Heat shield for gas turbine engine

Abstract

A heat shield for a combustor of a gas turbine engine. The heat shield includes an annular ring having an axial direction, a radial direction, and a circumferential direction. The annular ring includes a plurality of circumferential segments. Each of the plurality of circumferential segments is disconnected from adjacent ones of the plurality of circumferential segments to allow each circumferential segment to thermally grow during operation of the combustor.

Description

Technical Field

[0001] The present disclosure relates to heat shields, and in particular to heat shields for fuel nozzles in combustors, such as those used in gas turbine engines for aircraft. Background Art

[0002] A gas turbine engine, especially one used in an aircraft, is a rotary engine having a turbine in which working air flows continuously through a compressor section, a combustor section, and a turbine section. The working air is compressed in the compressor section. The compressed working air is then mixed with fuel and combusted in a combustor in the combustor section, producing hot combustion products. The combustion products are then used to drive a turbine in the turbine section. Fuel may be injected into the combustor using a fuel nozzle, and the fuel nozzle may include a heat shield to protect the fuel nozzle from the hot combustion products. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] Features and advantages of the present disclosure will become apparent from the following description of various exemplary aspects as illustrated in the drawings, wherein like reference numerals generally indicate identical, functionally similar, and / or structurally similar elements.

[0004] Figure 1 is a schematic diagram of an aircraft having a gas turbine engine according to an embodiment of the present disclosure.

[0005] Figure 2 is along Figure 1 The line 2-2 in Figure 1 Schematic cross-sectional view of a gas turbine engine of an aircraft shown in .

[0006] Figure 3 According to the embodiment of the present disclosure Figure 2 A schematic cross-sectional view of a combustor of a gas turbine engine is shown. Figure 3 Yes Display Figure 2 Detailed view of the details in 3.

[0007] Figure 4 is a schematic isometric view of a fuel nozzle (fuel nozzle tip) with a heat shield according to an embodiment of the present disclosure.

[0008] Figure 5 is along Figure 4 The line 5-5 in Figure 4 A partial cross-sectional view of a fuel injector is shown.

[0009] Figure 6 is along Figure 4 The line 6-6 in Figure 4 Another partial cross-sectional view of a fuel injector is shown.

[0010] Fig. 7A is along Figure 4 The line 7A-7A in Figure 4 A further cross-sectional view of a fuel injector is shown. Figure 7B From Fig. 7A A cross-sectional view at a similar angle shows the variation in protrusions.

[0011] FIG. 8A to FIG. 8C Alternative examples of the shape of the radial grooves and the corresponding protrusions of the heat shield are shown. Fig. 8A A first shape is shown, Figure 8B A second shape is shown, Figure 8C A third shape is shown.

[0012] Fig. 9 is a schematic isometric view of a fuel nozzle with a heat shield according to another embodiment of the present disclosure.

[0013] Fig.10 is along Fig. 9 The line 10-10 in Fig. 9 A partial cross-sectional view of a fuel nozzle is shown.

[0014] FIG. 11A to FIG. 11F Shows the manufacturing Fig. 9 Heat shield method shown. Fig.11A is the first step. Fig. 11B It is the second step. Fig. 11C It is the third step. Fig.11D It is the fourth step. Fig.11E It is the fifth step. Fig.11F It is the sixth step.

[0015] Fig.12 Can be used to form Fig. 9 A partial cross-sectional view of an alternative geometry for a panel of a heat shield is shown in FIG. DETAILED DESCRIPTION

[0016] Features, advantages and aspects of the present disclosure are set forth or apparent by considering the following detailed description, drawings and claims.In addition, the following detailed description is exemplary and is intended to provide further explanation without limiting the scope of the present disclosure as claimed.

[0017] Various aspects are discussed in detail below. Although specific aspects are discussed, this is for illustrative purposes only. Those skilled in the relevant art will recognize that other components and configurations may be used without departing from the spirit and scope of the present disclosure.

[0018] The terms "upstream" and "downstream" refer to relative directions relative to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing.

[0019] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, front refers to a position closer to the engine inlet, while rear refers to a position closer to the engine nozzle or exhaust.

[0020] The terms “coupled,” “fixed,” “attached,” “connected” and the like refer to both direct coupling, fixing, attachment, or connection as well as indirect coupling, fixing, attachment, or connection through one or more intermediate components or features, unless otherwise specified herein.

[0021] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0022] Here and throughout the specification and claims, range limitations are combined and interchanged, and unless context or language dictates otherwise, such ranges are identified and include all sub-ranges contained therein. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0023] As used herein, the terms "axial" and "axially" refer to directions and orientations extending substantially parallel to a centerline (such as, for example, the centerline of a turbine engine or the axis of a fuel nozzle). Additionally, the terms "radial" and "radially" refer to directions and orientations extending substantially perpendicular to these centerlines. Additionally, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations extending in an arc around these centerlines.

[0024] When radial directions are discussed herein, references to "inner" and "outer" refer to positions relative to the longitudinal centerline of the component. When referring to a portion of a component and otherwise specified, the "inner portion" preferably refers to the inner half of the component, more preferably the inner third of the component, and even more preferably the inner quarter of the component. Likewise, the "outer portion" preferably refers to the outer half of the component, more preferably the outer third of the component, and even more preferably the outer quarter of the component.

[0025] As used herein, the "hot side" is the side of a component of a turbine engine that is exposed to or otherwise oriented to face the combustion chamber of the combustion section. In the following discussion, the hot side of a component may be the rear side or the rearward side, and thus the rearward side may be used interchangeably with the hot side herein.

[0026] As used herein, the "cold side" is the side of the combustion section of a turbine engine that is not exposed to the combustion chamber or otherwise oriented to face the combustion chamber. In the following discussion, the cold side of a component may be the front side or forward side, and thus the forward side may be used interchangeably with the cold side herein.

[0027] As described above, fuel can be injected into the burner using a fuel nozzle, and the fuel nozzle may include a heat shield to protect the fuel nozzle from the hot combustion products generated by the combustion of the fuel in the burner. These heat shields may include an annular metal ring or flange positioned on the downstream (facing the burner) end of the fuel nozzle. The metal flange may be coated with a ceramic thermal barrier coating (TBC) to isolate the substrate below. During burner operation, the thermal expansion caused by high temperature deforms the heat shield and generates hoop stress in the flange. If the hoop stress becomes too large, these hoop stresses will produce radial cracks in the TBC. The cracks may cause TBC spalling and subsequent degradation of the thermal insulation provided by the TBC, resulting in further deformation of the heat shield and a reduction in the thermal protection provided by the heat shield. Deterioration of the heat shield may affect the durability and flame dynamics of the fuel nozzle.

[0028] The heat shield discussed herein is segmented circumferentially. This circumferential segmentation reduces (or may even eliminate) thermal deformation and hoop stress in the flange of the heat shield, thereby reducing cracking and spalling of the TBC. With the circumferentially segmented heat shield discussed herein, the durability and flame dynamics of the nozzle in the combustor can be maintained.

[0029] The heat shields discussed herein are particularly applicable to engines, such as gas turbine engines used on aircraft. Figure 1 is a perspective view of an aircraft 10 in which various advantageous embodiments may be implemented. Aircraft 10 includes a fuselage 12, wings 14 attached to fuselage 12, and an empennage 16. Aircraft 10 also includes a propulsion system that generates the propulsive thrust required to propel aircraft 10 during flight, during taxi operations, and the like. Figure 1 The propulsion system of the illustrated aircraft 10 includes a pair of engines 100. In this embodiment, each engine 100 is attached to one of the wings 14 via a pylon 18 in an underwing configuration. Figure 1 1 is attached to the wing 14 in the underwing configuration shown in FIG. 1 , but in other embodiments, the engine 100 may have alternative configurations and be coupled to other portions of the aircraft 10. For example, the engine 100 may additionally or alternatively include one or more aspects that are coupled to other portions of the aircraft 10, such as, for example, the tail 16 and the fuselage 12.

