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 The intermediate 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 6 yes Figure 3 Schematic side view of the cooling circuit of the blade assembly.

[0013] Figure 7 Describes the cooling circuit Figure 3 Cross-sectional view of the blade assembly.

[0014] Figure 8 yes Figure 6 Cross-section of the separator of the cooling circuit.

[0015] Figure 9 It is shown in Figure 5 intercepted at multiple planes of Figure 3 Schematic diagram of the cross-sectional shape of the cooling duct portion of the intermediate inlet channel.

[0016] Figure 10 It is used for calculation Figure 3 Schematic diagram of the varying extent of the intermediate inlet channel of the blade assembly along the suction side. DETAILED DESCRIPTION

[0017] 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.

[0018] 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.

[0019] The blade assembly may be a blade assembly 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 subjected to the highest thermal and mechanical stresses.

[0020] 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.

[0021] High engine temperatures and operating forces impose relatively high thermal and mechanical stresses on the blade assembly. Furthermore, the cooling ducts and inlet passages within the blade assembly cause stress redistribution. For example, the dimensions of the cooling ducts and inlet passages affect the thickness of the airfoil and shank walls, which in turn influences the stress distribution within the blade assembly. These relatively high stresses can lead to unplanned or premature component replacement. Therefore, there is a need for a blade assembly with greater durability to increase on-wing time.

[0022] 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.

[0023] 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.

[0024] 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.

[0025] 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.

[0026] As used herein, a "set" of elements or a group of elements may include any number of the elements, including one.

[0027] 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.

[0028] 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 fluid flow path and to a reference position along the fluid flow path.

[0029] 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.

[0030] As used herein, the forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0031] As used herein, a radial direction (denoted as "R") is a direction perpendicular to a base plane on the shank of a blade assembly.

[0032] 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.

[0033] As used herein, the tangential direction (denoted as "T") is a direction perpendicular to the radial direction and the axial direction.

[0034] 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.

[0035] 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.

[0036] 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.

[0037] As used herein, "cooling ducts" refer to flow paths formed in a blade assembly that carry a cooling fluid.

[0038] 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.

[0039] As used herein, "inlet passage" refers to a cooling duct formed in the shank of a bucket assembly.

[0040] 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.

[0041] 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.

[0042] As used herein, "intermediate inlet channel" refers to an inlet channel between the leading edge inlet channel and the trailing edge inlet channel.

[0043] 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.

[0044] The number of cooling ducts supplied by the intermediate inlet channel (denoted as “ N1 ”) refers to the number of different cooling ducts located downstream of the intermediate inlet channel.

[0045] The number of cooling duct sections spanning pressure surface to suction surface (denoted as “N2”) refers to the number of cooling duct sections spanning from the pressure surface to the suction side of the bucket assembly and fed by the intermediate inlet passage of the bucket assembly.

[0046] As used herein, the maximum normalized cooling duct section area (denoted as “A max ”) refers to the maximum normalized area among the cross-sectional areas of a set of cooling duct sections of the intermediate inlet channel.

[0047] The degree of variation of the cooling duct portion along the suction side of the airfoil of the bucket assembly (denoted as "DCCPASS") refers to the angle measured between lines along the suction side of the cooling duct portion of the intermediate inlet passage at two cross-sectional planes.

[0048] All measurements referred to herein were performed on the blade assembly prior to use or as a cooling component.

[0049] 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.

[0050] 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 each other. Turbine section 16 includes an HP turbine 26 and an LP turbine 28 fluidly coupled in series with each other. A drive shaft 18 operably couples LP compressor 22, HP compressor 24, HP turbine 26, and LP turbine 28 to each other. 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.

[0051] 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.

[0052] 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.

[0053] 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.

[0054] 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 gases are ultimately exhausted from the gas turbine engine 10 via an exhaust section (not shown) downstream of the turbine section 16. The drive of 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.

[0055] 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.

[0056] 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.

