Turbine engine with blade assembly having set of cooling conduits

Abstract

A gas turbine engine having a blade assembly having a platform, an airfoil, and a shank. The airfoil has a plurality of cooling conduits, and the shank has a plurality of inlet channels to provide cooling fluid to the cooling conduits in the airfoil. Cooling fluid exits along the airfoil through the plurality of cooling holes. The blade assembly has a particular geometry that improves durability.

Description

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS

[0002] This patent claims the benefit of U.S. Provisional Patent Application No. 63 / 562,542, filed on March 7, 2024, entitled "Turbine Engine with Blade Assembly Having a Set of Cooling Ducts." Priority is hereby claimed to U.S. Provisional Patent Application No. 63 / 562,542. Technical Field

[0003] The present invention generally relates to a blade assembly for a turbine engine and, more particularly, to a turbine engine with a blade assembly having a set of cooling ducts. Background Art

[0004] In some embodiments, a gas turbine engine typically includes a turbine with a fan. The turbine typically includes a compressor, a combustor, and a turbine arranged in series flow. The compressor compresses air that is directed to the combustor, where the air is mixed with fuel. The mixture is then ignited to produce hot combustion gases. The combustion gases are directed to the turbine, which extracts energy from the combustion gases to power the compressor and fan (if used), as well as to generate useful work to propel the aircraft in flight or to power a load such as a generator.

[0005] During operation of a gas turbine engine, various systems generate relatively large amounts of heat and stress. For example, during operation of the thrust generation system, lubrication system, electric motors and / or generators, hydraulic system, or other systems, significant amounts of heat or stress may be generated. Designs that reduce the thermal loads and / or stresses on engine components are advantageous. BRIEF DESCRIPTION OF THE DRAWINGS

[0006] A full and enabling disclosure of the present disclosure, including the best mode thereof, is set forth in this specification to those skilled in the art, and reference is made to the accompanying drawings, in which:

[0007] Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0008] Figure 2 According to an exemplary embodiment of the present disclosure Figure 1 Schematic cross-sectional view of the turbine section of a gas turbine engine.

[0009] Figure 3 is a perspective view of a blade assembly according to an exemplary embodiment of the present disclosure, the blade assembly including a Figure 1 Leading edge inlet passage in a gas turbine engine.

[0010] Figure 4It is used for calculation Figure 3 Schematic diagram of the stator rotor seal radius of the blade assembly.

[0011] Figure 5 According to an exemplary embodiment of the present disclosure Figure 3 A side view of a blade assembly showing multiple planes.

[0012] Figure 6A It is along Figure 5 The first plane of the plurality of planes Figure 3-5 A top cross-sectional view of the shank of a blade assembly.

[0013] Figure 6B It is along Figure 5 The second plane of the plurality of planes Figure 3-5 A top cross-sectional view of the shank of a blade assembly.

[0014] Figure 6C It is along Figure 5 The third plane of multiple planes Figure 3-5 A top cross-sectional view of the shank of a blade assembly.

[0015] Figure 6D It is along Figure 5 The fourth plane of the multiple planes Figure 3-5 A top cross-sectional view of the shank of a blade assembly.

[0016] Figure 6E It is along Figure 5 The fifth plane of the multiple planes Figure 3-5 A top cross-sectional view of the shank of a blade assembly.

[0017] Figure 6F It is along Figure 5 The sixth plane of the multiple planes Figure 3-5 A top cross-sectional view of the shank of a blade assembly.

[0018] Figure 6G It is along Figure 5 The seventh plane of multiple planes Figure 3-5 A top cross-sectional view of the shank of a blade assembly.

[0019] Figure 7 It is along Figure 5 The eighth plane of multiple planes Figure 3-5 A top cross-sectional view of an airfoil of a blade assembly.

[0020] Figure 8 It is along Figure 5 The ninth plane of the multiple planes Figure 3-5 A top cross-sectional view of an airfoil of a blade assembly. DETAILED DESCRIPTION

[0021] Reference will now be made in detail to embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numbers and letters to refer to features in the drawings. The same or similar reference numerals have been used in the drawings and the description to refer to the same or similar parts of the present disclosure.

[0022] Aspects of the present disclosure generally relate to a blade assembly having ducts within the blade assembly. Specifically, the blade assembly includes an airfoil having a plurality of cooling ducts and a shank having a plurality of inlet passages. The airfoil also includes cooling holes fluidly coupled to the plurality of cooling ducts within the airfoil.

[0023] The blade assembly may be in a turbine section of a gas turbine engine. For example, the blade assembly may be a first stage blade assembly of a high pressure turbine, which is typically subject to the highest thermal and mechanical stresses.

[0024] The shank can be used to attach the blade assembly to the turbine disk. In some embodiments, the shank is formed as a dovetail that is received in the turbine disk. The blade assembly platform, along with other circumferentially arranged platforms and seals of other blade assemblies, defines a substantially continuous annular ring that limits (e.g., prevents or reduces) hot gas leakage from the flow path into the turbine disk cavity. The airfoil extends radially from the platform, away from the turbine disk, while the shank extends radially from the platform, toward the turbine disk.

[0025] High engine temperatures and operating forces impose relatively large thermal and mechanical stresses on the blade assembly. Furthermore, cooling ducts and inlet passages in the blade assembly cause stress redistribution (i.e., the cooling ducts and inlet passages can act as stress concentrators, etc.). For example, the dimensions of the cooling ducts and inlet passages affect the thickness of the airfoil and shank walls, which affects the stress distribution in the blade assembly. Relatively large stresses can lead to unexpected or premature component replacement. Therefore, a blade assembly with greater durability is needed to increase on-wing time.

[0026] Unless otherwise specified, connection references (e.g., connect, couple, attach, and join) are to be interpreted broadly and may include intermediate structural elements between a collection of elements as well as relative movement between elements. Thus, connection references do not necessarily infer that two elements are directly connected and in a fixed relationship to each other. The exemplary figures are for illustrative purposes only, and the size, position, order, and relative sizes reflected in the accompanying drawings may vary.

[0027] As used herein, a "stage" of a gas turbine engine's compressor or turbine is a set of blade assemblies and an adjacent set of vane assemblies, both of which are arranged circumferentially about the engine's centerline. A pair of circumferentially adjacent vanes in the set of vane assemblies is called a nozzle. The blade assembly rotates relative to the engine's centerline. In one example, the blade assembly is mounted to a rotating structure, such as a disk.

