WO2016133513A1 - Aube de turbine dotée d'une paroi interne segmentée - Google Patents

Aube de turbine dotée d'une paroi interne segmentée Download PDF

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Publication number
WO2016133513A1
WO2016133513A1 PCT/US2015/016482 US2015016482W WO2016133513A1 WO 2016133513 A1 WO2016133513 A1 WO 2016133513A1 US 2015016482 W US2015016482 W US 2015016482W WO 2016133513 A1 WO2016133513 A1 WO 2016133513A1
Authority
WO
WIPO (PCT)
Prior art keywords
airfoil
wall
turbine
internal wall
connected segments
Prior art date
Application number
PCT/US2015/016482
Other languages
English (en)
Inventor
Evan C. LANDRUM
Jan H. Marsh
Paul A. SANDERS
Original Assignee
Siemens Energy, Inc.
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Siemens Energy, Inc. filed Critical Siemens Energy, Inc.
Priority to PCT/US2015/016482 priority Critical patent/WO2016133513A1/fr
Publication of WO2016133513A1 publication Critical patent/WO2016133513A1/fr

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Classifications

    • 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
    • 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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall

Definitions

  • This invention is directed generally to turbine airfoils, and more particularly to cooling systems in hollow turbine airfoils.
  • gas turbine engines typically include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power.
  • Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit.
  • Typical turbine combustor configurations expose turbine airfoils (such as turbine vanes and blades) to high temperatures.
  • turbine airfoils must be made of materials capable of withstanding such high temperatures, or must include cooling features to enable the component to survive in an environment which exceeds the capability of the material.
  • typical airfoils may include internal cooling systems for reducing the temperature of the airfoils. Such typical internal cooling systems, however, may be deficient.
  • An airfoil for a gas turbine engine in which the airfoil includes an internal cooling system formed in part by a segmented internal wall that distributes mechanical loads is disclosed.
  • the segmented internal wall may include a plurality of connected segments.
  • at least a portion of the plurality of connected segments may be non-parallel to a camber line of the generally elongated hollow airfoil.
  • all of the plurality of connected segments may be non-parallel to the camber line of the generally elongated hollow airfoil.
  • the plurality of connected segments may be angled in an alternating fashion with respect to the camber line of the generally elongated hollow airfoil.
  • a turbine airfoil may include a generally elongated hollow airfoil formed from an outer wall, and may have a leading edge, a trailing edge, a pressure side, and a suction side.
  • the turbine airfoil may further include a cooling system positioned within an interior portion of the generally elongated hollow airfoil.
  • the cooling system may be formed by the outer wall, a segmented internal wall, and a plurality of ribs positioned in-between the outer wall and the segmented internal wall. Cooling fluids may flow through the cooling system.
  • the turbine airfoil may be a turbine blade, or a turbine vane.
  • the segmented internal wall may include a plurality of connected segments. At least a portion of the plurality of connected segments may be non-parallel to a camber line of the generally elongated hollow airfoil. Furthermore, all of the plurality of connected segments may be non-parallel to the camber line of the generally elongated hollow airfoil. The plurality of connected segments may be angled in an alternating fashion with respect to the camber line of the generally elongated hollow airfoil. Additionally, each two successive connected segments of the plurality of connected segments may be connected at an angle between 10 degrees and 170 degrees.
  • the cooling system may include an aft flowing serpentine cooling channel and further include a forward flowing serpentine cooling channel.
  • the aft flowing serpentine cooling channel may be a five-pass aft flowing serpentine cooling channel.
  • the aft flowing serpentine cooling channel may be an aft flowing spiral serpentine cooling channel, and the forward flowing serpentine cooling channel may be a forward flowing spiral serpentine cooling channel.
  • the plurality of ribs may connect the segmented internal wall to the outer wall. Additionally, at least a portion of the plurality of ribs may connect the segmented internal wall to either the pressure side of the outer wall or the suction side of the outer wall in an alternating fashion.