[0030] The following will refer to Figure 2 Further description, Figure 1The illustrated engines 100 are gas turbine engines each of which is capable of selectively producing propulsive thrust for the aircraft 10. The amount of propulsive thrust may be controlled based at least in part on the amount of fuel provided to the engines 100 via the fuel system 150 (see Figure 2 ). The aviation turbine fuel used in the embodiments discussed herein may be a combustible hydrocarbon liquid fuel having a desired carbon number, such as a kerosene-type fuel, for example, JetA fuel. However, other suitable fuels may also be used, including, for example, sustainable aviation fuel (SAF), including biofuels, hydrogen-based fuels (H2), etc. The fuel is stored in a fuel tank 151 of the fuel system 150. Figure 1 As shown, at least a portion of the fuel tank 151 is located in each wing 14, and a portion of the fuel tank 151 is located in the fuselage 12 between the wings 14. However, the fuel tank 151 may be located at other suitable locations in the fuselage 12 or the wing 14. The fuel tank 151 may also be located entirely within the fuselage 12 or the wing 14. The fuel tank 151 may also be separate tanks rather than a single entity, such as two tanks each located within a respective wing 14.

[0031] although Figure 1 The aircraft 10 shown is an airplane, but the embodiments described herein may also be applicable to other aircraft 10, including, for example, helicopters and unmanned aerial vehicles (UAVs). Preferably, the aircraft discussed herein is a fixed-wing aircraft or a rotary-wing aircraft that generates lift through aerodynamic forces acting on, for example, fixed wings (e.g., wings 14) or rotary wings (such as, the rotors of a helicopter), and is a heavier-than-air aircraft rather than a lighter-than-air aircraft (such as a spacecraft). In addition, although not depicted herein, in other embodiments, the gas turbine engine may be any other suitable type of gas turbine engine, such as an industrial gas turbine engine incorporated into a power generation system, a marine gas turbine engine, etc.

[0032] Figure 2 yes Figure 1 A schematic cross-sectional view of one of the engines 100 used in the propulsion system of an aircraft 10 is shown. Figure 2 The cross-sectional view is along Figure 1 The line 2-2 in is intercepted. Figure 2 In the embodiment shown, the engine 100 is a high bypass turbofan engine. The engine 100 may also be referred to herein as a turbofan engine. The engine 100 has an axial direction (engine axial direction Ae) (reference Figure 2 , extending parallel to the longitudinal centerline 101), radial direction (engine radial direction R e ) and circumferential direction (engine circumferential direction C e ). Engine circumferential direction C eIn the axial direction of the engine A e The turbofan engine (engine 100 ) includes a fan section 102 and a turbine 104 disposed downstream of the fan section 102 .

[0033] Figure 2 The illustrated turbomachine 104 includes a tubular outer casing 106 (also referred to as a casing or nacelle) defining an inlet 108. In this embodiment, the inlet 108 is annular. The outer casing 106 surrounds an engine core, which includes a compressor section 110 in series flow relationship, the compressor section 110 including a supercharger or low pressure (LP) compressor 111 and a high pressure (HP) compressor 112; a combustion section 114; a turbine section 115, the turbine section 115 including a high pressure (HP) turbine 116 and a low pressure (LP) turbine 118; and an ejection exhaust nozzle section 120. The compressor section 110, the combustion section 114, and the turbine section 115 together at least partially define a core air flow path 121 extending from the inlet 108 to the ejection exhaust nozzle section 120. The turbofan engine further includes one or more drive shafts. More specifically, the turbofan engine includes a high pressure (HP) shaft or spool 122 drivingly connecting HP turbine 116 to HP compressor 112 , and a low pressure (LP) shaft or spool 124 drivingly connecting LP turbine 118 to LP compressor 111 .

[0034] Figure 2 The fan section 102 shown includes a fan 126 having a plurality of fan blades 128 coupled to a disk 130. The plurality of fan blades 128 and the disk 130 can rotate together about the longitudinal centerline (axis) 101 via the LP shaft 124. The LP shaft 111 can also be driven directly by the LP shaft 124, such as Figure 2 104. The disk 130 is covered by a rotatable front hub 132 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 128. Further, an annular fan casing or outer nacelle 134 is disposed circumferentially around the fan 126 and / or at least a portion of the turbine 104. The nacelle 134 is supported relative to the turbine 104 by a plurality of circumferentially spaced outlet guide vanes 136. A downstream section 138 of the nacelle 134 extends over an outer portion of the turbine 104 to define a bypass airflow passage 140 therebetween.

[0035] The engine 100 may operate with and receive a fuel flow from a fuel system 150. The fuel system 150 includes a fuel delivery assembly 153 that provides a fuel flow from a fuel tank 151 to the engine 100, and more specifically, for example, to a plurality of fuel injectors 200 that inject fuel to the burners 170 of the combustion section 114 (see Figure 3, further discussed below). Fuel injectors may also be generally referred to as fuel nozzles. As discussed in more detail below, components of the fuel system 150, and more specifically, the fuel tank 151, are examples of a fuel source that provides fuel to the fuel injectors 200. The fuel delivery assembly 153 includes pipes, tubes, conduits, etc. to fluidly connect the various components of the fuel system 150 to the engine 100. The fuel tank 151 is configured to store fuel, and the fuel is supplied from the fuel tank 151 to the fuel delivery assembly device 153. The fuel delivery assembly 153 is configured to transport fuel between the fuel tank 151 and the engine 100, and thus provides a flow path (fluid path) for the fuel from the fuel tank 151 to the engine 100.

[0036] The fuel system 150 includes at least one fuel pump that is fluidly connected to the fuel delivery assembly 153 to induce a flow of fuel through the fuel delivery assembly 153 to the engine 100. One such pump is a main fuel pump 155. The main fuel pump 155 is a high pressure pump that is the primary source of pressure increase in the fuel delivery assembly 153 between the fuel tank 151 and the engine 100. The main fuel pump 155 can be configured to increase the pressure in the fuel delivery assembly 153 to a pressure that is greater than the pressure within the combustion chamber 172 of the combustor 170.

[0037] The fuel system 150 also includes a fuel metering unit 157 in fluid communication with the fuel delivery assembly 153. Any suitable fuel metering unit 157 may be used, including, for example, a metering valve. The fuel metering unit 157 is positioned downstream of the main fuel pump 155 and upstream of a fuel manifold 159 configured to distribute fuel to the fuel injectors 200. The fuel system 150 is configured to provide fuel to the fuel metering unit 157, and the fuel metering unit 157 is configured to receive fuel from the fuel tank 151. The fuel metering unit 157 is further configured to provide a fuel flow to the engine 100 in a desired manner. More specifically, the fuel metering unit 157 is configured to meter fuel and provide a desired amount of fuel to the fuel manifold 159 of the engine 100 at, for example, a desired flow rate. The fuel manifold 159 is fluidly connected to the fuel injectors 200 and distributes (provides) the received fuel to a plurality of fuel injectors 200, where the fuel is injected into the combustion chamber 172 and combusted. Adjusting fuel metering unit 157 changes the amount of fuel provided to combustion chamber 172 , and therefore changes the amount of propulsive thrust produced by engine 100 to propel aircraft 10 .

[0038] The engine 100 also includes various auxiliary systems to assist in the operation of the engine 100 and / or an aircraft including the engine 100. For example, the engine 100 may include a main lubrication system 162, a compressor cooling air (CCA) system 164, an active thermal clearance control (ATCC) system 166, and a generator lubrication system 168, each of which is configured to operate in a controlled manner. Figure 2 Schematically shown in FIG. The main lubrication system 162 is configured to provide lubricant to various bearings and gear meshes in, for example, the compressor section 110, the turbine section 115, the HP spool 122, and the LP shaft 124. The lubricant provided by the main lubrication system 162 can increase the service life of these components, and a certain amount of heat can be removed from these components by using one or more heat exchangers. The compressor cooling air (CCA) system 164 provides air from one or both of the HP compressor 112 or the LP compressor 111 to one or both of the HP turbine 116 or the LP turbine 118. The active thermal clearance control (ATCC) system 166 is used to minimize the clearance between the turbine blade tip and the casing wall as the casing temperature changes during the flight mission. The generator lubrication system 168 provides lubrication for the electronic generator (not shown) and provides cooling / heat removal for the electronic generator. The electronic generator can provide power, for example, to a starter motor for the engine 100 and / or various other electronic components of the engine 100 and / or the aircraft including the engine 100. The lubrication systems for engine 100 (eg, main lubrication system 162 and generator lubrication system 168 ) may use a hydrocarbon fluid, such as oil, for lubrication, where the oil circulates through the inner surface of the oil removal line.