[0028] As used herein, the word "exemplary" means "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, all embodiments described herein should be considered exemplary unless expressly stated otherwise.

[0029] As used herein, the terms “first,” “second,” “third,” and “fourth” are used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of each component.

[0030] As used herein, a "set" of elements or a group of elements may include any number of the elements, including one.

[0031] As used herein, the terms "fore" and "aft" refer to relative positions within a gas turbine engine and refer to the normal operating attitude or direction of travel of the gas turbine engine. For example, with respect to a gas turbine engine, the forward position refers to a position relatively close to the nose of an aircraft, and the aft position refers to a position relatively close to the tail of the aircraft.

[0032] As used herein, the terms "upstream" and "downstream" refer to positions along a fluid flow path relative to the direction of fluid flow along the flow path and to a reference position along the fluid flow path.

[0033] As used herein, the term "fluid" refers to a gas or a liquid, and "fluid coupling" means that a fluid can flow between coupled areas.

[0034] As used herein, the forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0035] As used herein, a radial direction (denoted as "R") is a direction perpendicular to a base plane on the shank of a blade assembly.

[0036] As used herein, the axial direction (denoted as "A") is the direction perpendicular to the shank leading edge plane on the shank of the blade assembly and perpendicular to the radial direction.

[0037] As used herein, the tangential direction (denoted as "T") is a direction perpendicular to the radial direction and the axial direction.

[0038] The stator rotor seal radius (denoted as "SRSR") is the radius of curvature of the upper edge of the stator rotor seal on the blade assembly.

[0039] The term redline exhaust gas temperature (referred to herein as "redline EGT") refers to the maximum allowable takeoff temperature as recorded in a Federal Aviation Administration ("FAA") type certificate data sheet. For example, in certain exemplary embodiments, the term redline EGT may refer to the maximum allowable takeoff temperature of the airflow after the first stage stator downstream of the HP turbine of the engine, which is the temperature the engine is rated to withstand. The term redline EGT is also sometimes referred to as indicated turbine exhaust temperature or indicated turbine temperature.

[0040] The term redline core speed (referred to herein as "redline CS") refers to the maximum rotational speed of a gas turbine engine permitted at takeoff, as recorded in the FAA type certificate data sheet. For example, the redline core speed CS is the rate at which the drive shaft for the gas turbine engine rotates, which defines the rotational speed of a particular blade assembly within the gas turbine engine.

[0041] As used herein, "cooling ducts" refer to flow paths formed in a blade assembly that carry a cooling fluid.

[0042] As used herein, a "cooling duct portion" refers to a discrete portion of a cooling duct in a cross-sectional plane of a bucket assembly.

[0043] As used herein, "inlet passage" refers to a cooling duct formed in the shank of a bucket assembly.

[0044] As used herein, "leading edge inlet passage" refers to an inlet passage in the shank of a blade assembly adjacent the leading edge of the blade assembly.

[0045] As used herein, "trailing edge inlet passage" refers to an inlet passage in the shank of a blade assembly adjacent the trailing edge of the blade assembly.

[0046] As used herein, "intermediate inlet channel" refers to an inlet channel between the leading edge inlet channel and the trailing edge inlet channel.

[0047] As used herein, a "base plane" of a blade assembly refers to the plane defined by the base of the shank of the blade assembly.

[0048] As used herein, the "second normalized area" (denoted as ) refers to the ratio of (1) the cross-sectional area of ​​the leading edge inlet passage at a second plane spaced 0.0182 meters radially from the base plane of the blade assembly to (2) the cross-sectional area of ​​the leading edge inlet passage at a first plane spaced 0.0167 meters radially from the base plane.

[0049] As used herein, the "third normalized area" (denoted as ) refers to the ratio of (1) the cross-sectional area of ​​the leading edge inlet passage at a third plane spaced at a radial distance of 0.0212 meters from the base plane of the blade assembly to (2) the cross-sectional area of ​​the leading edge inlet passage at a first plane spaced at a radial distance of 0.0167 meters from the base plane.

[0050] As used herein, the "fourth normalized area" (denoted as ) refers to the ratio of (1) the cross-sectional area of ​​the leading edge inlet passage at a fourth plane spaced at a radial distance of 0.0227 meters from the base plane of the blade assembly to (2) the cross-sectional area of ​​the leading edge inlet passage at a first plane spaced at a radial distance of 0.0167 meters from the base plane.

[0051] As used herein, the "fifth normalized area" (denoted as ) refers to the ratio of (1) the cross-sectional area of ​​the leading edge inlet passage at a fifth plane spaced 0.0277 meters radially from the base plane of the blade assembly to (2) the cross-sectional area of ​​the leading edge inlet passage at a first plane spaced 0.0167 meters radially from the base plane.

[0052] As used herein, the "sixth normalized area" (denoted as ) refers to the ratio of (1) the cross-sectional area of ​​the leading edge inlet passage at a sixth plane spaced at a radial distance of 0.0295 meters from the base plane of the blade assembly to (2) the cross-sectional area of ​​the leading edge inlet passage at a first plane spaced at a radial distance of 0.0167 meters from the base plane.

[0053] As used herein, the "seventh normalized area" (denoted as ) refers to the ratio of (1) the cross-sectional area of ​​the leading edge inlet passage at a seventh plane spaced 0.0341 meters radially from the base plane of the blade assembly to (2) the cross-sectional area of ​​the leading edge inlet passage at a first plane spaced 0.0167 meters radially from the base plane.

[0054] The suction side cross-sectional area (denoted as "SSA") is the average of the sum of the cross-sectional areas of three cooling ducts located along the suction side of the airfoil of the blade assembly, taken at (1) an eighth plane spaced at a radial distance of 0.0400 meters from the base plane of the blade assembly and (2) a ninth plane spaced at a radial distance of 0.0476 meters from the base plane of the blade assembly.

[0055] All measurements referred to herein were performed on the blade assembly prior to use or as a cold part.

