TURBINE ENGINE AERODYNAMIC PROFILE

Abstract

The invention relates to an airfoil comprising a first end, a second end, a leading edge, a trailing edge, a pressure side, and a suction side. The leading and trailing edges are joined by the pressure and suction sides to provide an outer airfoil surface extending in a span direction from the first end of the airfoil to the second end. The outer airfoil surface is formed according to a plurality of cross-sectional profiles of the airfoil described by a set of Cartesian coordinates shown in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location.The local axial chord corresponds to the width of the airfoil between the leading edge and the trailing edge at the span location. Selected figure: Fig. 1.

Description

Title of the invention: AERODYNAMIC PROFILE OF A TURBINE ENGINE Technical field

[0001] This invention relates generally to a turbine engine, and more particularly to an aerodynamic profile for the turbine engine. Context of the invention

[0002] A turbine section in a gas turbine engine typically comprises one or more stator blade groups for conditioning (e.g., guiding, rotating, etc.) combustion products flowing in a flow path. Various airfoil designs are known in the art for such turbine stator blade group applications. Although these known airfoil designs have various advantages, there is always room for improvement in the art. Summary

[0003] According to one aspect of this disclosure, an apparatus is provided for a turbine engine. This apparatus comprises an airfoil, and the airfoil has a first end, a second end, a leading edge, a trailing edge, a pressure side, and a suction side. The leading and trailing edges are joined by the pressure side and the suction side to provide an outer airfoil surface extending in a span direction from the first end of the airfoil to the second end of the airfoil. The outer airfoil surface is formed according to a plurality of cross-sectional profiles of the airfoil described by a set of Cartesian coordinates shown in Table 1.The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading edge and the trailing edge at the span location.

[0004] According to another aspect of this disclosure, a stator blade structure is provided for a turbine engine. This stator blade structure comprises a first platform, a second platform, and a plurality of stator blades arranged circumferentially around an axis in a group. Each stator blade has an airfoil. The airfoil has a leading edge, a trailing edge, a pressure side, and a suction side. The leading and trailing edges are joined by the pressure and suction sides. to provide an outer airfoil surface extending in a span direction from the first platform to the second platform. The outer airfoil surface is formed according to a plurality of cross-sectional airfoil profiles defined by a set of Cartesian coordinates shown in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading and trailing edges at the span location.

[0005] According to yet another aspect of this disclosure, a turbine engine is provided comprising a flow path, a compressor section, a combustion device section, and a turbine section. The flow path extends through the compressor section, the combustion device section, and the turbine section from an inlet in the flow path to an outlet in the flow path. The turbine section comprises a plurality of turbine blades arranged circumferentially around an axis in a group. Each turbine blade has an airfoil located in the flow path. The airfoil has a first end, a second end, a leading edge, a trailing edge, a pressure side, and a suction side.The leading and trailing edges are joined by the pressure and suction sides to provide an outer airfoil surface extending in a span direction from the first end of the airfoil to the second end. The outer airfoil surface is formed according to a plurality of cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates shown in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location. The local axial chord corresponds to the width of the airfoil between the leading and trailing edges at the span location.

[0006] The turbine section may comprise a high-pressure turbine section and a low-pressure turbine section. The low-pressure turbine section may comprise a plurality of turbine blades.

[0007] The turbine blades can be part of a first stage of the low-pressure turbine section.

[0008] The set of Cartesian coordinates shown in Table 1 may have a tolerance of + / - 0.050 inch.

[0009] The turbine engine may also include a coating applied to the outer aerodynamic profile surface.

[0010] The set of Cartesian coordinates shown in Table 1 may have a tolerance of + / - 0.050 inch.

[0011] The span location can correspond to a distance from the axis.

[0012] The stator blades can be turbine blades.

[0013] The outer aerodynamic profile surface can be partially or totally covered with a coating.

[0014] The stator blades may only have thirty-six stator blades.

[0015] The set of Cartesian coordinates shown in Table 1 may have a tolerance of + / - 0.050 inch.

[0016] The span location can correspond to a distance from an axis of rotation of the turbine engine.

[0017] The device may also include an inner platform and an outer platform. The inner platform may be connected to the airfoil at the first end of the airfoil. The outer platform may be connected to the airfoil at the second end of the airfoil.

[0018] The device can be configured as, or otherwise include, a turbine blade.

[0019] The turbine blade can be a low pressure turbine blade.

[0020] The device may also include a coating applied to the outer aerodynamic profile surface.

[0021] The aerodynamic profile can be configured without an internal cooling passage.

[0022] The aerodynamic profile can be one of thirty-six aerodynamic profiles arranged circumferentially around an axis in a ring group.

[0023] This disclosure may include any or more of the individual features disclosed above and / or below alone or in any combination thereof.

[0024] The aforementioned features and the operation of the invention will become more evident in the light of the following description and the accompanying drawings. Brief description of the drawings

[0025] Fig. 1 is a schematic illustration of an aircraft powerplant.