  • an advantage of the internal cooling system formed by the segmented internal wall is that the segmented internal wall may more easily deform relative to the outer wall. As such, the segmented internal wall may mitigate thermal/mechanical stresses caused by the outer wall being heated to a higher temperature than inner portions of the airfoil.
  • a further advantage of the internal cooling system formed by the segmented internal wall is that the segmented internal wall may reduce the cross-sectional area of the cooling cavities, and may also increase the length of the cooling cavities, in one embodiment. As such, the cooling fluids may more efficiently cool the airfoil, even for low flow designs.
  • Figure 1 is a perspective view of a turbine airfoil according to one
  • Figure 2 is a cross-sectional view of an example of the turbine airfoil shown in Figure 1 taken along section line 2-2.
  • Figure 3 is a cross-sectional view of another example of the turbine airfoil shown in Figure 1 taken along section line 2-2.
  • an airfoil 12 for a gas turbine engine in which the airfoil 12 includes an internal cooling system 10 formed by an outer wall 38, a segmented internal wall 54, and a plurality of ribs 58 positioned in-between the outer wall 38 and the segmented internal wall 54, is disclosed.
  • the segmented internal wall 54 may include a plurality of connected segments 62.
  • at least a portion of the plurality of connected segments 62 may be non-parallel to a camber line 66 of the generally elongated hollow airfoil 32.
  • all of the plurality of connected segments 62 may be non-parallel to the camber line 66 of the generally elongated hollow airfoil 32.
  • the plurality of connected segments 62 may be angled in an alternating fashion with respect to the camber line 66 of the generally elongated hollow airfoil 32.
  • a turbine airfoil (such as a turbine blade or a turbine vane) may be cooled by an internal cooling system that allows cooling fluids to pass through the turbine airfoil, cooling the outer wall of the turbine airfoil.
  • Such internal cooling systems may be typically formed by a continuous internal wall that is parallel to a camber line of the turbine airfoil.
  • this typical internal wall may be deficient because it may be subject to high stresses which may cause it to fail.
  • a typical internal wall may have a lower temperature than the outer wall of the airfoil. This temperature difference may cause the outer wall to deform at a different rate than the typical internal wall, which may subject the internal wall to high stress, and possible failure.
  • airfoil 12 of Figures 1 -3 may include an internal cooling system 10 that provides various advantages.
  • an advantage of the internal cooling system 10 formed by the segmented internal wall 54 is that the segmented internal wall 54 may more easily deform relative to the outer wall 38. As such, the segmented internal wall 54 may mitigate thermal/mechanical stresses caused by the outer wall 32 being heated to a higher temperature than inner portions of the airfoil 12.
  • a further advantage of the internal cooling system 10 formed by the segmented internal wall 54 is that the segmented internal wall 54 may reduce the cross-sectional area of the cooling cavities, and may also increase the length of the cooling cavities. As such, the cooling fluids may more efficiently cool the airfoil 12, even for low flow designs.
  • FIG. 1 illustrates a turbine airfoil 12 according to one embodiment.
  • the airfoil 12 may be any type of turbine airfoil.
  • the airfoil 12 may be a turbine blade, a turbine vane, or any other turbine airfoil.
  • the airfoil 12 is a turbine blade for a gas turbine engine.
  • the airfoil 12 may include a generally elongated hollow airfoil 32 formed from an outer wall 38 adapted for use, for example, in a first stage of an axial flow turbine engine.
  • the generally elongated hollow airfoil 32 may have a leading edge 46, a trailing edge 48, a generally concave shaped portion forming pressure side 40, and a generally convex shaped portion forming suction side 42. As illustrated, the generally elongated hollow airfoil 32 may be coupled to a root 34 at a platform 36.
  • the root 34 may couple the airfoil 12 (such as a turbine blade) to a disc (not shown) of the turbine engine.
  • the airfoil 12 such as a turbine blade
  • the airfoil 12 may be a turbine vane with a first end coupled to the inner diameter of the turbine section of the turbine engine and a second end coupled to the outer diameter of the turbine section of the turbine engine.