[0056] Referring now to the accompanying drawings, Figure 1is a schematic diagram of a gas turbine engine 10. As a non-limiting example, the gas turbine engine 10 may be used in an aircraft. The gas turbine engine 10 includes an engine core extending along an engine centerline 20 and includes at least a compressor section 12, a combustor 14, and a turbine section 16 arranged in series flow. In some examples, the gas turbine engine 10 includes a fan (not shown) that is driven by the engine core to provide thrust and to supply air to the compressor section 12. The gas turbine engine 10 includes a drive shaft 18 that rotationally couples the fan, the compressor section 12, and the turbine section 16 so that rotation of one affects rotation of the other and defines an axis of rotation for the gas turbine engine 10 along the engine centerline 20.

[0057] In the illustrated example, compressor section 12 includes a low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24 fluidly coupled in series with one another. Turbine section 16 includes a high-pressure (HP) turbine 26 and a low-pressure (LP) turbine 28 fluidly coupled in series with one another. A drive shaft 18 operably couples LP compressor 22, HP compressor 24, HP turbine 26, and LP turbine 28 to one another. In some embodiments, drive shaft 18 includes an LP spool (not shown) and an HP spool (not shown), wherein the LP spool couples LP compressor 22 to LP turbine 28, and the HP spool couples HP compressor 24 to HP turbine 26.

[0058] The compressor section 12 includes a plurality of axially spaced stages. Each stage includes a set of circumferentially spaced rotating blade assemblies and a set of circumferentially spaced stationary vane assemblies. In one configuration, the compressor blade assemblies for one stage of the compressor section 12 are mounted to a disk that is mounted to the drive shaft 18. Each set of blade assemblies for a given stage may have its own disk. In one embodiment, the vane assemblies of the compressor section 12 are mounted to a casing that extends circumferentially around the turbine engine 10. In a counter-rotating gas turbine engine, the vane assemblies are mounted to a hub similar to the casing, except that the hub rotates in a direction opposite to the blade assemblies, while the casing is stationary. It should be understood that the representation of the compressor section 12 is merely schematic. The number of stages may vary.

[0059] Similar to the compressor section 12, the turbine section 16 includes a plurality of axially spaced stages, each stage having a set of circumferentially spaced rotating blade assemblies and a set of circumferentially spaced stationary vane assemblies. In one configuration, the turbine blade assemblies for a stage of the turbine section 16 are mounted to a disk that is mounted to the drive shaft 18. Each set of blade assemblies for a given stage may have its own disk. In one embodiment, the turbine section's vane assemblies are circumferentially mounted to the casing.

[0060] Combustor 14 is disposed in series between compressor section 12 and turbine section 16. Combustor 14 is fluidly coupled to at least a portion of compressor section 12 and turbine section 16 such that combustor 14 at least partially fluidly couples compressor section 12 to turbine section 16. As a non-limiting example, combustor 14 is fluidly coupled to an HP compressor 24 at an upstream end of combustor 14 and is fluidly coupled to an HP turbine 26 at a downstream end of combustor 14.

[0061] During operation of the gas turbine engine 10, ambient or atmospheric air is drawn into the compressor section 12 via a fan located upstream of the compressor section 12, where it is compressed to form pressurized air. The pressurized air then flows into the combustor 14, where it is mixed with fuel and ignited, thereby producing hot combustion gases. Some work is extracted from these combustion gases by the HP turbine 26 that drives the HP compressor 24. The combustion gases are exhausted into the LP turbine 28, which extracts additional work to drive the LP compressor 22, and the exhaust gas is ultimately exhausted from the gas turbine engine 10 via an exhaust section (not shown) downstream of the turbine section 16. The LP turbine 28 drives the LP spool to rotate the fan and the LP compressor 22. Together, the pressurized airflow and the combustion gases define the working airflow that flows through the fan, compressor section 12, combustor 14, and turbine section 16 of the gas turbine engine 10.

[0062] Go to Figure 2 , schematically illustrates a portion of a turbine section 16. The turbine section 16 includes a plurality of sets of blade assemblies 30 circumferentially mounted to respective disks 32. The number of individual blade assemblies in a set of blade assemblies 30 mounted to each disk 32 may vary. Although in Figure 2 Although shown schematically in FIG. 1 , it should be understood that the turbine section 16 may be a single stage turbine, or may include additional stages as shown.

[0063] Stationary vane assemblies 34 are mounted to a stator ring 36 located outside the distal end of each disk 32. Nozzles 38 are defined by the spaces between circumferentially adjacent pairs of vane assemblies 34. The number of nozzles 38 provided on the stator ring 36 may vary.

[0064] During operation of gas turbine engine 10, hot gases or a heated fluid flow (denoted as "HF") exit combustor 14 and enter turbine section 16. Heated fluid flow HF is directed through nozzle 38 and impinges upon blade assembly 30, which causes blade assembly 30 to rotate circumferentially about engine centerline 20 and induces rotation of drive shaft 18. The engine core is configured to produce a redline exhaust gas temperature (EGT) in the range of 988 degrees Celsius (°C) to 1120°C.

[0065] Figure 3It is used for gas turbine engine 10( Figure 1 ). FIG. 2 is a perspective view of a single blade assembly 30 of an HP turbine 26. Blade assembly 30 may correspond to a first-stage blade assembly of HP turbine 26. Blade assembly 30 includes a shank 40, a platform 50, and an airfoil 60 (also referred to as a blade or blade portion). Blade assembly 30 may be constructed as a single, unitary part or component (e.g., a monolithic structure). In other examples, shank 40, platform 50, and / or airfoil 60 may be constructed as separate parts or components that are coupled together to form blade assembly 30.

[0066] exist Figure 3 A directional reference system is shown in . The shank 40 extends between the base 42 and the platform 50. The base 42 of the shank 40 is a flat surface defining a plane, which is referred to herein as the base plane (denoted as "BP"). The radial direction (denoted as "R") of the blade assembly 30 is a direction perpendicular to the base plane BP. In addition, the shank 40 extends between the shank leading edge 44 and the shank trailing edge 46. The shank leading edge 44 is a flat surface defining a plane, which is referred to herein as the shank leading edge plane (denoted as "SLEP"). The axial direction (denoted as "A") of the blade assembly 30 is a direction perpendicular to the shank leading edge plane SLEP. The tangential direction (denoted as "T") is a direction perpendicular to the radial direction R and the axial direction A.