[0026] Figure 2 is a partial schematic side view illustration of a turbine section in the aircraft powerplant.

[0027] Fig. 3 is a partial schematic cross-sectional illustration of a turbine stator blade structure.

[0028] Fig. 4 is a partial schematic cross-sectional illustration of the turbine stator blade structure taken along line 4-4 in Fig. 3.

[0029] Fig. 5 is a perspective illustration of an aerodynamic profile of a turbine stator blade. Detailed description

[0030] Figure 1 illustrates a power unit 20 for an aircraft. The aircraft may be a helicopter, an airplane, a drone (e.g., an unmanned aerial vehicle (UAV)), or any other manned or unmanned aerial vehicle or system. The power unit 20 may be configured as, or part of, a propulsion and / or lift system for the aircraft. The power unit 20 may also, or alternatively, be configured as, or part of, a power supply system for the aircraft. However, this disclosure is not limited to aeronautical applications. The power unit 20, for example, may alternatively be configured as, or part of, a power supply system for ground operation (e.g., an industrial power plant), or other applications.However, for ease of description, the powerplant 20 is described below as an aircraft powerplant.

[0031] The aircraft powerplant 20 of [Fig. 1] comprises a mechanical load 22 and a core 24 of a gas turbine engine 26, where the engine core 24 is configured to power the operation of the mechanical load 22. The mechanical load 22 may be configured as, or alternatively comprise, a rotor 28 mechanically driven by the engine core 24. This driven rotor 28 may be a bladed propeller rotor for the aircraft propulsion and / or lift system. The propeller rotor may be an open propeller rotor (for example, an unfaired propeller rotor) or a faired propeller rotor. For example, when the turbine engine 26 is a turboshaft engine, the open propeller rotor may be a rotor of a gyroplane such as a helicopter main rotor or a helicopter tail rotor. When the turbine engine 26 is a turboprop, the open propeller rotor can be a propeller rotor.When the turbine engine 26 is a turboshaft engine, the shrouded propulsion rotor may be a fan rotor. Alternatively, the driven rotor 28 may be configured as a generator rotor for an electrical power generator for the aircraft's electrical power system; for example, an auxiliary power unit (APU) system. However, this disclosure is not limited to the aforementioned exemplary mechanical loads or the aforementioned exemplary turbine engines. The turbine engine 26, for example, may alternatively be configured as a turbojet, a propfan engine, a tail fan engine, or any other type of turbine engine suitable for powering the operation of the mechanical load 22.

[0032] The turbine motor 26 extends axially along an axis 30 from a forward, upstream end of the turbine motor 26 to a rear, downstream end of the turbine motor 26. In short, this axis 30 can be a centerline axis of the turbine motor 26 and / or of its motor core 24. The axis 30 can also be an axis of rotation of one or more components of the turbine motor 26 and of its motor core 24. The turbine motor 26 of FIG. 1 includes a compressor section 32, a combustion device section 33 and a turbine section 34. The turbine section 34 of [Fig.1] includes a high pressure turbine section (HPT) 34A and a low pressure turbine section (LPT) 34B, which LPT section 34B of [Fig.1] is a working turbine section (PT) to power the operation of the mechanical load 22.

[0033] The compressor section 32 comprises a compressor rotor 36. The HPT section 34A comprises a high-pressure turbine (HPT) rotor 38. The LPT section 34B comprises a low-pressure turbine (LPT) rotor 40. The compressor rotor 36, the HPT rotor 38, and the LPT rotor 40 each comprise one or more groups (e.g., stages) of rotor blades, where the rotor blades in each group are arranged circumferentially around a rotor and are connected to a respective rotor disk or hub. The rotor blades in each group, for example, may be formed as a single piece with the rotor or may be mechanically fastened, welded, brazed, and / or otherwise attached to the respective rotor disk and / or hub.

[0034] The compressor rotor 36 is coupled to and rotates with the HPT rotor 38. The compressor rotor 36 of FIG. 1, for example, is connected to the HPT rotor 38 by a high-speed shaft 42. At least (or only) the compressor rotor 36, the HPT rotor 38, and the high-speed shaft 42 collectively form a high-speed rotating assembly 44; for example, a high-speed housing of the turbine motor 26. The LPT rotor 40 of FIG. 1 is connected to a low-speed shaft 46. At least (or only) the LPT rotor 40 and the low-speed shaft 46 collectively form a low-speed rotating assembly 48; for example, a low-speed housing / working turbine housing of the turbine motor 26. This low-speed rotating assembly 48 is further coupled to the driven rotor 28 via a kinematic chain 50.This kinematic chain 50 can be configured as a geared kinematic chain, where a gear train 52 (e.g., a transmission, a gear-changing device, an epicyclic gear train, etc.) is arranged between and functionally couples the driven rotor 28 to the low-speed rotating assembly 48 and its LPT rotor 40. With this arrangement, the driven rotor 28 can rotate at a different (e.g., slower) speed than the low-speed rotating assembly 48 and its LPT rotor 40. However, the kinematic chain 50 can alternatively be configured as a direct-drive kinematic chain, where the gear train 52 is omitted. With such an arrangement, the rotor 28... driven can rotate at a common (identical) rotational speed with the low-speed rotating assembly 48 and its LPT rotor 40. With reference again to [Fig.1], each rotating assembly 44, 48 and its components can be rotated about the axis 30, and the axis 30 can be a midline axis of each rotating assembly 44, 48 and its components.

