System and apparatus for reducing bow waves in gas turbine engine

Abstract

A gas turbine engine includes a compressor section, a combustion section, and a turbine section, wherein the combustion section includes an inner liner and an outer liner spaced apart from the inner liner. The inner liner and the outer liner at least partially define a combustion chamber. The turbine section includes an inner band extending between an upstream side and a downstream side opposite the upstream side, and an outer band spaced apart from the inner band and extending between the upstream side and the downstream side. The inner band and the outer band at least partially define a working gas flow path. One or both of the inner and outer bands include a step portion adjacent the upstream side, and a body portion extending from the step portion to the downstream side. The step portion extends beyond the body portion in a radial direction.

Description

Technical Field

[0001] The present disclosure relates to gas turbine engines and reducing bow wave effects at components of gas turbine engines. Background Art

[0002] A turbine engine generally comprises a fan and a core section arranged in fluid communication with each other. A combustor is provided in the core section for generating combustion gases for driving a turbine in the core section of the turbine engine.

[0003] As combustion gases approach the interface between the combustor and turbine, a pressure wave, or bow wave, may cause a portion of the combustion gases to reflect upstream, creating pressure variations and uneven pressure distribution between the combustor and turbine. These pressure variations can cause the combustion gases to heat unwanted components of the turbine engine, which can reduce engine durability and lead to decreased engine performance. Therefore, there is a need for systems and apparatus for reducing bow wave effects within gas turbine engines. BRIEF DESCRIPTION OF THE DRAWINGS

[0004] A full and enabling disclosure of this description, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0005] Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0006] Figure 2 According to an exemplary embodiment of the present disclosure Figure 1 Cross-sectional side view of the compressor section, combustion section, and turbine section of a gas turbine engine in FIG.

[0007] Figure 3A According to an exemplary embodiment of the present disclosure Figure 2 Detailed top view of the interface between the combustion section and the turbine section.

[0008] Figure 3B According to an exemplary embodiment of the present disclosure Figure 3A Cross-sectional view along line III–III of the interface between the combustion section and the turbine section.

[0009] Figure 4A According to an exemplary embodiment of the present disclosure Figure 2 Detailed top view of the interface between the combustion section and the turbine section.

[0010] Figure 4B According to an exemplary embodiment of the present disclosure Figure 4A Cross-sectional view along line IV–IV of the interface between the combustion section and the turbine section.

[0011] Figure 5 According to an exemplary embodiment of the present disclosure Figures 3A to 4B Diagrammatic representation of a portion of the inner band of the turbine section. DETAILED DESCRIPTION

[0012] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. In the drawings and the description, the same or similar designations have been used to refer to the same or similar parts of the present disclosure.

[0013] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Furthermore, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0014] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

[0015] For example, in the context of “at least one of A, B, and C,” the term “at least one of” means only A, only B, only C, or any combination of A, B, and C.

[0016] The term "turbine engine" refers to an engine that uses a turbine as all or part of its power source. Example turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc., as well as hybrid versions of one or more of these engines.

[0017] The term "combustion section" refers to any heat addition system of a turbomachine. For example, the term "combustion section" may refer to a section that includes one or more of a deflagration combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assemblies. In certain exemplary embodiments, the combustion section may include an annular combustor, a can combustor, a tube combustor, a trapped vortex combustor (TVC), or other suitable combustion systems, or a combination thereof.

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

[0019] As used herein, the terms "axial" and "axially" refer to directions and orientations extending generally parallel to the centerline of a gas turbine engine. Furthermore, the terms "radial" and "radially" refer to directions and orientations extending generally perpendicular to the centerline of a gas turbine engine. Furthermore, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations extending in an arc about the centerline of a gas turbine engine.

[0020] As used herein, the "bypass ratio" of a turbine engine refers to the ratio of bypass air passing through the bypass path of the turbine engine to the core air passing through the core inlet of the turbine engine's turbine. For example, the bypass ratio is the ratio of bypass air 62 entering the bypass airflow passage 56 to the core air 64 entering the turbine 16.

[0021] The terms "coupled," "fixed," "attached," and the like refer to both direct coupling, fixing, or attachment, as well as indirect coupling, fixing, or attachment through one or more intermediate components or features, unless otherwise specified herein.

[0022] As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0023] For purposes of the following description, the terms "upper," "lower," "right," "left," "vertical," "horizontal," "top," "bottom," "lateral," "longitudinal," and their derivatives shall relate to the orientation of the embodiments in the accompanying drawings. However, it should be understood that, unless expressly specified to the contrary, the embodiments may assume various different variations. It should also be understood that the specific arrangements shown in the drawings and described in the following specification are merely exemplary embodiments of the present disclosure. Accordingly, specific dimensions and other physical characteristics associated with the embodiments disclosed herein should not be considered limiting.

[0024] The term "adjacent" as used herein when referring to two walls and / or surfaces means that the two walls and / or surfaces are in contact with each other, or that the two walls and / or surfaces are separated only by one or more non-structural layers, and the two walls and / or surfaces are in continuous contact with the one or more non-structural layers (i.e., a first wall / surface is in contact with the one or more non-structural layers, and the one or more non-structural layers are in contact with the second wall / surface).

