IMPROVED TURBINE BLADE COOLING SYSTEM

Introduction

The present disclosure generally pertains to gas turbine engines. More particularly this application is directed toward a turbine blade with improved cooling capabilities.

Internally cooled turbine blades may include passages and vanes (air deflectors) within the blade. These hollow blades may be cast. In casting hollow gas turbine engine blades having internal cooling passageways, a fired ceramic core is positioned in a ceramic investment shell mold to form internal cooling passageways in the cast airfoil. The fired ceramic core used in investment casting of hollow airfoils typically has an airfoil-shaped region with a thin cross- section leading edge region and trailing edge region. Between the leading and trailing edge regions, the core may include elongated and other shaped openings so as to form multiple internal walls, pedestals, turbulators, ribs, and similar features separating and/or residing in cooling passageways in the cast airfoil.

The present disclosure is directed toward overcoming one or more of the problems discovered by the inventors.

Brief Description Of The Fi ures

The details of embodiments of the present disclosure, both as to their structure and operation, may be gleaned in part by study of the

accompanying drawings, in which like reference numerals refer to like parts, and in which:

FIG. 1 is a schematic illustration of an exemplary gas turbine engine;

FIG. 2 is an axial view of an exemplary turbine rotor assembly; FIG. 3 is an isometric view of one turbine blade of FIG. 2,

FIG. 4 is a cutaway side view of the turbine blade of FIG. 3;

FIG. 5 is a cross section of the cooled turbine blade taken along the line 5 - 5 of FIG. 4; FIG. 6 is a cross section of the cooled turbine blade taken along the line 6 - 6 of FIG. 4;

FIG. 7 is a cross section of the cooled turbine blade taken along the line 7 - 7 of FIG. 4,

FIG. 8 is a cross section of the cooled turbine blade taken along the line 8 - 8 of FIG. 4;

FIG. 9 is a cutaway perspective view of a portion of the turbine blade of FIG. 3;

FIG. 10 is a cutaway perspective view of a portion of the turbine blade of FIG. 3;

FIG. 1 1 is a cutaway perspective view of a portion of the turbine blade of FIG. 3;

FIG. 12 is a cutaway perspective view of a portion of the turbi ne blade of FIG. 3; and

FIG 13 is a cutaway perspective view' of a portion of the turbine blade of FIG 3.

Detailed Description

The detailed description set forth below, in connection with the accompanying drawings, is intended as a description of various embodiments and is not intended to represent the only embodiments in which the disclosure may be practiced. The detailed description includes specific details for the purpose of providing a thorough understanding of the embodiments. However, it will be apparent to those skilled in the art that the disclosure without these specific details. In some instances, well-known structures and components are shown in simplified form for brevity of description.

FIG. 1 is a schematic illustration of an exemplary gas turbine engine. Some of the surfaces have been left out or exaggerated (here and in other figures) for clarity and ease of explanation. Also, the disclosure may reference a forward and an aft direction. Generally, all references to“forward” and“aft” are associated with the flow direction of primary air (i.e., air used in the combustion process), unless specified otherwise. For example, forward is“upstream” relative to primary air flow, and aft is“downstream” relative to primary air flow .

In addition, the disclosure may generally reference a center axis 95 of rotation of the gas turbine engine, which may be generally defined by the longitudinal axis of its shaft 120 (supported by a plurality of bearing assemblies 150). The center axis 95 may be common to or shared with various other engine concentric components. All references to radial, axial, and circumferential directions and measures refer to center axis 95, unless specified otherwise, and terms such as“inner” and“outer” generally indicate a lesser or greater radial distance from, wherein a radial 96 may be in any direction perpendicular and radiating outward from center axis 95.

Structurally, a gas turbine engine 100 includes an inlet 110, a gas producer or“compressor” 200, a combustor 300, a turbine 400, an exhaust 500, and a pov er output coupling 600. The compressor 200 includes one or more compressor rotor assemblies 220. The combustor 300 includes one or more injectors 350 and includes one or more combustion chambers 390. The turbine 400 includes one or more turbine rotor assemblies 420. The exhaust 500 includes an exhaust diffuser 520 and an exhaust collector 550.

As illustrated, both compressor rotor assembly 220 and turbine rotor assembly 420 are axial flow rotor assemblies, where each rotor assembly includes a rotor disk that is circumferentially populated with a plurality of airfoils (“rotor blades”). When installed, the rotor blades associated with one rotor disk are axially separated from the rotor blades associated with an adjacent disk by stationary vanes (“stator vanes” or“stators”) 250, 450 circumferentially distributed in an annular casing.

Functionally, a gas (typically air 10) enters the inlet 110 as a “working fluid”, and is compressed by the compressor 200. In the compressor 200, the working fluid is compressed in an annular flow path 115 by the series of compressor rotor assemblies 220. In particular, the air 10 is compressed in numbered“stages”, the stages being associated with each compressor rotor assembly 220. For example,“4th stage air” may be associated with the 4th compressor rotor assembly 220 in the downstream or“aft” direction -going from the inlet 110 towards the exhaust 500). Likewise, each turbine rotor assembly 420 may be associated with a numbered stage For example, first stage turbine rotor assembly 421 is the forward most of the turbine rotor assemblies 420. However, other numbering/naming conventions may also be used.

Once compressed air 10 leaves the compressor 200, it enters the combustor 300, where it is diffused and fuel 20 is added. Air 10 and fuel 20 are injected into the combustion chamber 390 via injector 350 and ignited. After the combustion reaction, energy is then extracted from the combusted fuel/air mixture via the turbine 400 by each stage of the series of turbine rotor assemblies 420. Exhaust gas 90 may then be diffused in exhaust diffuser 520 and collected, redirected, and exit the system via an exhaust collector 550. Exhaust gas 90 may also be further processed (e.g., to reduce harmful emissions, and/or to recover heat from the exhaust gas 90).

