US20140093386A1 - Cooled turbine blade with inner spar - Google Patents
Cooled turbine blade with inner spar Download PDFInfo
- Publication number
- US20140093386A1 US20140093386A1 US13/631,258 US201213631258A US2014093386A1 US 20140093386 A1 US20140093386 A1 US 20140093386A1 US 201213631258 A US201213631258 A US 201213631258A US 2014093386 A1 US2014093386 A1 US 2014093386A1
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- United States
- Prior art keywords
- inner spar
- skin
- airfoil
- turbine blade
- cooling fins
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/20—Specially-shaped blade tips to seal space between tips and stator
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
- F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
Definitions
- the present disclosure generally pertains to gas turbine engines, and is more particularly directed toward a cooled turbine blade.
- High performance gas turbine engines typically rely on increasing turbine inlet temperatures to increase both fuel economy and overall power ratings. These higher temperatures, if not compensated for, oxidize engine components and decrease component life. Component life has been increased by a number of techniques. Said techniques include internal cooling with air bled from an engine compressor section. Bleeding air results in efficiency loss however. In addition, stationary gas turbine engines typically may have less available compressed air than moving gas turbine engines.
- U.S. Pat. No. 7,690,894 issued to Liang on Apr. 6, 2010 shows a ceramic core assembly for a serpentine flow circuit in a turbine blade.
- the disclosure of Liang is directed toward a turbine blade for use in a gas turbine engine having an internal serpentine flow cooling circuit with pin fins and trip strips to promote heat transfer for obtaining a thermally balanced blade sectional temperature distribution.
- the turbine blade is cooled by a 7-pass serpentine flow cooling circuit that extends from the leading edge and along the pressure side wall of the airfoil, into the trailing edge and then flows along the suction side wall ending just downstream from the leading edge where the 7-pass serpentine flow circuit started.
- Leading edge film cooling holes are supplied from the first leg of the serpentine while a row of trailing edge exit holes is supplied from the third leg which extends across both walls of the airfoil in the trailing edge.
- the present disclosure is directed toward overcoming one or more of the problems discovered by the inventors.
- a cooled turbine blade having a base and an airfoil, the base including cooling air inlet and an internal cooling air passageway, and the airfoil including an internal heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil.
- the airfoil also includes a skin that encompasses a tip wall, an inner spar, a plurality of inner spar cooling fins extending from the inner spar to the skin, a plurality of trailing edge cooling fins extending from the pressure side of the skin to the lift side of the skin aft of the inner spar, and a leading edge chamber.
- a cooled turbine blade similar to the above but wherein the inner spar cooling fins are substantially parallel to each other along the mean camber line of the airfoil, is also disclosed herein.
- a cooled turbine blade similar to the above but wherein the cooling air follows a single bend heat exchange path, is also disclosed herein.
- 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 an exemplary turbine blade.
- FIG. 4 is a cutaway side view of an exemplary turbine blade.
- FIG. 5 is a sectional top view of an exemplary turbine blade, as taken along line 5 - 5 of FIG. 4 .
- 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.
- primary air i.e., air used in the combustion process
- 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 .
- 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 power 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 includes an exhaust diffuser 520 and an exhaust collector 550 .
- 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”).
- 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.
- a gas enters the inlet 110 as a “working fluid”, and is compressed by the compressor 200 .
- the working fluid is compressed in an annular flow path 115 by the series of compressor rotor assemblies 220 .
- the air 10 is compressed in numbered “stages”, the stages being associated with each compressor rotor assembly 220 .
- “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 ).
- each turbine rotor assembly 420 may be associated with a numbered stage.
- first stage turbine rotor assembly 421 is the forward most of the turbine rotor assemblies 420 .
- other numbering/naming conventions may also be used.
- 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 may be made from stainless steel and/or durable, high temperature materials known as “superalloys”.
- a superalloy, 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.
- 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 .
- 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 .
- the blade root 480 may incorporate “fir tree”, “bulb”, or “dove tail” roots, to list a few.
- 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 .
- the blade attachment grooves 432 may be configured to mate with the blade root 480 , both having a reciprocal shape with each other.
- the blade attachment grooves 432 may be slideably engaged with the blade attachment grooves 432 , for example, in a forward-to-aft direction.
- the first stage turbine rotor assembly 421 may incorporate active cooling.
- compressed cooling air may be internally supplied to each cooled turbine blade 440 as well as predetermined portions of the turbine rotor disk 430 .
- 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 .
- other stages of the turbine may incorporate active cooling as well.
- 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.
- each damper 426 may be configured to fit this under-platform cavity.
- the damper 426 may be omitted entirely.
- 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 .
- 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.
- the damper 426 may form at least a partial seal with the adjacent cooled turbine blades 440 .
- 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 an isometric view of the turbine blade of FIG. 2 .
- 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.
- 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.
- airfoil 474 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 .
- the inward direction is generally radially inward toward the center axis 95 ( FIG. 1 ), with its associated end called the “root end” 444 .
- the outward direction is generally radially outward from the center axis 95 ( FIG. 1 ), with its associated end called the “tip end” 445 .
- 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.
- 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).
- the inward and outward directions are generally measured in the radial direction relative to the center axis 95 ( FIG. 1 ).
- the inward and outward directions are generally measured in a plane perpendicular to a radial 96 of center axis 95 ( FIG. 1 ) with “inward” being toward the mean camber line 474 and “outward” being toward the “skin” 460 of the airfoil 441 .
- the airfoil 441 (along with the entire cooled turbine blade 440 ) may be made as a single metal casting, the outer surface 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 .
- the cooled turbine blade 440 of FIG. 3 is shown here with sections of 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.
- a section of the base 442 has been removed to expose portions of a cooling air passageway 482 , internal to the base 442 .
- 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.
- the cooling air inlet 481 is located in the blade root 480 .
- cooling air 15 may be received in a shank area radially outward from the blade root 480 but radially inward from the platform 443 .
- the cooled turbine blade 440 include 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 .
- the cooling air passageway 482 may be configured to translate the cooling air 15 in two dimensions (i.e., not merely in the plane of the figure) as it travels radially up (i.e., generally in the direction of a radial 96 of the center axis 95 ( FIG. 1 )) towards the airfoil 441 .
- 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 .
- the cooling air passageway 482 may be subdivided into a plurality of subpassages. As illustrated, the subdivisions may be evenly spaced, for example.
- airfoil 441 may include a tip wall 461 , an inner spar 462 , a leading edge chamber 463 , one or more section divider(s) 464 , one or more rib(s) 465 , one or more air deflector(s) 466 , and a plurality of inner spar cooling fins 467 .
- airfoil 441 may include a perforated trailing edge rib 468 and a plurality of trailing edge cooling fins 469 . Together with the skin 460 , these structures may form a single-bend heat exchange path 470 within the airfoil 441 .
- the internal structures making up the single-bend heat exchange path 470 may subdivide the single-bend heat exchange path 470 into multiple discrete sub-passageways or “sections”. For example, although single-bend heat exchange path 470 is shown by a representative path of cooling air 15 , three completely separated sections are illustrated (i.e., separated by section dividers 464 ) here on the pressure side 448 of cooled turbine blade 440 . Furthermore, in the particular embodiment illustrated, a total of six sub-passageways (including leading edge chamber 463 ) are identifiable.
- 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 .
- the tip end 445 is the tip wall 461 .
- 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 to the tip wall 461 , between the pressure side 448 ( FIG. 3 ) and the lift side 449 ( FIG. 3 ) of the skin 460 .
- 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 .
- the inner spar 462 may be configured 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 .
- the inner spar 462 may be solid (non-perforated) or substantially solid, such that cooling air 15 cannot pass.
- the inner spar 462 may extend less than the entire length of the mean camber line 474 .
- 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.
