Anti-icing system for an aircraft

Abstract

An anti-icing system for an exterior surface of an aircraft. The exterior surface is configured to have an airflow over the exterior surface and to have a plurality of zones. At least one heat source is thermally coupled to the exterior surface in each of the plurality of zones. A controller is configured to selectively control the at least one heat source in each of the plurality of zones based on operating conditions related to the airflow over the exterior surface.

Description

TECHNICAL FIELD

[0001] The present disclosure relates to ice protection systems, and in particular to ice protection systems for aircraft. BACKGROUND

[0002] Ice buildup on aircraft wings, propellers, engine inlets, and the like can present problems for the aircraft. Ice buildup can add considerable weight and change the airfoil or inlet configuration, affecting the controlled airflow over these surfaces and making the aircraft more difficult to fly. In the case of jet aircraft, ice chunks that fall off the leading edge of the engine inlet casing can damage the rotating fan and turbine blades or other internal engine components. BRIEF DESCRIPTION OF DRAWINGS

[0003] The features and advantages of the present disclosure will be apparent from the following description of various exemplary embodiments, as illustrated in the accompanying drawings, in which like reference numerals generally refer to like, functionally similar, and / or structurally similar elements.

[0004] Figure 1 is a schematic perspective view of an aircraft that can use the ice protection system of the present disclosure.

[0005] Figure 2 is a schematic cross-sectional view of one engine of the aircraft shown in Figure 1 along line 2-2 in Figure 1 .

[0006] Figure 3 shows an ice protection system according to an embodiment of the present disclosure. Figure 3 is a detail view of the nacelle showing detail 3 of Figure 2 .

[0007] Figure 4 shows another ice protection system according to an embodiment of the present disclosure. Figure 4 is a detail view of the nacelle showing detail 3 of Figure 2 .

[0008] Figure 5 shows another valve arrangement of the ice protection system shown in Figure 4 .

[0009] Figure 6 shows another configuration of the ice protection system shown in Figure 4 .

[0010] Figure 7A , 7B and 7C are perspective views of a lip of the nacelle showing different operating configurations of the heaters in the nacelle. Figure 7A shows an operating configuration of the heaters in which the engine is operating at high power conditions of takeoff or climb. Figure 7BAn operational configuration of the heater is shown where the engine is operating in a partial power condition. Figure 7C An operational configuration of the heater is shown where the engine is operating in a low power (idle) condition for descent.

[0011] Figure 8A and 8B is an isometric view of the lip of the nacelle showing different operational configurations of the heater in the nacelle. Figure 8A An operational configuration of the heater is shown where the engine is operating in a partial power condition and the aircraft (engine) has a high angle of attack. Figure 8B An operational configuration of the heater is shown where the engine is operating in a partial power condition and the engine is exposed to a right side slip condition.

[0012] Figure 9 is a schematic flow diagram of a method of controlling the heater.

[0013] Figure 10 is a schematic flow diagram of another method of controlling the heater. DETAILED DESCRIPTION

[0014] The features, advantages, and embodiments of the present disclosure are set forth with particularity in the detailed description that follows, accompanied by the accompanying drawings and the claims. Furthermore, it should be understood that the detailed description is intended for purposes of illustration only and is not intended to limit the scope of the present disclosure.

[0015] Various embodiments are discussed in detail below. While specific implementations are discussed, this is simply for illustration. One skilled in the relevant art will recognize from the following discussion that other components and configurations can be used without departing from the spirit and scope of the present disclosure.

[0016] Aircraft engine nacelles are subjected to icing conditions, particularly the nacelle leading edge at the inlet lip, when the engine is on the ground, particularly in flight conditions. Anti-icing systems that can be suitably used on the inlet lip of a nacelle or other suitable aircraft surface are discussed herein. These anti-icing systems can be used to remove ice accumulation (de-icing) and prevent ice accumulation (anti-icing). In some embodiments, the anti-icing systems include resistive heating elements. Such resistive heating elements can consume a considerable amount of power, and the high duty cycle of these heating elements can reduce the life (on-wing time) of the heating elements. The anti-icing systems discussed herein control the heating elements based on the conditions under which the engine and / or aircraft is operating (current operating conditions). In certain conditions, a surface, such as a portion of the lip of an engine nacelle, can be subjected to icing conditions, but in other conditions, the surface is less likely to be subjected to icing. By controlling the anti-icing system based on the current operating conditions, the heating elements can be operated with a targeted approach such that the duty cycle of the heating elements for a particular surface is reduced when the surface is less likely to be iced compared to conditions in which the surface is more likely to be iced. By using this approach, the life (on-wing time) of the heating elements can be increased, and the power consumption of the anti-icing system can be reduced.

[0017] The anti-icing systems discussed herein are suitable for use on an aircraft. Figure 1 is a perspective view of an aircraft 10 in which various preferred embodiments can be implemented. The aircraft 10 includes a fuselage 12, a pair of wings 14 attached to the fuselage 12, and a tail 16. The aircraft 10 also includes a propulsion system that generates propulsive thrust needed to propel the aircraft 10 during flight, taxi operations, and the like. Figure 1 The propulsion system of the aircraft 10 shown in Figure 1 is shown attached to the wings 14 in a wing- under configuration, in other embodiments, the engines 100 can have alternative configurations and be coupled to other portions of the aircraft 10. For example, the engines 100 can additionally or alternatively include one or more aspects coupled to other portions of the aircraft 10, such as the tail 16 (as shown in Figure 3 is shown attached to the wings 14 in a wing- under configuration, in other embodiments, the engines 100 can have alternative configurations and be coupled to other portions of the aircraft 10. For example, the engines 100 can additionally or alternatively include one or more aspects coupled to other portions of the aircraft 10, such as the tail 16 (as shown in Figure 1 The aircraft 10 shown in

[0018] As will be described further below with reference to Figure 2 the engines 100 shown in Figure 1 are gas turbine engines that are each capable of selectively generating propulsive thrust for the aircraft 10. The amount of propulsive thrust can be based at least in part on the amount of fuel provided to the engines 100 via a fuel system 130 (see Figure 2The amount of fuel provided to the engine 100 is controlled. The aviation turbine fuel in the embodiments discussed herein is a combustible hydrocarbon liquid fuel having a desired carbon number, such as a kerosene-type fuel. The fuel is stored in a fuel tank 131 of a fuel system 130. As shown in Figure 1 at least a portion of the fuel tank 131 is located in each wing 14 and a portion of the fuel tank 131 is located in the fuselage 12 between the wings 14. However, the fuel tank 131 can be located in other suitable locations in the fuselage 12 or wings 14. The fuel tank 131 can also be entirely located within the fuselage 12 or wings 14. The fuel tank 131 can also be separate tanks rather than a single unit, for example, two tanks, each located within a corresponding wing 14.