[0035] The turbine engine 26 of FIG. 1 has a core flow path 54 (for example, annular). The core flow path 54 extends longitudinally in the turbine engine 26 and its engine core 24 from an air flow inlet 56 in the core flow path 54 to a combustion product exhaust 58 from the core flow path 54. More particularly, the core flow path 54 extends from the core inlet 56, sequentially through the compressor section 32, the combustion device section 33, the HPT section 34A and the LPT section 34B, to the core exhaust 58.

[0036] During the operation of the turbine engine 26, air is directed into the engine core 24 through the core inlet 56. This air entering the core flow path 54 can be called core air. This core air is compressed by the compressor rotor 36 and directed into a combustion chamber 60 (for example, an annular combustion chamber) in a combustion device 62 (for example, an annular combustion device) of the combustion device section 33. Fuel is injected into the combustion chamber 60 by one or more fuel injectors 64 and mixed with the compressed core air to provide a fuel-air mixture. This fuel-air mixture is ignited, and its combustion products flow through the core, sequentially driving the rotation of the HPT rotor 38 and the LPT rotor 40.The rotation of the HPT rotor 38 drives the rotation of the compressor rotor 36, and consequently the compression of the air received from the core inlet 56. The rotation of the LPT rotor 40 drives the rotation of the driven rotor 28. When the driven rotor 28 is configured as the propulsion rotor, its rotation propels additional air (e.g., from the engine core 24 and its core flow path 54) to provide aircraft thrust and / or lift. When the driven rotor 28 is configured as the generator rotor, its rotation can facilitate the generation of electricity.

[0037] [Fig.2] illustrates a section 66 of the turbine engine 26. For ease of description, this engine section 66 is described below as the LPT 34B section in the turbine engine 26. However, it is envisaged that the engine section 66 could alternatively be the HPT 34A section in the turbine engine 26.

[0038] The motor section 66 of [Fig.2] comprises a motor rotor 68 (for example the LPT rotor 40) with a plurality of rotor stages 70A-C (generally indicated by "70"). This motor section 66 also includes a plurality of stator blade structures 72A-C (usually indicated by "72") interspersed with the rotor stages 70. The first stator blade structure 72A of [Fig. 2], for example, is located longitudinally beside and upstream of the first rotor stage 70A. The second stator blade structure 72B is located longitudinally beside and between the first rotor stage 70A and the second rotor stage 70B. The third stator blade structure is located longitudinally beside and between the second rotor stage 70B and the third rotor stage 70C. Here, motor section 66 is represented as a three-stage section of the turbine engine 26; for example, a three-stage LPT 34B section. However, it is envisaged that the engine section 66 and its engine rotor 68 could be configured, alternatively, with a single stage, two stages or more than three stages.

[0039] Each rotor stage 70 comprises a rotor disk 74 and a plurality of rotor blades 76 connected to the rotor disk 74. The rotor blades 76 are arranged circumferentially around the rotor disk 74 and the axis 30 in an annular group. Each of these rotor blades 76 projects in the span direction (for example, radially) from the rotor disk 74 into the core flow path 54.

[0040] With reference to [Fig. 3], each stator blade structure 72 comprises an inner platform 78, an outer platform 80, and a plurality of stator blades 82; for example, low-pressure turbine (LPT) blades. The inner platform 78 extends longitudinally along the core flow path 54 and axially along the axis 30. The inner platform 78 extends circumferentially around the axis 30, thus giving the inner platform 78 a complete loop geometry (for example, tubular). This inner platform 78 forms an inner peripheral boundary of the core flow path 54 longitudinally through the respective stator blade structure 72. The outer platform 80 is spaced radially outwards relative to the inner platform 78. The outer platform 80 extends longitudinally along the core flow path 54 and axially along the axis 30.The outer platform 80 extends circumferentially around the axis 30, thus giving the outer platform 80 a complete loop geometry (e.g., tubular). This outer platform 80 forms an outer peripheral boundary of the core flow path 54 longitudinally through the respective stator blade structure 72. The stator blades 82 are arranged circumferentially around the axis 30 in an annular group. These stator blades 82 are positioned between and connected to (e.g., formed as a single unit with or fixed to) the inner platform 78 and the outer platform 80. Each stator blade 82 extends in the span direction (e.g., radially) through the core flow path 54 of the platform. interior 78 to exterior platform 80. With such an arrangement, with reference to [Fig.2], each stator blade structure 72 is configured to condition (e.g. guide, rotate, etc.) the air that is evacuated from a respective rotor stage 70 and / or to condition (e.g. guide, rotate, etc.) the air that is directed to a respective rotor stage 70.