[0025] The present disclosure generally relates to reducing the bow wave effect at the interface between components within a gas turbine engine. Conventional gas turbine engines use an extended band or rearward-facing step between the combustion section and the turbine section to prevent the ingestion of combustion gases. However, the airflow stagnates at the leading edge of the airfoil, creating a bow wave that reverses the direction of the airflow. This reversal of airflow can cause gas to be ingested into the gap and can unduly increase the temperature of the gas turbine engine components. The present disclosure uses forward-facing steps on one or both of the inner and outer bands of the turbine section adjacent to the combustion section to form a stagnation area. This stagnation area increases the static pressure within the cavity adjacent to the forward-facing step to suppress the formation of the bow wave and the reversal of airflow. As a result, the ingestion of combustion gases can be prevented, which reduces the need for additional cooling and improves the durability of the gas turbine engine components.

[0026] Referring now to the accompanying drawings, Figure 1 is a schematic cross-sectional view of a turbine engine 10 according to an exemplary embodiment of the present disclosure.

[0027] like Figure 1 As shown, the turbine engine 10 has an axial direction A (extending parallel to the longitudinal centerline axis 12 ) and a radial direction R, which is perpendicular to the axial direction A. Generally speaking, the turbine engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14 .

[0028] The turbine 16 includes a generally tubular outer casing 18 defining an annular core inlet 20. Figure 1 As shown schematically, outer casing 18 encloses, in serial flow relationship, a compressor section 21 including a supercharger or low-pressure compressor ("LPC") 22 followed downstream by a high-pressure compressor ("HPC") 24; a combustion section 26; a turbine section 27 including a high-pressure turbine ("HPT") 28 followed downstream by a low-pressure turbine ("LPT") 30; and one or more core exhaust nozzles 32. A high-pressure ("HP") shaft 34, or rotating shaft, drivingly connects HPT 28 to HPC 24 for synchronous rotation with HPT 28. HPT 28 is drivingly coupled to HP shaft 34 for rotation when HPT 28 rotates. A low-pressure ("LP") shaft 36 drivingly connects LPT 30 to LPC 22 for synchronous rotation with LPT 30. LPT 30 is drivingly coupled to HP shaft 36 to rotate LP shaft 36 when LPT 30 rotates. The compressor section 21 , combustion section 26 , turbine section 27 , and one or more core exhaust nozzles 32 collectively define a working gas flow path 33 .

[0029] for Figure 1 In the illustrated embodiment, fan section 14 includes a fan 38 (eg, a variable pitch fan) having a plurality of fan blades 40 coupled to a disk 42 in a spaced-apart manner. Figure 1 As shown, fan blades 40 extend outwardly from disk 42 generally along a radial direction R. Since fan blades 40 are operably coupled to actuator 44, each fan blade 40 can rotate relative to disk 42 about pitch axis P, and actuator 44 is configured to collectively and synchronously change the pitch of fan blades 40. Fan blades 40, disk 42, and actuator 44 rotate together about longitudinal centerline axis 12 via fan shaft 45, which is powered by LP shaft 36 through a power gearbox (also referred to as gearbox assembly 46). Gearbox assembly 46 is Figure 1 The gearbox assembly 46 includes a plurality of gears for adjusting the speed of the fan shaft 45 and thereby adjusting the speed of the fan 38 relative to the LP shaft 36.

[0030] Still refer to Figure 1 In the exemplary embodiment, the disk 42 is covered by a rotatable fan hub 48 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 40. In addition, the fan section 14 includes an annular fan casing or nacelle 50 that circumferentially surrounds the fan 38 and / or at least a portion of the turbine 16. The nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 52. In addition, a downstream section 54 of the nacelle 50 extends over the exterior of the turbine 16 to define a bypass airflow passage 56 therebetween. One or more core exhaust nozzles 32 may extend through the nacelle 50 and be formed within the nacelle 50. In the exemplary embodiment, the one or more core exhaust nozzles 32 include one or more discrete nozzles that are spaced circumferentially about the nacelle 50. Other arrangements of the core exhaust nozzles 32 may also be used, such as including a single core exhaust nozzle in an annular or partially annular configuration about the nacelle 50.

[0031] During operation of turbine engine 10, a volume of air 58 enters turbine engine 10 through inlet 60 of nacelle 50 and / or fan section 14. As volume of air 58 passes over fan blades 40, a first portion of air (bypass air 62) is directed or delivered into bypass airflow passage 56 and a second portion of air (core air 64) is directed or delivered to an upstream section of working gas flow path 33, or more specifically, into annular core inlet 20. The ratio between the first portion of air (bypass air 62) and the second portion of air (core air 64) is referred to as the bypass ratio. In some embodiments, the bypass ratio is greater than 18:1. The pressure of core air 64 is then increased by LPC 22 to produce compressed air 65 ( Figure 2), compressed air 65 is routed through the HPC 24 and further compressed before being directed to the combustion section 26, where it is mixed with fuel and combusted to produce combustion gases 66 (combustion products). One or more stages may be used in each of the LPC 22 and the HPC 24, with each subsequent stage further compressing the compressed air 65. The HPC 24 has a compression ratio greater than 20:1, preferably in the range of 20:1 to 40:1. The compression ratio is the ratio of the pressure of the last stage of the HPC 24 to the pressure of the first stage of the HPC 24. The compression ratio may be greater than 20:1.