[0067] As a non-limiting example, the handle 40 is configured to be mounted to the pan 32 ( Figure 2 ) to rotatably drive the blade assembly 30. The shank 40 includes a plurality of inlet channels 48 (shown in phantom) for receiving a cooling fluid (denoted as "CF") (e.g., bleed air) to cool the blade assembly 30. In the illustrated example, the plurality of inlet channels 48 include a leading edge inlet channel 48l, a middle inlet channel 48m, and a trailing edge inlet channel 48t. Each of the inlet channels 48l, 48m, 48t extends between the base 42 and one or more cooling ducts in the airfoil 60, which are disclosed in further detail herein. The inlet channels 48l, 48m, 48t receive the cooling fluid CF at the base 42. The cooling fluid CF flows through the inlet channels 48l, 48m, 48t and enters the one or more cooling ducts in the airfoil 60. Although there are three inlet channels in this example, in other examples, the shank 40 may include more or fewer inlet channels.

[0068] Airfoil 60 extends radially outward from platform 50 to define a root 61 connected to platform 50 and a tip 62 opposite root 61. Furthermore, airfoil 60 includes an outer wall 63 defining an exterior surface 59 that defines a pressure side 64 and a suction side 65 opposite pressure side 64. Airfoil 60 extends between an airfoil leading edge 66 and an airfoil trailing edge 67 downstream of airfoil leading edge 66. Airfoil leading edge 66 and airfoil trailing edge 67 separate pressure side 64 from suction side 65. In the illustrated example, blade assembly 30 has a plurality of cooling ducts 70 (shown in phantom) formed within airfoil 60. Furthermore, blade assembly 30 has one or more cooling holes 69 formed in outer wall 63 of airfoil 60 to fluidly couple the plurality of cooling ducts 70 within airfoil 60 to the exterior of blade assembly 30. In the illustrated example, cooling holes 69 are located along pressure side 64 near airfoil trailing edge 67. In other examples, the cooling holes 69 may be provided at other locations. The plurality of cooling ducts 70 may include a plurality of ducts extending radially through the airfoil 60. In some examples, one or more cooling ducts 70 are fluidly coupled to certain inlet passages 481, 48m, 48t.

[0069] The platform 50 has a first surface 51, referred to as an upper surface, and a second surface 52, referred to as a lower surface, opposite the upper surface 51. The airfoil 60 is coupled to the upper surface 51 and extends radially outward from the upper surface 51, and the shank 40 is coupled to the lower surface 52 and extends radially inward from the lower surface 52. The platform 50 extends in the axial direction A between a platform leading edge 53 and a platform trailing edge 54 opposite the platform leading edge 53. The platform 50 further extends in the tangential direction T between a first slash face 55 and a second slash face 56 opposite the first slash face 55. When assembled, the continuous blade assembly 30 is rotated about the engine centerline 20 ( Figure 1 ) are arranged in the circumferential direction, with the consecutive side surfaces 55, 56 facing each other.

[0070] During operation of the gas turbine engine 10, a heating fluid flow HF flows along the blade assembly 30. The airfoil leading edge 66 is defined by a stagnation point relative to the heating fluid flow HF. The heating fluid flow HF generally flows in an axial direction from front to rear, while the local directionality may vary as the heating fluid flow HF is driven or rotated within the gas turbine engine 10. A cooling fluid flow CF is supplied to a plurality of inlet passages 48 and flows into a plurality of cooling ducts 70 to cool the airfoil 60. The cooling fluid flow CF is provided over the entire airfoil 60 and is discharged from the plurality of cooling ducts 70 via cooling holes 69 as a cooling film. The plurality of blade assemblies 30 are arranged circumferentially so that the platforms 50 of the blade assemblies 30 form a continuous ring. The platforms 50 help to radially contain the heating fluid flow HF to protect the disks 32. The platforms 50 are used to seal the platform 50 in the flow path of the heating fluid flow H and the disks 32 ( Figure 2 ) The disks 32 require considerable cooling to ensure durability of the HP turbine 26 components.

[0071] Materials used to form the blade assembly 30 include, but are not limited to, steel, refractory metals such as titanium, or high-temperature alloys based on nickel, cobalt, or iron, ceramic matrix composites, or combinations thereof. In non-limiting examples, the structure can be formed by a variety of methods, including additive manufacturing, casting, electroforming, or direct metal laser melting.

[0072] like Figure 3 As shown, the platform 50 has a stator rotor seal 57 extending axially forward from the platform leading edge 53. The stator rotor seal 57 helps to seal the forward buffer cavity (not shown) defined within the rotor assembly. The stator rotor seal 57 has an upper surface 79, a lower surface 81 opposite the upper surface 79, and a forward surface 82 between the upper surface 79 and the lower surface 81. The stator rotor seal 57 has an upper edge 83 between the upper surface 79 and the forward surface 82. The upper edge 83 is curved or arcuate. In particular, the upper edge 83 is curved between a first endpoint 84 at the first side 55 and a second endpoint 85 at the second side 56. The upper edge 83 of the stator rotor seal 57 has a center point 86 that forms a peak of the arc. The upper edge 83 of the stator rotor seal 57 has a radius of curvature, referred to herein as the stator rotor seal radius (denoted as "SRSR"). The center of the radius of curvature of the SRSR is the engine centerline 20 ( Figure 1 ).

[0073] like Figure 4As shown, the SRSR (i.e., the radius of curvature of the upper edge 83 of the stator-rotor seal 57) can be calculated using the straight-line distance (S) between two end points 84, 85 and the maximum offset (D) in the radial direction R between the two end points 84, 85 and the center point 86 of the arc. The SRSR can be calculated using SRSR = (D / 2) + (S 2 / (8xD)).

