer is configured to provide resistive heating to a component; as well as The heating layer is coupled to an acoustic layer, wherein the perforations extend to the acoustic layer.

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