[0019] Figure 2 is a schematic cross-sectional view of one of the engines 100 of the propulsion system of the aircraft 10 shown in Figure 1 Figure 2 is a cross-sectional view taken along line 2-2 in Figure 1 Figure 2 For the embodiment shown in Figure 2 , the engine 100 is a high-bypass turbofan engine that includes a fan section 102 and a turbine machine 104 disposed downstream from the fan section 102. The engine 100 has an axial direction A (extending parallel to a longitudinal centerline 101, shown in Figure 2 for reference), a radial direction R, and a circumferential direction. The circumferential direction (not shown in

[0020] Figure 2 The turbine machine 104 shown in includes a tubular outer casing 106 (also referred to as an outer case or nacelle) that defines an inlet 108. In this embodiment, the inlet 108 is annular. The casing 106 encloses an engine core that includes, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor 110 and a high pressure (HP) compressor 112, a combustion section 114, a turbine section including a high pressure (HP) turbine 116 and a low pressure (LP) turbine 118, and an ejection exhaust nozzle section 120. The compressor section, the combustion section 114, and the turbine section together at least partially define a core air flowpath 121 that extends from the inlet 108 to the ejection exhaust nozzle section 120. The turbine machine 104 further includes one or more drive shafts. More specifically, the turbine machine 104 includes a high pressure (HP) shaft or spool 122 that drivingly connects the HP turbine 116 to the HP compressor 112, and a low pressure (LP) shaft or spool 124 that drivingly connects the LP turbine 118 to the LP compressor 110.

[0021] ​The engine 100, and more specifically the turbine 104, can operate with and receive a flow of fuel from the fuel system 130. The fuel system 130 includes a fuel delivery assembly 133 that provides a flow of fuel from a fuel tank 131 to the engine 100, and more specifically, to a plurality of fuel nozzles 142 that inject fuel into the combustion chambers of the combustors 140 of the combustion section 114. The fuel delivery assembly 133 includes tubes, pipes, conduits, and the like to fluidly connect the various components of the fuel system 130 to the engine 100. The fuel tank 131 is configured to store a hydrocarbon fuel, and the hydrocarbon fuel is supplied from the fuel tank 131 to the fuel delivery assembly 133. The fuel delivery assembly 133 is configured to deliver the hydrocarbon fuel between the fuel tank 131 and the engine 100, and thus provide a flow path (fluid passageway) for the hydrocarbon fuel from the fuel tank 131 to the engine 100.

[0022] The fuel system 130 includes at least one fuel pump in fluid connection with the fuel delivery assembly 133 to direct fuel through the fuel delivery assembly 132 to the engine 100. One such pump is a main fuel pump 135. The main fuel pump 135 is a high pressure pump that is the primary source of pressure rise in the fuel delivery assembly 133 between the fuel tank 131 and the engine 100. The main fuel pump 135 can be configured to increase the pressure in the fuel delivery assembly 133 to a pressure that is greater than the pressure within the combustion chambers of the combustors 140.

[0023] The fuel system 130 also includes a fuel metering unit 137 in fluid communication with the fuel delivery assembly 133. Any fuel metering unit 137 can be used, including for example, a metering valve. The fuel metering unit 137 is positioned downstream of the main fuel pump 135 and upstream of a fuel manifold 139 that is configured to distribute fuel to the fuel nozzles 142. The fuel system 130 is configured to provide fuel to the fuel metering unit 137, and the fuel metering unit 137 is configured to receive fuel from the fuel tank 131. The fuel metering unit 137 is further configured to provide a flow of fuel to the engine 100 in a desired manner. More specifically, the fuel metering unit 137 is configured to meter fuel and provide a desired volume of fuel to the fuel manifold 139 of the engine 100 at, for example, a desired flow rate. The fuel manifold 139 is fluidly connected to the fuel nozzles 142 and distributes (provides) the received fuel into the plurality of fuel nozzles 142, where the fuel is injected into the combustion chambers and combusted. Adjusting the fuel metering unit 137 changes the volume of fuel provided to the combustion chambers, and thus the amount of propulsive thrust generated by the engine 100 to propel the aircraft 10.

[0024] The engine 100 also includes various auxiliary systems to aid in the operation of the engine 100 and / or the aircraft 10. For example, the engine 100 can include a main lubrication system 152, a compressor cooling air (CCA) system 154, an active thermal clearance control (ATCC) system 156, and a generator lubrication system 158, each of which is schematically shown in Figure 2 The main lubrication system 152 is configured to provide lubricant to various bearings and gear meshes in, for example, the compressor section, the turbine section, the HP spool 122, and the LP spool 124. The lubricant provided by the main lubrication system 152 can increase the useful life of these components, and can remove an amount of heat from these components through the use of one or more heat exchangers. The compressor cooling air (CCA) system 154 provides air from one or both of the HP compressor 112 or the LP compressor 110 to one or both of the HP turbine 116 or the LP turbine 118. The active thermal clearance control (ATCC) system 156 is used to minimize the clearance between the tips of the turbine blades and the casing wall during a flight mission as the casing temperature changes. The generator lubrication system 158 provides lubrication for an electronic generator (not shown), as well as cooling / heat dissipation for the electronic generator. The electronic generator can provide power to, for example, a starter motor of the engine 100 and / or various other electronic components of the engine 100 and / or the aircraft 10. The lubrication systems for the engine 100 (e.g., the main lubrication system 152 and the generator lubrication system 158) can use hydrocarbon fluids (e.g., oil) for lubrication, where the oil is circulated through the inner surfaces of the oil return tubes.

[0025] The engine 100 also includes an engine controller 170 configured to operate various systems of the engine 100, including, for example, the ice protection system 260 discussed below. In this embodiment, the controller 170 is a computing device having one or more processors 172 and one or more memories 174. The processor 172 can be any suitable processing device, including but not limited to a microprocessor, a microcontroller, an integrated circuit, a logic device, a programmable logic controller (PLC), an application-specific integrated circuit (ASIC), and / or a field programmable gate array (FPGA). The memory 174 can include one or more computer-readable media, including but not limited to non-transitory computer-readable media, computer-readable non-volatile media (e.g., flash memory), RAM, ROM, hard drives, flash drives, and / or other storage devices.