[0074] Figure 5 is a side view of the blade assembly 30. As described above, the base plane BP is the plane defined by the base 42 and is perpendicular to the radial direction R. The first plane (denoted as "P1") is parallel to the base plane BP and is located at a first radial distance (denoted as "R1") from the base plane BP. The second plane (denoted as "P2") is parallel to the base plane BP and is located at a second radial distance (denoted as "R2") from the base plane BP. The third plane (denoted as "P3") is parallel to the base plane BP and is located at a third radial distance (denoted as "R3") from the base plane BP. The fourth plane (denoted as "P4") is parallel to the base plane BP and is located at a fourth radial distance (denoted as "R4") from the base plane. The fifth plane (denoted as "P5") is parallel to the base plane BP and is located at a fifth radial distance (denoted as "R5") from the base plane. The sixth plane (denoted as "P6") is parallel to the base plane BP and is located at a sixth radial distance (denoted as "R6") from the base plane. The seventh plane (denoted as "P7") is parallel to the base plane BP and is located at a seventh radial distance (denoted as "R7") from the base plane. The eighth plane (denoted as "P8") is parallel to the base plane BP and is located at an eighth radial distance (denoted as "R8") from the base plane. The ninth plane (denoted as "P9") is parallel to the base plane BP and is located at a ninth radial distance (denoted as "R9") from the base plane.

[0075] Each of the planes P1, P2, P3, P4, P5, P,6, P7, P8, P9 extends perpendicular to the radial direction R. The first radial distance R1 is less than the second radial distance R2, the second radial distance R2 is less than the third radial distance R3, the third radial distance R3 is less than the fourth radial distance R4, the fourth radial distance R4 is less than the fifth radial distance, the fifth radial distance is less than the sixth radial distance: R1 < R2 < R3 < R4 < R5 < R6 < R7 < R8 < R9. Table 1 includes the radial distances R1, R2, R3, R4, R5, R6, R7, R8, R9 of the planes P1, P2, P3, P4, P5, P6, P7, P8, P9 from the base plane BP.

[0076] Table 1

[0077] flat Radial distance Distance from base plane P1 R1 0.0167 meters P2 R2 0.0182 meters P3 R3 0.0212 meters P4 R4 0.0227 meters P5 R5 0.0277 meters P6 R6 0.0295 meters P7 R7 0.0341 meters P8 R8 0.0400 meters P9 R9 0.0476 meters

[0078] The first plane P1, the second plane P2, the third plane P3, the fourth plane P4, the fifth plane P5, the sixth plane P6 and the seventh plane P7 extend through the shank 40. The eighth plane P8 and the ninth plane P9 extend through the airfoil 60. The platform 50, the root 61, the tip 62, the airfoil leading edge 66 and the inter-airfoil trailing edge 67 are Figure 5 The winning bid was awarded.

[0079] Figure 6A-8 Several cross-sectional top views of the blade assembly 30 are shown, showing the Figure 5 The inlet passages 481, 48m, 48t and the cooling duct 70 of the planes P1-P9 are shown. Figure 6A-8 The relative sizes and shapes of the various parts of the blade assembly 30 and the cooling ducts in each view are not necessarily accurate and / or to scale. Figure 6A-8 Portions of the cross-sectional view are depicted without hatching and are bounded by dashed lines. It should be understood that these portions may be solid and / or may contain other cooling conduits.

[0080] Figure 6A It is along Figure 5 FIG. 4 is a cross-sectional top view of the shank 40 taken along a first plane P1 that is a first radial distance R1 from the base plane BP. The shank 40 includes a leading edge inlet passage 481, a middle inlet passage 48m, and a trailing edge inlet passage 48t within the shank 40. A plurality of inlet passages 48 extend through the shank 40 to provide internal fluid communication with the cooling duct of the airfoil 60. In particular, the plurality of inlet passages 48 provide internal fluid communication with the cooling duct 70 ( Figure 3 ) provides cooling fluid flow CF( Figure 3 ). The leading edge inlet channel 481 is the channel closest to the shank leading edge 44. In the illustrated example, the set of inlet channels 48 includes three inlet channels, but in other examples, it may include more or fewer channels. Within the first plane P1, the leading edge inlet channel 481 defines a first cooling duct portion 90a having a first cross-sectional area (designated "A1").

[0081] Figure 6B It is along Figure 5 A cross-sectional top view of the shank 40 is shown taken along a second plane P2. In the second plane, the leading edge inlet passage 481 defines a second cooling duct portion 90b having a second cross-sectional area (denoted as "A2"). As used herein, the normalized second area (denoted as ) refers to the ratio of the second cross-sectional area A2 to the first cross-sectional area A1. The standardized second area can be calculated by the following equation:

[0082]

[0083] Figure 6C It is along Figure 5 A cross-sectional top view of the shank 40 is taken along a third plane P3. In the third plane, the leading edge inlet passage 481 defines a third cooling duct portion 90c having a third cross-sectional area (denoted as "A3"). As used herein, the normalized third area (denoted as ) refers to the ratio of the third cross-sectional area A3 to the first cross-sectional area A1. The standardized third area can be calculated by the following equation:

[0084]

[0085] Figure 6D It is along Figure 5 A cross-sectional top view of the shank 40 is taken along a fourth plane P4. In the fourth plane, the leading edge inlet passage 481 defines a fourth cooling duct portion 90d having a fourth cross-sectional area (denoted as "A4"). As used herein, the standardized fourth area (denoted as ) refers to the ratio of the fourth cross-sectional area A4 to the first cross-sectional area A1. The standardized fourth area can be calculated by the following equation:

[0086]

[0087] Figure 6E It is along Figure 5 A cross-sectional top view of the shank 40 is taken along a fifth plane P5. In the fifth plane, the leading edge inlet passage 481 defines a fifth cooling duct portion 90e having a fifth cross-sectional area (denoted as "A5"). As used herein, the standardized fifth area (denoted as ) refers to the ratio of the fifth cross-sectional area A5 to the first cross-sectional area A1. The standardized fifth area can be calculated by the following equation:

[0088]

[0089] Figure 6F It is along Figure 5 A cross-sectional top view of the shank 40 is shown taken along a sixth plane P6. Within the sixth plane (P6), the leading edge inlet passage 481 defines a sixth cooling duct portion 90f having a sixth cross-sectional area (denoted as "A6"). As used herein, the standardized sixth area (denoted as ) refers to the ratio of the sixth cross-sectional area A6 to the first cross-sectional area A1. The standardized sixth area can be calculated by the following equation:

[0090]

[0091] Figure 6G It is along Figure 5 FIG. 4 is a cross-sectional top view of the shank 40 taken along a seventh plane P7. In the seventh plane (P7), the leading edge inlet passage 481 defines a seventh cooling duct portion 90g having a sixth cross-sectional area (denoted as "A7"). As used herein, the standardized seventh area (denoted as ) refers to the ratio of the seventh cross-sectional area A7 to the first cross-sectional area A1. The standardized seventh area can be calculated by the following equation:

[0092]

[0093] Figure 7 It is along Figure 5 The airfoil 60 extends between an airfoil leading edge 66 and an airfoil trailing edge 67 to define a chord length CL between the airfoil leading edge 66 and the airfoil trailing edge 67. Figure 7 , the plurality of cooling ducts 70 include a first cooling duct portion 71, a second cooling duct portion 72, and a third cooling duct portion 73, which define a set of suction-side cooling ducts 80 proximate the suction side 65 and are numbered sequentially as they approach the airfoil leading edge 66. In some examples, the airfoil 60 may include one or more additional cooling ducts closer to the trailing edge 67.

