WO2017119898A1 - Aube de turbine à bout aminci d'aube multicouche à plusieurs hauteurs - Google Patents

Aube de turbine à bout aminci d'aube multicouche à plusieurs hauteurs Download PDF

Info

Publication number
WO2017119898A1
WO2017119898A1 PCT/US2016/012637 US2016012637W WO2017119898A1 WO 2017119898 A1 WO2017119898 A1 WO 2017119898A1 US 2016012637 W US2016012637 W US 2016012637W WO 2017119898 A1 WO2017119898 A1 WO 2017119898A1
Authority
WO
WIPO (PCT)
Prior art keywords
side rib
blade
pressure side
suction side
tip
Prior art date
Application number
PCT/US2016/012637
Other languages
English (en)
Inventor
Wen-Lung FU
Ching-Pang Lee
Zhihong Gao
Gm Salam Azad
Original Assignee
Siemens Aktiengesellschaft
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 Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Priority to PCT/US2016/012637 priority Critical patent/WO2017119898A1/fr
Publication of WO2017119898A1 publication Critical patent/WO2017119898A1/fr

Links

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
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/182—Two-dimensional patterned crenellated, notched
    • 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
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the present invention relates to turbine blades for a gas turbine and, more particularly, to multi-layer multi-height blade squealers.
  • Turbine inlet temperature is limited by the material properties and cooling capabilities of the turbine parts.
  • a combustion system receives air from a compressor and raises it to a high energy level by mixing in fuel and burning the mixture, after which products of the combustor are expanded through the turbine.
  • a turbine blade is formed from a root portion coupled to a rotor disc and an elongated portion forming a blade that extends outwardly from a platform coupled to the root portion at an opposite end of the turbine blade.
  • the blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge.
  • the tip of a turbine blade often has a tip feature to reduce the size of the gap between ring segments and blades in the gas path of the turbine to prevent tip flow leakage, which reduces the amount of torque generated by the turbine blades.
  • the tip features are often referred to as squealer tips and are frequently incorporated onto the tips of blades to help reduce aerodynamic losses between turbine stages. These features are designed to minimize the leakage between the blade tip and the ring segment.
  • a blade for a turbine engine comprising: a generally elongated blade comprising a leading edge, a trailing edge, a tip at a first end, a root coupled to the elongated blade at a second end generally opposite the first end supporting the blade and for coupling the blade to a disc, and an internal cooling system formed from at least one cavity positioned within the generally elongated blade; at least one outer pressure side rib extending radially from an outer surface of the tip; at least one inner pressure side rib extending radially from the outer surface of the tip, wherein the at least one outer pressure side rib has a smaller radial height than the at least one inner pressure side rib, wherein at least one open trench surface is formed between the at least one outer pressure side rib and the at least one inner pressure side rib; at least one film cooling hole positioned in between the at least one outer pressure side rib and the at least one inner pressure side rib with an outlet in an outer surface in the at least one open
  • FIG 1 is a cross-sectional view detailing portions of an exemplary embodiment of the present invention
  • FIG 2 is a cross-sectional view detailing coolant and hot gas streamlines of an exemplary embodiment of the present invention
  • FIG 3 is a perspective view of a turbine blade with a squealer tip according to an exemplary embodiment of the present invention
  • FIG 4 is a detailed view of the squealer tip at the leading edge of the turbine blade shown in FIG 3;
  • FIG 5 is a partial cross-sectional view of the turbine blade tip taken at section line 5-5 in FIG 3;
  • FIG 6 is a perspective view of a turbine blade with a squealer tip according to an exemplary embodiment of the present invention.
  • FIG 7 is a detailed view of the squealer tip at the leading edge of the turbine blade shown in FIG 6;
  • FIG 8 is a partial cross-sectional view of the turbine blade tip taken at section line 8-8 in FIG 6.
  • an embodiment of the present invention provides a blade for a turbine engine that includes a squealer tip formed from at least one outer pressure side rib, at least one inner pressure side rib, and at least one outer suction side rib extending radially outward from a tip of the turbine blade is disclosed.
  • the at least one outer pressure side rib has a smaller radial height than the at least one inner pressure side rib.
  • At least one open trench surface is formed between the at least one outer pressure side rib and the at least one inner pressure side rib.
  • At least one film cooling hole having an outlet and an inlet is positioned in the at least one open trench surface. The inlet of the at least one film cooling hole couples the at least one film cooling hole with at least one cavity forming an internal cooling system.
  • the at least one inner pressure side rib and the at least one outer suction side rib are separated by at least a linear tip surface.
  • a gas turbine engine may comprise a compressor section, a combustor and a turbine section.
  • the compressor section compresses ambient air.
  • the combustor combines the compressed air with a fuel and ignites the mixture creating combustion products comprising hot gases that form a working fluid.
  • the working fluid travels to the turbine section.
  • Within the turbine section are circumferential alternating rows of vanes and blades, the blades being coupled to a rotor. Each pair of rows of vanes and blades forms a stage in the turbine section.
  • the turbine section comprises a fixed turbine casing, which houses the vanes, blades and rotor.
  • a turbine blade 20 is formed from a root portion 44 coupled to a rotor disc (not shown) and an elongated portion forming a blade that extends outwardly from a platform 42 coupled to the root portion 44 at an opposite end of the turbine blade 20.
  • the blade 20 is composed of a tip 10 opposite the root section 44, a leading edge 48, and a trailing edge 46.
  • the tip 10 of a turbine blade 20 or tip feature is in position to reduce the size of the gap between ring segments 32 and blades 20 in a gas path of the turbine to prevent tip flow leakage, which reduces the amount of torque generated by the turbine blades 20.
  • the tip feature is referred to as a squealer or squealer tip 10 and is incorporated onto the tips of blades to help reduce aerodynamic losses between turbine stages. These features are designed to minimize the leakage between the blade tip and the ring segment 32.