[0032] Combustion gases 66 are delivered to and expanded through HPT 28, where a portion of the heat and / or kinetic energy from the combustion gases 66 is extracted via successive stages of HPT stator vanes 68 coupled to outer casing 18 and HPT rotor blades 70 coupled to HP shaft 34, thereby rotating high pressure shaft 34 and thereby supporting operation of HPC 24. Combustion gases 66 are then delivered to and expanded through LPT 30. Here, a second portion of the heat and / or kinetic energy is extracted from the combustion gases 66 via successive stages of LPT stators 72 coupled to outer casing 18 and LPT rotor blades 74 coupled to LP shaft 36, thereby rotating LP shaft 36 and thereby supporting operation of LPC 22 and rotation of fan 38 via gearbox assembly 46. One or more stages may be used in each of HPT 28 and LPT 30. The compression ratio of the HPC 24 is in the range of 20:1 to 40:1, enabling a pressure expansion ratio of the HPT 28 in the range of 1.5:1 to 4:1, and a pressure expansion ratio of the LPT 30 in the range of 4.5:1 to 28:1.

[0033] The combustion gases 66 are then routed through one or more core exhaust nozzles 32 of the turbine 16 to provide propulsive thrust. Simultaneously with the core air 64 flowing through the working gas flow path 33, the bypass air 62 is routed through the bypass flow passage 56 before being discharged from the fan bypass nozzle 76 of the turbine engine 10, also providing propulsive thrust. The HPT 28, the LPT 30, and the one or more core exhaust nozzles 32 at least partially define a hot gas path 78 for routing the combustion gases 66 through the turbine 16.

[0034] As described above, compressed air 65 (core air 64) is mixed with fuel in combustion section 26 to form a mixture of fuel and air, which is then burned to produce combustion gases 66 (combustion products). Fuel can include any type of fuel for a turbine engine, for example, sustainable aviation fuel (SAF), jet fuel A (JetA) or other hydrocarbon fuels including biofuels. Fuel can also be a hydrogen-based fuel (H2). Although hydrogen-based fuels may include a mixture with hydrocarbon fuels, the fuel used herein is preferably unmixed and is referred to herein as hydrogen fuel. In some embodiments, hydrogen fuel can include substantially pure hydrogen molecules (i.e., diatomic hydrogen). Fuel can also be a cryogenic fuel. For example, when using hydrogen fuel, hydrogen fuel can be stored in a liquid phase at low temperatures.

[0035] Figure 1 The turbine engine 10 shown in FIG is for example only. In other exemplary embodiments, the turbine engine 10 can have any other suitable configuration. For example, in other exemplary embodiments, the fan 38 can be configured in any other suitable manner (e.g., as a fixed pitch fan) and can also be supported using any other suitable fan frame configuration. Furthermore, in other exemplary embodiments, any other suitable number or configuration of compressors, turbines, shafts, or combinations thereof can be provided. In other exemplary embodiments, aspects of the present disclosure can be incorporated into any other suitable turbine engine, such as a turbofan engine, a propfan engine, a turboprop engine, a ground power generation device, or combinations thereof.

[0036] Figure 2 According to an exemplary embodiment of the present disclosure Figure 1 1 is a cross-sectional side view of the compressor section 21, combustion section 26, and turbine section 27 of the turbine engine 10. More specifically, the aft end of the HPC 24, the combustion section 26, and the forward end of the HPT 28 are shown.

[0037] The compressed air 65 exits the HPC 24 through an annular diffuser 200 located at the rear end or outlet of the HPC 24 and diffuses into the combustion section 26. The combustion section 26 of the turbine 16 is annularly surrounded by an inner combustor casing 205 and an outer combustor casing 210 radially spaced apart from the inner combustor casing 205. The radially spaced inner and outer combustor casings 205, 210 each extend generally in the axial direction A1 and surround the combustor assembly 215 in an annular ring. The inner and outer combustor casings 205, 210 are joined together at the front end of the combustion section 26 at the annular diffuser 200.

[0038] As shown, the combustor assembly 215 generally includes an inner liner 225 extending generally along an axial direction A1 between an aft end 201 and a forward end 203, and an outer liner 230 also extending generally along an axial direction A1 between an aft end 206 and a forward end 208. The inner liner 225 and the outer liner 230 together at least partially define a combustion chamber 235 therebetween. The inner liner 225 and the outer liner 230 are each attached to or integrally formed with an annular dome. More specifically, the annular dome includes an inner dome segment 220 integrally formed with the forward end 203 of the inner liner 225, and an outer dome segment 223 substantially integrally formed with the forward end 208 of the outer liner 230. Furthermore, the inner dome segment 220 and the outer dome segment 223 can each be integrally formed (or, alternatively, can be formed from multiple components attached in any suitable manner) and can each extend along a circumferential direction C1 to form an annular shape. However, it should be understood that in other example embodiments, the combustor assembly 215 may not include the inner and / or outer dome segments 220, 223; may include separately formed inner and / or outer dome segments 220, 223 attached to respective inner and outer liners 225, 230; or may have any other suitable configuration.

[0039] In at least one example embodiment, the turbine section 27 includes an inner band 270 and an outer band 275 radially spaced from the inner band 270. The inner band 270 and the outer band 275 define the working gas flow path 33 ( Figure 1 The HPT stator vanes 68 may extend from the inner band 270 , the outer band 275 , or both the inner band 270 and the outer band 275 into the working gas flow path.

[0040] Still refer to Figure 2 The combustor assembly 215 further includes a plurality of fuel-air mixers 238 spaced apart along a circumferential direction C1 and at least partially located within the annular dome. More specifically, the plurality of fuel-air mixers 238 are at least partially located between the outer dome segment 223 and the inner dome segment 220 along a radial direction R1. Compressed air 65 from the compressor section 21 of the turbine engine 10 flows into or through the plurality of fuel-air mixers 238, where it is mixed with fuel and ignited to produce combustion gases 66 within the combustion chamber 235. The inner dome segment 220 and the outer dome segment 223 are configured to facilitate the flow of compressed air 65 from the compressor section 21 into or through the plurality of fuel-air mixers 238.