- the inner spar 462 may extend from the leading edge chamber 463 , downstream to the plurality of trailing edge cooling fins 469 .
- 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 .
- the inner spar 462 may have a thickness approximately that of other internal structures.
- the inner spar 462 may have a wall thickness plus or minus 20% that of the one or more section dividers 464 , one or more ribs 465 .
- the inner spar 462 may be kept with 1.2 times the wall thickness of the skin 460 .
- the inner spar 462 may include one or more inner spar pass-through hole(s) 473 .
- the inner spar 462 may include perforations such that pressure is equalized between the pressure side 448 ( FIG. 5 ) and the lift side 449 ( FIG. 5 ) of the inner spar 462 .
- an inner spar pass-through hole 473 may be made in each discrete sub-passageway or “section” of the single-bend heat exchange path 470 .
- a single section may include more than one inner spar pass-through hole(s) 473 .
- the inner spar pass-through hole(s) 473 may be located throughout the inner spar 462 .
- the inner spar 462 may include inner spar pass-through hole(s) 473 near the platform 443 , near the tip wall 461 , and/or near the single bend.
- each section divider 464 may extend from the base 442 to the trailing edge 447 , generally including a ninety degree turn and including a smooth transition.
- each section divider 464 may extend outward from the inner spar 462 to the skin 460 on each of the pressure side 448 ( FIG. 3 ) or the lift side 449 ( FIG. 3 ). Accordingly, cooling air 15 may be constrained within a sub-passageway or “section” of the single-bend heat exchange path 470 defined by the inner spar 462 , either of the pressure side 448 ( FIG. 3 ) or the lift side 449 ( FIG. 3 ) of the skin 460 , a section divider 464 , and one of: an adjacent section divider 464 , the tip wall 461 , and the base 442 .
- each section divider 464 on one side of inner spar 462 may run parallel with each other.
- a section divider 464 on the pressure side 448 ( FIG. 3 ) of the inner spar 462 may minor another section divider 464 on the lift side 449 ( FIG. 3 ) of the inner spar 462 .
- two “mirrored” section dividers 464 may merge into a single section divider 464 downstream of the inner spar 462 such that the “merged” section divider 464 extends from the pressure side 448 ( FIG. 3 ) of the skin 460 directly to the lift side 449 ( FIG. 3 ) of the skin 460 .
- each rib 465 may extend radially from the base 442 toward the tip end 445 , terminating prior to reaching the tip wall 461 .
- each rib 465 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 ) (i.e., in and out of plane).
- a rib 465 may also include a single bend at its distal end, relative to the base 442 . The single bend may be approximately ninety degrees and include a smooth transition.
- the rib 465 may run parallel with an adjacent structure (e.g., section divider 464 ).
- a rib 465 on the pressure side 448 ( FIG. 3 ) of the inner spar 462 may mirror another rib 465 on the lift side 449 ( FIG. 3 ) of the inner spar 462 .
- the airfoil 441 may include a leading edge rib 472 .
- the leading edge rib 472 may extend radially from the base 442 toward the tip end 445 , terminating prior to reaching the tip wall 461 .
- 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 .
- the leading edge chamber 463 may form part of the single-bend heat exchange path 470 .
- each air deflector 466 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 ).
- Each air deflector 466 may include a single bend, which is configured to redirect cooling air 15 approximately ninety degrees. Accordingly, the single bend may be approximately ninety degrees and include a smooth transition.
- the single bend of the air deflector 466 may start from a radial/vertical direction and smoothly transition to a horizontal direction aimed toward the trailing edge 447 .
- the single bend of the air deflector 466 may run parallel with the single bend of an adjacent section divider 464 or rib 465 .
- an air deflector 466 on the pressure side 448 ( FIG. 3 ) of the inner spar 462 may mirror another air deflector 466 on the lift side 449 ( FIG. 3 ) of the inner spar 462 .
- the airfoil 441 may include a leading edge air deflector 475 .
- the leading edge air deflector 475 may include a single bend, which is configured to redirect cooling air 15 approximately ninety degrees. Accordingly, the single bend may be approximately ninety degrees and include a smooth transition.
- the leading edge air deflector 475 may be located so as to redirect cooling air 15 leaving the leading edge chamber 463 .
- the leading edge air deflector 475 may be radially located between and the leading edge rib 472 and the tip wall 461 .
- the leading edge air deflector 475 may physically interact with the inner spar 462 .
- the leading edge air deflector 475 may extend from the pressure side 448 ( FIG.
- 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 ).
- 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 .
- 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 .
- the inner spar cooling fins 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.
- 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.
- the inner spar cooling fins 467 may be long and thin.
- inner spar cooling fins 467 traversing less than half the thickness of the airfoil 467 , may use a round “pin” fin.
- pin fins having a height-to-diameter ratio of 2-7 may be used.
- the inner spar cooling fins 467 may be pin fins having a diameter of 0.017-0.040 inches, and a length off the inner spar 467 of 0.034-0.240 inches.
- the inner spar cooling fins 467 may also be densely packed.
- inner spar cooling fins 467 may be within two diameters of each other.
- a greater number of inner spar cooling fins 467 may be used for increased cooling.
- 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 be in the range of 50 to 400 fins per square inch per side of the inner spar 462 .
- the fin density may be in the range of 40 to 70 fins per square inch per side of the inner spar 462 .
- the fin density may be in the range of 215 to 385 fins per square inch per side of the inner spar 462 .
- the fin density may be in the range of 70 to 385 fins per square inch per side of the inner spar 462 .
- the fin density may be in the range of 40 to 215 fins per square inch per side of the inner spar 462 .
- the trailing edge rib 468 may extend radially from the base 442 toward the tip end 445 .
- the trailing edge rib 468 may radially extend between the base 442 and the section divider 464 that defines the subdivision of the single-bend heat exchange path that exhausts nearest the platform 443 .
- 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 .
- the cooling air passageway 482 and the single-bend heat exchange path 470 may be coordinated.
- the cooling air passageway 482 may be sub-divided into a plurality of flow paths.
- the subdivided cooling air passageway 482 may be coordinated with the one or more section divider(s) 464 and the one or more rib(s) 465 above, in the airfoil 441 .
- each subdivision within the base 442 may be aligned with and include a cross sectional shape (not shown) corresponding to the areas bounded by the skin 460 and each section divider 464 and rib 465 .
- cooling air passageway 482 may maintain the same overall cross sectional area (i.e., constant flow rate and pressure) in each subdivision, as between the cooling air inlet 481 and the airfoil 441 .
- the cooling air passageway 482 may vary the cross sectional area of individual subdivisions where differing performance parameters are desired for each section, in a particular application.
- the cooling air passageway 482 and the single-bend heat exchange path 470 may each include asymmetric divisions for reflecting localized thermodynamic flow performance requirements.
- the cooled turbine blade 440 may have two or more sections divided by the one or more section divider(s) 464 . Accordingly, there will be a section on each side of the section divider 464 . As with the cooling air passageway 482 , each section may maintain the same overall cross sectional area. Alternately, each section divider 464 may be located such that each section varies where different performance parameters are desired for each section, in a particular application. For example, by moving the horizontal arm of section divider 464 radially outward, and a larger section is created on its inward side, and vis versa.
- the individual inner spar cooling fins 467 and the trailing edge cooling fins 469 may also include localized thermodynamic structural variations.
- 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 .
- 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.
- the individual cooling fins may be modified in shape, size, positioning, spacing, and grouping.
- 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.
- the pin fins or pedestals may be arranged as a staggered array, a linear array, or an irregular array.