[0041] With reference to [Fig.4], each stator blade 82 comprises an aerodynamic profile 84; for example, a turbine blade aerodynamic profile. This airfoil 84 extends along the chord from a leading edge 86, upstream, of the airfoil 84 to a trailing edge 88, downstream, of the airfoil 84. The airfoil 84 extends laterally from a concave pressure side 90 of the airfoil 84 to a convex suction side 92 of the airfoil 84. The pressure side 90 and the suction side 92 extend along the chord between and meet at the leading edge 86 and the trailing edge 88. With reference to [Fig. 5], each airfoil element 86, 88, 90, and 92 extends along the span (e.g., radially) from a base end 94 of the airfoil 84 to a tip end 96 of the airfoil 84. With reference to [Fig.[3] The airfoil 84 is connected to the inner platform 78 at (for example on, adjacent to or near) the base end 94. The airfoil 84 is connected to the outer platform 80 at the tip end 96. The airfoil 84 is thus provided with an outer airfoil surface 98 which extends in the span direction from the base end 94 / the inner platform 78 to the tip end 96 / the outer platform 80. This outer airfoil surface 98 is formed by the leading edge 86, the trailing edge 88, the pressure side 90 and the suction side 92 of the airfoil 84. The outer airfoil surface 98 guides the combustion products flowing through the core flow path 54.

[0042] With reference to [Fig. 3], a geometry of the outer airfoil surface 98 is described below in terms of Cartesian coordinates defined on an x-axis, a y-axis, and a z-axis. The x-axis can be an axial direction parallel to the 30-axis. The y-axis can be a circumferential direction around the 30-axis (see also [Fig. 4]), where the y-axis is perpendicular to the x-axis. The z-axis can be a radial direction out of the 30-axis, where the z-axis is perpendicular to both the x-axis and the y-axis. More specifically, the geometry of the outer airfoil surface 98 is formed according to a plurality of cross-sectional profiles of the airfoil 84 as described by a set of Cartesian coordinates presented in Table 1 below. In Table 1, cross-sectional profiles are provided for three positions Z1-Z3 in the span direction (e.g. z coordinates) along the airfoil 84.The ZI position is . Position Z1 is located one-quarter (1 / 4) wingspan upwards from the base end 94 / inner platform 78 along the z-axis, where the wingspan coordinate (AZ1) is the radial distance from axis 30 to position Z1. Position Z2 is located one-half (1 / 2) wingspan upwards from the base end 94 / inner platform 78 along the z-axis, where the wingspan coordinate (AZ2) is the radial distance from axis 30 to position Z2. Position Z3 is located three-quarters (3 / 4) wingspan upwards from the base end 94 / inner platform 78 along the z-axis, where the wingspan coordinate (AZ3) is the radial distance from axis 30 to position Z3.

[0043] The axial (x) and circumferential (y) coordinates in Table 1 for each of the cross-sectional profiles are normalized by a local axial chord (Bx) for the cross-sectional profiles at the respective span coordinate (AZ1, AZ2, AZ3). By way of example, the local axial chord (Bx) for the axial (x) and circumferential (y) coordinates associated with the quarter-span coordinate (AZ1) corresponds to a width of the airfoil 84 between the leading edge 86 and the trailing edge 88 at the location of one quarter (1 / 4) of the span ZL