[0026] Memory 174 can store information accessible by processor 172, including computer-readable instructions that can be executed by processor 172. The instructions can be any set of instructions that, when executed by processor 172, cause processor 172 and controller 170 to perform operations. In some embodiments, the instructions can be executed by processor 172 to cause processor 172 to complete any operations and functions for which controller 170 is configured, which will be further described below. The instructions can be software written in any suitable programming language or can be implemented in hardware. Additionally, and / or alternatively, the instructions can be executed in logically and / or virtually separate threads on processor 172. Memory 174 can further store data that can be accessed by processor 172.

[0027] The technology discussed herein makes reference to computer-based systems, actions performed by the computer-based systems, and information sent and received by the computer-based systems. A person having ordinary skill in the art will recognize that the inherent flexibility of computer-based systems allows for a great variety of possible configurations, combinations, and divisions of tasks and functionality between and among components. For instance, processes discussed herein can be implemented using a single computing device or multiple computing devices working in combination. Databases, memories, programs, and applications can be implemented on a single system or distributed across multiple systems. Distributed components can operate sequentially or in parallel.

[0028] Figure 2 The fan section 102 shown in FIG. 1 includes a fan 160 having a plurality of fan blades 162 coupled to a disk 164. The fan blades 162 and disk 164 are rotatable together about an axis of rotation 166 by the LP shaft 124. In this embodiment, the axis of rotation 166 coincides with the longitudinal centerline (axis) 101 of the turbomachine 104. The disk 164 is covered by a rotatable front hub 168 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 162. Further, an annular fan casing or nacelle 200 is provided circumferentially around at least a portion of the fan 160 and / or turbomachine 104. The nacelle 200 will be referred to herein as the nacelle 200. The nacelle 200 is annular and defines an inlet 182 of the fan section 102. Although the nacelle 200 can be symmetrical, the nacelle 200 and inlet 182 can be asymmetrical, for example having asymmetries between the top and bottom, as well as between the left and right sides. The nacelle 200 is supported relative to the turbomachine 104 by a plurality of circumferentially spaced outlet guide vanes 184. A downstream section 202 of the nacelle 200 extends over an exterior of the turbomachine 104 so as to define a bypass airflow passage 186 therebetween.

[0029] Air flows from Figure 2 the left side of Figure 2right side of the core air flowpath 121 and into the inlet 182. A portion of the airflow can flow through the bypass airflow passage 186 past the fan blades 162 and the outlet guide wheel vanes 184. As discussed above, a portion of the airflow can enter the outer casing 106 through the annular inlet 108 as air to flow through the core air flowpath 121 to mix with fuel for combustion in the combustor 140 and exit through the injection exhaust nozzle section 120. The nacelle 200 helps to direct the airflow into the fan blades 162 of the fan 160.

[0030] Figure 3 is a detail view of the nacelle 200 showing Figure 2 Detail 3 of the embodiment. The nacelle 200 of this embodiment includes an inner barrel 210 having an inner barrel surface 212. The inner barrel 210 can include acoustic panels 216. The nacelle 200 also includes an outer barrel 220 having an outer barrel surface 222. Both the inner barrel 210 and the outer barrel 220 are annular, with the outer barrel 220 positioned radially outward in the radial direction R (see Figure 2 ). The outer barrel surface 222 encloses the inner barrel 210 and forms a barrel cavity 232 between the inner barrel 210 and the outer barrel 220. The inner barrel 210 and the outer barrel 220 are connected to each other by a plurality of bulkheads, such as an inlet front bulkhead 234 and an inlet rear bulkhead 236. The inlet front bulkhead 234 and the inlet rear bulkhead 236 provide structural support for the nacelle 200, particularly for the inner barrel 210 and the outer barrel 220.

[0031] The nacelle 200 includes a lip 240 formed at the front end of the nacelle 200. The lip 240 is attached to the front edge 214 of the inner barrel 210 and the front edge 224 of the outer barrel 220. The lip 240 has an aerodynamic profile to facilitate airflow into the inlet 182 of the fan section 102 and airflow over the outer barrel surface 222. In this embodiment, the lip 240 has a U-shape with a front (or leading) portion 242, an inner portion 244, and an outer portion 246. The lip 240 defines a lip cavity 248, which can be referred to as a D-shaped duct. In this embodiment, the lip cavity 248 is annular. The lip 240 also includes an outer surface 252 and an inner surface 254. Air flows over the outer surface 252, with the inner surface 254 facing the lip cavity 248.

[0032] The nacelle 200 includes an ice protection system 260. The ice protection system 260 of this embodiment is configured to selectively heat portions of the lip 240 and prevent ice from forming thereon or to de-ice the lip 240 if ice has already formed thereon. Heat is selectively applied to the regions (or zones) using heat from a heat source. Any suitable heat source can be used. In Figure 3In the illustrated embodiment, the heat sources are heating elements, and each zone (or region) includes at least one heating element 262 to selectively heat the region. The heating elements 262 of this embodiment are resistive heating elements that provide heat when an electrical current is provided thereto. The heat generated by the heating elements 262 can be controlled by controlling the electrical current provided to the heating elements 262. Any suitable heating elements 262 can be used, including, for example, graphite-based resistive elements or metal mesh heating elements. The heating elements 262 are thermally coupled to the lip 240 such that they heat the corresponding outer surface 252 of the region. The heating elements 262 can be attached to the inner surface 254 of the lip 240 within the lip cavity 248 using, for example, a thermally conductive adhesive or fasteners. In other embodiments, the heating elements 262 can be integrally formed with the lip 240, for example, embedded within the lip 240.

[0033] The lip 240 is divided into a plurality of regions, and heat can be selectively applied to each region to heat the outer surface 252 of each of the regions. The regions can also be referred to as zones. Figure 3 The example lip 240 illustrated has three zones, each corresponding to the leading portion 242, the inner portion 244, and the outer portion 246. In other embodiments, the lip 240 can have more or fewer zones, and each zone can correspond to a different portion of the lip 240. Figure 3 In the illustrated embodiment, one or more heating elements 262 are attached to each of the zones and are configured to selectively apply heat to the zones. In this embodiment, the heating elements 262 are operated by a controller, such as the engine controller 170 (see also FIG. 1). Other suitable controllers can be used, including, for example, a dedicated controller or a controller that is part of the flight control system (flight controller) of the aircraft 10. Figure 2 ) In other embodiments, the heating elements 262 can be integrally formed with the lip 240, for example, embedded within the lip 240.