  • a cavity 14 may be positioned in an inner portion of the airfoil 12 for directing one or more gases, which may include air received from a compressor (not shown), through the airfoil 12 and out one or more exhaust orifices 44 in the airfoil 12 to reduce the temperature of the airfoil 12.
  • the exhaust orifices 44 may be positioned in the leading edge 46, the trailing edge 48, the tip 50 in close proximity to the leading and trailing edges 46, 48, or any combination thereof, and have various configurations.
  • the leading edge 46 may include a plurality of orifices 44 that collectively form a showerhead for cooling the leading edge 46 of the airfoil 12.
  • Figure 1 illustrates orifices 44 as positioned at particular locations, the orifices 44 may be positioned anywhere along pressure side 40 and/or suction side 42.
  • the cavity 14 may be arranged in various configurations and is not limited to a particular flow path.
  • a cooling system 10 may be located in cavity 14, as shown in Figures 2-3. As such, the cooling system 10 may be positioned within an interior portion of the generally elongated hollow airfoil 32.
  • the cooling system 10 may be formed by a combination of the outer wall 38, a segmented internal wall 54, and a plurality of ribs 58 positioned in-between the outer wall 38 and the segmented internal wall 54.
  • the segmented internal wall 54 may include a plurality of connected segments 62.
  • the segmented internal wall 54 may have any number of connected segments 62.
  • the segmented internal wall 54 may have two connected segments 62, three connected segments 62, four connected segments 62, six connected segments 62, eight connected segments 62, or any other number of connected segments 62.
  • a connected segment 62 may have any shape.
  • a connected segment 62 may be a straight segment, a curved segment, an irregular segment, any other shaped segment, or any combination of the preceding.
  • a connected segment 62 may have any length and/or any width.
  • a portion of the connected segments 62 may be non- parallel to a camber line 66 halfway between the upper and lower surfaces of the generally elongated hollow airfoil 32.
  • one connected segment 62 may be non-parallel to the camber line 66
  • two connected segments 62 may be non- parallel to the camber line 66
  • three connected segments 62 may be non-parallel to the camber line 66
  • four connected segments 62 may be non-parallel to the camber line 66
  • any other number of connected segments 62 may be non-parallel to the camber line 66.
  • all of the connected segments 62 may be non- parallel to the camber line 66.
  • half of the connected segments 62 (or any other portion of the connected segments 62, such as more than half or less than half) may be non-parallel to the camber line 66.
  • a connected segment 62 may be non-parallel to the camber line 66 as a result of having an angle 68 with respect to the camber line 66.
  • Angle 68 may be any angle.
  • angle 68 may be 5 - 85 degrees.
  • the angle 68 of the connected segments 62 with respect to the camber line 66 may cause two successive connected segments 62 to be connected at an angle 70.
  • Angle 70 may be any angle.
  • angle 70 may be 10 - 170 degrees.
  • the connected segments 62 may be angled in an alternating fashion with respect to the camber line 66.
  • a first connected segment 62 may be angled with respect to the camber line 66 so as to cause the first connected segment 62 to extend towards the pressure side 40 of the outer wall 38
  • the next connected segment 62 may be angled with respect to the camber line 66 so as to cause the next connected segment 62 to extend towards the suction side 42 of the outer wall 38.
  • this may cause a zig-zag type of shape for the segmented internal wall 54.
  • this alternating fashion may continue along the entire length of the segmented internal wall 54, or only a portion of the segmented internal wall 54.
  • the segmented internal wall 54 may include one or more connected segments 62 that are not angled in an alternating fashion with respect to the camber line 66.
  • a first connected segment 62 may be angled with respect to the camber line 66 so as to cause the first connected segment 62 to extend towards the pressure side 40 of the outer wall 38
  • a next connected segment 62 (or more than one next connected segment 62) may also be angled so as to cause the next connected segment(s) 62 to extend towards the pressure side 40 of the outer wall 38 (or may be angled parallel to the camber line 66)
  • the subsequent connected segment may be angled with respect to the camber line 66 so as to cause the subsequent connected segment 62 to extend towards the suction side 42 of the outer wall 38.
  • the cooling system 10 may be formed by a