[0041] As previously described, combustion gases 66 flow from combustion chamber 235 and through turbine section 27 of turbine engine 10, where a portion of the thermal and / or kinetic energy from the combustion gases 66 is extracted via successive stages of turbine stator vanes and turbine rotor blades within HPT 28 and LPT 30. More specifically, as Figure 2 As shown, combustion gases 66 from the combustion chamber 235 flow into the HPT 28 immediately downstream of the combustion chamber 235 , where thermal and / or kinetic energy from the combustion gases 66 is extracted via successive stages of HPT stator vanes 68 and HPT rotor blades 70 .

[0042] like Figure 2 As shown, not all of the compressed air 65 flows into or directly through the plurality of fuel-air mixers 238 and into the combustion chamber 235. Some of the compressed air 65 is discharged into a plenum 240 surrounding the combustor assembly 215. The plenum 240 is generally defined between the inner and outer combustor casings 205, 210, and the inner and outer liners 225, 230. The outer combustor casing 210 and the outer liner 230 define an outer plenum 245 generally disposed radially outward from the combustion chamber 235. The inner combustor casing 205 and the inner liner 225 define an inner plenum 250 generally disposed radially inward relative to the combustion chamber 235. As the compressed air 65 is diffused by the annular diffuser 200, some of the compressed air 65 flows radially outward into the outer plenum 245, and some of the compressed air 65 flows radially inward into the inner plenum 250.

[0043] The compressed air 65 flowing radially outward into the outer chamber 245 flows generally axially toward the turbine section 27. Specifically, the compressed air 65 flows over the HPT stator vanes 68 and the HPT rotor blades 70. The outer chamber 245 may also extend to the LPT 30 (e.g., Figure 1 shown).

[0044] like Figure 2 As further shown, for the depicted embodiment, the HPT 28 includes a first stage 255 of the HPT stator vanes 68 and a second stage 260 of the HPT stator vanes 68 (as well as a first stage and a second stage of the HPT rotor blades 70). Furthermore, for the depicted embodiment, the second stage 260 of the HPT stator vanes 68 has a variable configuration such that the second stage 260 of the HPT stator vanes 68 includes a plurality of variable guide vane assemblies 265.

[0045] Figure 3A According to an exemplary embodiment of the present disclosure Figure 2 Detailed top view of the interface between the combustion section 26 and the turbine section 27. Figure 3B According to an exemplary embodiment of the present disclosure Figure 3AA cross-sectional view of the interface between the combustion section 26 and the turbine section 27 is shown along line III-III. More specifically, Figures 3A to 3B The interface 280 between the inner sleeve 225 and the inner band 270 is shown (eg, Figure 2 It will be appreciated that in some example embodiments, the interface between outer sleeve 230 and outer band 275 may be similar or analogous to interface 280 .

[0046] In at least one example embodiment, the interface 280 defines a gap, such as a chamber 300, between the downstream side 305 of the inner liner 225 and the upstream side 310 of the inner band 270. Figure 3B , the downstream side 305 of the inner band 270 and the upstream side 310 of the inner band 270 may be oriented toward the inner combustor casing 205 (e.g. Figure 2 20). Thus, chamber 300 also extends toward inner combustor casing 205. Furthermore, a seal 315 may be provided between inner liner 225 and inner band 270. Seal 315 may define at least a portion of chamber 300. In at least one example embodiment, seal 315 is flexible.

[0047] refer to Figure 3B In at least one example embodiment, at least a portion of the upstream side 310 of the inner band 270 extends at least partially into the working gas flow path 33 ( Figure 1 ), for example, extending to Figure 2 As shown, the inner band 270 includes a stepped portion 320 adjacent to the inner liner 225 and extending into the working gas flow path 33. Figure 3B As shown, at least a portion of the stepped portion 320 may have a generally concave shape. Additionally, or alternatively, the outer band 275 ( Figure 2 ) may include a stepped portion 320 adjacent the outer liner 230 and extending into the working gas flow path 33. In such an embodiment, the stepped portion 320 of the outer band 275 may be similar or analogous to the stepped portion 320 discussed herein.

[0048] In at least one example embodiment, the combustion gases 66 ( Figure 2 ) stagnates at the upstream side 310 of the stepped portion 320. This stagnation at the stepped portion 320 forms a stagnation region 330 near the upstream side 310 of the stepped portion 320. The stagnation region 330 is a high pressure region that pressurizes the chamber 300. Therefore, the stagnation region 330 pressurizes the chamber 300 so that the chamber 300 defines a high pressure region 333, as shown in FIG. Figure 3B For example, the overall static pressure within chamber 300 increases.