[0034] The controller 170 is operatively and communicatively coupled to a power distribution module 264 that receives power from a power source 266. The power source 266 can be any suitable power source, including, for example, a power source of the engine 100, such as an electric generator driven by rotation of one of the shafts (e.g., the HP shaft 122 or the LP shaft 124) of the engine 100. The power distribution module 264 is configured to selectively control application of power to each of the heating elements 262 based on instructions (signals) provided by the controller 170, such as by providing power to or disconnecting power from the heating elements 262. Thus, the controller 170 is able to selectively control operation of the heating elements 262 via the power distribution module 264. The controller 170 is configured to selectively control at least one heating element in each of the plurality of zones to one of a plurality of heating levels. In some embodiments, the plurality of heating levels can include a high heating level and a reduced heating level. The amount of heat generated over a set time interval for one of the plurality of zones is less when the zone is operated at the reduced heating level than when the zone is operated at the high heating level. As will be discussed further below, the controller 170 can selectively control the at least one heating element to one of the plurality of heating levels by controlling a duty cycle of the at least one heating element 262, where, for example, the duty cycle of the at least one heating element in the reduced heating level is less than the duty cycle of the at least one heating element in the high heating level.

[0035] Air in contact with the outer surface 252 of the lip 240 will have a stagnation point. In icing conditions, ice will tend to form at the stagnation point, and thus heat should be applied to the outer surface 252 at which the stagnation point is located to prevent ice formation. However, the stagnation point will change based on operating conditions of the aircraft 10, and more particularly, operating conditions of the engine 100. Under different operating conditions, the stagnation point can be located on the outer surface 252 of any of the leading portion 242, the inner portion 244, or the outer portion 246. Rather than heating the entire lip 240, the controller 170 is configured to selectively heat only a portion of the lip 240. More particularly, in this embodiment, the controller 170 is configured to determine a location of the stagnation point based on received inputs (as will be discussed further below with reference to FIG. 4), and selectively heat certain zones of the lip 240. In some embodiments, the controller 170 will apply heat to all of the zones of the lip 240, but some of the zones will have a reduced duty cycle and will be heated less than other zones. Figure 9 and Figure 10 Further discussion), and selectively heat certain zones of the lip 240. In some embodiments, the controller 170 will apply heat to all of the zones of the lip 240, but some of the zones will have a reduced duty cycle and will be heated less than other zones.

[0036] Figures 4-6 Hot air 304 is shown being used as a heat source for the lip 240, in addition to or in lieu of the heating elements 262 described above. Figures 4-6 is a detail view of the nacelle 200 showing Figure 2Detail 3. As mentioned above, any suitable heat source can be used, and Figures 4-6 In the illustrated embodiment, pneumatic heating (pneumatic heating system 300) is used in the anti-icing system 260. Hot air 304 is supplied from a hot air source 306 via a hot air supply duct 308. The hot air source 306 can be any suitable hot air source, including, for example, a source of hot air from the engine 100. Such engine hot air sources include, for example, compressor exhaust air or other exhaust air, such as fan exhaust air or precooler exhaust air. These sources can include air heated by other engine heat sources, including, for example, the main lubrication system 152, the active thermal gap control (ATCC) system 156, and the generator lubrication system 158.

[0037] exist Figure 4 In the illustrated embodiment, the lip cavity 248 is divided into multiple cavities. At least one of the multiple cavities corresponds to each region. The lip cavity 248 may include, for example, a leading cavity 312 corresponding to the leading portion 242, an inner cavity 314 corresponding to the inner portion 244, and an outer cavity 316 corresponding to the outer portion 246. These cavities are spaced apart from each other by cavity walls 318. A hot air supply conduit 308 is fluidly connected to each of the leading cavity 312, the inner cavity 314, and the outer cavity 316 via a corresponding cavity conduit 322. Each of the leading cavity 312, the inner cavity 314, and the outer cavity 316 also includes a hot air distribution manifold 324, which is fluidly connected to the cavity conduit 322 and configured to distribute hot air 304 into the corresponding cavity 312, 314, 316. Any suitable hot air distribution manifold 324 may be used, including, for example, a piccolo tube or a swirl nozzle. The type of hot air distribution manifold 324 can be the same or different among the lead cavity 312, inner cavity 314, and outer cavity 316. When a piccolo is used as the hot air distribution manifold 324, the hot air distribution manifold 324 can be a 360-degree piccolo, but the hot air distribution manifold 324 can also be only on a portion of the circumferential direction, for example when the lip 240 is divided into multiple zones in the circumferential direction.

[0038] At least one valve 330 is used to selectively supply hot air to each of the leading cavity 312, the inner cavity 314, and the outer cavity 316, thereby selectively heating each of the leading portion 242, the inner portion 244, and the outer portion 246. Figure 4In the illustrated embodiment, the valve 330 is a pneumatic on / off valve 332 configured to selectively provide hot air or stop air flow to each of the chamber conduits 322. The pneumatic on / off valve 332 is shown positioned in the barrel chamber 232, but it can be located in other locations, including more upstream in the hot air supply conduit 308 and closer to the hot air source 306. The pneumatic on / off valve 332 provides hot air 304 to only one of the front lead chamber 312, the inner chamber 314, and the outer chamber 316 at a time. The controller 170 is operably and communicably coupled to the pneumatic on / off valve 332 in order to provide hot air 304 to the front lead chamber 312, the inner chamber 314, and the outer chamber 316 for a period of time in the same manner as the controller 170 controls the heating elements 262 discussed herein.

[0039] Figure 5 Another valve arrangement for the pneumatic heating system 301 is shown. Figure 5 The illustrated pneumatic heating system 301 is the same as the Figure 4 air heating system 300 shown, but instead of using a single valve 330, multiple valves 330 are used. One valve 330 is located in each of the chamber conduits 322 to control the flow of hot air 304 in each of the chamber conduits 322, and thus control the flow of each of the hot air distribution manifolds 324 into each of the front lead chamber 312, the inner chamber 314, and the outer chamber 316. These valves 330 can be on / off valves that are operably and communicably coupled to the controller 170 and controlled based on a duty cycle in the same manner as the heating elements 262 discussed herein. Other suitable valves can be used, including, for example, flow control valves that can be throttled so that each of the chamber conduits 322 can receive different amounts of hot air 304. Alternatively, the valves 330 can be part of one valve with multiple discharge ports. The controller 170 can be configured to reduce the amount of heat in one of the front lead chamber 312, the inner chamber 314, and the outer chamber 316, and thus the amount of heat in the front lead portion 242, the inner portion 244, and the outer portion 246, respectively. The amount of heat can be reduced by the controller 170 in the manner discussed herein for the heating elements 262.

[0040] Figure 6 Another configuration of the pneumatic heating system 302 is shown. In Figure 4 and Figure 5 the lip chamber 248 is divided into multiple chambers by the chamber wall 318. However, the chamber wall 318 can be omitted and the hot air distribution manifold 324 is configured to direct air to a portion of the inner surface 254 of the lip 240 to selectively heat the corresponding area, as shown in Figure 6 In this embodiment, the hot air distribution manifold 324 can preferably be a short flute tube. Although shown using a pneumatic on / off valve 332, the configuration with a valve 330 in each of the chamber conduits 322 can also be used with Figure 6The configurations shown are used together.