[0094] Each cooling duct portion 71, 72, 73 defines a cross-sectional area in the eighth plane P8. The first cooling duct portion 71 has an eighth cross-sectional area (denoted as "A8"), the second cooling duct portion 72 has a ninth cross-sectional area (denoted as "A9"), and the third cooling duct portion 73 has a tenth cross-sectional area (denoted as "A 10 ”). The eighth cross-sectional area, the ninth cross-sectional area, and the tenth cross-sectional area together define a first sum of cross-sectional areas (denoted as “S1”): S1=A 10 +A 11 +A 12 .

[0095] Figure 8 It is along Figure 5 The cross-sectional view of the airfoil 60 is taken along the ninth plane P9, just above (radially outward from) the third plane P8. The first cooling duct portion 71 has an eleventh cross-sectional area (denoted as “A 11 ”), the second cooling duct portion 72 has a twelfth cross-sectional area (denoted as “A 12 ”), the third cooling duct portion 73 has a thirteenth cross-sectional area (denoted as “A 13 ”). The eleventh cross-sectional area, the twelfth cross-sectional area, and the thirteenth cross-sectional area together define a second sum of cross-sectional areas (denoted as “S2”): S2=A11 +A 12 +A 13 As used herein, the average of the first sum S1 and the second sum S2 of the set of suction-side cooling ducts 80 is referred to herein as the suction-side area (denoted as "SSA"). SAA can be calculated by the following equation:

[0096]

[0097] The blade assembly 30 of the HP turbine 26, and in particular the first stage blade assembly 30 of the HP turbine 26, has the highest flow path temperature of any blade group. This first stage blade assembly also rotates at an extremely high angular velocity. The extreme temperature environment and high rotational speed exert large forces on the first stage blade assembly 30, which can cause creep and fatigue, particularly along the suction side of the airfoil. Creep and fatigue can lead to unexpected or premature component replacement, which limits the engine time on wing (TOW). Therefore, there is a need for a blade assembly with high durability that can withstand these large centrifugal stresses and reduce (e.g., minimize) creep and fatigue.

[0098] The inventors have created a blade assembly with relatively high durability (e.g., creep and fatigue resistance) for a limited engine operating environment. The inventors developed multiple blade assemblies and determined that it would be beneficial to change the geometry of the cooling ducts by changing the duct size near the leading edge to locally increase the wall thickness while remaining within the constraints of the existing engine system. Further geometric changes to the inlet passages enable control of the amount of cooling fluid provided to the cooling ducts. The blade assembly disclosed herein includes cooling ducts and inlet passages sized to ensure sufficient cooling capacity while also improving creep and fatigue resistance.

[0099] The inventors have determined that the size of the cooling ducts in the airfoil and shank affects creep and fatigue in the airfoil 60. More specifically, the normalized second area Standardized third area Standardized fourth area Standardized fifth area Standardized sixth area Standardized seventh area and suction side area, SSA, for a specific set of operating characteristics.

[0100] The inventors determined that the standardized second region Standardized third area Standardized fourth area Standardized fifth area Standardized sixth area Standardized seventh area The suction side area SSA and the stator rotor seal radius SRSR have a significant impact on durability. Reducing the suction side area SSA increases the local thickness of the outer wall 63, thereby improving the load-bearing capacity at the leading edge 66 of the airfoil.

[0101] Table 2 below shows 14 examples (denoted as Ex. 1-14) of blade assemblies 30 developed by the present inventors. Thus, the present inventors generated a variety of new designs in their blade assembly design process that provide blade assemblies 30 with improved creep and fatigue resistance. As described above, the present inventors created solutions (Examples 1-10 of Table 2) with relatively high blade durability (e.g., reduced creep and fatigue, no crack formation or growth after multiple engine cycles) for a limited engine environment, compared to solutions without such high blade durability (Examples 11-14 of Table 2). Table 2 includes the following examples for each of these examples: Figure 6B The second area of ​​standardization Figure 6C The standardized third area Figure 6D The standardized fourth area Figure 6E The standardized fifth area Figure 6F The standardized sixth area Figure 6G The standardized seventh area Figure 7 and 8 The suction side area SSA and the stator rotor sealing radius SRSR.

[0102] Table 2

[0103]

[0104] The inventors have found that blade assembly designs having the parameters defined in Examples 1-10 exhibit relatively high structural integrity and durability while remaining within current engine constraints. In contrast, Examples 11-14 have relatively low durability for the particular engine environment.

[0105] The embodiments developed by the inventors shown in Table 1 can be characterized by an expression (EQ) that can be used to distinguish those designs in Examples 1-10 that meet the performance (durability) requirements from those designs in Examples 11-14 that do not meet the performance requirements. Thus, expression (EQ) can be used to identify improved blade assembly designs that are more suitable for a particular engine operating environment and take into account the constraints imposed on blade assembly designs with cooling holes used in such systems.