[0039] The turbofan engine (engine 100) discussed herein is provided by way of example only. In other embodiments, any other suitable engine may be used with various aspects of the present disclosure. For example, in other embodiments, the engine 100 may be any other suitable gas turbine engine, such as a turboshaft engine, a turboprop engine, a turbojet engine, a non-ducted single fan engine, etc. In this manner, it should also be understood that in other embodiments, the gas turbine engine may have other suitable configurations, such as other suitable numbers or arrangements of shafts, compressors, turbines, fans, etc. Further, although the turbofan engine (engine 100) is shown as a direct drive, fixed pitch turbofan engine, in other embodiments, the gas turbine engine may be a geared gas turbine engine (i.e., including a gearbox between the fan 126 and the shaft driving the fan (such as the LP shaft 124)) may be a variable pitch gas turbine engine (i.e., including a fan 126 having a plurality of fan blades 128 rotatable about their respective pitch axes), etc. In addition, still in an optional embodiment, aspects of the present invention may be incorporated into any other type of engine, or otherwise used with any other type of engine (e.g., a reciprocating engine). Additionally, in other exemplary embodiments, the exemplary turbofan engine (engine 100) may include or be operably connected to any other suitable accessory system. Additionally or alternatively, the exemplary engine 100 may not include or be operably connected to one or more of the accessory systems 162, 164, 166, and 168 described above.

[0040] like Figure 3 is a schematic cross-sectional view of a combustor 170 of a combustion section 114 according to an embodiment of the present disclosure. Specifically, Figure 3 It is shown Figure 2 Detailed view of detail 3 in FIG. The combustor 170 is an annular combustor that includes a combustion chamber 172 defined between an inner liner 174 and an outer liner 176. Each of the inner liner 174 and the outer liner 176 is circumferentially disposed about the longitudinal centerline 101 ( Figure 2 ) is annular and can therefore be arranged along the engine circumferential direction C e The combustor 170 also includes a combustor casing 178 that is also annular about the longitudinal centerline 101 of the engine 100. The combustor casing 178 extends circumferentially around the inner liner 174 and the outer liner 176, and the inner liner 174 and the outer liner 176 are located radially inward of the combustor casing 178. The combustor 170 also includes a dome 180 mounted to the forward end of each of the inner liner 174 and the outer liner 176. The dome 180 defines the upstream (or forward end) of the combustion chamber 172.

[0041] Multiple mixer assemblies 210 ( Figure 3 The plurality of mixer assemblies 210 are spaced circumferentially around the longitudinal centerline 101 of the engine 100. Figure 3 In the illustrated embodiment, each mixer assembly 210 is a twin annular premixing swirler (TAPS) that includes a main mixer 212 and a pilot mixer 214. The pilot mixer 214 is supplied with fuel from the fuel injector 200 during the entire engine operating cycle, and the main mixer 212 is supplied with fuel from the fuel injector 200 only during power increase conditions of the engine operating cycle (e.g., during takeoff and climb). As discussed in more detail below, the fuel injector 200 includes an aft heat shield 300. The TAPS mixer assembly 210 is provided by way of example, and the aft heat shield 300 herein may be used with other mixer assembly designs and other combustor designs, including, for example, rich burner combustors and mixer assemblies.

[0042] As described above, compressor section 110, including HP compressor 112 ( Figure 2 ), pressurizes the air, and the combustor 170 receives an annular flow of the pressurized air from the discharge outlet (compressor discharge outlet 216) of the HP compressor 112. This air may be referred to as compressor discharge air. A portion of the compressor discharge air flows into the mixer assembly 210. Fuel is injected into the air in the mixer assembly 210 to mix with the air and form a fuel-air mixture. The fuel-air mixture is provided from the mixer assembly 210 to the combustion chamber 172 for combustion. Ignition of the fuel-air mixture is accomplished by a suitable igniter 182, and the resulting combustion gases are directed in the engine axial direction A. e The first stage turbine nozzle 184 is defined by an annular flow passage and includes a plurality of radially extending, circularly spaced nozzle vanes 186 that turn the gases so that they flow at an angle and impact the HP turbine 116 ( Figure 2 ) of a first stage turbine blade (not shown) of a first turbine (not shown).

[0043] The fuel injector 200 is secured to the combustor housing 178 by a nozzle mount. In this embodiment, the nozzle mount is a flange 202 integrally formed with a stem 204 of the fuel injector 200. The flange 202 is secured to the combustor housing 178 and sealed to the combustor housing 178. The stem 204 includes a flow passage through which fuel flows, and the stem 204 extends radially inward from the flange 202. The fuel injector 200 also includes a fuel nozzle tip 220 through which fuel is injected into the combustion chamber 172 as part of the mixer assembly 210.

[0044] Figure 42 is a schematic isometric view showing a fuel nozzle tip 220 (more generally, a fuel nozzle) of a fuel injector 200. According to an embodiment, the fuel nozzle tip 220 includes a fuel nozzle body 222 and a rear heat shield 300. The rear heat shield 300 is attached to the fuel nozzle body 222. The fuel nozzle body 222 is mounted to an inlet cowling 224. The inlet cowling 224 is connected to or integral with the stem 204. The fuel nozzle body 222 includes a main fuel nozzle 230 and a dual orifice pilot fuel injector tip 240 (see Figure 3 ) has a main pilot fuel orifice and an auxiliary pilot fuel orifice.

[0045] like Figure 3 As shown, the dual orifice pilot fuel injector tip 240 may be substantially centered on the annular pilot inlet 241. The main fuel nozzle 230 surrounds the pilot inlet 241, and the pilot inlet 241 is located between the main fuel nozzle 230 and the dual orifice pilot fuel injector tip 240. In this embodiment, the fuel nozzle tip 220 is circular about an axis (referred to herein as the fuel nozzle axis 201) that extends through the center of the main pilot fuel orifice. In the following discussion, various features of the fuel nozzle tip 220 may be discussed relative to this axis (fuel nozzle axis 201).

[0046] Fuel is provided to the main pilot fuel orifice and the auxiliary pilot fuel orifice of the dual orifice pilot fuel injector tip 240 through the rod 204. The pilot mixer 214 includes a pilot swirler, which causes the air flowing therethrough to swirl. A portion of the compressor discharge air flows into the mixer assembly pilot inlet 241 and then flows into the pilot swirler. As described above, fuel and air are always provided to the pilot mixer 214 during the engine operating cycle, so that a main combustion zone is generated in the central portion of the combustion chamber 172. Each of the main pilot fuel orifice and the auxiliary pilot fuel orifice generally injects fuel into the compressed air flowing through the pilot swirler in a downstream direction. The pilot mixer 214 is supported by an annular pilot housing 243. The pilot housing 243 includes a tapered wall section 245 surrounding a tapered pilot mixing chamber 247, which is in flow communication with the pilot mixer 214 and is located downstream of the pilot mixer 214. The fuel and air mixture flows through the pilot mixing chamber 247, where the fuel and air are further mixed, passes through the outlet 249 of the pilot mixing chamber 247, and enters the combustion chamber 172 (see Figure 3 ).

[0047] Return to reference Figure 4, fuel is also provided to a main fuel nozzle 230 through the rod 204, and more specifically, to an annular main fuel passage of the annular main fuel ring. The main fuel nozzle 230 includes a circular array of main fuel injection orifices 232 (or an annular array of main fuel injection orifices 232) extending radially outward from the annular main fuel passage and through the wall of the annular main fuel ring. The main fuel nozzle 230 and the annular main fuel ring are spaced radially outward of the main pilot fuel orifices and the auxiliary pilot fuel orifices. The main fuel nozzle 230 injects fuel in a radially outward direction through the circular array of main fuel injection orifices 232.

[0048] As described above, the fuel nozzle tip 220 includes a fuel nozzle axis 201, and various features of the fuel nozzle tip 220 may be discussed relative to the fuel nozzle axis 201. The fuel nozzle (fuel nozzle tip 220) has an axial direction (nozzle axial direction A). n ) (extending parallel to the fuel nozzle axis 201), radial direction (nozzle radial direction R n ) and the circumferential direction (nozzle circumferential direction C n ). Nozzle circumferential direction C n Along the axial direction A around the nozzle n The direction of rotation extends.