One or more of the above components (or their subcomponents) may be made from stainless steel and/or durable, high temperature materials known as“superalloys”. A superal!oy, or high-performance alloy, is an alloy that exhibits excellent mechanical strength and creep resistance at high temperatures, good surface stability, and corrosion and oxidation resistance Superalloys may include materials such as HASTELLOY, INCONEL, WASPALOY, RENE alloys, HAYNES alloys, INCOLOY, MP98T, TMS alloys, and CMSX single crystal alloys.

FIG. 2 is an axial view of an exemplary turbine rotor assembly. In particular, first stage turbine rotor assembly 421 schematically illustrated in FIG.

1 is shown here in greater detail, but in isolation from the rest of gas turbine engine 100 First stage turbine rotor assembly 421 includes a turbine rotor disk 430 that is circumferentially populated with a plurality of turbine blades configured to receive cooling air (“cooled turbine blades” 440) and a plurality of dampers 426. Here, for illustration purposes, turbine rotor disk 430 is shown depopulated of all but three cooled turbine blades 440 and three dampers 426.

Each cooled turbine blade 440 may include a base 442 including a platform 443 and a blade root 480. For example, the blade root 480 may incorporate“fir tree”,“bulb”, or“dove tail” roots, to list a few. Correspondingly, the turbine rotor disk 430 may include a plurality of circumferentially distributed slots or“blade attachment grooves” 432 configured to receive and retain each cooled turbine blade 440. In particular, the blade attachment grooves 432 may be configured to mate with the blade root 480, both having a reciprocal shape with each other. In addition the blade attachment grooves 432 may be slideably engaged with the blade attachment grooves 432, for example, in a forward-to-aft direction.

Being proximate the combustor 300 (FIG. 1), the first stage turbine rotor assembly 421 may incorporate active cooling. In particular, compressed cooling air may be internally supplied to each cooled turbine blade 440 as well as predetermined portions of the turbine rotor disk 430. For example, here turbine rotor disk 430 engages the cooled turbine blade 440 such that a cooling air cavity 433 is formed between the blade attachment grooves 432 and the blade root 480. In other embodiments, other stages of the turbine may incorporate active cooling as well.

When a pair of cooled turbine blades 440 is mounted in adjacent blade attachment grooves 432 of turbine rotor disk 430, an under-platform cavity may be formed above the circumferential outer edge of turbine rotor disk 430, between shanks of adjacent blade roots 480, and below their adjacent platforms 443, respectively. As such, each damper 426 may be configured to fit this under platform cavity. Alternately, where the platforms are flush with circumferential outer edge of turbine rotor disk 430, and/or the under-platform cavity is sufficiently small, the damper 426 may be omitted entirely.

Here, as illustrated, each damper 426 may be configured to constrain received cooling air such that a positive pressure may be created within under-platform cavity to suppress the ingress of hot gases from the turbine. Additionally, damper 426 may be further configured to regulate the flow of cooling air to components downstream of the first stage turbine rotor assembly 421. For example, damper 426 may include one or more aft plate apertures in its aft face. Certain features of the illustration may be simplified and/or differ from a production part for clarity. Each damper 426 may be configured to be assembled with the turbine rotor disk 430 during assembly of first stage turbine rotor assembly 421, for example, by a press fit. In addition, the damper 426 may form at least a partial seal with the adjacent cooled turbine blades 440 Furthermore, one or more axial faces of damper 426 may be sized to provide sufficient clearance to permit each cooled turbine blade 440 to slide into the blade attachment grooves 432, past the damper 426 without interference after installation of the damper 426.

FIG. 3 is a perspective view of the turbine blade of FIG. 2. As described above, the cooled turbine blade 440 may include a base 442 having a platform 443 and a blade root 480. Each cooled turbine blade 440 may further include an airfoil 441 extending radially outward from the platform 443 The airfoil 441 may have a complex, geometry that varies radially. For example the cross section of the airfoil 441 may lengthen, thicken, twist, and/or change shape as it radially approaches the platform 443 inward from the tip end 445. The overall shape of airfoil 441 may also vary from application to application.

The cooled turbine blade 440 is generally described herein with reference to its installation and operation. In particular, the cooled turbine blade 440 is described with reference to both a radial 96 of center axis 95 (FIG. 1) and the aerodynamic features of the airfoil 441. The aerodynamic features of the airfoil 441 include a leading edge 446, a trailing edge 447, a pressure side 448, a lift side 449, and its mean camber line 474. The mean camber line 474 is generally defined as the line running along the center of the airfoil from the leading edge 446 to the trailing edge 447. It can be thought of as the average of the pressure side 448 and lift side 449 of the airfoil shape. As discussed above, airfoil 441 also extends radially between the platform 443 and the tip end 445. Accordingly, the mean camber line 474 herein includes the entire camber sheet continuing from the platform 443 to the tip end 445.

Thus, when describing the cooled turbine blade 440 as a unit, the inward direction is generally radially inward toward the center axis 95 (FIG. 1), with its associated end called the“root end” 444. Likewise is the outward direction is generally radially outward from the center axis 95 (FIG. 1), with its associated end called the“tip end” 445. When describing the platform 443, the forward edge 484 and the aft edge 485 of the platform 443 are associated the forward and aft axial directions of the center axis 95 (FIG. 1), as described above.

In addition, when describing the airfoil 441, the forward and aft directions are generally measured between its leading edge 446 (forward) and its trailing edge 447 (aft), along the mean camber line 474 (artificially treating the mean camber line 474 as linear). When describing the flow features of the airfoil 441, the inward and outward directions are generally measured in the radial direction relative to the center axis 95 (FIG. 1). However, when describing the thermodynamic features of the airfoil 441 (particularly those associated with the inner spar 462 (FIG. 4)), the inward and outward directions are generally measured in a plane perpendicular to a radial 96 of center axis 95 (FIG. 1) with imvard being toward the mean camber line 474 and outward being toward the “skin” 460 of the airfoil 441.