[0049] In at least one example embodiment, at least a second portion 335 of the working gas flow path 33 is stagnant at an upstream edge of a plurality of airfoils 340 extending from the inner band 270. The plurality of airfoils 340 may include the airfoils described above with reference to FIG. Figures 1 to 2 The HPT stator vane 68 discussed. Without the stepped portion 320, the second portion 335 of the working gas flow path 33 would likely stagnate on the airfoil 340 and generate a bow wave that would drive the second portion 335 toward a low pressure region, for example, into the chamber 300. However, the stagnant region 330 and the high pressure region 333 prevent the second portion 335 of the working gas flow path 33 from flowing from the plurality of airfoils 340 into the chamber 300. In contrast, Figure 3A As shown, a second portion 335 of the working gas flow path 33 is directed away from the chamber 300 around the airfoil 340 through the stagnation region 330 and the high pressure region 333, such that the second portion 335 continues along the working gas flow path 33. As a result, hot air from the working gas flow path 33, such as the combustion gases 66, is prevented from entering the chamber 300 and flowing toward other components of the turbine engine 10. If such hot air from the working gas flow path 33 were drawn into the chamber, the temperature of components within the turbine engine 10 could unduly increase, which would require additional cooling.

[0050] In at least one example embodiment, the chamber 300 may be cooler at higher pressures because hot fluids or gases, such as combustion gases 66, cannot enter the chamber 300. Therefore, the cooling air required to cool the chamber 300 (e.g., Figure 3B The amount of (indicated by arrow 345 in) may be reduced. For example, Figure 2 As shown, in some example embodiments, cooling air may enter chamber 300 from inner plenum 250 via interface 280. Reducing the amount of cooling air required may improve engine performance while maintaining component durability of turbine engine 10 because less air bypasses combustion chamber 235.

[0051] Figure 4A According to an exemplary embodiment of the present disclosure Figure 2 Detailed top view of the interface 280 between the combustion section 26 and the turbine section 27 . Figure 4B According to an exemplary embodiment of the present disclosure Figure 4A 26 and the turbine section 27. More specifically, the dimensions associated with the interface 280 are shown. Figure 2 As shown and referenced above Figures 3A to 3BIt will be appreciated that in some example embodiments, the interface between outer sleeve 230 and outer band 275 may be similar or analogous to interface 280 .

[0052] refer to Figure 4A , the plurality of airfoils 340 include an upstream end 400 and a downstream end 405 opposite the upstream end 400. Each of the plurality of airfoils 340 may include an airfoil length 408 extending between the upstream end 400 and the downstream end 405. In at least one example embodiment, the airfoil length 408 is equal to an airfoil height 465 between the inner band 270 and the outer band 275 (hereinafter referred to as Figure 4B In at least one further example embodiment, airfoil length 408 may be greater than or equal to 0.66 times the difference between the radius of outer band 275 and the radius of inner band 270. In other example embodiments, airfoil length 408 may be less than 0.66 times the difference between the radius of outer band 275 and the radius of inner band 270.

[0053] In at least one example embodiment, the plurality of airfoils 340 are spaced apart from the adjacent plurality of airfoils 340 by an airfoil spacing 410. For example, the plurality of airfoils 340 may include a first airfoil 415 and a second airfoil 420 spaced apart from the first airfoil 415. The upstream end 400 of the first airfoil 415 may be spaced apart from the same point on the upstream end 400 of the second airfoil 420 by the airfoil spacing 410. In at least one example embodiment, the result of dividing the airfoil spacing 410 by the airfoil length 408 may be between 1 and 3.

[0054] Still refer to Figure 4A Each of the plurality of airfoils 340 includes a first surface 425 and a second surface 430 opposite the first surface 425. The first surface 425 and the second surface 430 extend between the upstream end 400 and the downstream end 405 of each of the plurality of airfoils 340. In at least one example embodiment, the first surface 425 of each of the plurality of airfoils 340 includes a tangent point 435 between the upstream end 400 and the downstream end 405. The tangent point 435 of each of the plurality of airfoils 340 may be spaced apart from the upstream side 310 of the inner band 270 by a tangent point distance 440. In at least one example embodiment, the tangent point distance 440 is greater than or equal to 0.3 times the airfoil length 408 or less than or equal to 0.9 times the airfoil length 408. In other example embodiments, the tangent point distance 440 may be less than 0.3 times the airfoil length 408. In still other example embodiments, the tangent point distance 440 may be greater than 0.9 times the airfoil length 408.

[0055] Now refer to Figure 4B, the inner band 270 extends between the upstream side 310 and the downstream side 312. The inner band includes a step portion 320 adjacent to the upstream side 310 and a main body portion 451 extending from the step portion to the downstream side 312. The main body portion 451 extends from the step portion to the downstream side 312 of the inner band 270, for example, to the trailing edge 505. The step portion 320 extends along the inner band 270 between the leading edge 500 of the inner band 270 and an endpoint 452. The endpoint 452 is spaced between the leading edge 500 and the trailing edge 505 of the inner band 270. For example, the endpoint 452 may be adjacent to the airfoil 340 between the leading edge 500 and the trailing edge 505. The endpoint 452 represents the point along the inner band 270 where the step portion 320 ends. For example, the radius of curvature at the endpoint 452 may be zero, which will be referred to below. Figure 5 Have a discussion.

[0056] The step portion 320 is in the radial direction R (eg Figure 1 The stepped portion 320 may extend beyond the main body portion 451 (as shown). Furthermore, the stepped portion 320 includes a first bend 445 and a second bend 450 adjacent to the first bend 445. The first bend 445 may be adjacent to the upstream side 310 of the inner band 270, and the second bend 450 may be opposite the upstream side 310 and adjacent to the plurality of airfoils 340. Furthermore, the first bend 445 extends from the leading edge 500 to an inflection point 503 between the leading edge 500 and the endpoint 452. For example, the inflection point 503 may be located at any point along the inner band 270 between the leading edge 500 and the endpoint 452. Furthermore, the second bend 450 extends from the inflection point 503 to the endpoint 452 and may further extend toward a trailing edge 505 adjacent to the downstream side 312 of the inner band 270 opposite the upstream side 310.