[0057] 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.

[0058] 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.

[0059] 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.

[0060] As a non-limiting example, the handle 40 is configured to be mounted to the pan 32 ( Figure 2 ), thereby rotatably driving 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.

[0061] 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.

[0062] 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.

[0063] 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.

[0064] 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.

[0065] 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 front 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 front 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 front surface 82. The upper edge 83 is curved or arcuate. In particular, the upper edge 83 is curved between a first end point 84 at the first side 55 and a second end point 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 is the engine centerline 20 ( Figure 1 ).

[0066] 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)).

[0067] 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.

[0068] Each of the planes P1, P2, P3, P4, P5, P6 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, and the fifth radial distance is less than the sixth radial distance: R1 < R2 < R3 < R4 < R5 < R6. Table 1 includes the radial distances R1, R2, R3, R4, R5, R6 of the planes P1, P2, P3, P4, P5, P6 from the base plane BP.

[0069] Table 1

[0070] flat Radial distance Distance from base plane P1 R1 0.0167 meters P2 R2 0.0197 meters P3 R3 0.0215 meters P4 R4 0.0234 meters P5 R5 0.0245 meters P6 R6 0.0449 meters

[0071] The first plane P1, the second plane P2, the third plane P3, the fourth plane P4, and the fifth plane P5 extend through the shank 40. The sixth plane P6 extends through the airfoil 60. The platform 50, the root 61, the tip 62, the leading edge 66 of the airfoil, and the trailing edge 67 of the airfoil are marked in Figure 5 as shown.

[0072] Figure 6 is Figure 3-5A schematic side view of a cooling circuit 90 of a blade assembly 30 is shown. The cooling circuit 90 includes Figure 3 The front inlet channel 481, the middle inlet channel 48m and the rear inlet channel 48t. Figure 6 Inlet passages 481, 48m, 48t extend to Figure 3 The cooling duct 70. Figure 6 In the embodiment, the middle inlet passage 48m leads to the middle duct 92 of the cooling duct 70. It should be understood that the front inlet passage 481 and the rear inlet passage 48t also lead to the cooling ducts in the cooling duct 70, which are Figure 6 Not shown in order to avoid obscuring the intermediate conduit 92. Figure 6 middle, Figure 5 The planes P1, P2, P3, P4, P5, and P6 are marked.

[0073] exist Figure 6 , the intermediate inlet passage 48m includes a separator section 100. The intermediate duct 92 includes a front serpentine cooling duct 102 and a rear serpentine cooling duct 104, which branch from the intermediate inlet passage 48m at the separator section 100. The cooling flow (CF) flowing through the intermediate inlet passage 48m is divided into a first flow F1 passing through the front serpentine cooling duct 102 and a second flow F2 passing through the rear serpentine cooling duct 104 at the separator section 100.

[0074] The front serpentine cooling duct 102 includes a first front section 110 defined from the separator section 100 to a first front bend 112, a second front section 114 defined from the first front bend 112 to a second front bend 116, and a third front section 118 defined from the second front bend 116 to a front end 120. The rear serpentine cooling duct 104 includes a first rear section 130 defined from the separator section 100 to a first rear bend 132, a second rear section 134 defined from the first rear bend 132 to a second rear bend 136, and a third rear section 138 defined from the second rear bend 136 to a rear end 140. The amount (i.e., number) of ducts supplied by the intermediate inlet passage 48m (denoted as "N1") refers to the number of different cooling ducts located downstream of the intermediate inlet passage 48m. In Figure 6 In FIG. 4 , the middle inlet passage 48 m supplies two different cooling ducts, namely the front serpentine cooling duct 102 and the rear serpentine cooling duct 104 .