[0049] The fuel nozzle tip 220 extends into the combustion chamber 172 and is adjacent to the main combustion zone. The pilot housing 243 and the rear heat shield 300 are exposed to high temperatures. For example, the conical wall section 245 of the pilot housing 243 and the rear heat shield 300 may be exposed to a gas temperature of six hundred degrees Fahrenheit (600°F) to three thousand eight hundred degrees Fahrenheit (3,800°F). The pilot housing 243 and the rear heat shield 300 are made of materials suitable for use in these high temperature environments, including corrosion-resistant alloys such as stainless steel, nickel and chromium, and high-strength nickel-based alloys. Therefore, the pilot housing 243 and the rear heat shield 300 may be formed of a metal alloy selected from a combination consisting of an iron-based alloy, a nickel-based alloy, and a chromium-based alloy. In some embodiments, cobalt-based alloys may also be used.

[0050] The rear heat shield 300 includes an inner wall 310. The inner wall 310 may be annular and define a heat shield hole 312. The inner wall 310 may be formed along the nozzle axial direction A. nThe inner wall 310 may form a portion of the conical wall section 245, and the front end of the inner wall 310 may abut the rear end of the conical wall section 245 to provide a substantially continuous conical section of the pilot mixing chamber 247, wherein the heat shield hole 312 forms a portion of the pilot mixing chamber 247. The inner wall 310 may be attached to the pilot housing 243 to connect the aft heat shield 300 to the fuel nozzle tip 220. The inner wall 310 may be attached to the pilot housing 243 by any suitable means, including, for example, a fusion bonding process, such as brazing, welding, etc. The aft heat shield 300 also includes a shroud flange 320 located at the rear end of the inner wall 310. The shroud flange 320 is annular and extends radially outward from the fuel nozzle axis 201, and may be referred to herein as an annular ring. The shroud flange 320 may include a rearward surface 321. The rearward facing surface 321 is the surface that faces the combustion chamber 172, and is therefore a hot side surface. The rearward facing surface 321 of this embodiment is a generally flat surface.

[0051] In this embodiment, the rearward surface 321 is orthogonal to the fuel nozzle axis 201 and may be coated with a thermal barrier coating (TBC) 302 to insulate the fuel nozzle tip 220 from the heat from combustion. To illustrate the features of the rear heat shield 300, the TBC 302 is only Figure 4 302 is shown on a portion of the rearward surface 321. Preferably, the TBC 302 is a coating having a high thermal resistivity, such as a ceramic material. The ceramic can be a stable ceramic that can withstand relatively high temperature gradients so that the coated metal component can operate at a gas temperature above the melting point of the metal. For example, the TBC material can be one or more of yttria-stabilized zirconia (YSZ) and other rare earth-stabilized zirconia compositions, mullite (3Al2O3-2SiO2), alumina (Al2O3), ceria (CeO2), rare earth zirconates (e.g., La2Zr2O7), rare earth oxides (e.g., La2O3, Nb2O5, Pr2O3, CeO2), and metal-glass composites, including combinations thereof (e.g., alumina and YSZ or ceria and YSZ). A particularly suitable TBC material is, for example, yttria-stabilized zirconia (YSZ).

[0052] As described above, a portion of the compressor discharge air flows into the mixer assembly 210 ( Figure 3). While some of the compressor discharge air mixes with the fuel to form a fuel and air mixture that is discharged from the outlet 249, another portion of the compressor discharge air may be used as cooling air for the aft heat shield 300. The cooling air may be discharged from a plurality of cooling holes 332 formed in the aft surface 226 of the fuel nozzle body 222. The aft surface 226 may be referred to herein as the distal end of the fuel nozzle tip 220. For clarity, the aft surface 226 may be referred to herein as the distal end of the fuel nozzle tip 220. Figure 4 In the following discussion, only a portion of the cooling holes 332 are indicated by reference numerals.

[0053] Figure 5 is along Figure 4 5. A partial cross-sectional view of the fuel nozzle tip 220 taken along line 5-5 in FIG. The fuel nozzle body 222 may include an annular cooling air passage 334 formed in front of the rearward surface 226 and through which cooling air flows. The cooling holes 332 are fluidly connected to the cooling air passage 334, and the cooling air discharged from the cooling air passage 334 is directed toward the forward surface 323 of the shroud flange 320. The forward surface 323 is a surface on the opposite side of the shroud flange 320 from the rearward surface 321 and is therefore a cold side surface. Although the forward surface 323 is Figure 5 336 and the outer array 338. The cooling holes 332 in each of the inner array 336 and the outer array 338 are arranged in the circumferential direction C. n The fuel nozzles are aligned in the direction and circumferentially spaced apart, preferably evenly spaced apart, about the fuel nozzle axis 201 (see Figure 4 The cooling holes 332 of the inner array 336 may be arranged along the nozzle axial direction A parallel to the fuel nozzle axis 201. n The cooling holes 332 of the outer array 338 may be angled relative to the fuel nozzle axis 201 and oriented to impinge cooling air on or near the radially outer annular flange tip 325 of the shroud flange 320. The flange tip 325 of this embodiment is the radially outer edge, and more specifically, the radially outermost edge of the shroud flange 320.

[0054] The cover flange 320 of this embodiment extends from the heat shield hole 312 along the radial direction R of the fuel nozzle tip 220. n More specifically, in this embodiment, the normal of the rearward surface 321 is parallel to the nozzle axial direction A of the fuel nozzle tip 220. n The shroud flange 320 is annular and extends in a circumferential direction C of the fuel nozzle tip 220. nExtension. During operation, due to thermal expansion, the hood flange 320 heats up and grows (i.e., increases in physical size). In order to relieve hoop stress, the hood flange 320 includes a plurality of radial grooves 342, forming a plurality of circumferential segments 340 of the hood flange 320. Each circumferential segment 340 of the plurality of circumferential segments 340 is disconnected from an adjacent circumferential segment 340 of the plurality of circumferential segments 340. In this embodiment, adjacent circumferential segments 340 of the plurality of circumferential segments 340 are disconnected from each other by one of the radial grooves 342. The radial grooves 342 reduce the constraints of the hood flange 320 on itself and allow each circumferential segment 340 of the plurality of circumferential segments 340 to move independently of each other. More specifically, the radial grooves 342 can allow axial and radial growth of each circumferential segment 340, thereby reducing hoop stress.

[0055] In this embodiment, each radial groove 342 cuts (extends) inwardly from the flange end 325 toward the inner wall 310. The shroud flange 320 includes an inner ring portion 329, which in this embodiment is the inner portion of the shroud flange 320 that is not affected by the radial grooves 342. Each circumferential segment 340 extends radially outwardly from the inner ring portion 329, wherein the radial grooves 342 separate the circumferential segments 340.

[0056] The hood flange 320 has a width W f , width W f The width W is the distance from the flange end 325 to the inner diameter edge 314 of the shroud flange 320. f The radial grooves 342 may be referred to herein as the flange width or radial width. The radial grooves 342 may extend all the way through the hood flange 320 to the heat shield holes 312, but preferably, the radial grooves 342 do not extend all the way through the hood flange 320 to the heat shield holes 312. The longer the radial grooves 342, the more the hoop stress can be reduced, so the hoop stress reduction is greatest when the radial grooves 342 extend all the way through the hood flange 320. However, these longer grooves reduce the structural stiffness of the hood flange 320 and the discontinuity at the inner wall 310 may affect flame dynamics, so, preferably, the radial grooves 342 do not extend all the way through the hood flange 320 to the heat shield holes 312. In addition, a certain minimum amount of hoop stress needs to be reduced to avoid cracks in the TBC 302 (see Figure 4 ). The reduction of this hoop stress may depend on the thermal gradient across the shroud flange 320. Balancing these considerations, the radial groove 342 may therefore be reduced from the flange width W f More preferably, from ten percent (10%) to one hundred percent (100%), and more preferably, from twenty percent (20%) to eighty percent (80%).

[0057] The hood flange 320 also includes a flange thickness at locations where radial grooves 342 are formed. The radial grooves 342 are preferably formed through the entire flange thickness. Each radial groove 342 also includes a groove width W s (width in the circumferential direction). Narrow grooves with tight tolerances are difficult to manufacture, so in this embodiment, the groove width W s At least 50 thousandths of an inch (50 mils). Preferably, the groove width W s From 50 thousandths of an inch (50 mils) to 200 thousandths of an inch (200 mils), and more preferably, the groove width W s From 85 thousandths of an inch (85 mils) to 175 thousandths of an inch (175 mils).