- FIG. 5 is a sectional top view of the turbine blade of FIG. 4 , as taken along plane indicated by broken line 5 - 5 of FIG. 4 . From this view, inner spar 462 and the relationship with the above features and structures within the airfoil 441 are shown. For clarity, only the nearest row of internal structures within the airfoil 441 is shown. In addition, some of the cutaway internal structures are illustrated with alternating hatching for convenience and clarity, however, as discussed herein, in different embodiments they may be made from the same or different materials.
- airfoil 441 may have a varying profile in the radial direction.
- airfoil 441 may have a greater thickness near the platform 443 of base 442 than near the tip end 445 ( FIG. 3 ), as can be seen viewing both FIG. 3 (showing the airfoil 441 at the tip end 445 ) and FIG. 5 (showing the airfoil 441 closer to the base 442 ).
- the illustrated shape of the airfoil 441 is merely representative, and may vary from application to application.
- airfoil 441 may retain its aerodynamic features (i.e., leading edge 446 , trailing edge 447 , pressure side 448 , lift side 449 , and mean camber line 474 ) independent of its particular shape. Also, the illustrated thickness of the skin 460 and the structures residing within are also representative and not limiting.
- inner spar 462 may be located in between the pressure side 448 of the skin 460 and the lift side 449 the skin 460 .
- the inner spar 462 may substantially coincide with the mean camber line 474 of the airfoil 441 .
- inner spar 462 may bifurcate the single-bend heat exchange path 470 into a cavity associated with the pressure side 448 of the airfoil 441 and a cavity associated with the lift side 449 of the airfoil 441 .
- each section divider 464 and each rib 465 may further sub-divide the single-bend heat exchange path 470 .
- each section divider 464 and each rib 465 may extend outward from the inner spar 462 to the skin 460 on both the pressure side 448 and the lift side 449 , limiting cross flow within the single-bend heat exchange path 470 and subdividing the cavity on the pressure side 448 on the lift side 449 into a series of generally parallel cavities/flow passages.
- inner spar 462 may extend between the leading edge chamber 463 , at the leading edge rib 472 , and the trailing edge rib 468 .
- leading edge rib 472 and the trailing edge rib 468 may each extend from the pressure side 448 of the skin 460 directly to the lift side 449 of the skin 460 .
- the forward and aft ends of the inner spar 462 may be bound along the mean camber line 474 by the leading edge rib 472 and the trailing edge rib 468 , respectively.
- the origination of the inner spar 462 at the leading edge rib 472 provides for an increased cross section of the leading edge chamber 463 .
- the inner spar 462 may extend at least seventy-five percent the length of the mean camber line 474 .
- inner spar 462 may support the extension of the one or more section dividers 464 , the one or more ribs 465 , the one or more air deflectors 466 , and the plurality of inner spar cooling fins 467 .
- each structure/feature may extend from the inner spar 462 to the pressure side 448 or the lift side 449 of the airfoil 441 .
- each structure/feature may run parallel to each other.
- each structure/feature may be oriented perpendicular to the forward edge 484 (of aft edge 485 ) of the platform 443 , which may also be viewed as perpendicular to the center axis 95 ( FIG. 1 ).
- each structure/feature having a mirror structure/feature opposite the inner spar 462 may be equally treated or referred to as a single member or as two separate members.
- section dividers 464 on both sides of the inner spar 462 may equally be described as two separated members (i.e., as a first section divider 464 extending from the inner spar 462 to the lift side 449 of the skin 460 and a second section divider 464 extending from the inner spar 462 to the pressure side 449 of the skin 460 ) or as a single member that passes through or includes the corresponding section of the inner spar 462 (i.e., as a section divider 464 extending between the skin 460 on the lift side 449 and to the skin 460 on the pressure side 448 ).
- each structure/feature may include a “mirror image” on the opposite side of the inner spar 462 .
- each section divider 464 may extend to the trailing edge 447 , and two “mirrored” section dividers 464 may merge into a single section divider 464 downstream of the inner spar 462 such that the “merged” section divider 464 extends from the pressure side 448 of the skin 460 directly to the lift side 449 of the skin 460 .
- Both the inner spar cooling fins 467 and the trailing edge cooling fins 469 may be oriented for thermal performance, structural performance, and/or manufacturability.
- the plurality of inner spar cooling fins 467 may be oriented substantially parallel to each other and perpendicular to the center axis 95 .
- plurality of inner spar cooling fins 467 may populate at least ten percent of the volume of the single-bend heat exchange path 470 .
- the plurality of first inner spar cooling fins 467 may have a length at least twenty-five percent longer than the thickness of the inner spar 462 , as measured between the inner spar 462 and the pressure side 448 or the lift side 449 of the airfoil 441 .
- the structures/features toward the trailing edge 447 of the airfoil 441 may extend directly from the pressure side 448 to the lift side 449 of the skin 460 .
- both the trailing edge rib 468 and the plurality of trailing edge cooling fins 469 may extend skin-to-skin.
- the plurality of trailing edge cooling fins 469 may be oriented substantially parallel to each other.
- trailing edge cooling fins 469 may also be oriented so as to reduce the distance of the span between the pressure side 448 and the lift side 449 of the skin 460 .
- the plurality of trailing edge cooling fins 469 may be oriented substantially perpendicular to the mean camber line 474 .
- the plurality of trailing edge cooling fins 469 may be oriented substantially perpendicular to the skin 460 of the airfoil 441 as averaged between the pressure side 448 and the lift side 449 .
- 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.
- embodiments of the presently disclosed 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.
- 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.
- the entire cooled turbine blade may be cast formed.
- the cooled turbine blade 440 may be made from an investment casting process.
- the entire cooled turbine blade 440 may be cast from stainless steel and/or a superalloy using a ceramic core or fugitive pattern.
- the inclusion of the inner spar is amenable to the manufacturing process.
- 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.
- certain structures/features e.g., skin 460
- 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.
- the single bend provides for less turning losses, compared to serpentine configurations.
- the inner spar and copious cooling fin population provides for substantial heat exchange during the single pass.
- the inner spar itself may serve as a heat exchanger.
- 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 single-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 in a generally radial direction.
- the single-bend heat exchange path 470 is configured such that cooling air 15 will pass between, along, and around the various internal structures, but will generally flow in a ninety degree path as viewed from the side view (conceptually treating the camber sheet as a plane). Accordingly, the single-bend heat exchange path 470 may include some negligible lateral travel (i.e., into the plane) associated with the general curvature of the airfoil 441 .
- the single-bend heat exchange path 470 is illustrated by a single representative flow line traveling through a single section for clarity, the single-bend heat exchange path 470 includes the entire flow path carrying cooling air 15 through the airfoil 441 . Moreover, unlike other internally cooled turbine blades, the single-bend heat exchange path 470 is not serpentine, but rather has a single bend that efficiently redirects the cooling air 15 to the cooling air outlet 471 at the trailing edge 447 with a single turn.
Abstract
Description
- The present disclosure generally pertains to gas turbine engines, and is more particularly directed toward a cooled turbine blade.
- High performance gas turbine engines typically rely on increasing turbine inlet temperatures to increase both fuel economy and overall power ratings. These higher temperatures, if not compensated for, oxidize engine components and decrease component life. Component life has been increased by a number of techniques. Said techniques include internal cooling with air bled from an engine compressor section. Bleeding air results in efficiency loss however. In addition, stationary gas turbine engines typically may have less available compressed air than moving gas turbine engines.
- U.S. Pat. No. 7,690,894 issued to Liang on Apr. 6, 2010 shows a ceramic core assembly for a serpentine flow circuit in a turbine blade. In particular, the disclosure of Liang is directed toward a turbine blade for use in a gas turbine engine having an internal serpentine flow cooling circuit with pin fins and trip strips to promote heat transfer for obtaining a thermally balanced blade sectional temperature distribution. The turbine blade is cooled by a 7-pass serpentine flow cooling circuit that extends from the leading edge and along the pressure side wall of the airfoil, into the trailing edge and then flows along the suction side wall ending just downstream from the leading edge where the 7-pass serpentine flow circuit started. Leading edge film cooling holes are supplied from the first leg of the serpentine while a row of trailing edge exit holes is supplied from the third leg which extends across both walls of the airfoil in the trailing edge.