[0075] Figure 7 It is along Figure 5 A cross-sectional view of the airfoil 60 of the P6 blade assembly 34. Figure 7 In FIG, the pressure side 64, the suction side 65, the airfoil leading edge 66 and the airfoil trailing edge 67 of the airfoil 60 are indicated. Figure 7 In, marked Figure 6 The sections 110 , 114 , 118 and Figure 5 The sections 130, 134, 138 of the rear serpentine cooling duct 104. Figure 7 In FIG, sections 110, 114, 118 are adjacent to the suction side 65 and offset from the pressure side 64. Figure 7 In FIG, sections 130, 134, 138 are adjacent to the pressure side 64 and the suction side 65. Figure 7 In FIG, the portion of the cross-section of the airfoil 60 includes the dashed portion. In some examples, some or all of the area of ​​the dashed portion includes other conduits of the cooling circuit 90, such as cooling conduits extending from the front inlet passage 481 and / or the rear inlet passage 48t.

[0076] exist Figure 7 , the segments 130, 134, 138 in the aft segment have a first cooling duct portion 142, a second cooling duct portion 144, and a third cooling duct portion 146, respectively. As used herein, the amount (e.g., number) of cooling duct portions spanning from the pressure surface 64 to the suction surface 65 (denoted as "N2") refers to the number of cooling duct portions fed from the intermediate inlet passage that span at least 60% of the thickness of the airfoil as measured perpendicular to the arc of the airfoil 60 and from the geometric center of the respective cooling duct portion. Figure 7 , the bucket assembly 34 includes three cooling duct sections, ie, cooling duct sections 142 , 144 , 146 , spanning from the pressure surface 64 to the suction surface 65 .

[0077] Figure 8 is Figure 6 At P6, Figure 6 A cross-sectional view of the blade assembly 30 at the separator section 100 of FIG. Figure 8 , the first section 110 of the front serpentine cooling duct 102 and the first section 130 of the rear serpentine cooling duct 104 are marked. At P6, the first section 110 of the front serpentine cooling duct 102 has a first cross-sectional area 147A and a first major axis 148A. At P6, the first section 130 of the rear serpentine cooling duct 104 has a second cross-sectional area 147B and a second major axis 148B. Figure 8 In FIG, the major axes 148A, 148B form a generally L-shape (e.g., a dart shape). Figure 8 In the embodiment, the first cross-sectional area 147A is smaller than the second cross-sectional area 147B.

[0078] Figure 9 It is shown in Figure 5 Schematic diagram of the cross-sectional shape of the intermediate inlet channel 48m taken at P1, P2, P3, P4 and P5. Figure 9, the first intermediate inlet passage 48m defines a first cooling duct portion 150 at P1, a second cooling duct portion 152 at P2, a third cooling duct portion 154 at P3, a fourth cooling duct portion 156 at P4, and a fifth cooling duct portion 158 at P5. The first cooling duct portion 150 has a first cross-sectional area 162, the second cooling duct portion 152 has a second cross-sectional area 164, the third cooling duct portion 154 has a third cross-sectional area 166, the fourth cooling duct portion 156 has a fourth cross-sectional area 168, and the fifth cooling duct portion 158 has a fifth cross-sectional area 170. As used herein, the maximum normalized cooling duct portion area (denoted as "A") is the maximum normalized cooling duct portion area. max ”) refers to the largest area among the cross-sectional areas 164, 166, 168, 170 normalized by the first cross-sectional area 162. max It can be calculated by the following equation:

[0079]

[0080] A1 is the first cross-sectional area 162 , A2 is the second cross-sectional area 164 , A3 is the third cross-sectional area 166 , A4 is the fourth cross-sectional area 168 , and A5 is the fifth cross-sectional area 170 .