[0058] like Figure 4 As shown, the radial grooves 342 are evenly spaced from each other. In this embodiment, the hood flange 320 has eight circumferential segments 340, wherein each radial groove 342 is positioned at forty-five degrees (45 degrees) from each other. The hood flange 320 preferably includes at least two circumferential segments 340, wherein the radial grooves 342 are circumferentially spaced one hundred eighty degrees (180 degrees) from each other, however, in other embodiments, the hood flange 320 may include four circumferential segments 340, wherein the radial grooves 342 are circumferentially spaced ninety degrees (90 degrees) from each other. In other embodiments, the plurality of circumferential segments 340 may be greater than eight, wherein the radial grooves 342 have corresponding even circumferential spacing.

[0059] Figure 6 is along Figure 46-6 in FIG. 3 . Another partial cross-sectional view of the fuel nozzle tip 220 taken along line 6-6 in FIG. As described above, the radial slots 342 are relatively wide, which facilitates manufacturing. With such wide radial slots 342, heat can pass through the slots, and the rear heat shield 300 of this embodiment includes a slot filler 356 positioned in each radial slot 342 to prevent heat from passing through the radial slots 342. The slot filler 356 may be attached to a structure on the cold side of the hood flange 320. In this embodiment, the slot filler 356 is a protrusion 350 extending from the rearward surface 226 of the fuel nozzle body 222. The protrusion 350 is positioned on the rearward surface 226 to extend into the radial slot 342, and therefore, as described above, may be arranged in a manner similar to the radial slots 342. The protrusion 350 may be integrally formed with the fuel nozzle body 222, or may be a separate component attached to the fuel nozzle body 222 by a suitable means (e.g., a fusion bonding process such as welding). The projection 350 may be the same material as the fuel nozzle body 222, such as a metal, or a metal that is compatible with the material of the fuel nozzle body 222 (e.g., weldable to the fuel nozzle body 222 and having similar thermal properties). Thus, the projection 350 may be formed of a metal alloy selected from the group consisting of an iron-based alloy, a nickel-based alloy, and a chromium-based alloy. In some embodiments, a cobalt-based alloy may also be used.

[0060] Fig. 7A It is along Figure 4 7A-7A in FIG. 7A shows another partial cross-sectional view of the shroud flange 320. Each protrusion 350 includes a rearward surface 352, and in this embodiment, the protrusion 350 extends into the radial slot 342 so that the rearward surface 352 of the protrusion 350 is substantially coplanar with the rearward surface 321 of the shroud flange 320. The rearward surface 352 of the protrusion 350 can be flush with the rearward surface 321 of the shroud flange 320, and these surfaces can be machined to achieve such a condition. Because the rearward surface 352 of the protrusion 350 is substantially coplanar with the rearward surface 321 of the shroud flange 320, a suitable method such as chemical vapor deposition can be used to apply the TBC 302 to form a uniform layer of ceramic of the TBC 302 on the combustor-facing surface of each of the shroud flange 320 and the protrusion 350. The protrusion 350 includes a side surface 354. In this embodiment, the side surface 354 extends perpendicular to the rearward facing surface 226 of the fuel nozzle body 222 , and therefore, the projection 350 has a rectangular cross-section in a circumferential plane.

[0061] As described above, cooling air is directed from the cooling holes 332 toward the forward surface 323 of the shroud flange 320. To increase the surface area exposed to the cooling air, and thus, the cooling effect, a plurality of fins 327 are formed on the forward surface 323 of each circumferential segment 340. The fins 327 may have any suitable shape, including, for example, an axially tapered shape.

[0062] Figure 7B is with Fig. 7A Instead of the slot filler 356 being a protrusion 350 having a rectangular cross-section in the circumferential plane, the slot filler 356 may have other cross-sectional shapes. Figure 7B Another projection 360 is shown which is axially tapered in the circumferential plane rather than having a rectangular cross-section. Figure 7B The protrusion 350 shown in FIG. 1 is otherwise similar to Fig. 7A The protrusion 350 shown in FIG. 1 and the discussion of the features of the protrusion 350 and the hood flange 320 also apply here. Figure 7B The protrusion 360 and shroud flange 320 are shown as the same or similar components, and the same reference numerals are used as the protrusion 350 and shroud flange 320 discussed above. The protrusion 360 is axially tapered, with its width becoming narrower as it moves from the rearward surface 226 of the fuel nozzle body 222 toward the rearward surface 321 of the shroud flange 320. The protrusion 360 may have a proximal portion 362 of the protrusion 360 that is proximal to the rearward surface 226 of the fuel nozzle body 222 and a distal portion 364 of the protrusion 360 that is distal to the rearward surface 226 of the fuel nozzle body 222. The protrusion 360 may be tapered such that the proximal portion 362 of the protrusion 360 is wider than the distal portion 364 of the protrusion. The sidewall 346 of the radial slot 342 also has a corresponding axial taper, with the width of the radial slot 342 at the rearward surface 321 being narrower than the width of the radial slot 342 at the forward surface 323. With the taper, the projection 360 resists movement in the forward direction of each circumferential segment 340. When the circumferential segment 340 moves forward, the side wall 346 of the radial groove 342 contacts the side surface 354 of the projection 360, thereby resisting the movement.

[0063] Fig. 8AThe rearward surface 321 of the hood flange 320 and the rearward surface 352 of the protrusion 350 are shown. The shape of the protrusion 350 within the radial groove 342 preferably corresponds to the shape of the radial groove 342 to fill the radial groove 342. Therefore, the radial groove 342 and the rearward surface 352 of the protrusion 350 preferably have the same shape in the assembled state (e.g., ambient temperature) and have an appropriate gap 358. In the assembled state, the gap 358 between each side surface 354 of the protrusion 350 and each side wall 346 of the radial groove 342 can be, for example, from one thousandth of an inch (1 mil) to ten thousandths of an inch (10 mils), and more preferably from one thousandth of an inch (1 mil) to five thousandths of an inch (5 mils). A gap 358 less than one thousandth of an inch (1 mil) may cause assembly problems, such as interference. As described above, the TBC 302 is applied on the combustor-facing surface of each of the shroud flange 320 and the protrusion 350, and thus the TBC 302 preferably bridges the gap 358 to form a continuous coating. Therefore, the maximum size of the gap 358 is preferably set to a distance that can be bridged by the TBC 302, which may be based on the thickness of the TBC 302.

[0064] The radial groove 342 and the rearward surface 352 of the protrusion 350 can have any suitable shape, but in some embodiments, the shape is selected for ease of manufacturing. Fig. 8A As shown, the radial groove 342 and the rearward surface 352 of the protrusion 350 have a rectangular shape. However, since square corners may be stress concentrators, the radial groove 342 and the rearward surface 352 of the protrusion 350 may have rounded corners (or chamfers), such as Figure 8B shown. Figure 8B The illustrated shape of the radial groove 342 and the rearwardly facing surface 352 of the protrusion 350 may be referred to as a U-shaped groove.

[0065] Figure 8C Another shape of the radial groove 342 and the rearward surface 352 of the protrusion 350 is shown. The radial groove 342 and the rearward surface 352 of the protrusion 350 have a V-shape, with the widest part of the groove at the flange end 325. To avoid stress concentration in this embodiment, the point (or narrowest part) of the radial groove 342 and the rearward surface 352 of the protrusion 350 can be rounded.

[0066] Fig. 9is a schematic isometric view of a rear heat shield 400 according to another embodiment of the present disclosure. The rear heat shield 400 of this embodiment is similar to the rear heat shield 300 discussed above and can be positioned on the fuel injector 200 in a manner similar to the rear heat shield 300 discussed above. The same reference numerals will be used for components of the rear heat shield 400 of this embodiment that are the same as or similar to the components of the rear heat shield 300 described above. The above description of these components also applies to this embodiment, and a detailed description of these components and other features (such as positioning on the fuel injector 200 and the material of the rear heat shield) is omitted here.

[0067] The rear heat shield 400 of this embodiment is similar to the rear heat shield 300 described above, and includes a plurality of circumferential segments that are able to move and grow (i.e., increase in size) independently of one another, thereby alleviating or preventing hoop stresses. As will be described in more detail below, in this embodiment, each of the plurality of circumferential segments is a plate 410. For clarity, in Fig. 9 In the following discussion, only a portion of the plates 410 is indicated by a reference numeral. A plurality of the plates 410 are stacked (arranged) together to form an annular ring, and then the inner portion of each plate 410 is bonded together to form the rear heat shield 400. Each plate 410 includes a rearward surface (referred to herein as a plate rearward surface 411), and when the plates 410 are bonded together, the plate rearward surfaces 411 collectively form the rearward surface 321 of the rear heat shield 400. The plates 410 are therefore preferably stacked such that the plate rearward surfaces 411 are substantially coplanar with each other, such as flush with each other.