- The present disclosure is directed toward overcoming one or more of the problems discovered by the inventors.
- A cooled turbine blade is disclosed herein. The cooled turbine blade having a base and an airfoil, the base including cooling air inlet and an internal cooling air passageway, and the airfoil including an internal heat exchange path beginning at the base and ending at a cooling air outlet at the trailing edge of the airfoil. The airfoil also includes a skin that encompasses a tip wall, an inner spar, a plurality of inner spar cooling fins extending from the inner spar to the skin, a plurality of trailing edge cooling fins extending from the pressure side of the skin to the lift side of the skin aft of the inner spar, and a leading edge chamber. According to one embodiment, a cooled turbine blade, similar to the above but wherein the inner spar cooling fins are substantially parallel to each other along the mean camber line of the airfoil, is also disclosed herein. According to another embodiment, a cooled turbine blade, similar to the above but wherein the cooling air follows a single bend heat exchange path, is also disclosed herein.
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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 an exemplary turbine blade. -
FIG. 4 is a cutaway side view of an exemplary turbine blade. -
FIG. 5 is a sectional top view of an exemplary turbine blade, as taken along line 5-5 ofFIG. 4 . -
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). Thecenter 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 tocenter 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 fromcenter axis 95. - Structurally, a
gas turbine engine 100 includes aninlet 110, a gas producer or “compressor” 200, acombustor 300, aturbine 400, anexhaust 500, and apower output coupling 600. Thecompressor 200 includes one or morecompressor rotor assemblies 220. Thecombustor 300 includes one ormore injectors 350 and includes one ormore combustion chambers 390. Theturbine 400 includes one or moreturbine rotor assemblies 420. The exhaust includes anexhaust diffuser 520 and anexhaust collector 550. - As illustrated, both
compressor rotor assembly 220 andturbine 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 thecompressor 200. In thecompressor 200, the working fluid is compressed in anannular flow path 115 by the series ofcompressor rotor assemblies 220. In particular, theair 10 is compressed in numbered “stages”, the stages being associated with eachcompressor rotor assembly 220. For example, “4th stage air” may be associated with the 4thcompressor rotor assembly 220 in the downstream or “aft” direction—going from theinlet 110 towards the exhaust 500). Likewise, eachturbine rotor assembly 420 may be associated with a numbered stage. For example, first stageturbine rotor assembly 421 is the forward most of theturbine rotor assemblies 420. However, other numbering/naming conventions may also be used. - Once compressed
air 10 leaves thecompressor 200, it enters thecombustor 300, where it is diffused andfuel 20 is added.Air 10 andfuel 20 are injected into thecombustion chamber 390 viainjector 350 and ignited. After the combustion reaction, energy is then extracted from the combusted fuel/air mixture via theturbine 400 by each stage of the series ofturbine rotor assemblies 420.Exhaust gas 90 may then be diffused inexhaust diffuser 520 and collected, redirected, and exit the system via anexhaust 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 superalloy, 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.
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FIG. 2 is an axial view of an exemplary turbine rotor assembly. In particular, first stageturbine rotor assembly 421 schematically illustrated inFIG. 1 is shown here in greater detail, but in isolation from the rest ofgas turbine engine 100. First stageturbine rotor assembly 421 includes aturbine 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 ofdampers 426. Here, for illustration purposes,turbine rotor disk 430 is shown depopulated of all but three cooledturbine blades 440 and threedampers 426. - Each cooled
turbine blade 440 may include abase 442 including aplatform 443 and ablade root 480. For example, theblade root 480 may incorporate “fir tree”, “bulb”, or “dove tail” roots, to list a few. Correspondingly, theturbine rotor disk 430 may include a plurality of circumferentially distributed slots or “blade attachment grooves” 432 configured to receive and retain each cooledturbine blade 440. In particular, theblade attachment grooves 432 may be configured to mate with theblade root 480, both having a reciprocal shape with each other. In addition theblade attachment grooves 432 may be slideably engaged with theblade attachment grooves 432, for example, in a forward-to-aft direction. - Being proximate the combustor 300 (
FIG. 1 ), the first stageturbine rotor assembly 421 may incorporate active cooling. In particular, compressed cooling air may be internally supplied to each cooledturbine blade 440 as well as predetermined portions of theturbine rotor disk 430. For example, hereturbine rotor disk 430 engages the cooledturbine blade 440 such that a coolingair cavity 433 is formed between theblade attachment grooves 432 and theblade 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 adjacentblade attachment grooves 432 ofturbine rotor disk 430, an under-platform cavity may be formed above the circumferential outer edge ofturbine rotor disk 430, between shanks ofadjacent blade roots 480, and below theiradjacent platforms 443, respectively. As such, eachdamper 426 may be configured to fit this under-platform cavity. Alternately, where the platforms are flush with circumferential outer edge ofturbine rotor disk 430, and/or the under-platform cavity is sufficiently small, thedamper 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 stageturbine 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 theturbine rotor disk 430 during assembly of first stageturbine rotor assembly 421, for example, by a press fit. In addition, thedamper 426 may form at least a partial seal with the adjacent cooledturbine blades 440. Furthermore, one or more axial faces ofdamper 426 may be sized to provide sufficient clearance to permit each cooledturbine blade 440 to slide into theblade attachment grooves 432, past thedamper 426 without interference after installation of thedamper 426. -
FIG. 3 is an isometric view of the turbine blade ofFIG. 2 . As described above, the cooledturbine blade 440 may include a base 442 having aplatform 443 and ablade root 480. Each cooledturbine blade 440 may further include anairfoil 441 extending radially outward from theplatform 443. Theairfoil 441 may have a complex, geometry that varies radially. For example the cross section of theairfoil 441 may lengthen, thicken, twist, and/or change shape as it radially approaches theplatform 443 inward from thetip end 445. The overall shape ofairfoil 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 cooledturbine blade 440 is described with reference to both a radial 96 of center axis 95 (FIG. 1 ) and the aerodynamic features of theairfoil 441. The aerodynamic features of theairfoil 441 include aleading edge 446, a trailingedge 447, apressure side 448, alift side 449, and itsmean camber line 474. Themean camber line 474 is generally defined as the line running along the center of the airfoil from theleading edge 446 to the trailingedge 447. It can be thought of as the average of thepressure side 448 andlift side 449 of the airfoil shape. As discussed above,airfoil 474 also extends radially between theplatform 443 and thetip end 445. Accordingly, themean camber line 474 herein includes the entire camber sheet continuing from theplatform 443 to thetip end 445. - Accordingly, 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 theplatform 443, theforward edge 484 and theaft edge 485 of theplatform 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 themean camber line 474 as linear). When describing the flow features of theairfoil 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. 5)), the inward and outward directions are generally measured in a plane perpendicular to a radial 96 of center axis 95 (FIG. 1 ) with “inward” being toward themean camber line 474 and “outward” being toward the “skin” 460 of theairfoil 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 surface 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 ofFIG. 3 . In particular, the cooledturbine blade 440 ofFIG. 3 is shown here with sections of theskin 460 removed from thepressure side 448 of theairfoil 441, exposing its internal structure and cooling paths. For example, theairfoil 441 may include a composite flow path made up of multiple subdivisions and cooling structures. Similarly, a section of thebase 442 has been removed to expose portions of a coolingair passageway 482, internal to thebase 442. - As described above, the cooled