[0057] In at least one example embodiment, the stepped portion 320 is axially spaced in the direction A (eg, Figure 1 As shown) on the upstream side 310 of the inner band 270 and the first curved portion 445 in the radial direction R (as shown) Figure 1310 ). Peak 463 of stepped portion 320 may be between leading edge 500 of inner band 270 and upstream end 400 of plurality of airfoils 340. In at least one example embodiment, peak distance 460 is greater than or equal to 0.01 times airfoil length 408 and less than or equal to 0.4 times airfoil length 408. When peak distance 460 is 0.01 times airfoil length 408, this results in a substantially vertical surface at leading edge 500 and upstream side 310 of inner band 270. In some further example embodiments, peak distance 463 is greater than or equal to 0.1 times airfoil length 408 and less than or equal to 0.2 times airfoil length 408. In such embodiments, peak 463 may be positioned upstream of airfoil 340, which allows stepped portion 320 to more smoothly form or curve from upstream side 310 toward downstream side 312.

[0058] In other exemplary embodiments, peak distance 460 may be less than 0.01 times airfoil length 408 . In still other exemplary embodiments, peak distance 460 may be greater than 0.4 times airfoil length 408 .

[0059] Additionally, the peak distance 460 may be greater than or equal to 0.05 times and less than or equal to 1.0 times the distance between the leading edge 500 of the inner band 270 and the upstream end 400 of the plurality of airfoils 340. Furthermore, the ratio of the peak distance 460 to the first bend length 453 may be greater than or equal to 0.1 and less than or equal to 1.

[0060] In other exemplary embodiments, peak distance 460 may be less than 0.05 times the distance between leading edge 500 of inner band 270 and upstream end 400 of plurality of airfoils 340. In still other exemplary embodiments, peak distance 460 may be greater than 1.0 times the distance between leading edge 500 of inner band 270 and upstream end 400 of plurality of airfoils 340.

[0061] In at least one example embodiment, the stepped portion 320 includes a stepped length 455. The stepped length 455 includes the first bend 445 and the second bend 450 in the axial direction A (eg, Figure 1 In addition, the step length 455 includes the length in the axial direction A (as shown) between the leading edge 500 and the end point 452. Figure 1In at least one example embodiment, step length 455 may be less than or equal to tangent distance 440. For example, step length 455 may be 0.2 to 0.9 times tangent distance 440. Furthermore, step length 455 may be less than tangent distance 440 so that step portion 320 terminates before reaching the Mach region of gas turbine engine 10 to prevent airfoil loss. In some further example embodiments, step length 455 may be less than or equal to 0.9 times airfoil length 408.

[0062] Additionally, the first bend 445 includes a first bend length 453 that extends between the upstream side 310 of the inner band 270 and an end of the first bend 445 adjacent the second bend 450 (indicated by the inflection point 503). In at least one example embodiment, the first bend length 453 can be greater than or equal to 0.1 times the difference between the step length 455 and the peak distance 460 and less than or equal to 0.9 times the difference. Thus, the length of the second bend 450 can be the difference between the step length 455 and the first bend length 453. The first bend length 453 can be greater than the length of the second bend 450 such that the inflection point 503 of the step portion 320 is downstream of the peak 463 of the first bend 445. Additionally, the step length 455 is less than or equal to the tangent distance 440 (e.g., Figure 4A The ratio of (as shown) can be greater than or equal to 0.05 and less than or equal to 1.

[0063] In at least one example embodiment, each of the plurality of airfoils 340 includes an airfoil height 465. The airfoil height 465 may be at least partially within the inner band 270 and the outer band 275 (e.g., Figure 2 The airfoil height 465 may vary between the upstream end 400 and the downstream end 405 .

[0064] Figure 5 According to an exemplary embodiment of the present disclosure Figures 3A to 4B 270 of the turbine section 27. More specifically, Figure 5 The curvature of the stepped portion 320 of the inner band 270 is shown.

[0065] In at least one example embodiment, the inner band 270 has multiple curvatures between the leading edge 500 and the trailing edge 505 of the stepped portion 320. More specifically, the stepped portion 320 of the inner band 270 may have at least two curvatures between the leading edge 500 and the endpoint 452 between the leading edge 500 and the trailing edge 505 of the inner band 270. For example, the stepped portion 320 of the inner band 270 may have a change in curvature at an inflection point 503. The inflection point 503 may be a point along the stepped portion 320 of the inner band 270 where the curvature changes. For example, the curvature of the stepped portion 320 of the inner band 270 may change from negative concavity (i.e., convex) to positive concavity (i.e., concave) at the inflection point 503.

[0066] like Figure 4B As shown, the first curved portion 445 comprises a generally convex shape, and the second curved portion 450 comprises a generally concave shape. Figure 5 , which is manifested as a negative concavity of the first bend 445 and a positive concavity of the second bend 450. At the inflection point 503 between the first bend 445 and the second bend 450, the concavity changes. In at least one example embodiment, the curvature of at least the stepped portion 320 of the inner band 270 can be represented by the following formula:

[0067]

[0068] Positive concavity (i.e., concave curvature) occurs when k > 1, and negative concavity (i.e., convex curvature) occurs when k < 1. The value of k can be approximated by using the second derivative of the radial coordinate with respect to the axial coordinate, which is expressed by the following formula:

[0069]

[0070] In the above formula, y is the radius, for example Figure 1 The radius extending in the radial direction R, x is the axial distance, for example, Figure 1 The axial distance extending along the axial direction A.