[0041] Figure 7A 、 7B and 7C is a perspective view of the lip 240 showing different operational configurations of the plurality of zones, such as a leading portion 242, an inner portion 244, and an outer portion 246. The engine 100 can operate in a plurality of power conditions. As used herein, a “power condition” can refer to an amount of propulsive thrust produced by the engine 100. The plurality of power conditions can include a high power condition, a partial power condition, and a low power condition. In the high power condition, the engine 100 produces a greater amount of propulsive thrust than each of the partial power condition and the low power condition. In the partial power condition, the engine 100 produces a greater amount of propulsive thrust than the low power condition, but the engine 100 produces less propulsive thrust in the partial power condition than in the high power condition. The engine 100 produces less propulsive thrust in the low power condition than in either of the high power condition and the partial power condition. In some embodiments, the low power condition is an idle condition of the engine 100 or a minimum throttle setting for a particular operation of the aircraft 10. Such idle conditions can include, for example, a ground idle condition, a flight idle condition, and / or an approach idle condition.

[0042] Figure 7A shows an operational configuration in which the engine 100 is operating in the high power condition and at a high angle of attack for takeoff or climb. When the engine 100 is operating in such conditions, the engine 100 draws a large amount of air into the inlet 182 of the nacelle 200. In such conditions, the stagnation point can be located on the outer portion 246. The controller 170 controls the heat sources (e.g., heating elements 262 in Figure 3 ) of the outer portion 246 to have a high duty cycle. In some embodiments, this duty cycle can be a full power condition. In such conditions, water located on the outer portion 246 can flow back into the inner barrel 210, and thus, the controller 170 also controls the heat sources (e.g., heating elements 262 in Figure 3 ) of the leading portion 242 and the inner portion 244 to have a high duty cycle. Thus, there is no reduction in heater power in such conditions.

[0043] Figure 7B shows an operational configuration of the heat sources (e.g., heating elements 262 in Figure 3 ) in which the engine 100 is operating in the partial power condition. The partial power condition can be, for example, a power condition for cruising while the aircraft 10 is in a horizontal position. In such conditions, the stagnation point can be located on the leading portion 242, sometimes referred to as a high light. The controller 170 controls the heat sources (e.g., heating elements 262 in Figure 3The heating element 262 in the outer portion 246 has a high duty cycle. Water on the outer portion 246 is unlikely to flow into the inner cylinder 210, and the heat source of the outer portion 246 (e.g., Figure 3 The duty cycle of the heating element 262 in the middle can be relative to the heat source of the leading part 242 (e.g., Figure 3 The duty cycle of the heating element 262 in the middle is reduced. Therefore, the controller 170 reduces the duty cycle of the heat source of the external part 246 (e.g., Figure 3 The heating element 262 in the middle operates to a heat source smaller than the leading portion 242 (e.g., Figure 3 The duty cycle of the heating element 262 in the middle. In this embodiment, the controller 170 also controls the heat source of the internal part 244 (e.g., Figure 3 The heating element 262 in the middle has a heat source (e.g., with a larger heat source than the outer portion 246) Figure 3 The heating element 262 in the middle has a high duty cycle. Figures 7A-8B In the diagram, a high duty cycle is represented by a dark (jumping) dot, while a reduced duty cycle is represented by a bright dot.

[0044] Figure 7C A heat source (e.g., is shown) Figure 3 The operating configuration of the heating element 262 in the engine 100 is such that the engine 100 operates under low-power conditions. Low-power conditions can be, for example, idling conditions of the engine 100 used for descent of the aircraft 10. Under these conditions, the stall point can be located on the internal portion 244. The controller 170 controls the heat source of the internal portion 244 (e.g., Figure 3 The heating element 262 in the inner part has a high duty cycle. Only water on the inner part 244 can flow into the inner cylinder 210, and relative to the heat source of the inner part 244 (e.g., Figure 3 The duty cycle of the heating element 262 in the middle, the heat source of both the leading part 242 and the outer part 246 (e.g., Figure 3 The duty cycle of the heating element 262 in the middle can be reduced. Therefore, the controller 170 controls the heat source (e.g., in the leading part 242 and the outer part 246) of the heating element 262. Figure 3 The heating element 262 in the middle operates to a heat source smaller than the inner part 244 (e.g., Figure 3 The duty cycle of the heating element 262 in the middle.

[0045] Figure 8A and 8B Also a three-dimensional view of the lip margin 240, showing the heat source (e.g., Figure 3 Different operating configurations of the heating element 262 in the middle. Figures 7A-7C In this context, the power of engine 100 is a heat source (e.g., Figure 3The stagnation point is primarily driven by the operational configuration of the heating elements 262 in the lip 240, but the direction of airflow into the nacelle 200 can also affect the stagnation point. When the airflow into the nacelle 200 is at an angle relative to the longitudinal centerline 101 of the engine 100, the stagnation point can be asymmetrically located on the lip 240 of the nacelle 200. Figure 8A and 8B Such conditions are illustrated. To address this asymmetric airflow, the lip 240 of the nacelle 200 is also divided into multiple regions (or zones) in the circumferential direction. In this embodiment, the lip 240 has four circumferential zones corresponding to the four quadrants of the lip 240 of the nacelle 200, but any suitable number of zones can be used. The lip 240 includes a right upper zone 272, a left upper zone 274, a right lower zone 276, and a left lower zone 278, and the controller 170 is configured to selectively control the duty cycle of the heat sources (e.g., heating elements 262) in each of these zones. Figure 3

[0046] In Figure 8A and 8B the engine 100 is operating in a partial power condition, so the stagnation point would be located on the leading portion 242 if the angle of airflow into the nacelle 200 is not considered. Figure 8A conditions in which the aircraft 10 and the engine 100 have a high angle of attack. In this condition, the aircraft 10 is tilted upward such that the longitudinal centerline 101 of the engine 100 is tilted upward. In this condition, the right upper zone 272 and the left upper zone 274 of the leading portion 242 and the outer portion 246 are both areas of low air impingement. Accordingly, the duty cycle of the heat sources (e.g., heating elements 262) in the right upper zone 272 and the left upper zone 274, which are located on both the leading portion 242 and the outer portion 246, is decreased relative to the duty cycle of the heat sources (e.g., heating elements 262) in the other zones and regions. Figure 3 Figure 3