Finally, certain traditional aerodynamics terms may be used from time to time herein for clarity, but without being limiting. For example, while it will be discussed that the airfoil 441 (along with the entire cooled turbine blade 440) may be made as a single metal casting, the outer surfac e of the airfoil 441 (along with its thickness) is descriptively called herein the“skin” 460 of the airfoil 441.

FIG 4 is a cutaway side view of the turbine blade of FIG. 3. In particular, the cooled turbine blade 440 of FIG. 3 is shown here with the skin 460 removed from the pressure side 448 of the airfoil 441 , exposing its internal structure and cooling paths. The airfoil 441 may include a composite flow path made up of multiple subdivisions and cooling structures. Similarly, a section of the base 442 has been removed to expose portions of a cooling air passageway 482, internal to the base 442. The cooling air passageway 482 can have one or more channels 483 extending from the blade root 480 toward the tip end 445 as described below.

The cooled turbine blade 440 may include an airfoil 441 and a base 442. The base 442 may include the platform 443, the blade root 480, and one or more cooling air inlet(s) 481. The airfoil 441 interfaces with the base 442 and may include the skin 460, a tip wall 461, and the cooling air outlet 471. Compressed secondary air may be routed into one or more cooling air inlet(s) 481 in the base 442 of cooled turbine blade 440 as cooling air 15. The one or more cooling air inlet(s) 481 may be at any convenient location. For example, here the cooling air inlet 481 is located in the blade root 480.

Alternately, cooling air 15 may be received in a shank area radially outward from the blade root 480 but radially inward from the platform 443.

Within the base 442, the cooled turbine blade 440 includes the cooling air passageway 482 that is configured to route cooling air 15 from the one or more cooling air inlet(s) 481, through the base, and into the airfoil 441 via the channels 483. The cooling air passageway 482 may be configured to translate the cooling air 15 in three dimensions (e.g , not merely in the plane of the figure) as it travels radially up (e.g., generally in the along a radial 96 of the center axis 95 (FIG 1)) towards the airfoil 441 and along the multi-bend heat exchange path 470. For example, the cooling air 15 can travel radially and within the air foil 441. Further, the inner spar 462 effectively splits the cooling air 15 between pressure side 448 and the lift side 449 The multi-bend heat exchange path 470 is depicted as a solid line drawn as a weaving path through the airfoil 441, exiting through the tip flag cooling system 650 (FIG. 13) ending with an arrow.

Moreover, the cooling air passageway 482 may be structured to receive the cooling air 15 from a generally rectilinear cooling air inlet 481 and smoothly "reshape" it fit the curvature and shape of the airfoil 441. In addition, the cooling air passageway 482 may be subdivided into a plurality of subpassages or channels 483 that direct the cooling air in one or more paths through the airfoil 441.

Within the skin 460 of the airfoil 441, several internal structures are viewable. In particular, airfoil 441 may include the tip wall 461, an inner spar 462, a leading edge chamber 463, one or more turning vane(s) 465, one or more air deflector(s) 466, and a plurality of inner spar cooling fins 467. In addition, airfoil 441 may include a perforated trailing edge rib 468 allowing flow of the cooling air 15 to exit the trailing edge 447. Together with the skin 460, these structures may form the multi-bend heat exchange path 470 within the airfoil 441. The internal structures making up the multi-bend heat exchange path 470 may form multiple discrete sub-passageways or "sections". For example, although multi-bend heat exchange path 470 is shown by a

representative path of cooling air 15, multiple paths are possible as described more detail in the following sections

With regard to the airfoil structures, the tip wall 461 extends across the airfoil 441 and may be configured to redirect cooling air 15 from escaping through the tip end 445. In an embodiment, the tip end 445 may be formed as a shared structure, such as a joining of the pressure side 448 and the lift side 449 of the airfoil 441. The tip wall 461 may be recessed inward such that it is not flush with the tip of the airfoil 441 The tip wall 461 may include one or more perforations (not shown) such that a small quantity of the cooling air 15 may be bled off for film cooling of the tip end 445.

The inner spar 462 may extend from the base 442 radially outward toward the tip wall 461, between the pressure side 448 (FIG. 3) and the lift side 449 (FIG. 3) of the skin 460. In addition, the inner spar 462 may extend between the leading edge 446 and the trailing edge 447, parallel with, and generally following, the mean camber line 474 (FIG. 3) of the airfoil 441 , and terminating with inner spar trailing edge 476. Accordingly, the inner spar 462 may be confi gured to bifurcate a portion or all of the airfoil 441 generally along its mean camber line 474 (FIG. 3) and between the pressure side 448 and the lift side 449. Also, the inner spar 462 may be solid (non-perforated) or substantially solid (including some perforations), such that cooling air 15 cannot pass.

According to an embodiment, the inner spar 462 may extend less than the entire length of the mean camber line 474. In particular the inner spar 462 may extend less than ninety percent of the mean camber line 474 and may exclude the leading edge chamber 463 entirely. For example, the inner spar 462 may extend from an edge of the leading edge chamber 463 proximate the trailing edge 447, downstream to the plurality of trailing edge cooling fins 469. In addition, the inner spar 462 may have a length within the range of seventy to eighty percent, or approximately three quarters the length of, and along, the mean camber line 474. In some embodiments, the inner spar 462 may have a length within the range of fifty to sixty percent, or approximately three quarters the length of, and along, the mean camber line 474.

According to an embodiment, the airfoil 441 may include a leading edge rib 472. The leading edge rib 472 may extend radially from an area proximate the base 442 toward the tip end 445, terminating prior to reaching the tip wall 461. In addition, the leading edge rib 472 may extend directly from the pressure side 448 (FIG. 3) of the skin 460 to the lift side 449 (FIG. 3) of the skin 460. In doing so, the leading edge rib 472 may form the leading edge chamber 463 in conjunction with the skin 460 at the leading edge 446 of the airfoil 441. Additionally, at least a portion of the cooling air 15 leaving the leading edge chamber 463 may be redirected toward the trailing edge 447 by tip wall 461 and other cooling air 15 within the airfoil 441. Accordingly, the leading edge chamber 463 may form part of the multi-bend heat exchange path 470.