[0071] Although Figure 5 The concavity at the end point 452 is shown as zero, but it should be understood that in other example embodiments, the concavity at the end point 452 can be greater or less than zero, so that there is a curve or slope at the end point 452 along the inner band 270. In addition, the step portion 320 and the trailing edge 505 (such as Figure 4B The concavity of the main portion 451 between the main portions 451 and the main portions 451 may be 0. Additionally, or alternatively, there may be a bend or slope along the main portion 451 such that the concavity of the main portion is greater than or less than 0. Furthermore, the concavity along the main portion 451 may vary.

[0072] Therefore, including a forward-facing step on the inner band of the turbine section of a turbine engine adjacent to the combustion section can create a stagnation area for airflow, thereby preventing the formation of bow waves. This stagnation area can increase the static pressure within the chamber adjacent to the forward-facing step, thereby suppressing the formation of bow waves and preventing the reversal of airflow into the chamber caused by bow waves. Without the stagnation area, gases such as hot combustion gases may be drawn into the chamber, which can undesirably increase the temperature of turbine engine components. Therefore, the forward-facing step also reduces the need for additional cooling components of the turbine engine and improves the durability of these components.

[0073] The following topics provide further details:

[0074] A component for a gas turbine engine, the gas turbine engine including a compressor section, a combustion section, and a turbine section in a serial flow order, wherein the sections at least partially define a working gas flow path, the component including: an inner liner and an outer liner spaced apart from the inner liner, the inner liner and the outer liner defining a combustion chamber of the combustion section, the combustion chamber defining at least a portion of the working gas flow path; an inner band and an outer band spaced apart from the inner band, the inner band and the outer band defining at least a portion of the working gas flow path within the turbine section, wherein the inner band includes a first bend adjacent to an upstream end and a second bend adjacent to the first bend, the second bend being opposite to the upstream end; and a plurality of airfoils extending from the inner band, the outer band, or both the inner band and the outer band into the working gas flow path.

[0075] An assembly as claimed in any preceding clause, wherein the upstream end of the inner band extends at least partially into the working gas flow path.

[0076] An assembly as in any preceding clause, wherein at least a portion of the second bend is located between the first bend and the plurality of airfoils.

[0077] An assembly as claimed in any preceding clause, wherein: the first bend comprises a convex shape; and the second bend comprises a concave shape.

[0078] An assembly as claimed in any preceding clause, wherein: the first bend comprises a negative concavity; and the second bend comprises a positive concavity.

[0079] An assembly as in any preceding clause, wherein the inner liner and inner band define a chamber between the combustion section and the turbine section.

[0080] An assembly as in any preceding clause, wherein: the first portion of fluid flowing through the working gas flow path creates a stagnant region adjacent the upstream end of the inner band; and the stagnant region pressurizes the chamber such that the chamber defines a high pressure region.

[0081] An assembly as in any preceding clause, wherein: the stagnant area and the high pressure region prevent the second portion of the fluid flowing through the working gas flow path from entering the chamber; and the stagnant area and the high pressure region direct the second portion of the fluid away from the chamber and along the working gas flow path.

[0082] The assembly of any preceding clause, wherein: the first and second bends of the inner band include a step portion; and the step portion includes a step length; the plurality of airfoils include an upstream end and a downstream end opposite the upstream end; the plurality of airfoils include a tangent point between the upstream end and the downstream end; the tangent point is spaced apart from the upstream end of the inner band by a tangent point distance; and a ratio of the step length to the tangent point distance is greater than or equal to 0.05 and less than or equal to 1.

[0083] An assembly as described in any preceding clause, wherein the inner band includes: a peak distance between the upstream end of the inner band and the peak of the first bend; and a first bend length between the upstream end of the inner band and an end of the first bend adjacent to the second bend.

[0084] An assembly according to any preceding clause, wherein the ratio of the peak distance to the first bend length is greater than or equal to 0.1 and less than or equal to 1.

[0085] An assembly as claimed in any preceding clause, wherein the outer band comprises a first bend adjacent an upstream end of the outer band and a second bend adjacent the first bend.

[0086] A gas turbine engine comprises: a compressor section; a combustion section comprising: an inner liner and an outer liner spaced apart from the inner liner, the inner liner and the outer liner at least partially defining a combustion chamber; and a turbine section comprising: an inner band extending between an upstream side and a downstream side opposite the upstream side, an outer band spaced apart from the inner band and extending between the upstream side and the downstream side, the inner band and the outer band at least partially defining a working gas flow path, wherein one or both of the inner band and the outer band comprises a step portion adjacent to the upstream side and a main body portion extending from the step portion to the downstream side, and wherein the step portion extends beyond the main body portion in a radial direction, and a plurality of airfoils extending from the inner band, the outer band, or both the inner band and the outer band into the working gas flow path; wherein the compressor section, the combustion section, and the turbine section are in a serial flow order and define at least a portion of the working gas flow path.

[0087] A gas turbine engine as described in any preceding clause, wherein the inner band includes a stepped portion; and the stepped portion extends at least partially into the working gas flow path.

[0088] A gas turbine engine as claimed in any preceding clause, wherein at least a portion of the step portion comprises a concave shape.

[0089] A gas turbine engine as claimed in any preceding clause, wherein the step portion comprises a first bend and a second bend.