[0081] Figure 10 It is used for calculation Figure 3 Schematic diagram of the varying extent (denoted as "DCCPASS") of the intermediate inlet passage 48m of the blade assembly 34 along the suction side 176. The DCCPASS 176 is determined between the first cooling duct portion 150 and the fifth cooling duct portion 158. To determine the DCCPASS 176, a first line 172 and a second line 174 are marked on the first cooling duct portion 150 and the fifth cooling duct portion 158, respectively. The first line 172 is defined between a first point 178A and a second point 178B. To determine the positions of the points 178A, 178B, a first boundary circle 180 is drawn around the perimeter of the first cooling duct portion 150. The first boundary circle 180 is the smallest circle that completely defines the first cooling duct portion 150. The first point 178A is a point on the perimeter of the first cooling duct portion 150 that is 15 degrees apart from a line extending from the center of the first boundary circle 180 toward the suction side 65 (e.g., the 12:30 o'clock position on the first boundary circle 180, etc.). Second point 178B is a point on the boundary of first cooling duct portion 150 that is 45 degrees apart from a line extending from the center of first boundary circle 180 toward suction side 65 (eg, 1:30 o'clock on first boundary circle 1800 , etc.).

[0082] Second line 174 is defined between third point 184A and fourth point 184B. To determine the locations of points 184A and 184B, a second bounding circle 186 is drawn around the perimeter of fifth cooling duct portion 158. Second bounding circle 186 is the smallest circle that completely defines fifth cooling duct portion 158. Third point 184A is a point on the perimeter of first cooling duct portion 150 that is 15 degrees apart from a line extending from the center of second bounding circle 186 toward suction side 65 (e.g., at 12:30 o'clock on second bounding circle 186, etc.). Fourth point 184B is a point on the perimeter of fifth cooling duct portion 158 that is 45 degrees apart from a line extending from the center of second bounding circle 186 toward suction side 65 (e.g., at 1:30 o'clock on second bounding circle 186, etc.). DCCPASS 176 is the angle formed between first line 172 and second line 174.

[0083] 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.

[0084] The inventors have created a blade assembly with relatively high durability (e.g., creep and fatigue resistance) for a defined engine operating environment characterized by redline EGT and redline CS. The inventors developed multiple blade assemblies and determined that changing the geometry of the intermediate inlet passage 48m was beneficial. Further geometric changes to the inlet passage enabled control of the amount of cooling fluid provided to the cooling duct. The blade assembly disclosed herein includes cooling ducts and inlet passages sized to ensure adequate cooling capacity while also improving creep and fatigue resistance. The inventors determined that the size and shape of the intermediate inlet passage 48m were a significant contributor to creep and fatigue in the airfoil 60. More specifically, the shape and size of the intermediate inlet passage 48m, such as DCCPASS 176, the maximum standardized area of ​​the cooling duct sections 150, 152, 154, 156, 158, were used for a specific set of operating characteristics.

[0085] The inventors determined that the degree of variation along the suction side DCCPASS176, the maximum normalized area A max, which has a significant impact on durability, the number N1 of cooling ducts fed by the intermediate inlet passage, and the number N2 of cooling duct sections spanning the pressure surface to the suction surface provide values ​​that contribute to the relationship associated with improved creep and fatigue resistance of the blade assembly 30.

[0086] Table 2 below shows 14 examples (denoted as Ex. 1-14) of blade assemblies 30 developed by the inventors. As described above, the 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 examples of each. Figure 10 The value of DCCPASS varies along the suction side of the cooling duct section. Figure 4 The stator rotor seal radius SRSR value is determined by Figure 6 The number N1 of cooling duct sections supplied by the intermediate inlet passage 48m, and Figure 7 The number N2 of cooling duct sections that supply cooling from the pressure surface to the suction surface.

[0087] Table 2

[0088] parameter DCCPASS SRSR <![CDATA[N1]]> <![CDATA[N2]]> unit Degree(°) meter (m) - - Ex.1 21.00 0.224 2 3 Ex.2 33.00 0.239 2 3 Ex.3 26.84 0.232 2 3 Ex.4 24.60 0.237 2 3 Ex.5 32.94 0.235 2 3 Ex.6 22.64 0.236 2 3 Ex.7 22.90 0.225 2 3 Ex.8 32.81 0.225 2 3 Ex.9 27.71 0.237 2 3 Ex.10 25.72 0.236 2 3 Ex.11 12.20 0.235 2 2 Ex.12 9.70 0.236 2 2 Ex.13 13.70 0.225 2 2 Ex.14 8.70 0.237 2 2

[0089] 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.