turbine blade 440 may include anairfoil 441 and abase 442. The base 442 may include theplatform 443, theblade root 480, and one or more cooling air inlet(s) 481. Theairfoil 441 interfaces with thebase 442 and may include theskin 460, atip wall 461, and the coolingair outlet 471. - Compressed secondary air may be routed into one or more cooling air inlet(s) 481 in the
base 442 of cooledturbine blade 440 as coolingair 15. The one or more cooling air inlet(s) 481 may be at any convenient location. For example, here the coolingair inlet 481 is located in theblade root 480. Alternately, coolingair 15 may be received in a shank area radially outward from theblade root 480 but radially inward from theplatform 443. - Within the
base 442, the cooledturbine blade 440 include the coolingair passageway 482 that is configured to route coolingair 15 from the one or more cooling air inlet(s) 481, through the base, and into theairfoil 441. The coolingair passageway 482 may be configured to translate the coolingair 15 in two dimensions (i.e., not merely in the plane of the figure) as it travels radially up (i.e., generally in the direction of a radial 96 of the center axis 95 (FIG. 1 )) towards theairfoil 441. Moreover, the coolingair passageway 482 may be structured to receive the coolingair 15 from a generally rectilinearcooling air inlet 481 and smoothly “reshape” it fit the curvature and shape of theairfoil 441. In addition, the coolingair passageway 482 may be subdivided into a plurality of subpassages. As illustrated, the subdivisions may be evenly spaced, for example. - Within the
skin 460 of theairfoil 441, several internal structures are viewable. In particular,airfoil 441 may include atip wall 461, aninner spar 462, a leadingedge chamber 463, one or more section divider(s) 464, one or more rib(s) 465, one or more air deflector(s) 466, and a plurality of innerspar cooling fins 467. In addition,airfoil 441 may include a perforatedtrailing edge rib 468 and a plurality of trailingedge cooling fins 469. Together with theskin 460, these structures may form a single-bendheat exchange path 470 within theairfoil 441. - The internal structures making up the single-bend
heat exchange path 470 may subdivide the single-bendheat exchange path 470 into multiple discrete sub-passageways or “sections”. For example, although single-bendheat exchange path 470 is shown by a representative path of coolingair 15, three completely separated sections are illustrated (i.e., separated by section dividers 464) here on thepressure side 448 of cooledturbine blade 440. Furthermore, in the particular embodiment illustrated, a total of six sub-passageways (including leading edge chamber 463) are identifiable. - With regard to the airfoil structures, the
tip wall 461 extends across theairfoil 441 and may be configured to redirect coolingair 15 from escaping through thetip end 445. In addition, one embodiment of thetip end 445 is thetip wall 461. Moreover,tip end 445 may be formed as a shared structure, such as a joining of thepressure side 448 and thelift side 449 of theairfoil 441. According to one embodiment, thetip wall 461 may be recessed inward such that it is not flush with the tip of theairfoil 441. According to one embodiment, thetip wall 461 may include one or more perforations (not shown) such that a small quantity of the coolingair 15 may be bled off for film cooling of thetip end 445. - The
inner spar 462 may extend from thebase 442 radially outward to thetip wall 461, between the pressure side 448 (FIG. 3 ) and the lift side 449 (FIG. 3 ) of theskin 460. In addition, theinner spar 462 may extend between theleading edge 446 and the trailingedge 447, parallel with, and generally following, the mean camber line 474 (FIG. 3 ) of theairfoil 441, and terminating with innerspar trailing edge 476. Accordingly, theinner spar 462 may be configured to bifurcate a portion or all of theairfoil 441 generally along its mean camber line 474 (FIG. 3 ) and between thepressure side 448 and thelift side 449. Also, theinner spar 462 may be solid (non-perforated) or substantially solid, such that coolingair 15 cannot pass. - According to one embodiment, the
inner spar 462 may extend less than the entire length of themean camber line 474. In particular theinner spar 462 may extend less than ninety percent of themean camber line 474 and may exclude theleading edge chamber 463 entirely. For example, theinner spar 462 may extend from the leadingedge chamber 463, downstream to the plurality of trailingedge cooling fins 469. In addition, theinner spar 462 may have a length within the range of seventy to eighty percent, or approximately three quarters the length of, and along, themean camber line 474. - According to one embodiment, the
inner spar 462 may have a thickness approximately that of other internal structures. In particular, theinner spar 462 may have a wall thickness plus or minus 20% that of the one ormore section dividers 464, one ormore ribs 465. In addition, theinner spar 462 may be kept with 1.2 times the wall thickness of theskin 460. - According to one embodiment, the
inner spar 462 may include one or more inner spar pass-through hole(s) 473. In particular, theinner spar 462 may include perforations such that pressure is equalized between the pressure side 448 (FIG. 5 ) and the lift side 449 (FIG. 5 ) of theinner spar 462. For example, an inner spar pass-throughhole 473 may be made in each discrete sub-passageway or “section” of the single-bendheat exchange path 470. In addition, depending on the pressure profile of the particular cooledturbine blade 440, a single section may include more than one inner spar pass-through hole(s) 473. Furthermore, the inner spar pass-through hole(s) 473 may be located throughout theinner spar 462. For example, and as illustrated, theinner spar 462 may include inner spar pass-through hole(s) 473 near theplatform 443, near thetip wall 461, and/or near the single bend. - Within the
airfoil 441, eachsection divider 464 may extend from the base 442 to the trailingedge 447, generally including a ninety degree turn and including a smooth transition. In addition, eachsection divider 464 may extend outward from theinner spar 462 to theskin 460 on each of the pressure side 448 (FIG. 3 ) or the lift side 449 (FIG. 3 ). Accordingly, coolingair 15 may be constrained within a sub-passageway or “section” of the single-bendheat exchange path 470 defined by theinner spar 462, either of the pressure side 448 (FIG. 3 ) or the lift side 449 (FIG. 3 ) of theskin 460, asection divider 464, and one of: anadjacent section divider 464, thetip wall 461, and thebase 442. - According to one embodiment, each
section divider 464 on one side ofinner spar 462 may run parallel with each other. According to another embodiment, asection divider 464 on the pressure side 448 (FIG. 3 ) of theinner spar 462 may minor anothersection divider 464 on the lift side 449 (FIG. 3 ) of theinner spar 462. Furthermore two “mirrored”section dividers 464 may merge into asingle section divider 464 downstream of theinner spar 462 such that the “merged”section divider 464 extends from the pressure side 448 (FIG. 3 ) of theskin 460 directly to the lift side 449 (FIG. 3 ) of theskin 460. - Within the
airfoil 441, eachrib 465 may extend radially from the base 442 toward thetip end 445, terminating prior to reaching thetip wall 461. In addition, eachrib 465 may extend outward from theinner spar 462 to theskin 460 on either of the pressure side 448 (FIG. 3) or the lift side 449 (FIG. 3 ) (i.e., in and out of plane). According to one embodiment, arib 465 may also include a single bend at its distal end, relative to thebase 442. The single bend may be approximately ninety degrees and include a smooth transition. In addition, therib 465 may run parallel with an adjacent structure (e.g., section divider 464). Furthermore, and as above, arib 465 on the pressure side 448 (FIG. 3 ) of theinner spar 462 may mirror anotherrib 465 on the lift side 449 (FIG. 3 ) of theinner spar 462. - According to one embodiment, the