[0044] The axial (x) and circumferential (y) coordinates in Table 1 for each of the cross-sectional profiles at the respective span coordinate (AZ1, AZ2, AZ3) describe a contour of the outer airfoil surface 98 at that respective span coordinate (AZ1, AZ2, AZ3). This contour of the outer airfoil surface 98 is formed by joining adjacent points in Table 1 in a regular manner in the xy plane. The three-dimensional outer airfoil surface 98 is formed by joining adjacent cross-sectional profiles in a regular manner along the span – the z-axis. The manufacturing tolerance with respect to the specified coordinates is + / - 0.050 inch (+ / - 1.27 millimeter).The coordinates in Table 1 define points on a cold, uncoated, and fixed airfoil surface in a plane at the corresponding wingspan location. However, the airfoil 84 and its outer airfoil surface 98 may be coated after fabrication. Here, it is envisaged that additional features such as one or more cooling holes, protective coatings, fillets, sealing structures, and / or others may also be formed by, in, and / or on the outer airfoil surface 98 in other embodiments; but these additional features may not be defined by the standardized coordinates in Table 1. [Tables 1] TABLE 1 REFERENCE RADIUS: AZ1 CUTTING COORDINATES (X, Y) / Bxl 0.000 0.375 0.001 0.387 0.005 0.393 0.012 0.399 0.021 0.403 0.031 0.404 0.042 0.404 0.055 0.402 0.075 0.396 0.100 0.388 0.125 0.379 0.150 0.369 0.175 0.357 0.200 0.343 0.225 0.329 0.250 0.313 0.275 0.296 0.300 0.277 0.325 0.257 0.350 0.236 0.375 0.213 0.400 0.189 0.425 0.164 0.450 0.137 0.475 0.109 0.500 0.081 0.525 0.051 0.550 0.020 0.575 -0.010 0.600 -0.042 0.625 -0.074 0.650 -0.107 0.675 -0.141 0.700 -0.175 0.725 -0.210 0.750 -0.246 0.775 -0.282 0.800 -0.319 0.825 -0.356 0.850 -0.395 0.875 -0.434 0.900 -0.474 0.925 -0.515 0.943 -0.546 0.958 -0.573 0.971 -0.595 0.981 -0.613 0.987 -0.624 0.992 -0.633 0.996 -0.640 0.999 -0.645 1.000 -0.650 0.999 -0.655 0.996 -0.660 0.992 -0.663 0.987 -0.665 0.981 -0.664 0.971 -0.654 0.958 -0.633 0.943 -0.609 0.925 -0.582 0.900 -0.545 0.875 -0.510 0.850 -0.476 0.825 -0.443 0.800 -0.411 0.775 -0.379 0.750 -0.349 0.725 -0.319 0.700 -0.290 0.675 -0.261 0.650 -0.234 0.625 -0.206 0.600 -0.180 0.575 -0.154 0.550 -0.128 0.525 -0.103 0.500 -0.078 0.475 -0.054 0.450 -0.030 0.425 -0.007 0.400 0.016 0.375 0.038 0.350 0.061 0.325 0.083 0.300 0.105 0.275 0.127 0.250 0.148 0.225 0.169 0.200 0.190 0.175 0.211 0.150 0.232 0.125 0.252 0.100 0.273 0.075 0.293 0.055 0.310 0.042 0.320 0.031 0.329 0.021 0.338 0.012 0.348 0.005 0.360 0.001 0.369 REFERENCE RADIUS: AZ2 CUTTING COORDINATES (X, Y) / Bx2 0.000 0.364 0.001 0.369 0.005 0.375 0.012 0.382 0.021 0.386 0.031 0.388 0.042 0.388 0.055 0.386 0.075 0.383 0.100 0.378 0.125 0.372 0.150 0.364 0.175 0.354 0,200 0,344 0,225 0,332 0,250 0,318 0,275 0,303 0,300 0,287 0,325 0,269 0,350 0,250 0,375 0,230 0,400 0,208 0,425 0,184 0,450 0,159 0,475 0,133 0,500 0,105 0,525 0,077 0,550 0,047 0,575 0,016 0,600 -0,015 0,625 -0,048 0,650 -0,081 0,675 -0,115 0,700 -0,150 0,725 -0,186 0,750 -0,222 0,775 -0,259 0,800 -0,297 0,825 -0,336 0,850 -0,376 0,875 -0,417 0,900 -0,460 0,925 -0,503 0,943 -0,535 0,958 -0,564 0,971 -0,588 0,981 -0,607 0,987 -0,619 0,992 -0,628 0,996 -0,636 0,999 -0,641 1,000 -0,646 0,999 -0,651 0,996 -0,656 0,992 -0,659 0,987 -0,660 0,981 -0,658 0,971 -0,645 0,958 -0,622 0,943 -0,596 0,925 -0,567 0,900 -0,528 0,875 -0,491 0,850 -0,456 0,825 -0,421 0,800 -0,388 0,775 -0,356 0,750 -0,325 0,725 -0,295 0,700 -0,266 0,675 -0,237 0,650 -0,209 0,625 -0,182 0.600 -0.156 0.575 -0.130 0.550 -0.104 0.525 -0.079 0.500 -0.055 0.475 -0.031 0.450 -0.007 0.425 0.015 0.400 0.037 0.375 0.060 0.350 0.081 0.325 0.103 0.300 0.124 0.275 0.144 0.250 0.164 0.225 0.184 0.200 0.203 0.175 0.222 0.150 0.240 0.125 0.258 0.100 0.275 0.075 0.291 0.055 0.304 0.042 0.312 0.031 0.319 0.021 0.327 0.012 0.335 0.005 0.345 0.001 0.353 REFERENCE RADIUS: AZ3 CUTTING COORDINATES (X, Y) / Bx3 0.000 0.348 0.001 0.358 0.005 0.363 0.012 0.370 0.021 0.374 0.031 0.376 0.042 0.377 0.055 0.377 0.075 0.377 0.100 0.374 0.125 0.371 0.150 0.366 0.175 0.359 0.200 0.351 0.225 0.343 0.250 0.332 0.275 0.320 0.300 0.306 0.325 0.292 0.350 0.275 0.375 0.257 0.400 0.238 0.425 0.217 0.450 0.194 0.475 0.170 0.500 0.145 0.525 0.118 0.550 0.089 0.575 0.060 0,600 0,028 0,625 -0,003 0,650 -0,036 0,675 -0,071 0,700 -0,106 0,725 -0,143 0,750 -0,180 0,775 -0,219 0,800 -0,259 0,825 -0,300 0,850 -0,342 0,875 -0,386 0,900 -0,431 0,925 -0,478 0,943 -0,512 0,958 -0,543 0,971 -0,569 0,981 -0,590 0,987 -0,603 0,992 -0,613 0,996 -0,622 0,999 -0,627 1,000 -0,632 0,999 -0,637 0,996 -0,641 0,992 -0,644 0,987 -0,645 0,981 -0,642 0,971 -0,626 0,958 -0,601 0,943 -0,572 0,925 -0,541 0,900 -0,499 0,875 -0,459 0,850 -0,422 0,825 -0,386 0,800 -0,351 0,775 -0,318 0,750 -0,285 0,725 -0,254 0,700 -0,224 0,675 -0,195 0,650 -0,166 0,625 -0,139 0,600 -0,112 0,575 -0,086 0,550 -0,060 0,525 -0,035 0,500 -0,011 0,475 0,012 0,450 0,035 0,425 0,057 0,400 0,079 0,375 0,100 0,350 0,120 0,325 0,139 0,300 0,158 0,275 0,176 0,250 0,193 0,225 0,210 0,200 0,225 0,175 0,240 0,150 0,254 0,125 0,268 0,100 0,280 0,075 0,292 0,055 0,301 0,042 0,307 0,031 0,312 0,021 0,318 0,012 0,326 0,005 0,335 0,001 0,343