Within the airfoil 441, the plurality of inner spar cooling fins 467 may extend outward from the inner spar 462 to the skin 460 on either of the pressure side 448 (FIG. 3) or the lift side 449 (FIG 3). In contrast, the plurality of trailing edge cooling fins 469 may extend from the pressure side 448 (FIG. 3) of the skin 460 directly to the lift side 449 (FIG 3) of the skin 460 Accordingly, the plurality of inner spar cooling fins 467 are located forward of the plurality of trailing edge cooling fins 469, as measured along the mean camber line 474 (FIG. 3) of the airfoil 441.

Both the inner spar cooling fins 467 and the trailing edge cooling fins 469 may be disbursed copiously throughout the single-bend heat exchange path 470. In particular, the inner spar cooling tins 467 and the trailing edge cooling fins 469 may be disbursed throughout the airfoil 441 so as to thermally interact with the cooling air 15 for increased cooling. In addition, the distribution may be in the radial direction and in the direction along the mean camber line 474 (FIG. 3). The distribution may be regular, irregular, staggered, and/or localized.

According to an embodiment, the inner spar cooling fins 467 may be long and thin. In particular, inner spar cooling fins 467, traversing less than half the thickness of the airfoil 441, may use a round "pin" fin. Moreover, pin fins having a height -to-diameter ratio of 2-7 may be used. For example, the inner spar cooling fins 467 may be pin fins having a diameter of 0.017 - .040 inches, and a length off the inner spar 462 of 0.034 - 0.240 inches.

Additionally, according to one embodiment, the inner spar cooling fins 467 may also be densely packed. In particular, inner spar cooling fins 467 may be within two diameters of each other. Thus, a greater number of inner spar cooling fins 467 may be used for increased cooling. For example, across the inner spar 462, the fin density may be in the range of 80 to 300 fins per square inch per side of the inner spar 462. The fin density may also be higher, in the 40 to 200 fins per square inch per side of the inner spar 462.

Within the airfoil 441, the trailing edge rib 468 may extend radially from the base 442 toward the tip end 445. The trailing edge rib 468 may be located along the inner spar trailing edge 476 and between the inner spar cooling fins 467 and the trailing edge cooling fins 469.

The trailing edge rib 468 may be perforated to include one or more openings. This will allow cooling air 15 to pass through the trailing edge rib 468 toward the cooling air outlet 471 in the trailing edge 447, and thus complete the single-bend heat exchange path 470.

Taken as a whole the cooling air passageway 482 and the multi- bend heat exchange path 470 may be coordinated. In particular and returning to the base 442 of the cooled turbine blade 440, the cooling air passageway 482 may be sub-divided into a plurality of flow paths. These flow paths may be arranged in a serial arrangement as the air 15 enters the blade root 480 at the cooling air inlet 482, as shown in FIG. 5. The cooling air inlets 481 can funnel the cooling air 15 into multiple subpassageways or channels 483, labeled individually 483a, 483b, 483c, 483d chord-wise along the blade root 480. The serial arrangement may be advantageous given the limited amount of available surface area on the blade root 480. Other (e.g., parallel) arrangements may limit the flow ^? of cooling air 15 into the cooling air inlets 481

The flow ^? path of the cooling air passageway 482 may change from the serial arrangement to a parallel or a series-parallel arrangement as the air 15 continues through the channels 483 and the multi-bend heat exchange path 470. These arrangements are described in further detail in connection with FIG. 5 through FIG. 9. Each subdivision within the base 442 may be aligned with and include a cross sectional shape (see, FIG. 5) corresponding to the areas bounded by the skin 460. In addition, the cooling air passageway 482 may maintain the same overall cross sectional area (i.e., constant flow rate and pressure) in each subdivision (e.g., the channels 483), as between the cooling air inlet 481 and the airfoil 441. Alternately, the cooling air passageway 482 may vary the cross sectional area of the individual channels 483 where differing performance parameters are desired for each section, in a particular application.

According to one embodiment, the cooling air passageway 482 and the multi-bend heat exchange path 470 may each include asymmetric divisions for reflecting localized thermodynamic flow performance requirements. In particular, as illustrated, the cooled turbine blade 440 may have two or more sections divided by the one or more serial or parallel channels 483.

According an embodiment, the individual inner spar cooling fins 467 and the trailing edge cooling fins 469 may also include localized

thermodynamic structural variations. In particular, the inner spar cooling fins 467 and/or the trailing edge cooling fins 469 may have different cross sections/surface area and/or fin spacing at different locations of the inner spar 462. For example, the cooled turbine blade 440 may have localized "hot spots" that favor a greater thermal conductivity, or low internal flow areas that favor reduced airflow resistance. In which case, the individual cooling fins may be modified in shape, size, positioning, spacing, and grouping.

According to one embod iment, one or more of the inner spar cooling fins 467 and the trailing edge cooling fins 469 may be pin fins or pedestals. The pin fins or pedestals may include many different cross-sectional areas, such as: circular, oval, racetrack, square, rectangular, diamond cross- sections, just to mention only a few. As discussed above, the pin fins or pedestals may be arranged as a staggered array, a linear array, or an irregular array.