airfoil 441 may include aleading edge rib 472. Theleading edge rib 472 may extend radially from the base 442 toward thetip end 445, terminating prior to reaching thetip wall 461. In addition, the leadingedge rib 472 may extend directly from the pressure side 448 (FIG. 3 ) of theskin 460 to the lift side 449 (FIG. 3 ) of theskin 460. In doing so, the leadingedge rib 472 may form theleading edge chamber 463 in conjunction with theskin 460 at theleading edge 446 of theairfoil 441. Additionally, all or part of the coolingair 15 leaving theleading edge chamber 463 may be redirected toward the trailingedge 447 bytip wall 461 andother cooling air 15 within theairfoil 441. Accordingly, the leadingedge chamber 463 may form part of the single-bendheat exchange path 470. - Within the
airfoil 441, eachair deflector 466 may extend outward from theinner spar 462 to theskin 460 on either of the pressure side 448 (FIG. 3 ) or the lift side 449 (FIG. 3 ). Eachair deflector 466 may include a single bend, which is configured to redirect coolingair 15 approximately ninety degrees. Accordingly, the single bend may be approximately ninety degrees and include a smooth transition. Generally, the single bend of theair deflector 466 may start from a radial/vertical direction and smoothly transition to a horizontal direction aimed toward the trailingedge 447. In addition, the single bend of theair deflector 466 may run parallel with the single bend of anadjacent section divider 464 orrib 465. Furthermore, and as above, anair deflector 466 on the pressure side 448 (FIG. 3 ) of theinner spar 462 may mirror anotherair deflector 466 on the lift side 449 (FIG. 3 ) of theinner spar 462. - According to one embodiment, the
airfoil 441 may include a leadingedge air deflector 475. As above, the leadingedge air deflector 475 may include a single bend, which is configured to redirect coolingair 15 approximately ninety degrees. Accordingly, the single bend may be approximately ninety degrees and include a smooth transition. The leadingedge air deflector 475 may be located so as to redirect coolingair 15 leaving theleading edge chamber 463. In particular, the leadingedge air deflector 475 may be radially located between and theleading edge rib 472 and thetip wall 461. Additionally, the leadingedge air deflector 475 may physically interact with theinner spar 462. In particular, the leadingedge air deflector 475 may extend from the pressure side 448 (FIG. 3 ) of theskin 460 to the lift side 449 (FIG. 3 ) of theskin 460, wherein at least a portion of the leadingedge air deflector 475 is intersected by theinner spar 462 between the pressure side 448 (FIG. 3 ) of theskin 460 and the lift side 449 (FIG. 3 ) of theskin 460. - Within the
airfoil 441, the plurality of innerspar cooling fins 467 may extend outward from theinner spar 462 to theskin 460 on either of the pressure side 448 (FIG. 3 ) or the lift side 449 (FIG. 3 ). In contrast, the plurality of trailingedge cooling fins 469 may extend from the pressure side 448 (FIG. 3 ) of theskin 460 directly to the lift side 449 (FIG. 3 ) of theskin 460. Accordingly, the plurality of innerspar cooling fins 467 are located forward of the plurality of trailingedge cooling fins 469, as measured along the mean camber line 474 (FIG. 3 ) of theairfoil 441. - Both the inner
spar cooling fins 467 and the trailingedge cooling fins 469 may be disbursed copiously throughout the single-bendheat exchange path 470. In particular, the innerspar cooling fins 467 and the trailingedge cooling fins 469 may be disbursed throughout theairfoil 441 so as to thermally interact with the coolingair 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 one embodiment, the inner
spar cooling fins 467 may be long and thin. In particular, innerspar cooling fins 467, traversing less than half the thickness of theairfoil 467, may use a round “pin” fin. Moreover, pin fins having a height-to-diameter ratio of 2-7 may be used. For example, the innerspar cooling fins 467 may be pin fins having a diameter of 0.017-0.040 inches, and a length off theinner spar 467 of 0.034-0.240 inches. - According to one embodiment, the inner
spar cooling fins 467 may also be densely packed. In particular, innerspar cooling fins 467 may be within two diameters of each other. Thus, a greater number of innerspar cooling fins 467 may be used for increased cooling. For example, across theinner spar 462, the fin density may be in the range of 80 to 300 fins per square inch per side of theinner spar 462. Alternately, the fin density may be in the range of 50 to 400 fins per square inch per side of theinner spar 462. Alternately, the fin density may be in the range of 40 to 70 fins per square inch per side of theinner spar 462. Alternately, the fin density may be in the range of 215 to 385 fins per square inch per side of theinner spar 462. Alternately, the fin density may be in the range of 70 to 385 fins per square inch per side of theinner spar 462. Alternately, the fin density may be in the range of 40 to 215 fins per square inch per side of theinner spar 462. - Within the
airfoil 441, the trailingedge rib 468 may extend radially from the base 442 toward thetip end 445. In particular, the trailingedge rib 468 may radially extend between the base 442 and thesection divider 464 that defines the subdivision of the single-bend heat exchange path that exhausts nearest theplatform 443. In addition, the trailingedge rib 468 may be located along the innerspar trailing edge 476 and between the innerspar cooling fins 467 and the trailingedge cooling fins 469. - Unlike a
section divider 464 or arib 465, the trailingedge rib 468 may be perforated to include one or more openings. This will allow coolingair 15 to pass through the trailingedge rib 468 toward the coolingair outlet 471 in the trailingedge 447, and thus complete the single-bendheat exchange path 470. - Taken as a whole the cooling
air passageway 482 and the single-bendheat exchange path 470 may be coordinated. In particular and returning to thebase 442 of the cooledturbine blade 440, the coolingair passageway 482 may be sub-divided into a plurality of flow paths. As illustrated, the subdivided coolingair passageway 482 may be coordinated with the one or more section divider(s) 464 and the one or more rib(s) 465 above, in theairfoil 441. Accordingly, each subdivision within thebase 442 may be aligned with and include a cross sectional shape (not shown) corresponding to the areas bounded by theskin 460 and eachsection divider 464 andrib 465. In addition, the coolingair passageway 482 may maintain the same overall cross sectional area (i.e., constant flow rate and pressure) in each subdivision, as between the coolingair inlet 481 and theairfoil 441. Alternately, the coolingair passageway 482 may vary the cross sectional area of individual subdivisions where differing performance parameters are desired for each section, in a particular application. - According to one embodiment, the cooling
air passageway 482 and the single-bendheat exchange path 470 may each include asymmetric divisions for reflecting localized thermodynamic flow performance requirements. In particular, as illustrated and discussed above, the cooledturbine blade 440 may have two or more sections divided by the one or more section divider(s) 464. Accordingly, there will be a section on each side of thesection divider 464. As with the coolingair passageway 482, each section may maintain the same overall cross sectional area. Alternately, eachsection divider 464 may be located such that each section varies where different performance parameters are desired for each section, in a particular application. For example, by moving the horizontal arm ofsection divider 464 radially outward, and a larger section is created on its inward side, and vis versa. - Similarly, according one embodiment, the individual inner
spar cooling fins 467 and the trailingedge cooling fins 469 may also include localized thermodynamic structural variations. In particular, the innerspar cooling fins 467 and/or the trailingedge cooling fins 469 may have different cross sections/surface area and/or fin spacing at different locations of theinner spar 462. For example, the cooledturbine 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 embodiment, one or more of the inner