[0090] The examples developed by the inventors shown in Table 2 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. Therefore, expression (EQ) can be used to identify improved blade assembly designs that are more suitable for a specific engine operating environment and take into account the constraints imposed on the design of blade assemblies with cooling holes used in such systems.

[0091] The expression (EQ) is defined as:

[0092]

[0093] Where DCCPASS refers to the degree of variation of the intermediate inlet passage 48m along the suction side. SRSR refers to the stator rotor seal radius of the blade assembly 30. N1 refers to the number of cooling ducts supplied by the intermediate inlet passage 48m of the blade assembly 30. N2 refers to the number of cooling duct sections supplied by the intermediate inlet passage from the pressure side to the suction side. The values ​​of the expression (EQ) for each example of Table 2 are shown in Table 3.

[0094] Table 3

[0095] parameter DCCPASS SRSR <![CDATA[N 1 ]]> <![CDATA[N 2 ]]> EQ unit Degree(°) meter (m) - - - Ex.1 21.00 0.224 2 3 4.949 Ex.2 33.00 0.239 2 3 13.039 Ex.3 26.84 0.232 2 3 8.373 Ex.4 24.60 0.237 2 3 7.185 Ex.5 32.94 0.235 2 3 12.774 Ex.6 22.64 0.236 2 3 6.047 Ex.7 22.90 0.225 2 3 5.911 Ex.8 32.81 0.225 2 3 12.134 Ex.9 27.71 0.237 2 3 9.118 Ex.10 25.72 0.236 2 3 7.821 Ex.11 12.20 0.235 2 2 0.519 Ex.12 9.70 0.236 2 2 0.329 Ex.13 13.70 0.225 2 2 0.627 Ex.14 8.70 0.237 2 2 0.266

[0096] Based on the expression (EQ) values ​​for Examples 1-10 in Table 2, it was determined that blade assembly designs with EQ values ​​ranging from 4.949 to 13.039 (ie, 4.949≤EQ≤13.039) advantageously meet durability requirements while remaining within required tolerances and being usable with existing engine systems.

[0097] When manufactured components, including the blade assembly 30, have geometries where the expression (EQ) falls within the range of 4.949 to 13.039 (i.e., 4.949 ≤ EQ ≤ 13.039), various benefits are achieved. Such benefits include reduced stress in the airfoil and improved blade cooling, 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.

[0098] 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.

[0099] Table 4 below shows the degree of variation DCCPASS of the cooling duct portion of the blade assembly suitable for meeting durability requirements along the suction side, the number N1 of cooling ducts supplied by the intermediate inlet channel 48m, the number N2 of cooling duct portions supplied by the intermediate inlet channel spanning from the pressure side to the suction side, the stator rotor seal radius SRSR, and the value range of the minimum and maximum values ​​of expression (EQ).

[0100] Table 4

[0101] parameter: element: Minimum: Maximum value: unit: DCCPASS Variation along the suction side 21.0 33.0 Degree(°) <![CDATA[N1]]> Number of cooling ducts supplied by the intermediate circuit 3 3 - <![CDATA[N2]]> Number of cooling duct sections spanning from the pressure surface to the suction surface 2 2 - SRSR Stator rotor seal radius 0.224 0.239 meter (m) EQ expression 4.949 13.039 n / a

[0102] Additional benefits associated with the blade assembly 30 with the stator-rotor seal 57 and the intermediate inlet passage 48m 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 blade assemblies and vane assemblies for the engine design and specific blade design. Narrowing these various factors to a possible area saves time, money, and resources. The blade assembly 30 with the stator-rotor seal 57 and the intermediate inlet passage 48m 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.