[0047] Figure 8B conditions in which the engine 100 is exposed to a right side slip condition. In this condition, air is angled into the nacelle 200 from the right side, and the left upper zone 274 and the left lower zone 278 are areas of low air impingement. Accordingly, the duty cycle of the heat sources (e.g., heating elements 262) in the left upper zone 274 and the left lower zone 278, which are located on both the leading portion 242 and the outer portion 246, is decreased relative to the duty cycle of the heat sources (e.g., heating elements 262) in the other zones and regions. Figure 3 Figure 3 and the right lower zone 276 of the leading portion 242 and the outer portion 246, are areas of low air impingement, and the duty cycle of the heat sources (e.g., heating elements 262) in the right upper zone 270 and the right lower zone 276, which are located on both the leading portion 242 and the outer portion 246, is decreased relative to the duty cycle of the heat sources (e.g., heating elements 262) in the other zones and regions.​​​​Figure 3 the duty cycle of the heating elements 262 in the regions and zones relative to the heat sources (e.g., Figure 3 the duty cycle of the heating elements 262 in the regions and zones is reduced.

[0048] Figure 9 is a method of controlling the heat sources (e.g., Figure 3 the heating elements 262 in the regions and zones in the manner described above. The anti-icing system 260 is activated in step S305. In some embodiments, this step can be a manual step, such as a pilot of the aircraft 10 selecting an option to activate the anti-icing system 260, but in other embodiments, the controller 170 is configured to receive inputs and determine from those inputs that icing conditions exist. Such inputs include, for example, the airspeed of the aircraft 10, the temperature, the humidity, the pressure, and the altitude of the aircraft 10. The controller 170 can be directly communicatively coupled to sensors to determine these inputs. The controller 170 can also be indirectly coupled to such sensors and receive the inputs from another source, such as a flight controller of the aircraft 10. When the controller 170 determines that icing conditions exist, the controller 170 activates the anti-icing system 260.

[0049] In step S310, the controller 170 determines the operating conditions of the engine 100. The controller 170 is configured to receive inputs and determine from those inputs the conditions of the engine 100. Such inputs include, for example, the airspeed of the aircraft 10, the wind speed and direction (e.g., the crosswind speed), the angle of attack of the aircraft 10, and the mass flow of air through the engine 100. The controller 170 can be directly communicatively coupled to sensors to determine these inputs. The controller 170 can also be indirectly coupled to such sensors and receive the inputs from another source, such as a flight controller of the aircraft 10. In some embodiments, the controller 170 is configured to calculate these parameters. For example, the mass flow through the engine 100 and more specifically into the inlet 182 can not be directly measured. Instead, the mass flow can be calculated using, for example, the airspeed of the aircraft 10.

[0050] In step S315, the controller 170 then selectively operates the heat sources (e.g., Figure 3 the heating elements 262 in the regions and zones) based on the operating conditions of the aircraft 10 and the engine 100, such as in the manner discussed above with respect to Figures 7A-8B When the on and off times (timers) are set by the controller 170 in step S320 based on the operating conditions of the aircraft 10 and the engine 100, the heat sources (e.g., Figure 3 the heating elements 262 in the regions and zones) can be controlled using the on control method. The on and off times determine the duty cycle of the heating elements 262 and thus the temperature of the regions or zones of the lip 240. In other embodiments, other parameters can be used to control the heat sources (e.g.,Figure 3 the heating elements 262) and temperature, e.g., set the resistance of the heating elements 262. The heat source (e.g., the heating elements 262 (heaters)) is then activated in step S325 based on the timer in step S320.

[0051] Figure 10 is another method of controlling the heat source (e.g., the heating elements 262) in the lip 240 of the nacelle 200. Figure 3 is a schematic flowchart of another method of controlling the heat source (e.g., the heating elements 262) in the lip 240 of the nacelle 200. Figure 10 The method shown includes each of the steps S305-S325 described above, but instead of open-loop control, the method includes a feedback loop. Each zone or region can include a temperature sensor configured to measure the temperature of the outer surface 252 (or another suitable surface). In step S330, the temperature sensor measures the temperature of the outer surface 252 of each region. The temperature sensor is communicatively coupled to the controller 170, and in step S335, the controller 170 compares the measured temperature to a target temperature for each region. The target temperature for each region can be determined by the controller 170 in step S315 based on operating conditions of the aircraft 10 and the engine 100. The controller 170 can control the heat source (e.g., the heating elements 262) based on the target temperature by, for example, adjusting the duty cycle (timer). The controller 170 returns to step S320 to adjust the duty cycle (timer) based on the comparison between the measured temperature and the target temperature. For example, the duty cycle can be decreased if the measured temperature is greater than the target temperature, and the duty cycle can be increased if the measured temperature is less than the target temperature. Figure 3

[0052] In the embodiments discussed above, the anti-icing system 260 is applied to the lip 240 of the nacelle 200. However, the anti-icing system 260 can be applied to other suitable outer surfaces of the aircraft and engine, particularly, outer surfaces that are configured to have an airflow over the outer surface when the aircraft is flying in the air, and the anti-icing system 260 can be applied to other surfaces other than the aircraft. Such outer surfaces of the aircraft and engine can include, for example, various inlets and / or airfoils of the aircraft 10 or other surfaces inside the engine 100, such as a booster splitter or an inlet guide vane. Like the lip 240 of the nacelle 200, such surfaces can have a plurality of regions, and the controller 170 can be configured to selectively control at least one heating element 262 in each of the plurality of regions based on operating conditions related to the airflow flowing over the outer surface. Such operating conditions include those that affect the stagnation point of the air over the outer surface, and can include, for example, the angle of attack of the aircraft 10.

[0053] As Figure 1 ​As shown, the leading edge 20 of the wing 14 is another example of an outer surface on which an anti-icing system 260 discussed herein is implemented. The leading edge 20 of the wing 14 can include heating elements 262 and be divided into a plurality of zones in a similar manner to the lip 240 of the nacelle 200 described above. For example, the leading edge 20 can have a U-shape with a forward portion 22, a lower portion 24, and an upper portion 26. The heating elements 262 of the forward portion 22 of the wing 14 can operate in a similar manner to the forward portion 242 of the nacelle 200. The heating elements 262 of the lower portion 24 of the wing 14 can operate in a similar manner to the inner portion 244 of the nacelle 200. The heating elements 262 of the upper portion 26 of the wing 14 can operate in a similar manner to the outer portion 246 of the nacelle 200. In particular, the controller 170 can selectively control the heating elements 262 in each of the forward portion 22, the lower portion 24, and the upper portion 26 based on, for example, an angle of attack of the aircraft 10, and more specifically, an angle of attack of the wing 14. Similar to the wing 14, the anti-icing system 260 can also be implemented on the horizontal and vertical surfaces of the aft portion (tail 16).