spar cooling fins 467 and the trailingedge 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. -
FIG. 5 is a sectional top view of the turbine blade ofFIG. 4 , as taken along plane indicated by broken line 5-5 ofFIG. 4 . From this view,inner spar 462 and the relationship with the above features and structures within theairfoil 441 are shown. For clarity, only the nearest row of internal structures within theairfoil 441 is shown. In addition, some of the cutaway internal structures are illustrated with alternating hatching for convenience and clarity, however, as discussed herein, in different embodiments they may be made from the same or different materials. - As illustrated,
airfoil 441 may have a varying profile in the radial direction. In particular,airfoil 441 may have a greater thickness near theplatform 443 ofbase 442 than near the tip end 445 (FIG. 3 ), as can be seen viewing bothFIG. 3 (showing theairfoil 441 at the tip end 445) andFIG. 5 (showing theairfoil 441 closer to the base 442). The illustrated shape of theairfoil 441 is merely representative, and may vary from application to application. Moreover,airfoil 441 may retain its aerodynamic features (i.e., leadingedge 446, trailingedge 447,pressure side 448,lift side 449, and mean camber line 474) independent of its particular shape. Also, the illustrated thickness of theskin 460 and the structures residing within are also representative and not limiting. - As illustrated,
inner spar 462 may be located in between thepressure side 448 of theskin 460 and thelift side 449 theskin 460. In particular, theinner spar 462 may substantially coincide with themean camber line 474 of theairfoil 441. Accordingly,inner spar 462 may bifurcate the single-bendheat exchange path 470 into a cavity associated with thepressure side 448 of theairfoil 441 and a cavity associated with thelift side 449 of theairfoil 441. Moreover, eachsection divider 464 and eachrib 465 may further sub-divide the single-bendheat exchange path 470. In particular and as discussed above, eachsection divider 464 and eachrib 465 may extend outward from theinner spar 462 to theskin 460 on both thepressure side 448 and thelift side 449, limiting cross flow within the single-bendheat exchange path 470 and subdividing the cavity on thepressure side 448 on thelift side 449 into a series of generally parallel cavities/flow passages. - According to one embodiment,
inner spar 462 may extend between theleading edge chamber 463, at theleading edge rib 472, and the trailingedge rib 468. As above and as illustrated, leadingedge rib 472 and the trailingedge rib 468 may each extend from thepressure side 448 of theskin 460 directly to thelift side 449 of theskin 460. Accordingly, the forward and aft ends of theinner spar 462 may be bound along themean camber line 474 by theleading edge rib 472 and the trailingedge rib 468, respectively. Notably, the origination of theinner spar 462 at theleading edge rib 472 provides for an increased cross section of theleading edge chamber 463. Notwithstanding, according to one embodiment, theinner spar 462 may extend at least seventy-five percent the length of themean camber line 474. - As illustrated and discussed above,
inner spar 462 may support the extension of the one ormore section dividers 464, the one ormore ribs 465, the one ormore air deflectors 466, and the plurality of innerspar cooling fins 467. In particular, each structure/feature may extend from theinner spar 462 to thepressure side 448 or thelift side 449 of theairfoil 441. According to another embodiment, each structure/feature may run parallel to each other. Likewise, each structure/feature may be oriented perpendicular to the forward edge 484 (of aft edge 485) of theplatform 443, which may also be viewed as perpendicular to the center axis 95 (FIG. 1 ). - For convenience or clarity, and as the entire cooled
turbine blade 440 may be formed as a single casting, each structure/feature having a mirror structure/feature opposite theinner spar 462 may be equally treated or referred to as a single member or as two separate members. For example,section dividers 464 on both sides of theinner spar 462 may equally be described as two separated members (i.e., as afirst section divider 464 extending from theinner spar 462 to thelift side 449 of theskin 460 and asecond section divider 464 extending from theinner spar 462 to thepressure side 449 of the skin 460) or as a single member that passes through or includes the corresponding section of the inner spar 462 (i.e., as asection divider 464 extending between theskin 460 on thelift side 449 and to theskin 460 on the pressure side 448). - According to one embodiment and as illustrated each structure/feature may include a “mirror image” on the opposite side of the
inner spar 462. Notably, as the section cut is taken radially inward of the single bend of thesection dividers 464, only a portion is illustrated. As discussed above eachsection divider 464 may extend to the trailingedge 447, and two “mirrored”section dividers 464 may merge into asingle section divider 464 downstream of theinner spar 462 such that the “merged”section divider 464 extends from thepressure side 448 of theskin 460 directly to thelift side 449 of theskin 460. - Both the inner
spar cooling fins 467 and the trailingedge cooling fins 469 may be oriented for thermal performance, structural performance, and/or manufacturability. For example, the plurality of innerspar cooling fins 467 may be oriented substantially parallel to each other and perpendicular to thecenter axis 95. In addition, plurality of innerspar cooling fins 467 may populate at least ten percent of the volume of the single-bendheat exchange path 470. Also, the plurality of first innerspar cooling fins 467 may have a length at least twenty-five percent longer than the thickness of theinner spar 462, as measured between theinner spar 462 and thepressure side 448 or thelift side 449 of theairfoil 441. - With regard to the structures/features toward the trailing
edge 447 of theairfoil 441, having a narrower thickness, the structures/features may extend directly from thepressure side 448 to thelift side 449 of theskin 460. In particular, both the trailingedge rib 468 and the plurality of trailingedge cooling fins 469 may extend skin-to-skin. Like the innerspar cooling fins 467, the plurality of trailingedge cooling fins 469 may be oriented substantially parallel to each other. However, trailingedge cooling fins 469 may also be oriented so as to reduce the distance of the span between thepressure side 448 and thelift side 449 of theskin 460. For example, the plurality of trailingedge cooling fins 469 may be oriented substantially perpendicular to themean camber line 474. Alternately, the plurality of trailingedge cooling fins 469 may be oriented substantially perpendicular to theskin 460 of theairfoil 441 as averaged between thepressure side 448 and thelift side 449. - 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, embodiments of the presently disclosed 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 cooledturbine 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 single-bend
heat exchange path 470 begins at the base 442 wherepressurized cooling air 15 is received into theairfoil 441. The coolingair 15 is received from the coolingair passageway 482 in a generally radial direction. The single-bendheat exchange path 470 is configured such that coolingair 15 will pass between, along, and around the various internal structures, but will generally flow in a ninety degree path as viewed from the side view (conceptually treating the camber sheet as a plane). Accordingly, the single-bendheat exchange path 470 may include some negligible lateral travel (i.e., into the plane) associated with the general curvature of theairfoil 441. Also, as discussed above, although the single-bendheat exchange path 470 is illustrated by a single representative flow line traveling through a single section for clarity, the single-bendheat exchange path 470 includes the entire flow path carrying coolingair 15 through theairfoil 441. Moreover, unlike other internally cooled turbine blades, the single-bendheat exchange path 470 is not serpentine, but rather has a single bend that efficiently redirects the coolingair 15 to the coolingair outlet 471 at the trailingedge 447 with a single turn. - 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.