[0068] Each plate 410 also includes side walls 413 (see also Fig.11A ), and in this embodiment, the side walls 413 of adjacent plates 410 abut against each other. Optionally, a small gap 414 may be formed between the side walls 413 of adjacent plates 410, which may be similar to the gap 414 between each of the side surfaces 354 of the protrusion 350 and each of the side walls 346 of the radial slot 342. Each plate 410 is also axially tapered (narrowed in the forward direction), with the plate rearward surface 411 being wider than the plate forward surface 415, forming a fin 417 for use with the fin 327 ( Fig. 7A and 7B ) The rear heat shield 400 is cooled in a similar manner.

[0069] Fig.10 is along Fig. 910-10 in FIG. 10 is a partial cross-sectional view of the rear heat shield 400 of this embodiment. As described above, the inner portion of each plate 410 is bonded to the adjacent plate 410 to form the rear heat shield 400. In this embodiment, the rear heat shield 400 includes an inner ring 420, and the inner surface 419 of each plate 410 is bonded to the inner ring 420, and more specifically, to the outer surface 422 of the inner ring 420. As will be further discussed below, any suitable bonding process may be used, such as brazing. In this embodiment, the inner ring 420 also includes an inner surface 424, which can be bonded to other portions of the rear heat shield 400 to form the inner wall 310.

[0070] Each plate 410 has a circumferential width W p . Plate circumferential width W p is the distance between the side walls 413, and when the plate 410 is gradually narrowed in the radial direction, the plate circumferential width W p is the average width. The inventors observed that when TBC 302 ( Fig. 7A and 7B ) When the TBC 302 ruptures due to hoop stress, it usually occurs at four locations, so it is preferred to use at least four plates 410. A greater number of plates 410 provides additional benefits, such as ease of assembly and improved heat transfer when the plates 410 taper in the forward direction to form the fins 417, so thirty-two plates 410 or more may be used. A circumferential width W of 10 thousandths of an inch (10 mils) or less p The plates 410 may be difficult to assemble, so the maximum number of plates can be to maintain the circumferential width W p The number of plates 410 is at least 10 thousandths of an inch (10 mils). In this embodiment, the plate circumferential width W of each plate 410 is p Equal to flange width W f More preferably, the plate circumferential width W p is the flange width W f 10% to 50% of the plate circumferential width W, and more preferably 10% to 30% of the plate circumferential width W. p Smaller than the plate 410 along the nozzle axial direction A n Extended distance (plate axial depth D p ).

[0071] FIG. 11A to FIG. 11F A method of forming the rear heat shield 400 of this embodiment is shown. Fig.11A is a first step in which a plurality of plates 410 are stacked together to form an annular ring 430 of stacked plates 410 (see Fig. 11B). In some embodiments, each plate 410 may also be radially tapered, with the outer portion 432 (outer surface) of each plate 410 being wider than the inner surface 419 so as to be stackable in an annular configuration. Fig. 11B The second step is to hold the annular ring 430 of the stacked plates 410 together for further processing. Although the stacked plates 410 can be held together for further processing by mechanical means (such as a fixture), the stacked plates 410 can be held together by bonding. More specifically, the outer portions 432 of adjacent plates 410 can be bonded to each other using, for example, a fusion bonding process (such as welding or brazing). Since this is a temporary bonding process, a suitable temporary bonding, such as spot welding, can be used.

[0072] Fig. 11C This is the third step. After the outer portions 432 of the plates 410 are bonded to each other, the inner portions 434 of the plates 410 are machined to the inner diameter. Fig. 11B As schematically shown, the inner portion 434 of the plate 410 may extend into the hole 436 of the annular ring 430 at different distances. Machining the inner portion 434 establishes a uniform inner diameter of the annular ring 430. By machining the inner portion 434, the plates 410 may be stacked and joined with less precision in previous steps, thereby improving the manufacturability of the rear heat shield 400.

[0073] In the fourth step, if Fig.11D As shown, the inner ring 420 is placed in the hole 436, and then the inner surface 419 of each plate 410 is attached to the inner ring 420. In this embodiment, each plate 410 is fused to the inner ring 420 by a fusion bonding process, such as welding or brazing. Brazing is a process that avoids deformation of the plate 410, and is preferably used when the operating temperature is lower than the brazing temperature of the copper solder used to attach each plate 410 to the inner ring 420.

[0074] The outer portion 432 is machined in a fifth step, such as Fig.11E As shown. The outer portion 432 is machined to remove the Fig. 11B ). Removing the weld removes the constraint on the outer portion 432 of each plate 410, thereby allowing each plate 410 to move and grow (i.e., increase in size) independently of each other. As described above, since the plates 410 are not constrained and attached to each other at the outer portion 432, hoop stresses can be alleviated or prevented.

[0075] After machining, TBC 302 Fig.11F More specifically, the plate rearward surfaces 411 collectively form the rearward surface 321 of the rear heat shield 400, and the TBC 302 is applied using a suitable coating process. Suitable coating processes include, for example, chemical vapor deposition or thermal spray techniques.

[0076] Fig.12 is shown to be useful for forming Fig. 9 4 is a partial cross-sectional view of another plate 440 of the heat shield shown in FIG. The plate 440 of this embodiment is similar to the plate 410 discussed above. The same reference numerals will be used for components and features of the plate 440 of this embodiment that are the same as or similar to the components and features of the plate 410 described above. The above description of the components and features of the plate 410 also applies to the plate 440 of this embodiment, and a detailed description of these components and features is omitted here. The plate 410 shown and described above has a nozzle axial direction A. n and the nozzle radial direction R n The plate 440 of this embodiment is L-shaped with a substantially rectangular cross-section extending along the nozzle axial direction A. n Extended Leg 442. Leg 442 is located on the inner portion 434 of plate 440 and extends from the rectangular portion 444 of plate 440 toward the cold side (forward direction).

[0077] The leg 442 includes an outer surface (radial outer surface 446) in the radial direction. The ring 450 may be placed on the radial outer surface 446 of each plate 440, and the plates 440 are connected together to the ring 450 by fusion bonding (such as by brazing), and in this case, the ring 450 may be a brazed ring 450. More specifically, the leg 442 of the plate 440 is brazed to the ring 450.

[0078] The fuel injector 200 (fuel nozzle) discussed herein includes an annular rear heat shield (rear heat shield 300 or rear heat shield 400) having a TBC 302. These rear heat shields (rear heat shield 300 or rear heat shield 400) include a shroud flange 320, which includes a plurality of circumferential segments that allow thermal growth and deformation, thereby reducing hoop stresses formed in the shroud flange 320 during operation. By alleviating these hoop stresses, the rear heat shield (rear heat shield 300 or rear heat shield 400) discussed herein helps prevent the TBC 302 from cracking. Further aspects of the present disclosure are provided by the subject matter of the following clauses.

[0079] A heat shield for a combustor of a gas turbine engine, the heat shield comprising: an annular ring having an axial direction, a radial direction, and a circumferential direction, the annular ring comprising a plurality of circumferential segments, each of the plurality of circumferential segments being disconnected from adjacent circumferential segments of the plurality of circumferential segments to allow thermal growth of each circumferential segment during operation of the combustor.

[0080] The heat shield according to the preceding clause, wherein the annular ring comprises a plurality of radial grooves, adjacent circumferential segments of the plurality of circumferential segments being disconnected from each other by one of the plurality of radial grooves.

[0081] A heat shield according to any of the preceding clauses, wherein the plurality of circumferential segments is two circumferential segments and the plurality of radial slots is two radial slots.

[0082] A heat shield as described in any preceding clause wherein the radial slots are positioned one hundred eighty degrees from each other.

[0083] A heat shield as in any of the preceding clauses, wherein the plurality of circumferential segments is four circumferential segments and the plurality of radial slots is four radial slots.

[0084] A heat shield as described in any preceding clause wherein the radial slots are positioned at ninety degrees to each other.