[0054] In the embodiments discussed above, each of the lower portion 24 and the upper portion 26 of the wing 14 and the inner portion 244 and the outer portion 246 of the nacelle 200 are downstream of the forward portion 22 of the wing 14 or the forward portion 242 of the nacelle 200, respectively, relative to a direction of airflow over the outer surface. As used herein, the lower portion 24 of the wing 14 and the inner portion 244 of the nacelle 200 can be a first downstream portion, while the upper portion 26 of the wing 14 and the outer portion 246 of the nacelle 200 can be a second downstream portion.

[0055] Further aspects of the present disclosure are provided by the subject matter of the following clauses.

[0056] An anti-icing system for a surface of an aircraft. The anti-icing system includes an outer surface for an aircraft, the outer surface configured to have an airflow over the outer surface. The outer surface has a plurality of zones. At least one heat source is thermally coupled to the outer surface in each of the plurality of zones, a controller configured to selectively control the at least one heat source in each of the plurality of zones based on an operating condition related to the airflow over the outer surface.

[0057] The anti-icing system according to the preceding clause, wherein the heat source is one of hot air or an electric heating element.

[0058] The ice protection system of any preceding clause, wherein the controller is configured to selectively control the at least one heat source to one of a plurality of heating levels in each of the plurality of zones. The plurality of heating levels includes a high heating level and a reduced heating level. An amount of heat generated for one of the plurality of zones over a set time interval is less when the zone is operated at the reduced heating level than when the zone is operated at the high heating level.

[0059] The ice protection system of any preceding clause, wherein the controller selectively controls the at least one heat source to one of the plurality of heating levels by controlling a duty cycle of the at least one heat source. The duty cycle of the at least one heat source in the reduced heating level is less than the duty cycle of the at least one heat source in the high heating level.

[0060] The ice protection system of any preceding clause, wherein each of the plurality of zones includes a temperature sensor configured to measure a temperature of the outer surface. The controller selectively controls the at least one heat source to one of the plurality of heating levels based on a target temperature of the outer surface.

[0061] The ice protection system of any preceding clause, wherein the outer surface has a leading portion, a first downstream portion, and a second downstream portion. The first downstream portion is downstream of the leading portion in an airflow direction from the second downstream portion.

[0062] The ice protection system of any preceding clause, wherein the outer surface has a U-shape. The first downstream portion is an upper surface and the second downstream portion is a lower surface.

[0063] The ice protection system of any preceding clause, wherein the outer surface is a leading edge of a wing of the aircraft. The first downstream portion is an upper surface and the second downstream portion is a lower surface.

[0064] The ice protection system of any preceding clause, wherein the operating condition is an angle of attack of the wing.

[0065] A gas turbine engine comprising: the ice protection system of any preceding clause; and a nacelle defining an inlet. The nacelle includes a lip and the lip is the outer surface.

[0066] The gas turbine engine of the preceding clause, wherein the plurality of zones includes a plurality of regions in a circumferential direction of the nacelle.

[0067] The gas turbine engine according to any preceding clause, wherein the plurality of zones includes at least one upper zone and at least one lower zone. The controller is configured to selectively control the at least one heat source in each of the plurality of zones to one of a plurality of heating levels. The plurality of heating levels includes a high heating level and a reduced heating level. An amount of heat generated for one of the plurality of zones over a set time interval is less when the zone is operated at the reduced heating level than when the zone is operated at the high heating level. The operating condition is an angle of attack of the engine, and the controller is configured to operate the at least one lower zone at the high heating level and the at least one upper zone at the reduced heating level when the gas turbine engine has an upward angle of attack.

[0068] The gas turbine engine according to any preceding clause, wherein the plurality of zones includes at least one left zone and at least one right zone. The controller is configured to selectively control the at least one heat source in each of the plurality of zones to one of a plurality of heating levels. The plurality of heating levels includes a high heating level and a reduced heating level. An amount of heat generated for one of the plurality of zones over a set time interval is less when the zone is operated at the reduced heating level than when the zone is operated at the high heating level. The operating condition is a crosswind direction. The controller is configured to operate one of the at least one left zone and the at least one right zone at the high heating level and the other of the at least one left zone and the at least one right zone at the reduced heating level based on the crosswind direction.

[0069] The gas turbine engine according to any preceding clause, wherein the plurality of zones includes a forward portion, an inner portion, and an outer portion.

[0070] The gas turbine engine according to any preceding clause, wherein the gas turbine engine includes a plurality of power conditions, and the operating condition is the power condition of the engine.

[0071] The gas turbine engine according to any preceding clause, wherein the plurality of power conditions includes a high power condition, a partial power condition, and a low power condition. The engine produces a greater amount of propulsive thrust in the high power condition than in each of the partial power condition and the low power condition. The engine produces a greater amount of propulsive thrust in the partial power condition than in the low power condition. The controller is configured to selectively control the at least one heat source in each of the plurality of zones to one of a plurality of heating levels. The plurality of heating levels includes a high heating level and a reduced heating level. An amount of heat produced for one of the plurality of zones over a set time interval is less when the zone is operated at the reduced heating level than when the zone is operated at the high heating level.

[0072] The gas turbine engine according to any preceding clause, wherein when the engine is operated in the high power condition, the controller is configured to operate each of the forward portion, the inner portion, and the outer portion at the high heating level.

[0073] The gas turbine engine according to any preceding clause, wherein when the engine is operated in the partial power condition, the controller is configured to operate the outer portion at the reduced heating level.

[0074] The gas turbine engine according to any preceding clause, wherein when the engine is operated in the partial power condition, the controller is configured to operate each of the forward portion and the inner portion at the high heating level.

[0075] The gas turbine engine according to any preceding clause, wherein when the engine is operated in the low power condition, the controller is configured to operate each of the forward portion and the outer portion at the reduced heating level.

[0076] The gas turbine engine according to any preceding clause, wherein when the engine is operated in the partial power condition, the controller is configured to operate the inner portion at the high heating level.

[0077] While the above description has been made to the preferred embodiments, it is noted that other variations and modifications will be apparent to persons skilled in the art and can be made without departing from the spirit or scope of the present disclosure. Furthermore, features described in conjunction with one embodiment can be used in conjunction with other embodiments, even if not explicitly stated above.