  • a rib 58 may refer to a segment (or other element) that connects the segmented internal wall 54 to the outer wall 38.
  • the ribs 58 may connect the segmented internal wall 54 to the outer wall 38.
  • the airfoil 12 may include any number of ribs 58.
  • the airfoil 12 may include four ribs 58, five ribs 58, six ribs 58, eight ribs 58, nine ribs 58, ten ribs 58, or any other number of ribs 58.
  • a rib 58 may have any shape.
  • a rib 58 may be a straight segment, a curved segment, an irregular segment, any other shaped segment, or any combination of the preceding.
  • a rib 58 may have any length and/or any width.
  • a rib 58 may be made of the same material as the segmented internal wall 54.
  • the ribs 58 may connect the segmented internal wall 54 to the outer wall 38 in any manner.
  • a rib 58 may connect the segmented internal wall 54 to both the pressure side 40 of the outer wall 38 and the suction side 42 of the outer wall 38.
  • a rib 58 may connect the segmented internal wall 54 to either the pressure side 40 of the outer wall 38 or the suction side 42 of the outer wall 38. In such an example, such a connection may be done in an alternating fashion by at least a portion of the ribs 58, as is illustrated in Figure 2.
  • the portion of the ribs 58 that connect the segmented internal wall 54 to either the pressure side 40 of the outer wall 38 or the suction side 42 of the outer wall 38 in an alternating fashion may include any number of ribs 58, such as two ribs 58, three ribs 58, four ribs 58, five ribs 58, six ribs 58, or any other number of ribs 58.
  • the portion of the ribs 58 that connect the segmented internal wall 54 to either the pressure side 40 of the outer wall 38 or the suction side 42 of the outer wall 38 in an alternating fashion may include all of the ribs 58.
  • the portion of the ribs 58 that connect the segmented internal wall 54 to either the pressure side 40 of the outer wall 38 or the suction side 42 of the outer wall 38 in an alternating fashion may include half of the ribs 58 (or any other portion of the ribs 58, such as more than half or less than half).
  • the ribs 58 may be connected to the connected segments 62.
  • the ribs 58 may be connected to any location on the connected segments 62.
  • the ribs 58 may be connected to the end points of the connected segments 62, as is illustrated in Figure 2.
  • the ribs 58 may be connected to middle points (or any other points) of the connected segments 62.
  • the ribs 58 may be connected to inflection points of curved connected segments 62.
  • the ribs 58 may be connected to the connected segments 62 at an angle 76.
  • Angle 76 may be any angle.
  • angle 76 may be 30 - 270 degrees.
  • the ribs 58 may be connected to the outer wall 38 at an angle 78.
  • Angle 78 may be any angle.
  • angle 78 may be 30 - 1 50 degrees.
  • Cooling system 10 may include one or more cooling channels that allow cooling fluids to flow through the cooling system 10 in order to cool the airfoil 12.
  • the cooling channels may allow the cooling fluids to flow in the forward direction (i.e., a direction from the trailing edge 48 towards the leading edge 46), in the aft direction (i.e., a direction from the leading edge 46 towards the trailing edge 48), or both the forward direction and the aft direction.
  • the one or more cooling channels may allow the cooling fluids to flow through all or a portion of the cooling system 10 in order to cool the airfoil 12.
  • cooling system 10 may include an aft flowing serpentine cooling channel 72 and a forward flowing serpentine channel 74.
  • the aft flowing serpentine cooling channel 72 may extend from a position proximate the root 34 to the tip 50 of an airfoil 12 that is a turbine blade (or, for a turbine vane, from a position proximate the inner diameter of the turbine section to a position proximate the outer diameter of the turbine section).
  • the aft flowing serpentine cooling channel 72 may be formed from at least a two pass serpentine cooling channel, and, in particular embodiments, may be a five-pass serpentine cooling channel or greater.
  • the aft flowing serpentine cooling channel 72 may be an aft flowing spiral serpentine cooling channel having aft flowing cavities 82, 84, 86, 88, and 90 in contact with at least one of the pressure side 40 and the suction side 42.
  • Aft flowing cavity 82 may be in communication with a first cooling fluid supply inlet (not shown), and may be configured to pass the cooling fluids through the aft flowing serpentine cooling channel 72 (via aft flowing cavities 84, 86, 88, and 90) to be exhausted from the airfoil through the trailing edge 42, and in at least one embodiment, through a trailing edge exhaust orifice 80.