[0106] The expression (EQ) is defined as:

[0107]

[0108] in express Figure 6B The area of ​​the leading edge channel 481 at the second plane P2 is given by Figure 6A The area of ​​the leading edge channel 481 at the first plane P1 is normalized. express Figure 6C The area of ​​the leading edge channel 481 at the third plane P3 is given by Figure 6A The area of ​​the leading edge channel 481 at the first plane P1 is normalized. express Figure 6D The area of ​​the leading edge channel 481 at the fourth plane P4 is given by Figure 6A The area of ​​the leading edge channel 481 at the first plane P1 is normalized. express Figure 6E The area of ​​the leading edge channel 481 at the fifth plane P5 is given by Figure 6A The area of ​​the leading edge channel 481 at the first plane P1 is normalized. express Figure 6F The area of ​​the leading edge channel 481 at the sixth plane P6 is given by Figure 6A The area of ​​the leading edge channel 481 at the first plane P1 is normalized. express Figure 6E The area of ​​the leading edge channel 481 at the seventh plane P7 is given by Figure 6A The area of ​​the leading edge channel 481 at the first plane P1 is normalized. SSA represents Figure 7 and 8 The sum of the areas of the cooling ducts 71, 72, 73 at the eighth plane P8 and the ninth plane P9. SRSR stands for Figure 4 The values ​​of the expression (EQ) for each example of Table 2 are shown in Table 3.

[0109] Table 3

[0110]

[0111] Based on the expression (EQ) values ​​of Examples 1-10 in Table 3, it was determined that blade assembly designs having EQ values ​​within the range of 1211.444 to 4490.406 (ie, 1211.444≤EQ≤4490.406) advantageously meet durability requirements while remaining within required tolerances and being usable with existing engine systems.

[0112] When manufactured components, including the blade assembly 30, have geometries where the expression (EQ) falls within the range of 1211.444 to 4490.406 (i.e., 1211.444 ≤ EQ ≤ 4490.406), various benefits are achieved. Such benefits include reduced stresses at the suction side 65 of the airfoil 60, which increases the life of the blade assembly 30 and, therefore, extends the time between required part replacements. This provides increased durability for the blade assembly 30, thereby reducing required maintenance and costs while increasing overall engine reliability.

[0113] Furthermore, the benefits included herein provide for the blade assembly 30 to be assembled within an existing engine. For example, the values ​​of the expression (EQ) provided herein take into account existing engines, allowing for the replacement of the current blade assembly with a replacement blade assembly (or a new blade assembly) having the parameters of the blade assembly 30 described herein. This consideration provides for the replacement and improvement of current engine systems without the need to produce new engine parts capable of retaining the blade assembly 30. This can improve the durability of current engines without increasing the cost of producing a new engine or further improving an existing engine.

[0114] Table 4 below shows value ranges for the trailing edge area TEA, the second area, and the stator rotor seal radius SRSR of the blade assembly suitable for meeting the durability requirements, as well as the minimum and maximum values ​​of the expression (EQ).

[0115] Table 4

[0116]

[0117] Additional benefits associated with the blade assembly 30 with cooling ducts 71, 72, 73 and leading edge passages 481 described herein include rapid evaluation of design parameters in terms of blade assembly size and cooling duct geometry, engine operating conditions, and the number of blades and blade assemblies for the engine design and a particular blade design. Narrowing these various factors to a possible area saves time, money, and resources. The blade assembly 30 with cooling ducts 71, 72, 73 and leading edge passages 481 described herein enables the development and production of high-performance turbine engines and blade assemblies across multiple performance metrics within a given set of constraints.

[0118] As mentioned above, designs such as Examples 11-14 in Tables 2 and 3 were found to have relatively low durability for a particular engine environment. This is reflected in the relevant expression (EQ) values ​​outside the range of 1211.444 to 4490.406. Lower durability results in less time on wing (TOW) and greater maintenance costs.

[0119] Additionally or alternatively, designs outside the EQ range may attempt to increase durability by sacrificing in weight, aerodynamic performance, and efficiency. For example, a standard practice for addressing the problem of improving blade assembly durability is to utilize higher strength materials. However, such materials result in increased cost, system weight, and the total space occupied by the blade assembly. With material-driven solutions, overall engine efficiency may be reduced, and the associated components must be redesigned to compensate for the higher strength materials. In some cases, such redesign results in impractical or impossible results. Therefore, a solution is needed for reducing stresses in airfoils currently used in existing engines without requiring redesign of the associated components or sacrificing overall engine efficiency.

[0120] In other examples, increasing the size of the airfoil or related components, utilizing higher-strength materials, and / or providing additional cooling features can combat centrifugal and thermal stresses. However, such increased size, higher-strength materials, and additional cooling features can result in increased cost, system weight, the total space occupied by the blade assembly, and performance penalties (e.g., higher fuel consumption, etc.), as well as increased localized stresses in the cooling ducts due to the increased weight and size associated with centrifugal forces. The increased cooling features result in a relatively small amount of material being used, which can lead to increased localized stresses in the cooling ducts. Therefore, a solution is needed for reducing stresses in the cooling ducts without additionally increasing stress, weight, size, or reducing engine efficiency.

[0121] As described above, the present inventors have discovered that Examples 1-10 of Tables 2 and 3 provide successful solutions without increasing weight, strength, or the number of cooling features. Examples 1-10 of Tables 2 and 3 show that designs with expression (EQ) values ​​ranging from 1211.444 to 4490.406 (i.e., 1211.444 ≤ EQ ≤ 4490.406) achieve increased durability without compromising size, weight, strength, or stress. In other words, through Examples 1-10 of Tables 2 and 3, effective stress reduction can be achieved without changing the airfoil, using heavier, higher-strength materials, or adding additional cooling features.

[0122] As described above, the present inventors have created a blade assembly having relatively high durability (eg, creep and fatigue resistance) for a limited engine operating environment.

[0123] To the extent that one or more structures provided herein may be known in the art, it should be understood that the present disclosure may include combinations of structures not previously known, based at least in part on conflicting benefits, desired modes of operation, or other forms of teaching in the art.