Claims (20)
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US13/631,258 US20140093386A1 (en) | 2012-09-28 | 2012-09-28 | Cooled turbine blade with inner spar |
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US13/631,258 US20140093386A1 (en) | 2012-09-28 | 2012-09-28 | Cooled turbine blade with inner spar |
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US20140093386A1 true US20140093386A1 (en) | 2014-04-03 |
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US13/631,258 Abandoned US20140093386A1 (en) | 2012-09-28 | 2012-09-28 | Cooled turbine blade with inner spar |
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Cited By (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20170002662A1 (en) * | 2015-07-01 | 2017-01-05 | United Technologies Corporation | Gas turbine engine airfoil with bi-axial skin core |
FR3041989A1 (en) * | 2015-10-06 | 2017-04-07 | Snecma | DAWN COMPRISING A LEAK EDGE COMPRISING THREE SEPARATE COOLING REGIONS |
US20170306765A1 (en) * | 2016-04-25 | 2017-10-26 | General Electric Company | Airfoil with variable slot decoupling |
WO2019118110A1 (en) * | 2017-12-13 | 2019-06-20 | Solar Turbines Incorporated | Improved turbine blade cooling system |
CN111927563A (en) * | 2020-07-31 | 2020-11-13 | 中国航发贵阳发动机设计研究所 | Turbine blade suitable for high temperature environment |
WO2021019170A1 (en) * | 2019-08-01 | 2021-02-04 | Safran Aircraft Engines | Blade provided with a cooling circuit |
US10920610B2 (en) * | 2018-06-11 | 2021-02-16 | Raytheon Technologies Corporation | Casting plug with flow control features |
CN114215609A (en) * | 2021-12-30 | 2022-03-22 | 华中科技大学 | Blade inner cooling channel capable of enhancing cooling and application thereof |
US11339718B2 (en) * | 2018-11-09 | 2022-05-24 | Raytheon Technologies Corporation | Minicore cooling passage network having trip strips |
US11408289B2 (en) * | 2019-04-04 | 2022-08-09 | MAN Energy Solution SE | Moving blade of a turbo machine |
US20230250725A1 (en) * | 2021-07-02 | 2023-08-10 | Raytheon Technologies Corporation | Cooling arrangement for gas turbine engine component |
Citations (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3017159A (en) * | 1956-11-23 | 1962-01-16 | Curtiss Wright Corp | Hollow blade construction |
US5601399A (en) * | 1996-05-08 | 1997-02-11 | Alliedsignal Inc. | Internally cooled gas turbine vane |
US5772397A (en) * | 1996-05-08 | 1998-06-30 | Alliedsignal Inc. | Gas turbine airfoil with aft internal cooling |
US7210906B2 (en) * | 2004-08-10 | 2007-05-01 | Pratt & Whitney Canada Corp. | Internally cooled gas turbine airfoil and method |
US7530789B1 (en) * | 2006-11-16 | 2009-05-12 | Florida Turbine Technologies, Inc. | Turbine blade with a serpentine flow and impingement cooling circuit |
US20100226790A1 (en) * | 2009-03-04 | 2010-09-09 | Siemens Energy, Inc. | Turbine blade leading edge tip cooling system |
US8047789B1 (en) * | 2007-10-19 | 2011-11-01 | Florida Turbine Technologies, Inc. | Turbine airfoil |
US20140093388A1 (en) * | 2012-09-28 | 2014-04-03 | Solar Turbines Incorporated | Cooled turbine blade with leading edge flow deflection and division |
US20140093391A1 (en) * | 2012-09-28 | 2014-04-03 | Solar Turbines Incorporated | Cooled turbine blade with trailing edge flow metering |
US20140093390A1 (en) * | 2012-09-28 | 2014-04-03 | Solar Turbines Incorporated | Cooled turbine blade with leading edge flow redirection and diffusion |
-
2012
- 2012-09-28 US US13/631,258 patent/US20140093386A1/en not_active Abandoned
Patent Citations (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3017159A (en) * | 1956-11-23 | 1962-01-16 | Curtiss Wright Corp | Hollow blade construction |
US5601399A (en) * | 1996-05-08 | 1997-02-11 | Alliedsignal Inc. | Internally cooled gas turbine vane |
US5772397A (en) * | 1996-05-08 | 1998-06-30 | Alliedsignal Inc. | Gas turbine airfoil with aft internal cooling |
US7210906B2 (en) * | 2004-08-10 | 2007-05-01 | Pratt & Whitney Canada Corp. | Internally cooled gas turbine airfoil and method |
US7530789B1 (en) * | 2006-11-16 | 2009-05-12 | Florida Turbine Technologies, Inc. | Turbine blade with a serpentine flow and impingement cooling circuit |
US8047789B1 (en) * | 2007-10-19 | 2011-11-01 | Florida Turbine Technologies, Inc. | Turbine airfoil |
US20100226790A1 (en) * | 2009-03-04 | 2010-09-09 | Siemens Energy, Inc. | Turbine blade leading edge tip cooling system |
US20140093388A1 (en) * | 2012-09-28 | 2014-04-03 | Solar Turbines Incorporated | Cooled turbine blade with leading edge flow deflection and division |
US20140093391A1 (en) * | 2012-09-28 | 2014-04-03 | Solar Turbines Incorporated | Cooled turbine blade with trailing edge flow metering |
US20140093390A1 (en) * | 2012-09-28 | 2014-04-03 | Solar Turbines Incorporated | Cooled turbine blade with leading edge flow redirection and diffusion |
Cited By (28)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20170002662A1 (en) * | 2015-07-01 | 2017-01-05 | United Technologies Corporation | Gas turbine engine airfoil with bi-axial skin core |
US10767491B2 (en) | 2015-10-06 | 2020-09-08 | Safran Aircraft Engines | Blade comprising a trailing edge having three distinct cooling regions |
WO2017060613A1 (en) * | 2015-10-06 | 2017-04-13 | Safran Aircraft Engines | Blade comprising a trailing edge having three distinct cooling regions |
GB2558113B (en) * | 2015-10-06 | 2021-05-12 | Safran Aircraft Engines | Blade comprising a trailing edge having three distinct cooling regions |
FR3041989A1 (en) * | 2015-10-06 | 2017-04-07 | Snecma | DAWN COMPRISING A LEAK EDGE COMPRISING THREE SEPARATE COOLING REGIONS |
GB2558113A (en) * | 2015-10-06 | 2018-07-04 | Safran Aircraft Engines | Blade comprising a trailing edge having three distinct cooling regions |
US10156146B2 (en) * | 2016-04-25 | 2018-12-18 | General Electric Company | Airfoil with variable slot decoupling |
EP3260661A1 (en) * | 2016-04-25 | 2017-12-27 | General Electric Company | Airfoil with variable slot decoupling |
US20170306765A1 (en) * | 2016-04-25 | 2017-10-26 | General Electric Company | Airfoil with variable slot decoupling |
CN114961877A (en) * | 2017-12-13 | 2022-08-30 | 索拉透平公司 | Improved turbine bucket cooling system |
US11002138B2 (en) * | 2017-12-13 | 2021-05-11 | Solar Turbines Incorporated | Turbine blade cooling system with lower turning vane bank |
US10718219B2 (en) | 2017-12-13 | 2020-07-21 | Solar Turbines Incorporated | Turbine blade cooling system with tip diffuser |
US10830059B2 (en) | 2017-12-13 | 2020-11-10 | Solar Turbines Incorporated | Turbine blade cooling system with tip flag transition |
WO2019118110A1 (en) * | 2017-12-13 | 2019-06-20 | Solar Turbines Incorporated | Improved turbine blade cooling system |
CN114961879A (en) * | 2017-12-13 | 2022-08-30 | 索拉透平公司 | Improved turbine bucket cooling system |
US10920597B2 (en) | 2017-12-13 | 2021-02-16 | Solar Turbines Incorporated | Turbine blade cooling system with channel transition |
US10815791B2 (en) | 2017-12-13 | 2020-10-27 | Solar Turbines Incorporated | Turbine blade cooling system with upper turning vane bank |
CN111465751A (en) * | 2017-12-13 | 2020-07-28 | 索拉透平公司 | Improved turbine bucket cooling system |
US10920610B2 (en) * | 2018-06-11 | 2021-02-16 | Raytheon Technologies Corporation | Casting plug with flow control features |
US11339718B2 (en) * | 2018-11-09 | 2022-05-24 | Raytheon Technologies Corporation | Minicore cooling passage network having trip strips |
US11408289B2 (en) * | 2019-04-04 | 2022-08-09 | MAN Energy Solution SE | Moving blade of a turbo machine |
WO2021019170A1 (en) * | 2019-08-01 | 2021-02-04 | Safran Aircraft Engines | Blade provided with a cooling circuit |
CN114245841A (en) * | 2019-08-01 | 2022-03-25 | 赛峰航空器发动机 | Blade provided with cooling circuit |
FR3099523A1 (en) * | 2019-08-01 | 2021-02-05 | Safran Aircraft Engines | Blade fitted with a cooling circuit |
US11719102B2 (en) | 2019-08-01 | 2023-08-08 | Safran Aircraft Engines | Blade provided with a cooling circuit |
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US20230250725A1 (en) * | 2021-07-02 | 2023-08-10 | Raytheon Technologies Corporation | Cooling arrangement for gas turbine engine component |
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