[0103] 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 relative expression (EQ) values ​​outside the range of 4.949 to 13.039. Lower durability results in less time on wing (TOW) and greater maintenance costs.

[0104] 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.

[0105] In other examples, increasing the size of airfoils or related components, utilizing stronger materials, and / or providing additional cooling features can counteract centrifugal and thermal stresses. However, such increased size, stronger materials, and additional cooling features can result in increased cost, system weight, the total space occupied by the blade assembly, performance losses, and 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.

[0106] As described above, the inventors have discovered that Examples 1-10 of Tables 2 and 3 provide successful solutions without increasing thickness, 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 4.949 to 13.039 (i.e., 4.949 ≤ EQ ≤ 13.039) 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 making areas of the airfoil thicker, or using heavier, higher-strength materials, or adding additional cooling features.

[0107] 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.

[0108] To the extent that one or more structures provided herein may be known in the art, it is understood that the present disclosure may include combinations of structures not previously known, based at least in part on conflicts of interest and expense, desired modes of operation, or other forms of teaching in the art.

[0109] 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.

[0110] Further aspects of the present disclosure are provided by the subject matter of the following clauses:

[0111] 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 including a leading edge and a trailing edge; and a shank, the shank coupled to the lower surface, the shank having a base defining a base plane; and an intermediate inlet passage located within the shank, the intermediate inlet passage having: a first cooling conduit portion located at a first plane radially spaced 0.0167 meters from the base plane, the first cooling conduit portion having a first cross-sectional area, the first cooling portion defining a first boundary circle, the first cooling portion having: a first point at a first 12:30 o'clock position on the first boundary circle; and a second point at a first 1:30 o'clock position on the first boundary circle, the first point and the a second point defining a first line; a second cooling duct portion located at a second plane radially spaced 0.0197 meters from the base plane, the second cooling duct portion having a second cross-sectional area; a third cooling duct portion located at a third plane radially spaced 0.0215 meters from the base plane, the third cooling duct portion having a third cross-sectional area; a fourth cooling duct portion located at a fourth plane radially spaced 0.0234 meters from the base plane, the fourth cooling duct portion having a fourth cross-sectional area; a fifth cooling duct portion located at a fifth plane radially spaced 0.0245 meters from the base plane, the fifth cooling duct portion having a fifth cross-sectional area, the second cooling portion defining a second bounding circle, the second cooling portion having: a third point at a second 12:30 o'clock position on the second bounding circle and a fourth point at a third 1:30 o'clock position on the second bounding circle, the third point and the fourth point defining a second line, wherein the intermediate inlet passage defines a maximum normalized cooling duct portion area (A) within a range between 1.30 and 2.10 max ), A max Defined by the following formula: wherein the first line and the second line define a degree of variation of the cooling duct portion along the suction side (DCCPASS) ranging between 21.0 and 33.0 degrees (°); and a plurality of cooling ducts located within the airfoil, the plurality of cooling ducts including a first number (N1) of cooling ducts fed by the intermediate inlet passage, wherein the first number is 2, wherein the plurality of cooling ducts define a second number (N2) of cooling duct portions spanning between the suction side and the pressure side, wherein the second number is 3, and wherein:

[0112] The blade assembly of any preceding clause, further comprising a plurality of inlet channels within the shank, the plurality of inlet channels including the intermediate inlet channel.

[0113] 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.

[0114] The blade assembly of any preceding clause, wherein the plurality of inlet channels further comprises a leading edge inlet channel and a trailing edge inlet channel.

[0115] The blade assembly of any preceding clause, wherein the shank is configured as a dovetail.

[0116] The blade assembly of any preceding clause, wherein the blade assembly is a first stage blade assembly of a high pressure turbine of the gas turbine engine.

[0117] A blade assembly as described in any preceding clause, wherein said number (N1) of cooling ducts fed by said intermediate inlet passage comprises: a forward serpentine cooling duct in said airfoil; and an aft serpentine cooling duct in said airfoil.