  • the forward flowing serpentine cooling channel 74 may extend from a position proximate the root 34 to the tip 50 of an airfoil 12 that is a turbine blade (or, for a turbine vane, from a position proximate the inner diameter of the turbine section to a position proximate the outer diameter of the turbine section).
  • the forward flowing serpentine cooling channel 74 may be formed from at least a two pass serpentine cooling channel, and, in particular embodiments, may be a three-pass serpentine cooling channel or greater.
  • forward flowing serpentine cooling channel 74 may be a forward flowing spiral serpentine cooling channel having forward flowing cavities 94, 96, 98, and 100 in contact with at least one of the pressure side 40 and the suction side 42.
  • Forward flowing cavity 94 may be in communication with a second cooling fluid supply inlet (not shown), and may be configured to pass the cooling fluids through the forward flowing cooling channel 74 (via forward flowing cavities 96, 98, and 100) to be exhausted from the airfoil through the leading edge 46, and in at least one embodiment, through leading edge orifices 44.
  • cooling fluids flowing through forward flowing cavity 96 may be fed into forward flowing cavity 98 through one or more tangential holes 102, and the cooling fluids flowing through forward flowing cavity 98 may be fed into forward flowing cavity 100 through one or more tangential holes 104.
  • cooling fluids may be passed from a cooling fluid supply (not shown), such as but not limited to, a compressor, to the airfoil 12.
  • a first portion of the cooling fluids enter the aft flowing cavity 82 of the aft flowing serpentine cooling channel 72 via a first cooling fluid supply inlet (not shown).
  • the first portion of the cooling fluids pass through the aft flowing serpentine cooling channel 72 in a spiral, serpentine manner, absorbing heat from the surfaces of the pressure side 40 and suction side 42 of the outer wall 38.
  • the first portion of cooling fluids may flow up (e.g., out of the page in Figure 2) aft flowing cavity 82 and pass over to aft flowing cavity 84, flow down (e.g., into the page in Figure 2) aft flowing cavity 84 and pass over to aft flowing cavity 86, flow up aft flowing cavity 86 and pass over to aft flowing cavity 88, flow down aft flowing cavity 88 and pass over to aft flowing cavity 90, and flow up aft flowing cavity 90 before being exhausted from the airfoil 12 through the trailing edge 48.
  • a second portion of the cooling fluids also enter the forward flowing cavity 94 of the forward flowing serpentine cooling channel 74 via a second cooling fluid supply inlet (not shown).
  • the second portion of the cooling fluids pass through the forward flowing serpentine cooling channel 74 in a spiral, serpentine manner, absorbing heat from the surfaces of the pressure side 40 and suction side 42 of the outer wall 38.
  • the second portion of cooling fluids may flow up forward flowing cavity 94 and pass over to forward flowing cavity 96, flow down forward flowing cavity 96 and pass over to forward flowing cavity 98 through tangential holes 102.
  • the cooling fluids flowing through forward flowing cavity 98 may be fed into forward flowing cavity 100 through one or more tangential holes 104.
  • the second portion of cooling fluids may impinge on a backside surface of the leading edge 46 and may be exhausted through the orifices 44 forming a showerhead.
  • the second portion of the cooling fluids may flow up forward flowing cavity 94, pass over to forward flowing cavity 96, and flow down forward flowing cavity 96. Cooling fluids may be fed from cavity 96 into forward flowing cavity 98 through one or more cross over holes 102. Cooling fluids may impinge on a backside surface of the leading edge 46 and may be exhausted through the orifices 44 forming a showerhead.
  • the cooling system 10 may include one or more additional elements and/or modifications.
  • the cooling system 10 may include one or more impingement orifices, one or more trailing edge pin fins, one or more blade tip holes, one or more rib turbulators, one or more film or gill holes, one or more refresher feeds, one or more fillets at the connection points between ribs 58 and the outer wall 38, any other elements for cooling airfoils 12, or any combination of the preceding.
  • cooling fluids are described above as flowing through aft flowing serpentine cooling channel 72 and/or forward flowing serpentine cooling channel 74 in a spiral, serpentine manner, the cooling fluids may pass over and/or under any shared wall (such as a connected segment 62 or a rib 58) in the channel 72 and/or channel 74 while flowing through channel 72 and/or channel 74.
  • shared wall such as a connected segment 62 or a rib 58