[0124] This written description uses examples to disclose the present disclosure, including the best mode, and 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. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

[0125] Further aspects of the present disclosure are provided by the subject matter of the following clauses:

[0126] A blade assembly for a gas turbine engine, the blade assembly comprising: a platform having an upper surface and a lower surface, the platform having a stator rotor seal, the stator rotor seal having an upper edge, the upper edge having a radius of curvature defined as a stator rotor seal radius (SRSR), wherein the stator rotor seal radius (SRSR) is 0.224 meters to 0.239 meters; an airfoil extending from the upper surface of the platform, the airfoil having an outer surface, the outer surface defining a pressure side and a suction side, the airfoil comprising a leading edge and a trailing edge; a shank extending from the lower surface, the shank having a base defining a base plane; a leading edge inlet passage located within the shank, the leading edge inlet passage having: a first cross-sectional area at a first plane, the first plane being radially spaced 0.0167 meters from the base plane; a second cross-sectional area at a second plane, the second plane being radially spaced 0.0182 meters from the base plane, a first ratio of the second cross-sectional area to the first cross-sectional area defining a second normalized area in the range of 0.996 to 1.139 a third cross-sectional area at a third plane radially spaced 0.0212 meters from the base plane, a second ratio of the third cross-sectional area to the first cross-sectional area defining a third normalized area in a range of 1.485 to 1.620 a fourth cross-sectional area at a fourth plane radially spaced 0.0226 meters from the base plane, a third ratio of the fourth cross-sectional area to the first cross-sectional area defining a fourth normalized area in a range of 1.743 to 1.861 a fifth cross-sectional area at a fifth plane radially spaced 0.0277 meters from the base plane, a fourth ratio of the fifth cross-sectional area to the first cross-sectional area defining a fifth normalized area within a range of 1.155 to 1.262 a sixth cross-sectional area at a sixth plane radially spaced 0.0294 meters from the base plane, a seventh ratio of the sixth cross-sectional area to the first cross-sectional area defining a sixth normalized area within a range of 1.039 to 1.179 and a seventh cross-sectional area at a seventh plane radially spaced 0.0341 meters from the base plane, a sixth ratio of the seventh cross-sectional area to the first cross-sectional area defining a seventh normalized area within a range of 0.890 to 1.078 and a plurality of cooling ducts located within the airfoil, the plurality of cooling ducts including a group of suction-side cooling ducts proximate the suction side, the group of suction-side cooling ducts having a first sum of areas at an eighth plane radially spaced 0.0400 meters from the base plane, the group of suction-side cooling ducts having a second sum of areas at a ninth plane radially spaced 0.0476 meters from the base plane, the average of the first sum and the second sum defining a suction-side cross-sectional area (SSA), wherein the suction-side cross-sectional area (SSA) is 0.000005 m 2 to 0.00000945m 2 , and among them,

[0127] A blade assembly as described in any preceding clause, further comprising a plurality of inlet channels within the shank, the plurality of inlet channels including the leading edge inlet channel.

[0128] A blade assembly as described in any preceding clause, wherein each said inlet passage extends between said base and one or more of said plurality of cooling ducts.

[0129] The blade assembly of any preceding clause, wherein the plurality of inlet channels further comprises a mid-inlet channel and a trailing edge inlet channel.

[0130] The blade assembly of any preceding clause, wherein the shank is configured as a dovetail.

[0131] A blade assembly as described in any preceding clause, wherein the blade assembly is a first stage blade assembly of a high pressure turbine.

[0132] The blade assembly of any preceding clause, wherein the first cross-sectional area is substantially equal to the seventh cross-sectional area.

[0133] The following claims are incorporated into this detailed description by reference, with each claim standing on its own as a separate embodiment of the disclosure.

Claims

1. A blade assembly for a gas turbine engine, characterized in that The blade assembly comprises: a platform having an upper surface and a lower surface, the platform having a stator rotor seal, the stator rotor seal having an upper edge with a radius of curvature defined as a stator rotor seal radius (SRSR), wherein the stator rotor seal radius (SRSR) is 0.224 meters to 0.239 meters; an airfoil extending from the upper surface of the platform, the airfoil having an exterior surface defining a pressure side and a suction side, the airfoil including a leading edge and a trailing edge; a handle extending from the lower surface, the handle having a base defining a base plane; a leading edge inlet passage located within the shank, the leading edge inlet passage having: a first cross-sectional area at a first plane radially spaced 0.0167 meters from the base plane; a second cross-sectional area at a second plane radially spaced 0.0182 meters from the base plane, a first ratio of the second cross-sectional area to the first cross-sectional area defining a second normalized area in a range from 0.996 to 1.139 a third cross-sectional area at a third plane radially spaced 0.0212 meters from the base plane, a second ratio of the third cross-sectional area to the first cross-sectional area defining a third normalized area in a range of 1.485 to 1.620 a fourth cross-sectional area at a fourth plane radially spaced 0.0226 meters from the base plane, a third ratio of the fourth cross-sectional area to the first cross-sectional area defining a fourth normalized area in a range of 1.743 to 1.861 a fifth cross-sectional area at a fifth plane radially spaced 0.0277 meters from the base plane, a fourth ratio of the fifth cross-sectional area to the first cross-sectional area defining a fifth normalized area within a range of 1.155 to 1.262 a sixth cross-sectional area at a sixth plane radially spaced 0.0294 meters from the base plane, a seventh ratio of the sixth cross-sectional area to the first cross-sectional area defining a sixth normalized area within a range of 1.039 to 1.179 and a seventh cross-sectional area at a seventh plane radially spaced 0.0341 meters from the base plane, a sixth ratio of the seventh cross-sectional area to the first cross-sectional area defining a seventh normalized area within a range of 0.890 to 1.078 and a plurality of cooling ducts located within the airfoil, the plurality of cooling ducts including a group of suction-side cooling ducts proximate the suction side, the group of suction-side cooling ducts having a first sum of areas at an eighth plane radially spaced 0.0400 meters from the base plane, the group of suction-side cooling ducts having a second sum of areas at a ninth plane radially spaced 0.0476 meters from the base plane, the average of the first sum and the second sum defining a suction-side cross-sectional area (SSA), wherein the suction-side cross-sectional area (SSA) is 0.000005 m 2 to 0.00000945m 2 , and among them, 2. The blade assembly according to claim 1, wherein: Further included is a plurality of inlet channels within the shank, the plurality of inlet channels including the leading edge inlet channel.

3. The blade assembly according to claim 2, wherein: Each of the inlet passages extends between the base and one or more of the plurality of cooling conduits.

4. The blade assembly according to claim 2, wherein: The plurality of inlet channels further include a middle inlet channel and a trailing edge inlet channel.

5. The blade assembly according to claim 1, wherein: Wherein the shank is configured as a dovetail.

6. The blade assembly according to claim 1, wherein: The blade assembly is a first-stage blade assembly of a high-pressure turbine.

7. The blade assembly according to claim 1, wherein: The first cross-sectional area is substantially equal to the seventh cross-sectional area.

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