In some embodiments, the cooling air 15 can flow into the blade root 480 via the cooling air inlet 481 into the cooling air passageway 482 (e.g , the channels 483). The cooling air passageway 482 can be arranged in multiple sections with different geometries arranged chord-wise along the cooled turbine blade 440 The varying geometries are shown in FIG. 5, FIG. 6, FIG. 7, and FIG. 8

The multi-bend heat exchange path 470 can proceed as follows. The cooling air 15 can enter the blade root 480 at the cooling air inlet 481 , flowing through the channels 483. The channels 483 can begin in a series arrangement (FIG. 5) at the blade root 480. In some embodiments, at least the channels 483b, 483c can enter a series-to-parallel transition 490 (indicated in dashed lines) that twists and redirects the channels 483b, 483c from the series arrangement at the blade root 480 to a parallel arrangement. The channels 483b, 483c can be routed radially outward toward the tip end 445 and a. first turning vane bank 500 shown in dashed lines (FIG. 10). The first turning vane bank 500 can redirect the cooling air 15 back toward the base 442 and a second turning vane bank 550 shown in dashed lines (FIG. 11). The second turning vane bank 550 can redirect the cooling air 15 toward the tip end 445 and transition the parallel flow of the channels 483b, 483c into a single, serial channel of the leading edge chamber 463. The leading edge chamber 463 can direct at least a portion of the cooling air 15 back toward the tip end 445 and a tip diffuser 600 shown in dashed lines (FIG. 12). The tip diffuser 600 can diffuse the cooling air 15 from the single (e.g., series) leading edge channel 463 into two parallel tip flag channels 652 (FIG. 8) within a tip flag cooling system 650 shown in dashed lines (FIG. 13).

FIG. 5 is a cross section of the cooled turbine blade taken along the line 5 - 5 of FIG. 4. The channels 483 can have a serial arrangement 512 at the cooling air inlet 481 proximate the blade root 480 As the cooling air passageway 482 approaches the level of the platfomi 443, the channels 483 can redirect cooling air 15 within the multi-bend heat exchange path 470 via a transition arrangement 514 toward a parallel arrangement. The transition arrangement 514 is a portion of a series-to-parallel transition 540, described in connection with FIG. 9.

FIG. 6 is a cross section of the cooled turbine blade taken along the line 6 - 6 of FIG. 4. As the cooling air flows through the cooling air passageway 482 in the transition arrangement 514, the channels 483b, 483c redirect the cooling air 15 into a parallel arrangement 516, where the cooling air inlets 481a, 481b are a side-by-side between the pressure side 448 and the lift side 449

FIG. 7 is a cross section of the cooled turbine blade taken along the line 7 - 7 of FIG. 4 The parallel arrangement 516 provides side-by-side channels 483b, 483c, separated by the inner spar 462, to channel cooling air 15 radially outward in a trailing edge section 522 toward the tip end 445, for example. The cooling air 15 can be redirected within the cooling air passageway 482 in the first turning vane bank 500 (FIG. 10) proximate the tip end 445. The cooling air 15 can then flow radially inward in a leading edge section 524 within the airfoil 441 away from the tip end 445 toward the second turning vane bank 550 (FIG 1 1). The second turning vane bank 550 can redirect the cooling 15 radially outward toward the tip end 445 into the leading edge chamber 463 As described in more detail below, the second turning vane bank 550 can include a parallel -to-series transition, redirecting the channels 483b, 483c from two parallel channels to a single channel within the leading edge chamber 463.

FIG. 8 is a cross section of the cooled turbine blade taken along the line 8 - 8 of FIG. 4. As the cooling air 15 approaches the tip end 445 within the lead ing edge chamber 463, at least a portion of the cooling air 15 enters the tip diffuser 600. The tip diffuser 600 includes a series-to-parallel transition that redirects the cooling air 15 from the single flow path within the leading edge chamber 463 to two parallel tip flag channels 652 (labeled as tip flag channel 652a and tip flag channel 652b) within the tip flag cooling system 650 (FIG. 13).

FIG. 9 is a cutaway perspective view of a portion of the turbine blade of FIG 3. As shown in FIG. 4 and FIG. 5, the cooling air 15 can enter the blade root 480 through the cooling air inlet 481 into the channels 483. The channels 483 can have the series arrangement 512 (FIG. 5) at the beginning of the cooling air passageway 482. The“serial” disposition can be arranged generally along the blade root 480. This can also substantially coincide with the fore and aft direction of the central axis 95 when the cooled turbine blade is installed in a turbine engine, for example. The series arrangement 512 can gradually redirect the cooling air 15 via the transition arrangement 514 (FIG. 6) into the parallel arrangement 516 (FIG. 7), where the channels 483b, 483c are side by side when viewed from the leading edge 446 to the trailing edge 447. The cross section lines 6 - 6 and 7 - 7 are repeated in this figure showing the approximate locations of the transition arrangement 514 (FIG. 6) and the parallel arrangement 516 (FIG. 7) for the channels 483.

The series-to-parallel transition 490 twists or redirects the series flow of cooling air 15 at the cooling air inlet 481 into a parallel arrangement (e.g., the parallel arrangement 516). Given space constraints at the blade root 480, the channels 483 are disposed in series near the air inlet 481. However, the series-to- parallel transition 490 twists the channels to a parallel cooling flow in main core of the airfoil 441 and provides more rapid or efficient heat transfer than a single (series) cooling path. Hence, cooling air flows in series at the inlet 481, twists and redirects the cooling air 15 to form the parallel flow ⁷ that continues toward the tip end 445. An advantage of the embodiments using parallel flow of the cooli ng air within the airfoil 441 is reduced pressure loss and increased fatigue life of the blade 440.

FIG. 10 is a cutaway perspective view of a portion of the turbine blade of FIG. 3. The first turning vane bank 500 is shown in dashed lines in FIG. 4. The first turning vane bank 500 shown is related to the channel 483b. Only the first turning vane bank 500 for the channel 483b is shown in this view, as the first turning vane bank for the channel 483c (e.g., on the lift side 449) is obscured.

The first turning vane bank 500 can have a first turning vane 502, a second turning vane 504, a third turning vane 506, a first corner vane 508, and a second corner vane 510. The first turning vane 502, the second turning vane 504, and the third turning vane 506 can be the same or similar to the at least one turning vane 465 described above in connection with FIG. 4. Additionally, the first corner vane 508, and a second corner vane 510 can be the same or similar to the one or more air deflector(s) 466 described above in connection with FIG. 4.