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne une aube (12) pour une turbine à gaz, l'aube (12) comprenant un système de refroidissement interne (10) formé en partie par une paroi interne segmentée (54) qui répartit des charges mécaniques. La paroi interne segmentée (54) peut comprendre une pluralité de segments connectés (62). Dans un mode de réalisation, au moins une partie de la pluralité de segments connectés (62) peut être non parallèle à une ligne de cambrure (66) de l'aube creuse généralement allongée (32). Dans un autre mode de réalisation, la totalité de la pluralité de segments connectés (62) peut être non parallèle à la ligne de cambrure (66) de l'aube creuse généralement allongée (32). Dans un autre mode de réalisation, la pluralité de segments connectés (62) peut être inclinée de façon alternée par rapport à la ligne de cambrure (66) de l'aube creuse généralement allongée (32).
PCT/US2015/016482 2015-02-19 2015-02-19 Aube de turbine dotée d'une paroi interne segmentée WO2016133513A1 (fr)

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Application Number Priority Date Filing Date Title
PCT/US2015/016482 WO2016133513A1 (fr) 2015-02-19 2015-02-19 Aube de turbine dotée d'une paroi interne segmentée

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2015/016482 WO2016133513A1 (fr) 2015-02-19 2015-02-19 Aube de turbine dotée d'une paroi interne segmentée

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WO2016133513A1 true WO2016133513A1 (fr) 2016-08-25

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112105800A (zh) * 2018-05-29 2020-12-18 赛峰飞机发动机公司 包括配备有多个最佳布置的干扰元件的内部流体流动通道的涡轮机叶片

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2672338A1 (fr) * 1991-02-06 1992-08-07 Snecma Aube de turbine munie d'un systeme de refroidissement.
EP1008724A2 (fr) * 1998-12-09 2000-06-14 General Electric Company Configuration de refroidissement pour une aube d'une turbomachine
US7611330B1 (en) * 2006-10-19 2009-11-03 Florida Turbine Technologies, Inc. Turbine blade with triple pass serpentine flow cooling circuit
US7901181B1 (en) * 2007-05-02 2011-03-08 Florida Turbine Technologies, Inc. Turbine blade with triple spiral serpentine flow cooling circuits
US8070442B1 (en) * 2008-10-01 2011-12-06 Florida Turbine Technologies, Inc. Turbine airfoil with near wall cooling
US8342802B1 (en) * 2010-04-23 2013-01-01 Florida Turbine Technologies, Inc. Thin turbine blade with near wall cooling

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2672338A1 (fr) * 1991-02-06 1992-08-07 Snecma Aube de turbine munie d'un systeme de refroidissement.
EP1008724A2 (fr) * 1998-12-09 2000-06-14 General Electric Company Configuration de refroidissement pour une aube d'une turbomachine
US7611330B1 (en) * 2006-10-19 2009-11-03 Florida Turbine Technologies, Inc. Turbine blade with triple pass serpentine flow cooling circuit
US7901181B1 (en) * 2007-05-02 2011-03-08 Florida Turbine Technologies, Inc. Turbine blade with triple spiral serpentine flow cooling circuits
US8070442B1 (en) * 2008-10-01 2011-12-06 Florida Turbine Technologies, Inc. Turbine airfoil with near wall cooling
US8342802B1 (en) * 2010-04-23 2013-01-01 Florida Turbine Technologies, Inc. Thin turbine blade with near wall cooling

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112105800A (zh) * 2018-05-29 2020-12-18 赛峰飞机发动机公司 包括配备有多个最佳布置的干扰元件的内部流体流动通道的涡轮机叶片
CN112105800B (zh) * 2018-05-29 2023-04-07 赛峰飞机发动机公司 飞行器涡轮机叶片及其增材制造方法和飞行器发动机

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