[0090] A gas turbine engine as in any preceding clause, wherein at least a portion of the second curved portion is located between the first curved portion and the plurality of airfoils.

[0091] A gas turbine engine as claimed in any preceding clause, wherein the first bend is different from the second bend.

[0092] A gas turbine engine as claimed in any preceding clause, wherein: the first curvature comprises a convex shape; and the second curvature comprises a concave shape.

[0093] A gas turbine engine as claimed in any preceding clause, wherein: the first curvature comprises negative concavity; and the second curvature comprises positive concavity.

[0094] A gas turbine engine according to any preceding clause, wherein: the inner band includes a peak distance between an upstream side of the inner band and a peak of the first bend; the inner band includes a first bend length between an upstream side of the inner band and an end of the first bend adjacent to the second bend; and a ratio of the peak distance to the first bend length is greater than or equal to 0.1 and less than or equal to 1.

[0095] A gas turbine engine as claimed in any preceding clause, wherein the inner liner and the inner band define a chamber between the combustion section and the turbine section.

[0096] A gas turbine engine as claimed in any preceding clause, wherein the combustion section further comprises a seal disposed between the inner liner and the inner band, the seal defining at least a portion of the chamber.

[0097] A gas turbine engine as claimed in any preceding clause, wherein the first portion of fluid flowing through the working gas flow path creates a stagnant region adjacent the upstream end of the inner band.

[0098] A gas turbine engine as claimed in any preceding clause, wherein the stagnation region pressurises the chamber such that the chamber defines a high pressure zone.

[0099] A gas turbine engine as claimed in any preceding clause, wherein the stagnant region and the high pressure region prevent the second portion of the fluid flowing through the working gas flow path from entering the chamber.

[0100] A gas turbine engine as claimed in any preceding clause, wherein the stagnant region and the high pressure zone direct a second portion of the fluid flowing through the working gas flow path away from the chamber and along the working gas flow path.

[0101] A gas turbine engine according to any preceding clause, wherein: the step portion includes a step length; the plurality of airfoils include an upstream end and a downstream end opposite the upstream end; the plurality of airfoils include a tangent point between the upstream end and the downstream end; the tangent point is spaced apart from the upstream end of the inner band by a tangent point distance; and a ratio of the step length to the tangent point distance is greater than or equal to 0.05 and less than or equal to 1.

[0102] A gas turbine engine as claimed in any preceding clause, wherein the concavity of the main body portion is zero.

[0103] A gas turbine engine as claimed in any preceding clause, wherein the concavity of at least a portion of the body portion is greater than zero.

[0104] A gas turbine engine as claimed in any preceding clause, wherein the concavity of at least a portion of the body portion is less than zero.

[0105] A gas turbine engine as claimed in any preceding clause, wherein: the outer band includes a stepped portion; and the stepped portion extends at least partially into the working gas flow path.

[0106] This written description uses examples to disclose the present disclosure, including its best mode, and also to enable any person skilled in the art to practice the present disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the present disclosure is defined by the claims and may include other examples that occur to those skilled in the art. If these other examples include structural elements that do not differ from the literal language of the claims, or include equivalent structural elements with insubstantial differences from the literal language of the claims, then these other examples are intended to be within the scope of the claims.

Claims

1. A gas turbine engine, characterized in that: include: compressor section; A combustion section, the combustion section comprising: Inner bushing, and an outer liner spaced apart from the inner liner, the inner liner and the outer liner at least partially defining a combustion chamber; and A turbine section, the turbine section comprising: an inner belt extending between an upstream side and a downstream side opposite the upstream side, an outer band spaced apart from the inner band and extending between the upstream side and the downstream side, the inner band and the outer band at least partially defining a working gas flow path, wherein one or both of the inner band and the outer band include a step portion adjacent to the upstream side and a main body portion extending from the step portion to the downstream side, and wherein the step portion extends beyond the main body portion in a radial direction, and a plurality of airfoils extending from the inner band, the outer band, or both the inner band and the outer band into the working gas flow path; The compressor section, the combustion section, and the turbine section are in a serial flow order and define at least a portion of the working gas flow path.

2. The gas turbine engine according to claim 1, wherein: in: The inner belt includes the stepped portion; and The step portion at least partially extends into the working gas flow path.

3. The gas turbine engine according to claim 1, characterized in that: in, At least a portion of the stepped portion includes a concave shape.

4. The gas turbine engine according to claim 1, wherein: in, The stepped portion includes a first bent portion and a second bent portion.

5. The gas turbine engine according to claim 4, characterized in that in, At least a portion of the second curved portion is located between the first curved portion and the plurality of airfoils.

6. The gas turbine engine according to claim 4, characterized in that in, The first curved portion is different from the second curved portion.

7. The gas turbine engine according to claim 4, characterized in that in: The first curved portion comprises a convex shape; and The second curved portion includes a concave shape.

8. The gas turbine engine according to claim 4, characterized in that in: The first curved portion comprises a negative concavity; and The second curve includes a positive concavity.

9. The gas turbine engine according to claim 4, characterized in that in: the inner band including a peak distance between the upstream side of the inner band and a peak of the first bend; the inner band comprising a first bend length between the upstream side of the inner band and an end of the first bend adjacent to the second bend; and A ratio of the peak distance to the length of the first curved portion is greater than or equal to 0.1 and less than or equal to 1.

10. The gas turbine engine according to claim 1, wherein: in, The inner liner and the inner band define a chamber between the combustion section and the turbine section.

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