[0085] A heat shield as in any preceding clause, wherein the plurality of circumferential segments is eight circumferential segments and the plurality of radial slots is eight radial slots.

[0086] A heat shield as described in any of the preceding clauses, wherein the radial slots are positioned at forty-five degrees from each other.

[0087] A heat shield as described in any preceding clause, wherein the annular ring comprises a radially outer edge, each radial slot extending inwardly from the radially outer edge.

[0088] A heat shield according to any preceding clause, wherein the annular ring has a width in the radial direction, and each radial groove extends inwardly from the radial outer edge a distance that is less than the width of the annular ring in the radial direction.

[0089] A heat shield according to any preceding clause, wherein each radial slot extends inwardly from the radially outer edge a distance which is between 10% and 100% of the width of the annular ring in the radial direction.

[0090] A heat shield as described in any of the preceding clauses, wherein each radial slot has a width in the circumferential direction, the width in the circumferential direction being between fifty mils and two hundred mils.

[0091] A heat shield as described in any of the preceding clauses, wherein each radial slot has a width in the circumferential direction, the width in the circumferential direction being from eighty-five mils to one hundred seventy-five mils.

[0092] A heat shield as described in any preceding clause wherein the annular ring comprises an inner ring portion.

[0093] A heat shield according to any of the preceding clauses, wherein each circumferential segment of the plurality of circumferential segments extends radially outwardly from the inner ring portion.

[0094] A heat shield as described in any of the preceding clauses, wherein the inner ring portion is an inner portion of the annular ring not constrained by the plurality of radial grooves.

[0095] A heat shield as described in any preceding clause wherein the annular ring is part of a flange.

[0096] A heat shield according to any preceding clause, further comprising an inner wall to which the flange is attached.

[0097] A fuel nozzle for a gas turbine engine, the fuel nozzle comprising: a fuel nozzle tip including a distal end; and a heat shield according to any of the preceding clauses, the heat shield attached to the distal end of the fuel nozzle tip.

[0098] A fuel nozzle according to the preceding clause, wherein the distal end of the fuel nozzle tip includes a hot side surface having a plurality of protrusions extending therefrom, each protrusion of the plurality of protrusions extending into a respective one of the plurality of radial slots.

[0099] A fuel nozzle as described in any of the preceding clauses, wherein the annular ring includes a hot side surface.

[0100] The fuel nozzle of the preceding clause, wherein the hot side surface of the annular ring is a flat surface.

[0101] A fuel nozzle according to the preceding clause, wherein each of the plurality of protrusions includes a hot side surface, each of the plurality of protrusions extending into the corresponding one of the plurality of radial slots such that the hot side surface of the protrusion is substantially coplanar with the hot side surface of the annular ring.

[0102] A fuel nozzle according to the preceding clause, wherein each of the plurality of protrusions includes a hot side surface, each of the plurality of protrusions extending into the corresponding one of the plurality of radial slots such that the hot side surface of the protrusion is substantially flush with the hot side surface of the annular ring.

[0103] The fuel nozzle according to the preceding clause, further comprising a thermal barrier coating formed on the hot side surface of the annular ring and the hot side surfaces of the plurality of protrusions.

[0104] The fuel nozzle of the preceding clause, wherein the thermal barrier coating is ceramic.

[0105] A fuel nozzle as in the preceding clause, wherein each protrusion of the plurality of protrusions has a rectangular cross-section in a circumferential plane.

[0106] The fuel nozzle of the preceding clause, wherein each protrusion of the plurality of protrusions is axially tapered.

[0107] A fuel nozzle according to the preceding clause, wherein each of the plurality of protrusions comprises a proximal portion and a distal portion, the proximal portion being adjacent to the hot side surface of the fuel nozzle tip, and each of the plurality of protrusions being tapered such that the proximal portion is wider than the distal portion.

[0108] The fuel nozzle of the preceding clause, wherein each radial slot of the plurality of radial slots is tapered.

[0109] A fuel nozzle according to the preceding clause, wherein the annular ring includes a hot side surface and a cold side surface, and each radial groove of the plurality of radial grooves is tapered so that the width of the radial groove at the hot side surface of the annular ring is narrower than the width of the radial groove at the cold side surface of the annular ring.

[0110] A fuel nozzle according to the preceding clause, wherein each of the plurality of protrusions includes a hot side surface having a certain shape, and each of the plurality of radial grooves has a corresponding shape in the radial direction at the hot side surface of the annular ring.

[0111] The fuel nozzle of the preceding clause, wherein the shape of the hot side surface of each protrusion of the plurality of protrusions is rectangular.

[0112] The fuel nozzle of the preceding clause, wherein the shape of the hot side surface of each protrusion of the plurality of protrusions is U-shaped.

[0113] The fuel nozzle of the preceding clause, wherein the shape of the hot side surface of each protrusion of the plurality of protrusions is V-shaped.

[0114] A fuel nozzle as described in the preceding clause, wherein the narrowest portion of the V-shape is rounded.

[0115] A heat shield according to any of the preceding clauses, further comprising a plurality of plates arranged to form the annular ring, wherein each circumferential segment of the plurality of circumferential segments is a plate of the plurality of plates.

[0116] A heat shield according to any preceding clause, wherein each plate of the plurality of plates tapers in a radial direction.

[0117] A heat shield according to any preceding clause, wherein each panel of the plurality of panels comprises at least one side wall.

[0118] A heat shield according to any preceding clause, wherein the at least one side wall of each panel of the plurality of panels abuts the at least one side wall of an adjacent panel.

[0119] A heat shield as described in any preceding clause, wherein a gap is formed between the at least one side wall of each panel of the plurality of panels and the at least one side wall of an adjacent panel.

[0120] The heat shield according to the preceding clause, wherein the gap is from 1 mil to 10 mils.

[0121] A heat shield according to any preceding clause, wherein each plate of the plurality of plates is axially tapered.

[0122] A heat shield according to any preceding clause, wherein each plate of the plurality of plates comprises a plate hot side surface which collectively form a hot side surface of the annular ring.

[0123] A heat shield as described in any preceding clause wherein the plate hot side surfaces are substantially coplanar with each other.

[0124] A heat shield as described in any preceding clause wherein the plate hot side surfaces are flush with each other.

[0125] A heat shield as described in any of the preceding clauses, further comprising a thermal barrier coating formed on the hot side surface of the annular ring.

[0126] A heat shield as described in any preceding clause, wherein the thermal barrier coating is a ceramic.

[0127] A heat shield as described in any preceding clause, wherein each plate of the plurality of plates is shaped to form a fin on a side of the annular ring opposite the hot side surface of the annular ring.

[0128] A heat shield according to any preceding clause, wherein each panel of the plurality of panels comprises an inner portion and an outer portion, the outer portion of each panel being unattached to the outer portion of an adjacent panel.

[0129] A heat shield as described in any preceding clause wherein the inner portion of each panel of the plurality of panels is bonded.

[0130] A heat shield as described in any preceding clause, wherein each of the plurality of panels is L-shaped having a leg extending in the axial direction from the inner portion of each of the plurality of panels.

[0131] A heat shield as described in any preceding clause wherein the legs of each panel of the plurality of panels are bonded to a ring.

[0132] A heat shield as described in any preceding clause, wherein each leg of each panel of the plurality of panels comprises an upper surface, the upper surface of each leg being bonded to the ring.

[0133] A heat shield as described in any preceding clause wherein the upper surface of each leg is brazed to the ring.

[0134] A heat shield as described in any of the preceding clauses, further comprising an inner ring, wherein the inner portion of each panel of the plurality of panels is bonded to the inner ring.

[0135] A heat shield as described in any preceding clause, wherein each panel of the plurality of panels comprises an inner surface bonded to the inner ring.

[0136] A heat shield as described in any preceding clause, wherein the inner ring comprises an outer surface, the inner surface of each panel of the plurality of panels being bonded to the outer surface of the inner ring.

[0137] A heat shield as described in any preceding clause, wherein the inner surface of each plate of the plurality of plates is brazed to the outer surface of the inner ring.

[0138] A heat shield as in any preceding clause, wherein each of the plurality of plates has a circumferential width and the annular ring has a radial width, the circumferential width of each of the plurality of plates being less than the radial width of the annular ring.

[0139] A heat shield as in any of the preceding clauses, wherein the circumferential width of each plate of the plurality of plates is between ten percent (10%) and fifty percent (50%) of the radial width of the annular ring.