The first turning vane 502 and the second turning vane 504 can have a semi-circular shape that spans approximately 180 degrees. The third turning vane 506 can span an angle 513. The angle 513 can be approximately 120 degrees. Each of the first turning vane 502, the second turning vane 504, and the third turning vane 506 can have an even or symmetrical curvature. In some other embodiments, one or more of the first turning vane 502, the second turning vane 504, and the third turning vane 506 can have an asymmetrical curvature.

The first turning vane 502, the second turning vane 504, and the third turning vane 506 can each have a vane width 515. In the embodiment shown, the vane width 515 is a uniform width along the entire curvature of the first turning vane 502, the second turning vane 504, and the third turning vane 506. In some other embodiments, the first turning vane 502, the second turning vane 504, and the third turning vane 506 have non uniform vane width 515 The first turning vane 502 can be separated or displaced from the second turning vane 504 by a first vane spacing 517. The second turning vane 504 can be separated from the third turning vane 506 by a second vane spacing 519. In some embodiments, the first vane spacing 517 and the second vane spacing 519 can be approximately two times the vane width 515 (e.g., 2: 1 ratio). In some embodiments, the first vane spacing 517 can be different from the second vane spacing 519 For example, the first vane spacing 517 can be two times the vane width 515 and the second vane spacing 519 can be two to three times the vane width 515. In some embodiments, the spacing-to-width ratio can also be higher, for example having a 2: 1, 3: 1, or 4: 1 spacing-to-width ratio, for example. The first vane spacing 517 and the second vane spacing 519 do not have to be equivalent. The first vane spacing 517 and the second vane spacing 519 can also be the same, or equivalent.

The first corner vane 508 and the second corner vane 510 can be spaced approximately 90 degrees apart, with respect to the turning vanes. The first corner vane 508 and the second corner vane 510 can also have an aerodynamic shape having a chord length to width ratio of approximately 2: 1 to 3: 1 ratio. The first corner vane 508 and the second corner vane 510 have sizes and positions selected to maximize cooling in a leading comer 526 and a trailing corner 528. The first turning vane bank 500 can also have one or more turbulators 430. The turbulators 430 can be formed as ridges on the inner spar 462. The turbulators 430 can be positioned between the turning vanes 502, 504, 506 in various locations. The turbulators 430 can interrupt flow along the inner spar 462 and prevent formation of a boundary layer which can decrease cooling effects of the cooling air 15. The first turning vane bank 500 can have one or more turbulators 430 turbulators below the first turning vane 502. One turbulators 430 is shown below _' the first turning vane 502 in FIG. 10. Three turbulators are shown between the first turning vane 502 and the second turning vane 504. In some embodiments more or turbulators 430 may be present between the first turning vane 502 and the second turning vane 504. Two turbulators are shown between the second turning vane 504 and the third turning vane 506. However, in some embodiments more or fewer turbulators 430 may be present between the second turning vane 504 and the third turning vane 506.

The size, arrangement, shape of the turning vanes 502, 504, 506 and their respective separation or distance between the vanes, are selected to optimize cooling effectiveness of the cooling air 15 and increase fatigue life of the cooled turbine blade 440. The cooling air 15 can move through the first turning vane bank 500 with a minimum loss off pressure and a smooth manner.

FIG. I I is a cutaway perspective view' of a portion of the turbine blade of FIG. 3. The cooling air 15 flows radially inward (e.g., in the leading edge section 524 of FIG. 7) away from the first turning vane bank 500 in both the channel 483b and the channel 483c, separated by the inner spar 462. The cooling air 15 in both the channels 483b, 483c is then routed radially inward toward the second turning vane bank 550.

The two channels 483b, 483c in the leading edge section 524 are in a parallel arrangement, flowing radially inward toward the blade root 480. The second turning vane bank 550 can have at least one turning vane 552 that redirects the cooling air 15 into the leading edge chamber 463. Accordingly, the parallel arrangement of the channels 483b, 483c converges into the leading edge chamber 463 as a single, serial channel flowing radially outward tow'ard the tip end 445. The turning vane 552 can have a symmetrical curve, spanning approximately 180 degrees. In some embodiments, the turning vane 552 can alternatively have an asymmetrical curve. The second turning vane bank 550 can also have a second turning bank wall 554 that has a similar curvature as the turning vane 552. However, the curvature of the second turning bank wall 554 and the turning vane 552 do not have to be the same. The spacing between the turning vane 552 and the second turning bank wall 554 provides a smooth path for the cooling air 15. This can prevent hotspots on the second turning bank wall 554 and other adjacent components.

The turning vane 552 can be separated or otherwise decoupled from the inner spar 462 and the leading edge rib 472, for example. The inner spar 462 can further have a cutout 558 that provides a separation from the turning vane 552. The cutout 558 and separation between the turning vane 552 and the leading edge rib 472, for example, can prevent hotspots and increase fatigue life of the cooled turbine blade 440. The size, number, spacing, shape and

arrangement of the turning vanes 552 in the second turning vane hank 550 can vary and is not limited to the one shown. Multiple turning vanes 552 can be implemented.

FIG. 12 is a cutaway perspective view of a portion of the turbine blade of FIG. 3. The cooling air 15 can follow the multi-bend heat exchange path 470 past the second turning vane bank 550 and flow radially outward in the leading edge chamber 463. The leading edge chamber 463 can have a plurality of perforations 464 that provide a flow path for the cooling air 15. A portion of the cooling air 15 may flow through the perforations 464 and out cooling holes along the leading edge 446 of the cooled turbine blade 440

The cooling air 15 can then flow from the leading edge chamber 463 in a series flow into the tip diffuser 600. The tip diffuser 600 may refer to the area depicted in FIG. 12 proximate the tip end 445 and the leading edge 446. The tip diffuser 600 can receive the cooling air 15 from the leading edge chamber 463. The tip diffuser 600 can direct the cooling air through two diffuser outputs 602 and into two, parallel tip flag channels 652 (labeled individually tip flag channels 652a, 652b). The diffuser outputs 602 can be referred to as a first diffuser output 602a and a second diffuser output 602b. Similarly, the tip flag channels 652 may be referred to individually as a first tip flag channel 652a and a second tip flag channel 652b each coupled to a respective one of the diffuser outputs 602. The second tip flag channel 652b is not fully visible due to the aspect of the figure.