[0118] A blade assembly according to any of the preceding clauses, wherein the intermediate inlet passage is fluidly coupled to the front serpentine cooling duct and the rear serpentine cooling duct at a separator, the separator comprising: a first cooling duct portion, the first cooling duct portion being located at a sixth plane radially spaced 0.449 meters from the base plane, the first cooling duct portion being associated with the front serpentine cooling duct, the first cooling duct portion having a first major axis; and a second cooling duct portion at the sixth plane, the second cooling duct portion being associated with the rear serpentine cooling duct, the second cooling duct portion having a second major axis, wherein the first major axis and the second major axis define an L-shape.

[0119] 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 coupled to the lower surface, the handle having a base defining a base plane; and a central entryway within the handle, the central entryway having: a first cooling conduit portion, the first cooling conduit portion being located at a first plane radially spaced 0.0167 meters from the base plane, the first cooling conduit portion having a first cross-sectional area, the first cooling portion defining a first bounding circle, the first cooling portion having: a first point at a first 12:30 o'clock position on said first bounding circle; and a second point at a first 1:30 o'clock position on the first bounding circle, the first point and the second point defining a first line; a second cooling conduit portion located at a second plane radially spaced 0.0197 meters from the base plane, the second cooling conduit portion having a second cross-sectional area; a third cooling duct portion located at a third plane radially spaced 0.0215 meters from the base plane, the third cooling duct portion having a third cross-sectional area; a fourth cooling conduit portion located at a fourth plane radially spaced 0.0234 meters from the base plane, the fourth cooling conduit portion having a fourth cross-sectional area; a fifth cooling duct portion, the fifth cooling duct portion being located at a fifth plane radially spaced 0.0245 meters from the base plane, the fifth cooling duct portion having a fifth cross-sectional area, the fifth cooling portion defining a second boundary circle, the second cooling portion having: a third point at a second 12:30 o'clock position on the second boundary circle; and a fourth point at a third 1:30 o'clock position on the second boundary circle, the third point and the fourth point defining a second line; wherein the intermediate inlet passage defines a maximum normalized cooling duct partial area (A) in the range of 1.30 to 2.

10. max ), A max Defined by the following formula: wherein the first line and the second line define a degree of variation of the cooling duct portion along the suction side (DCCPASS), the degree of variation ranging between 21.0 and 33.0 degrees (°); and a plurality of cooling conduits located within the airfoil, the plurality of cooling conduits comprising a first number (N1) of cooling conduits fed by the intermediate inlet passage, wherein the first number is two, wherein the plurality of cooling conduits define a second number (N2) of cooling conduit sections spanning between the suction side and the pressure side, wherein the second number is three, and wherein:

2. The blade assembly according to claim 1, wherein: Further included are a plurality of inlet channels within the handle, the plurality of inlet channels including the middle 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 leading edge 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 of the gas turbine engine.

7. The blade assembly according to claim 1, wherein: wherein said number (N1) of cooling ducts supplied by said intermediate inlet passage comprises: A forward serpentine cooling duct in the airfoil; and A rear serpentine cooling duct in the airfoil.

8. The blade assembly according to claim 7, wherein: wherein the intermediate inlet passage is fluidly coupled to the front serpentine cooling duct and the rear serpentine cooling duct at a separator, the separator comprising: a first cooling duct portion, the first cooling duct portion being located at a sixth plane radially spaced 0.449 meters from the base plane, the first cooling duct portion being associated with the front serpentine cooling duct, the first cooling duct portion having a first major axis; and a second cooling duct portion at the sixth plane, the second cooling duct portion being associated with the rear serpentine cooling duct, the second cooling duct portion having a second major axis, wherein: The first major axis and the second major axis define an L-shape.

9. The blade assembly according to claim 1, wherein: wherein each of the second number of cooling duct sections spans at least 60% of the thickness of the airfoil.

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