[0140] A heat shield as in any of the preceding clauses, wherein the circumferential width of each plate of the plurality of plates is between ten percent (10%) and thirty percent (30%) of the radial width of the annular ring.

[0141] A heat shield as described in any of the preceding clauses, wherein each plate of the plurality of plates has an axial depth, the circumferential width of each plate of the plurality of plates being less than the axial depth.

[0142] A fuel nozzle for a gas turbine engine, the fuel nozzle comprising: a fuel nozzle tip including a distal end; and a heat shield according to any of the preceding clauses attached to the distal end of the fuel nozzle tip.

[0143] The fuel nozzle of any of the preceding clauses, wherein the distal end of the fuel nozzle tip includes a hot side surface, and the fuel nozzle further includes a plurality of cooling holes formed in the hot side surface of the fuel nozzle tip.

[0144] A fuel nozzle according to any of the preceding clauses, wherein the annular ring includes a cold side surface, and the plurality of cooling holes are arranged in the hot side surface of the fuel nozzle tip to direct cooling air discharged through each of the plurality of cooling holes toward the cold side surface of the annular ring.

[0145] A fuel nozzle as recited in any preceding clause, further comprising a plurality of fins formed on the cold side surface of the annular ring.

[0146] A fuel nozzle as in any of the preceding clauses, wherein at least a portion of the plurality of cooling holes are arranged in an inner array.

[0147] A fuel nozzle as in any preceding clause wherein the cooling holes of the inner array are aligned in the circumferential direction.

[0148] A fuel nozzle as in any of the preceding clauses, wherein the cooling holes of the inner array are oriented in the nozzle axial direction parallel to an axis of the fuel nozzle.

[0149] A fuel nozzle as in any of the preceding clauses, wherein at least a portion of the plurality of cooling holes are arranged in an outer array.

[0150] A fuel nozzle as in any preceding clause wherein the cooling holes of the outer array are aligned in the circumferential direction.

[0151] A fuel nozzle as in any of the preceding clauses, wherein the annular ring includes a radially outer edge, and the cooling holes of the outer array are oriented to direct cooling air towards the radially outer edge.

[0152] A method of forming an annular heat shield for a combustor. The method includes stacking a plurality of plates together to form an annular ring having a hole, machining an inner portion of the plurality of plates to form an inner diameter of the hole of the annular ring, placing an inner ring into the hole of the annular ring, and bonding the inner portion of each of the plurality of plates to the inner ring.

[0153] The method of the preceding clause, further comprising bonding outer portions of each plate to each other before machining the inner portions of the plurality of plates, and removing the bonding on the outer portion of each plate after machining the inner portions of the plurality of plates.

[0154] A method as in any preceding clause, wherein removing bonds on the outer portion of each plate comprises machining the outer portion of each plate to remove bonds.

[0155] A method as in any preceding clause wherein the bond on the outer portion of each panel is a fusion bond.

[0156] A method as in any preceding clause, wherein bonding the outer portions of each plate to one another comprises welding, and the fusion bonding is welding.

[0157] A method as in any of the preceding clauses, wherein bonding the inner portion of each plate of the plurality of plates to the inner ring comprises forming a fusion bond.

[0158] A method as in any of the preceding clauses, wherein bonding the inner portion of each plate of the plurality of plates to the inner ring comprises brazing.

[0159] A method according to any of the preceding clauses, wherein the inner portion of the plurality of plates is machined to form the inner surface of each of the plurality of plates, and joining the inner portion of each of the plurality of plates to the inner ring includes joining the inner surface of each of the plurality of plates to the inner ring.

[0160] A method according to any of the preceding clauses, wherein the inner ring includes an outer surface, and bonding the inner portion of each of the plurality of plates to the inner ring includes bonding the inner surface of each of the plurality of plates to the outer surface of the inner ring.

[0161] A method as in any of the preceding clauses, wherein bonding the inner portion of each plate of the plurality of plates to the inner ring comprises brazing the inner surface of each plate of the plurality of plates to the outer surface of the inner ring.

[0162] A method as in any preceding clause, wherein each plate of the plurality of plates tapers in a radial direction.

[0163] A method as in any preceding clause, wherein each plate of the plurality of plates comprises at least one side wall.

[0164] A method as in any of the preceding clauses, wherein stacking the plurality of plates together comprises positioning the plates such that the at least one side wall of each plate of the plurality of plates abuts the at least one side wall of an adjacent plate.

[0165] A method as in any of the preceding clauses, wherein stacking the plurality of plates together comprises positioning the plates such that a gap is formed between the at least one side wall of each plate of the plurality of plates and the at least one side wall of an adjacent plate.

[0166] The method of any of the preceding clauses, wherein the gap is from 1 mil to 10 mils.

[0167] A method as in any of the preceding clauses, wherein stacking the plurality of plates together comprises positioning a plate hot side surface of each plate of the plurality of plates such that the plate hot side surfaces collectively form a hot side surface of the annular ring.

[0168] The method of any of the preceding clauses, further comprising coating the hot side surface of the annular ring with a thermal barrier coating.

[0169] A method as in any preceding clause, wherein the thermal barrier coating is a ceramic.

[0170] Although the foregoing description is directed to preferred aspects, other variations and modifications will be apparent to those skilled in the art, and other variations and modifications may be made without departing from the spirit or scope of the present disclosure. In addition, the features described in conjunction with one aspect may be used in conjunction with other aspects, even if not explicitly stated above.

Claims

1. A heat shield for a combustor of a gas turbine engine, characterized in that: The heat shield comprises: An annular ring having an axial direction, a radial direction, and a circumferential direction, the annular ring including a plurality of circumferential segments, each circumferential segment of the plurality of circumferential segments being disconnected from an adjacent circumferential segment of the plurality of circumferential segments to allow thermal growth of each circumferential segment during operation of the combustor.

2. The heat shield according to claim 1, characterized in that: The annular ring includes a plurality of radial grooves, and adjacent circumferential segments of the plurality of circumferential segments are disconnected from each other by one radial groove of the plurality of radial grooves.

3. The heat shield according to claim 2, characterized in that: The annular ring has a width in the radial direction and includes a radial outer edge, and each radial groove extends inwardly from the radial outer edge by a distance smaller than the width of the annular ring in the radial direction.

4. The heat shield according to claim 3, characterized in that: The annular ring includes an inner ring portion, which is an inner portion of the annular ring not constrained by the plurality of radial grooves, and each of the plurality of circumferential segments extends radially outward from the inner ring portion.

5. A fuel nozzle for a gas turbine engine, characterized in that: The fuel nozzle comprises: a fuel nozzle tip including a distal end; and The heat shield of claim 2, attached to the distal end of the fuel nozzle tip, Wherein the distal end of the fuel nozzle tip includes a hot side surface having a plurality of protrusions extending therefrom, each protrusion of the plurality of protrusions extending into a respective one of the plurality of radial slots.

6. The fuel nozzle according to claim 5, characterized in that wherein each of the plurality of projections includes a proximal portion and a distal portion, the proximal portion being adjacent to the hot side surface of the fuel nozzle tip, each of the plurality of projections being tapered such that the proximal portion is wider than the distal portion, and The annular ring includes a hot side surface and a cold side surface, and each radial groove of the plurality of radial grooves is tapered so that the width of the radial groove at the hot side surface of the annular ring is narrower than the width of the radial groove at the cold side surface of the annular ring.

7. The fuel nozzle according to claim 5, characterized in that wherein the annular ring includes a hot side surface that is a flat surface, and Wherein each of the plurality of protrusions includes a hot side surface, each of the plurality of protrusions extending into the corresponding one of the plurality of radial slots such that the hot side surface of the protrusion is substantially coplanar with the hot side surface of the annular ring.

8. The fuel nozzle according to claim 7, characterized in that Further included is a thermal barrier coating formed on the hot side surface of the annular ring and the hot side surfaces of the plurality of protrusions.

9. The heat shield according to claim 1, characterized in that: Further included is a plurality of plates arranged to form the annular ring, wherein each circumferential segment of the plurality of circumferential segments is a plate of the plurality of plates.

10. The heat shield according to claim 9, characterized in that Each plate of the plurality of plates includes a plate hot side surface that collectively forms a hot side surface of the annular ring and are substantially coplanar with each other, a thermal barrier coating being formed on the hot side surface of the annular ring.

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