In some examples, other cooling mechanisms and the path of the cooling air 15 may not maximize cooling at the leading edge 446. In addition, discharge of the cooling 15 air to parallel tip flag channels can also be low. This can lead to pressure losses and decreased fatigue life of the b lade 440.

The tip diffuser 600 can act as a collector positioned at the leading edge chamber 463. The tip diffuser 600 can have diffuser box 660 having a U- shaped cross section as viewed along the mean camber line 474, with the bottom of the“U” disposed proxim ate the tip end 445. The U-shaped portion can accumulate maximum the cooling air 15 from the leading edge chamber 463.

This cooling air can be re-directed to the parallel tip flag channels 652 tip of the tip flag cooling system 650. The cooling air 15 can have radial flow and axial flow from two sources that combine at the tip diffuser 600. For example, the axial flow can be collected from the leading edge chamber 463 and the radial flow can be collected from the channel 483a, flowing directly through the leading edge. The curvature of the diffuser box 660 provides collecting of the cooling air 15, redirection to parallel axial flow to the tip flag channels 652, and impingement cooling of the tip end 445 at a tip edge 662 of the diffuser box 660. At the same time, the cooling air 15 can cool the area around the tip diffuser 600 and the flow through the diffuser outputs 602.

FIG. 13 is a cutaway perspective view ⁷ of a portion of the turbine blade of FIG. 3. The cooling air 15 can exit the tip diffuser 600 through the diffuser outputs 602 into the tip flag cooling system 650. The tip flag cooling system 650 can have the two parallel tip flag channels 652. However, only the tip flag channel 652a is shown in this view due to aspect. The features of the tip flag channel 652b are the same as the tip flag channel 652a. FIG. 8 show's the second tip flag channels 652b in a tip-down cross section of the parallel flow pattern of the tip flag channels 652. The tip flag channels 652 extend from the tip diffuser 600 along the pressure side 448 and the lift side 449 and join at a tip diffuser trailing edge 656. The tip flag channels 652a, 652b rejoin at the tip diffuser trailing edge 656 and form the tip flag output channel 658 (see also FIG. 8). This arrangement then forms a parallel-to-series flow as depicted in FIG. 8. The series flow through the tip flag output channel 658 can eject the cooling air 15 via the cooling air outlets 471 to the trailing edge 447.

The tip flag output channel 658 can decrease in camber width approaching an area proximate the trailing edge 447. In this sense, the camber width is a distance from the pressure side 448 to the lift side 449. The tip flag output channel 658 can also increase is height from the tip diffuser trailing edge 656 to the trailing edge 447. For example, the tip flag output channel 658 can have a height 664 proximate the tip diffuser trailing edge 656. The tip flag output channel 658 can have a height 666 proximate the trailing edge 447. The height 666 can be greater than the height 664. Thus, as the tip flag output channel 658 narrows from the pressure side 448 to the lift side 449 and the height increases, the mass flow of the cooling air 15 through the tip flag cooling system 650 can remain generally constant, except for film cooling holes (not shown) that penetrate the pressure side 448 in the area of the tip flag cooling system 650. The film cooling holes may allow some cooling air 15 to escape through the pressure side 448 which can subtract off some of the cooling air 15.

The design of tip cooling system includes parallel to series cooling paths. The parallel paths of cooling air are joined to form an expanded series flow path. So, there is an expanded trailing edge cooling path. Such a pattern of cooling paths provide effective and efficient cooling of tip of turbine blade.

Industrial Applicability

The present disclosure generally applies to cooled turbine blades, and gas turbine engines having cooled turbine blades. The described

embodiments are not limited to use in conjunction with a particular type of gas turbine engine, but rather may be applied to stationary or motive gas turbine engines, or any variant thereof. Gas turbine engines, and thus their components, may be suited for any number of industrial applications, such as, but not limited to, various aspects of the oil and natural gas industry (including include transmission, gathering, storage, withdrawal, and lifting of oil and natural gas), power generation industry, cogeneration, aerospace and transportation industry, to name a few examples.

Generally, embodiment s of the presently discl osed cooled turbine blades are applicable to the use, assembly, manufacture, operation, maintenance, repair, and improvement of gas turbine engines, and may be used in order to improve performance and efficiency, decrease maintenance and repair, and/or lower costs. In addition, embodiments of the presently disclosed cooled turbine blades may be applicable at any stage of the gas turbine engine’s life, from design to prototyping and first manufacture, and onward to end of life. Accordingly, the cooled turbine blades may be used in a first product, as a retrofit or enhancement to existing gas turbine engine, as a preventative measure, or even in response to an event. This is particularly true as the presently disclosed cooled turbine blades may conveniently include identical interfaces to be interchangeable with an earlier type of cooled turbine blades.

As discussed above, the entire cooled turbine blade may be cast formed. According to one embodiment, the cooled turbine blade 440 may be made from an investment casting process. For example, the entire cooled turbine blade 440 may be cast from stainless steel and/or a superalloy using a ceramic core or fugitive pattern. Accordingly, the inclusion of the inner spar is amenable to the manufacturing process. Notably, while the structures/features have been described above as discrete members for clarity, as a single casting, the structures/features may pass through and be integrated with the inner spar.

Alternately, certain structures/features (e.g., skin 460) may be added to a cast core, forming a composite structure.