  • the squealer 10 can be a sacrificial feature in a turbine blade to maintain a small tip clearance for better turbine efficiency and protect a blade internal cooling circuit, or also known as a cooling system 56, under a tip cap 64 in the event of tip rubbing against the ring segment 32 during the transient engine operation. Because the squealer 10 is away from the cooling circuit, it is very difficult to maintain an acceptable metal temperature. Because the squealer is subject to rubbing against the ring segment, no thermal barrier coating (TBC) 34 is applied on the squealer top surface.
  • TBC thermal barrier coating
  • Embodiments of the present invention provide a multilayer, multi-height blade squealer that may allow for the reduction of tip leakage and reduction of temperature along the squealer.
  • the squealer tip 10 formed from side ribs including at least one outer pressure side rib 12, at least one inner pressure side rib 14 and at least one outer suction side rib 16 extending radially outward from a tip cap upper surface 18 of a turbine blade 20 is disclosed.
  • the at least one outer pressure side rib 12 may be positioned along a pressure side 22 of the turbine blade 20.
  • the at least one inner pressure side rib 14 may be positioned along the tip cap upper surface 18 in proximity to the at least one outer pressure side rib 12.
  • At least one open trench surface 26 is formed between the at least one outer pressure side rib 12 and the at least one inner pressure side rib 14.
  • the at least one outer pressure side rib 12 may have a smaller radial height than the at least one inner pressure side rib 14.
  • the at least one outer suction side rib 16 may be positioned along a suction side 24 of the turbine blade 20.
  • the at least one inner pressure side rib 14 and the at least one outer suction side rib 16 may be separated by at least a linear tip surface.
  • the turbine blade 20 may be formed from a generally elongated blade 50 having a leading edge 48 and a trailing edge 46.
  • the generally elongated blade 50 may include the tip 10 at a first end 52 and a root 44 coupled to the elongated blade 50 at a second end 54 generally opposite the first end 52 for supporting the blade 20 and for coupling the blade 20 to a disc.
  • the internal cooling system 56 may be formed from at least one cavity 30 positioned within the generally elongated blade 50.
  • the cooling system 56 may have any appropriate configuration to cool the turbine blade 20 during use in an operating gas turbine engine.
  • the turbine blade 20 and its related components listed above may be formed from any appropriate material already known in the art or yet to be discovered or identified.
  • Each of the at least one open trench surface 26 may include at least one film cooling hole 28 having an exhaust outlet 60 (not shown in figures) positioned therein.
  • the exhaust outlet 60 may be in an outer surface 58 (not shown in figures) in the at least one open trench surface 26.
  • the at least one film cooling hole 28 may include an inlet 62 (not shown) that couples the film cooling hole 28 with the cavity 30 forming the internal cooling system 56.
  • the turbine blade 20 may include at least one inner suction side rib 40 extending radially from the outer surface of the tip cap upper surface 18.
  • the at least one inner suction side rib 40 may have a smaller radial height than the at least one outer suction side rib 16.
  • At least one film cooling hole 28 may be positioned in between the at least one inner suction side rib 40 and the at least one outer suction side rib 16.
  • the at least one film cooling hole 28 may include an outlet 60 in an outer surface 58 in at least one open trench surface 26 between the at least one outer suction side rib 16 and the at least one inner suction side rib 40 and an inlet 62 that couples the at least one film cooling hole 28 with the at least one cavity 30 forming the internal cooling system 56.
  • the at least one outer pressure side rib 12 may have a thermal barrier coating 34 on the top surface.
  • the at least one inner suction side rib 40 may have thermal barrier coating 34 on the top surface as well.
  • Thermal barrier coating 34 may be applied to the outer surfaces forming the pressure side 22 and the suction side 24 of the blade 20. Thermal barrier coating 34 may be also applied to the at least one trench surface 26 and tip cap upper surface 18.
  • the at least one outer suction side rib 16 may extend from the leading edge 48 toward the trailing edge 46 of the generally elongated blade 50.
  • the at least one outer suction side rib 16 may terminate at the trailing edge 46 and may be coupled to the at least one outer pressure side rib 12.
  • the at least one film cooling hole 28 may be drilled through the at least one open trench 26 and may be connected to the cooling passages inside the elongated blade 50.
  • the side ribs are of different radial heights.
  • the at least one outer pressure side rib 12 is shorter than the at least one inner pressure side rib 14.
  • the at least one outer suction side rib 16 is taller than the at least one inner suction side rib 40.
  • the squealer 10 may have side ribs directly above the tip cap 64 that may have direct conduction cooling from the tip cap 64.
  • the at least one inner pressure side rib 14 and in certain embodiments, the at least one inner suction side rib 40 may be directly above the tip cap 64.
  • cooling fluids 36 are passed into the internal cooling system 56.
  • the cooling fluids 36 may be passed into the at least one film cooling hole 28 in the tip 18 of the turbine blade 20.
  • the cooling fluids 36 may cool aspects of the squealer tip 10 by being exhausted through the outlets 60 into the at least one trench 26.
  • the at least one trench 26 may be along the pressure side 22 of the turbine blade 20 or along both the pressure side 22 and suction side 24 of the turbine blade 20.
  • the differences in radial height among the side ribs may allow for the addition of thermal barrier coating 34 to the at least one outer pressure side rib 12 or to the at least one outer pressure side rib 12 and at least one inner suction side rib 40 that may allow for a decrease in temperature along the top of the squealer 10. Since the lower radial height side ribs do not come in contact with the ring segment 32, the thermal barrier coating 34 is possible in that location. The thermal barrier coating 34 in that location allows for a decrease in the heat flux as the hot gas 38 crosses the squealer 10. The taller radial height side ribs may control the tip clearance against the ring segment 32.
  • the at least one open trench 26 provides access for cooling fluid 36 to flow out from the internal portion of the elongated blade 50 out into the squealer area.
  • the at least one open trench 26 may protect the film cooling flow from mixing with the hot gases and may provide convection cooling and film cooling effectiveness to the taller radial height side ribs which have the highest temperatures.
  • the at least one open trench 26 may be protected by the shorter radial height ribs for increased effectiveness in film cooling within the at least one open trench 26.