Claims

1. A gas turbine engine characterized by, including: a nacelle defining an inlet and including a lip having an outer surface configured to have an airflow thereon, the outer surface having a plurality of zones in a circumferential direction of the nacelle, the plurality of zones including an upper zone and a lower zone; and and an ice protection system for the outer surface of the lip, the ice protection system including: at least one heat source thermally coupled to the outer surface in each of the plurality of zones; and a controller configured to selectively control the at least one heat source in each of the plurality of zones based on an angle of attack of the gas turbine engine, wherein the controller is configured to selectively control the at least one heat source in each of the plurality of zones to one of a plurality of heating levels, the plurality of heating levels including a high heating level and a reduced heating level, an amount of heat generated in a set time interval for one of the plurality of zones is less when the zone is operated at the reduced heating level than when the zone is operated at the high heating level, and wherein the controller is configured to operate the lower zone at the high heating level and the upper zone at the reduced heating level when the gas turbine engine has an upward angle of attack.

2. The gas turbine engine of claim 1, wherein, wherein the heat source is one of hot air or an electrical heating element.

3. The gas turbine engine of claim 1, wherein, wherein the controller selectively controls the at least one heat source to one of the plurality of heating levels by controlling a duty cycle of the at least one heat source, the duty cycle of the at least one heat source in the reduced heating level being less than the duty cycle of the at least one heat source in the high heating level.

4. The gas turbine engine of claim 1, wherein, wherein each of the plurality of zones includes a temperature sensor configured to measure a temperature of the outer surface, and wherein the controller selectively controls the at least one heat source to one of the plurality of heating levels based on a target temperature of the outer surface.

5. The gas turbine engine of claim 1, wherein, wherein the plurality of zones includes a left zone and a right zone, and wherein the controller is configured to operate one of the left zone and the right zone at the high heating level and the other of the left zone and the right zone at the reduced heating level based on a crosswind direction.

6. The gas turbine engine of claim 1, wherein, wherein the upper zone is one of a plurality of upper zones, the plurality of upper zones including a leading portion, an inner portion, and an outer portion, the controller configured to operate at least one of the leading portion or the outer portion at the reduced heating level when the gas turbine engine has an upward angle of attack.

7. A gas turbine engine characterized by, including a plurality of power conditions, the plurality of power conditions including a high power condition, a partial power condition, and a low power condition, the gas turbine engine producing a greater amount of propulsive thrust in the high power condition than in each of the partial power condition and the low power condition, the gas turbine engine producing a greater amount of propulsive thrust in the partial power condition than in the low power condition, the gas turbine engine including: a nacelle defining an inlet and including a lip having an outer surface configured to have an airflow thereon, the outer surface having a plurality of zones including a leading portion, an inner portion, and an outer portion; and an anti-icing system for the outer surface of the lip, the anti-icing system including: at least one heat source thermally coupled to the outer surface in each of the plurality of zones; and a controller configured to selectively control the at least one heat source in each of the plurality of zones to one of a plurality of heating levels based on a current power condition of the plurality of power conditions of the gas turbine engine, the plurality of heating levels including a high heating level and a reduced heating level, an amount of heat generated for one of the plurality of zones over a set time interval being less when the zone is operated at the reduced heating level than when the zone is operated at the high heating level, and wherein the heating level of the outer portion is different for one of the high power condition, the partial power condition, or the low power condition than for another of the high power condition, the partial power condition, or the low power condition.

8. The gas turbine engine of claim 7, wherein, wherein when the gas turbine engine is operated in the high power condition, the controller is configured to operate each of the leading portion, the inner portion, and the outer portion at the high heating level.

9. The gas turbine engine of claim 7, wherein, wherein when the gas turbine engine is operated in the partial power condition, the controller is configured to operate the outer portion at the reduced heating level.

10. The gas turbine engine of claim 9, wherein, wherein when the gas turbine engine is operated in the partial power condition, the controller is configured to operate each of the leading portion and the inner portion at the high heating level.

11. The gas turbine engine of claim 7, wherein, wherein when the gas turbine engine is operated in the low power condition, the controller is configured to operate each of the leading portion and the outer portion at the reduced heating level.

12. The gas turbine engine of claim 11, wherein, wherein when the gas turbine engine is operated in the partial power condition, the controller is configured to operate the inner portion at the high heating level.

13. The gas turbine engine of claim 7, wherein, wherein the heat source is one of hot air or an electrical heating element.

14. The gas turbine engine of claim 7, wherein, wherein each of the plurality of zones includes a temperature sensor configured to measure a temperature of the outer surface, and wherein the controller selectively controls the at least one heat source to one of the plurality of heating levels based on a target temperature of the outer surface.

15. A gas turbine engine characterized by, including: a nacelle defining an inlet, the nacelle including a lip having an outer surface configured to have an airflow thereon, the outer surface having a plurality of zones in a circumferential direction of the nacelle, the plurality of zones including a left zone and a right zone; and an anti-icing system for the outer surface of the lip, the anti-icing system including: at least one heat source thermally coupled to the outer surface in each of the plurality of zones; and a controller configured to selectively control the at least one heat source in each zone of the plurality of zones based on a crosswind direction, wherein the controller is configured to selectively control the at least one heat source in each zone of the plurality of zones to one of a plurality of heating levels, the plurality of heating levels including a high heating level and a reduced heating level, an amount of heat generated for one of the plurality of zones over a set time interval is less when the zone is operated at the reduced heating level than when the zone is operated at the high heating level, and wherein the controller is configured to operate one of the left zone and the right zone at the high heating level and the other of the left zone and the right zone at the reduced heating level based on the crosswind direction.

16. The gas turbine engine of claim 15, wherein, wherein the left zone is one of a plurality of left zones, the plurality of left zones including a front portion, an inner portion, and an outer portion, the controller is configured to operate at least one of the front portion and the outer portion at the reduced heating level when the crosswind direction is from the right side.

17. The gas turbine engine of claim 15, wherein, wherein the right zone is one of a plurality of right zones, the plurality of right zones including a front portion, an inner portion, and an outer portion, the controller is configured to operate at least one of the front portion or the outer portion at the reduced heating level when the crosswind direction is from the left side.

18. The gas turbine engine of claim 15, wherein, wherein the heat source is one of hot air or an electric heating element.

19. The gas turbine engine of claim 15, wherein, wherein the controller selectively controls the at least one heat source to one of the plurality of heating levels by controlling a duty cycle of the at least one heat source, the duty cycle of the at least one heat source in the reduced heating level is less than the duty cycle of the at least one heat source in the high heating level.

20. The gas turbine engine of claim 15, wherein, wherein each zone of the plurality of zones includes a temperature sensor configured to measure a temperature of the outer surface, and wherein the controller selectively controls the at least one heat source to one of the plurality of heating levels based on a target temperature of the outer surface. wherein the controller selectively controls the at least one heat source to one of the plurality of heating levels based on a target temperature of the outer surface.

Images