Embodiments of the presently disclosed cooled turbine blades provide for a lower pressure cooling air supply, which makes it more amenable to stationary gas turbine engine applications. In particular, the single bend provides for less turning losses, compared to serpentine configurations. In addition, the inner spar and copious cooling fin population provides for substantial heat exchange during the single pass. In addition, besides structurally supporting the cooling fins, the inner spar itself may serve as a heat exchanger. Finally, by including subdivided sections of both the single-bend heat exchange path in the airfoil, and the cooling air passageway in the base, the cooled turbine blades may be tunable so as to be responsive to local hot spots or cooling needs at design, or empirically discovered, post-production.

The disclosed multi-bend heat exchange path 470 begins at the base 442 where pressurized cooling air 15 is received into the airfoil 441. The cooling air 15 is received from the cooling air passageway 482 and the channels 483 in a generally radial direction. The channels 483 are arranged serially at the blade root 480. As the cooling air enters the base 442 the channels 483 are redirected from a serial arrangement into a parallel arrangement near the end of the airfoil 441 proximate the root 480. A parallel arrangement provides increased cooling effects of the cooling air 15 as it passed through the multi-bend heat exchange path 470 and past the cooling fins 467.

The cooling air 15 follows the parallel channels 483b, 483c toward the first turning vane bank 500, which efficiently redirects the cooling air back toward the base 442 and the second turning vane bank 550. The second turning vane bank 550 has a turning vane 552 that redirects the cooling air 15 back in the direction of the tip end 445. The turning vane 552 also includes a parallel to series arrangement that directs the channels 483b, 483c into the leading edge chamber 463. The l eading edge chamber 463 carri es at least a porti on of the cooling air toward the tip end 445 while allowing a portion of the cooling air to escape through the perforations 464 to cool the leading edge 446 of the cooled turbine blade.

As the cooling air 15 approaches the tip end 445 within the leading edge chamber 463, all or part of the cooling air can enter the tip diffuser 600. The tip diffuser 600 receives the cooling air 15 from the leading edge chamber 463 and the channel 483a, or main body serpentine (main body). The tip diffuser 600 includes a series to parallel flow transition as the cooling air 15 leaves the leading edge chamber 463 and impinges on the U-shaped diffuser box 660. The cooling air 15 can then be redirected toward the trailing edge 447 by tip wall 461 via the tip flag channels.

The tip flag channels 562 are parallel flow channels that take advantage of increased surface area for cooling the internal surfaces of the airfoil 441. The tip flag cooling system 650 also implements a parallel to series transition at the tip diffuser trailing edge 656. The output of the tip flag cooling system narrows along the camber (e.g., from the pressure side 448 to the lift side 449) while increasing in height (measured span-wise) along the trailing edge 447. This can maintain a constant mass flow ⁷ rate and constant pressure as the cooling air 15 leaves the tip flag cooling system at the cooling air outlet 471.

The multi-bend heat exchange path 470 is configured such that cooling air 15 will pass between, along, and around the various internal structures, but generally flows in serpentine path as viewed from the side view from the blade root 480 back and forth toward and away from the tip end 445 (e.g., conceptually treating the camber sheet as a plane). Accordingly, the multi bend heat exchange path 470 may include some negligible lateral travel (e.g., into and out of the plane) associated with the general curvature of the airfoil 441.

Also, as discussed above, although the multi-bend heat exchange path 470 is illustrated by a single representative flow line traveling through a single section for clarity, the multi-bend heat exchange path 470 includes the entire flow path carrying cooling air 15 through the airfoil 441. With the implementation of the first turning vane bank 500, the second turning vane bank 550, the tip diffuser 600 and the tip flag cooling system 650, the multi-bend heat exchange path 470 makes use of the serpentine flow path with minimum flow losses otherwise associated with multiple bends. This provides for a lower pressure cooling air supply.

In rugged environments, certain superalloys may be selected for their resistance to particular corrosive attack. However, depending on the thermal properties of the superalloy, greater cooling may be beneficial. Without increasing the cooling air supply pressure, the described method of

manufacturing a cooled turbine blade provides for increasingly dense cooling fin arrays, as the fins may have a reduced cross section. In particular, the inner spar cuts the fin distance half, allowing for the thinner extremities, and thus a denser cooling fin array. Moreover, the shorter fin extrusion distance (i.e., from the inner spar to the skin rather than skin-to-skin) reduces challenges to casting in longer, narrow cavities. This is also complementary to forming the inner blade core with the inner blade pattern as shorter extrusions are used.

Although this invention has been shown and described with respect to detailed embodiments thereof, it will be understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the spirit and scope of the claimed invention. Accordingly, the preceding detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. In particular, the described embodiments are not limited to use in conjunction with a particular type of gas turbine engine. For example, the described embodiments may be applied to stationary or motive gas turbine engines, or any variant thereof.

Furthermore, there is no intention to be bound by any theory presented in any preceding section. It is also understood that the illustrations may include exaggerated dimensions and graphical representation to better illustrate the referenced items shown, and are not consider limiting unless expressly stated as such.

Although this invention has been shown and described with respect to detailed embodiments thereof, it will be understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the spirit and scope of the claimed invention. Accordingly, the preceding detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. In particular, the described embodiments are not limited to use in conjunction with a particular type of gas turbine engine. For example, the described embodiments may be applied to stationary or motive gas turbine engines, or any variant thereof.

Furthermore, there is no intention to be bound by any theory presented in any preceding section. It is also understood that the illustrations may include exaggerated dimensions and graphical representation to better illustrate the referenced items shown, and are not consider limiting unless expressly stated as such.

It will be understood that the benefits and advantages described above may relate to one embodiment or may relate to several embodiments. The embodiments are not limited to those that solve any or all of the stated problems or those that have any or all of the stated benefits and advantages.

Any reference to‘an’ item refers to one or more of those items. The term‘comprising’ is used herein to mean including the method blocks or elements identified, but that such blocks or elements do not comprise an exclusive list and a method or apparatus may contain additional blocks or elements.