[0045] The set of points defined by the coordinates above in Table 1 represents an original and unique aerodynamic profile well suited for use in the turbine section 34 of the turbine engine 26. More particularly, the set of points defined by the coordinates above in Table 1 represents an original and unique aerodynamic profile well suited for use in the LPT section 34B, for example in the first stage of the LPT section 34B; for example in the stator blade group 72A of [Fig. 2]. In the first stage of the LPT, the stator blade structure 72 can comprise a total of thirty-six (36) stator blades 82 / aerodynamic profiles 84 arranged circumferentially around the axis 30 in the group.

[0046] In general, the aerodynamic profile 84 described herein has a combination of axial sweep and tangential inclination. Depending on the specific configuration, the inclination and sweep angles sometimes vary by up to plus or minus ten degrees (+ / - 10°) or more. In addition, the stator blade 82 and its aerodynamic profile 84 can be rotated about a radial axis or a line normal to the surface of the inner platform 78 or the deflector, for example, by up to plus or minus ten degrees (+ / - 10°) or more.

[0047] Original aspects of the stator blade 82 and its outer aerodynamic profile surface 98 described herein are achieved by substantial conformity to specified geometries. Substantial conformity generally includes, or may include, a manufacturing tolerance of + / - 0.050 inch (+ / - 1.27 millimeter), in order to hold This tolerance accounts for variations in molding, cutting, shaping, surface finishing, and other manufacturing processes, and accommodates variability in coating thicknesses. This tolerance is generally constant or non-adjustable and applies to each specified stator blade surface, regardless of stator blade size.

[0048] Substantive conformity is based on point sets representing a three-dimensional surface with specific physical dimensions, for example in inches or millimeters, as determined by selecting particular values ​​of the scaling parameters. A substantially conforming airfoil or stator blade has surfaces that conform to the specified point sets, within the specified tolerance.

[0049] Alternatively, substantial conformity is based on a determination by a national or international regulatory body, for example, in a part certification or part manufacturing approval (PMA) process for the Federal Aviation Administration, the European Aviation Safety Agency, the Civil Aviation Administration of China, the Japan Civil Aviation Bureau, or the Federal Agency for Air Transport of the Russian Federation. In these configurations, substantial conformity encompasses a determination that a particular part or structure is identical, or sufficiently similar, to the specified airfoil or stator blade, or that the part or structure meets the airworthiness standards applicable to the specified stator blade or airfoil.In particular, substantial conformity encompasses any regulatory determination that a particular part or structure is sufficiently similar, identical, or alike to a specified stator blade or airfoil, such that certification or authorization for use is based at least in part on the determination of similarity.

[0050] Each stator blade 82 and its airfoil 84 can be constructed from a high-strength, heat-resistant material such as a nickel-based or cobalt-based superalloy, or from a ceramic or composite material resistant to high temperatures and stresses. One or more thermal barrier coatings, abrasion-resistant coatings, or other protective coatings may alternatively be applied to the airfoil 84. Although the airfoil 84 is generally described above as being configured without any internal cooling, it is envisaged that the airfoil 84 may alternatively be modified to include one or more internal cooling passages with or without one or more cooling holes in the outer airfoil surface 98.

[0051] Although various embodiments of this disclosure have been described, it will be obvious to those skilled in the art that many other embodiments and implementations are possible within the scope of the disclosure. For example, the disclosure as described herein has several aspects and embodiments that have particular features. Although these features can be described individually, it is within the scope of this disclosure that some or all of these features can be combined with any of the aspects and remain within the scope of the disclosure. Accordingly, this disclosure should not be restricted except in light of the accompanying claims and their equivalents.

Claims

1. Demands Apparatus for a turbine engine, comprising: an aerodynamic profile comprising a first end, a second end, a leading edge, a trailing edge, a pressure side and a suction side; the leading edge and trailing edge being joined by the pressure side and the suction side to provide an outer airfoil surface extending in a span direction from the first end of the airfoil to the second end of the airfoil; the outer aerodynamic profile surface being formed in accordance with a plurality of cross-sectional profiles of the aerodynamic profile described by a set of Cartesian coordinates presented in Table 1; [Tables 1] TABLE 1 REFERENCE RADIUS: AZ1 CUTTING COORDINATES (X, Y) / Bxl 0.000 0.375 0.001 0.387 0.005 0.393 0.012 0.399 0.021 0.403 0.031 0.404 0.042 0.404 0.055 0.402 0.075 0.396 0.100 0.388 0.125 0.379 0.150 0.369 0.175 0.357 0.200 0.343 0.225 0.329 0.250 0.313 0.275 0.296 0.300 0.277 0.325 0.257 0.350 0.236 0.375 0.213 0.400 0.189 0.425 0.164 0.450 0.137 0.475 0.109 0.500 0.081 0.525 0.051 0.550 0.020 0.575 -0.010 0.600 -0.042 0.625 -0.074 0.650 -0.107 0.675 -0.141 0.700 -0.175 0.725 -0.210 0.750 -0.246 0.775 -0.282 0.800 -0.319 0.825 -0.356 0.850 -0.395 0.875 -0.434 0.900 -0.474 0.925 -0.515 0.943 -0.546 0.958 -0.573 0.971 -0.595 0.981 -0.613 0.987 -0.624 0.992 -0.633 0.996 -0.640 0.999 -0.645 1.000 -0.650 0.999 -0.655 0.996 -0.660 0.992 -0.663 0.987 -0.665 0.981 -0.664 0.971 -0.654 0.958 -0.633 0.943 -0.609 0.925 -0.582 0.900 -0.545 0.875 -0.510 0.850 -0.476 0.825 -0.443 0.800 -0.411 0.775 -0.379 0.750 -0.349 0.725 -0.319 0.700 -0.290 0.675 -0.261 0.650 -0.234 0.625 -0.206 0.600 -0.180 0.575 -0.154 0.550 -0.128 0.525 -0.103 0.500 -0.078 0.475 -0.054 0.450 -0.030 0.425 -0.007 0.400 0.016 0.375 0.038 0.350 0.061 0.325 0.083 0.300 0.105 0.275 0.127 0.250 0.148 0.225 0.169 0.200 0.190 0.175 0.211 0.150 0.232 0.125 0.252 0.100 0.273 0.075 0.293 0.055 0.310 0.042 0.320 0.031 0.329 0.021 0.338 0.012 0.348 0.005 0.360 0.001 0.369 REFERENCE RADIUS: AZ2 CUTTING COORDINATES (X, Y) / Bx2 0.000 0.364 0.001 0.369 0.005 0.375 0.012 0.382 0.021 0.386 0.031 0.388 0.042 0.388 0.055 0.386 0.075 0.383 0.100 0.378 0.125 0.372 0.150 0.364 0.175 0.354 0,200 0,344 0,225 0,332 0,250 0,318 0,275 0,303 0,300 0,287 0,325 0,269 0,350 0,250 0,375 0,230 0,400 0,208 0,425 0,184 0,450 0,159 0,475 0,133 0,500 0,105 0,525 0,077 0,550 0,047 0,575 0,016 0,600 -0,015 0,625 -0,048 0,650 -0,081 0,675 -0,115 0,700 -0,150 0,725 -0,186 0,750 -0,222 0,775 -0,259 0,800 -0,297 0,825 -0,336 0,850 -0,376 0,875 -0,417 0,900 -0,460 0,925 -0,503 0,943 -0,535 0,958 -0,564 0,971 -0,588 0,981 -0,607 0,987 -0,619 0,992 -0,628 0,996 -0,636 0,999 -0,641 1,000 -0,646 0,999 -0,651 0,996 -0,656 0,992 -0,659 0,987 -0,660 0,981 -0,658 0,971 -0,645 0,958 -0,622 0,943 -0,596 0,925 -0,567 0,900 -0,528 0,875 -0,491 0,850 -0,456 0,825 -0,421 0,800 -0,388 0,775 -0,356 0,750 -0,325 0,725 -0,295 0,700 -0,266 0,675 -0,237 0,650 -0,209 0,625 -0,182 0.600 -0.156 0.575 -0.130 0.550 -0.104 0.525 -0.079 0.500 -0.055 0.475 -0.031 0.450 -0.007 0.425 0.015 0.400 0.037 0.375 0.060 0.350 0.081 0.325 0.103 0.300 0.124 0.275 0.144 0.250 0.164 0.225 0.184 0.200 0.203 0.175 0.222 0.150 0.240 0.125 0.258 0.100 0.275 0.075 0.291 0.055 0.304 0.042 0.312 0.031 0.319 0.021 0.327 0.012 0.335 0.005 0.345 0.001 0.353 REFERENCE RADIUS: AZ3 CUTTING COORDINATES (X, Y) / Bx3 0.000 0.348 0.001 0.358 0.005 0.363 0.012 0.370 0.021 0.374 0.031 0.376 0.042 0.377 0.055 0.377 0.075 0.377 0.100 0.374 0.125 0.371 0.150 0.366 0.175 0.359 0.200 0.351 0.225 0.343 0.250 0.332 0.275 0.320 0.300 0.306 0.325 0.292 0.350 0.275 0.375 0.257 0.400 0.238 0.425 0.217 0.450 0.194 0.475 0.170 0.500 0.145 0.525 0.118 0.550 0.089 0.575 0.060 0.600 0.028 0.625 -0.003 0.650 -0.036 0.675 -0.071 0.700 -0.106 0.725 -0.143 0.750 -0.180 0.775 -0.219 0.800 -0.259 0.825 -0.300 0.850 -0.342 0.875 -0.386 0.900 -0.431 0.925 -0.478 0.943 -0.512 0.958 -0.543 0.971 -0.569 0.981 -0.590 0.987 -0.603 0.992 -0.613 0.996 -0.622 0.999 -0.627 1.000 -0.632 0.999 -0.637 0.996 -0.641 0.992 -0.644 0.987 -0.645 0.981 -0.642 0.971 -0.626 0.958 -0.601 0.943 -0.572 0.925 -0.541 0.900 -0.499 0.875 -0.459 0.850 -0.422 0.825 -0.386 0.800 -0.351 0.775 -0.318 0.750 -0.285 0.725 -0.254 0.700 -0.224 0.675 -0.195 0.650 -0.166 0.625 -0.139 0.600 -0.112 0.575 -0.086 0.550 -0.060 0.525 -0.035 0.500 -0.011 0.475 0.012 0.450 0.035 0.425 0.057 0.400 0.079 0.375 0.100 0.350 0.120 0.325 0.139 0.300 0.158 0.275 0.176 0.250 0.193 0.225 0.210 0.200 0.225 0.175 0.240 0.150 0.254 0.125 0.268 0.100 0.280 0.075 0.292 0.055 0.301 0.042 0.307 0.031 0.312 0.021 0.318 0.012 0.326 0.005 0.335 0.001 0.343

2.

3.

4.

5. the Cartesian coordinates being provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by the local axial chord, and a span location; and the local axial chord corresponding to a width of the aerodynamic profile between the leading edge and the trailing edge at the span location. Apparatus according to claim 1, wherein the set of Cartesian coordinates shown in Table 1 has a tolerance of + / - 0.050 inch. Apparatus according to claim 1 or 2, further comprising: an inner platform connected to the aerodynamic profile at the first end of the aerodynamic profile; and an outer platform connected to the aerodynamic profile at the second end of the aerodynamic profile. An apparatus according to any one of claims 1 to 3, wherein the turbine blade is a low-pressure turbine blade. An apparatus according to any one of the preceding claims, further comprising a coating applied to the outer airfoil surface.

6. Apparatus according to any prior claim, wherein the aerodynamic profile is configured without internal cooling passage.

7. Apparatus according to any prior claim, wherein the aerodynamic profile is one of thirty-six aerodynamic profiles arranged circumferentially around an axis in an annular group.

8. Stator blade structure for a turbine engine, comprising: a first platform; a second platform; and a plurality of stator blades arranged circumferentially around an axis in a group, each of the plurality of stator blades comprising an apparatus according to any preceding claim.

9. Turbine engine, comprising: a flow path, a compressor section, a combustion device section and a turbine section; the flow path extending through the compressor section, the combustion device section and the turbine section from an inlet in the flow path to an outlet in the flow path; the turbine section comprising a plurality of turbine blades arranged circumferentially around an axis in a group, each of the plurality of turbine blades comprising an apparatus according to any one of claims 1 to 7, the aerodynamic profile of each apparatus being located in the flow path.

10. Turbine engine according to claim 9, wherein the turbine section comprises a high-pressure turbine section and a low-pressure turbine section, the low-pressure turbine section comprises the plurality of turbine blades, and the plurality of turbine blades are part of a first stage of the low-pressure turbine section.

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