Landscapes

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

Abstract

L'invention concerne une aube (20) destinée à une turbomachine qui comprend un bout aminci (10) formé d'au moins une nervure (12) côté pression extérieure, d'au moins une nervure (14) côté pression intérieure et d'au moins une nervure (16) côté aspiration extérieure s'étendant radialement vers l'extérieur depuis une surface (18) supérieure du capuchon de pointe de l'aube (20) de turbine. La hauteur radiale de ladite nervure (12) côté pression extérieure est inférieure à celle de la nervure (14) côté pression intérieure. Au moins une surface (26) de tranchée ouverte, pourvue d'au moins un trou (28) de refroidissement lamellaire, est formée entre ladite nervure (12) côté pression extérieure et ladite nervure (14) côté pression intérieure. Une entrée (62) dudit trou (28) de refroidissement lamellaire relie ledit trou (28) de refroidissement lamellaire à au moins une cavité (30) formant un système (56) de refroidissement interne.
PCT/US2016/012637 2016-01-08 2016-01-08 Aube de turbine à bout aminci d'aube multicouche à plusieurs hauteurs WO2017119898A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PCT/US2016/012637 WO2017119898A1 (fr) 2016-01-08 2016-01-08 Aube de turbine à bout aminci d'aube multicouche à plusieurs hauteurs

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2016/012637 WO2017119898A1 (fr) 2016-01-08 2016-01-08 Aube de turbine à bout aminci d'aube multicouche à plusieurs hauteurs

Publications (1)

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WO2017119898A1 true WO2017119898A1 (fr) 2017-07-13

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Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3354853A1 (fr) * 2017-01-30 2018-08-01 United Technologies Corporation Aube de turbine à refroidissement pelliculaire à fente
EP3578759A1 (fr) * 2018-06-07 2019-12-11 United Technologies Corporation Profil aérodynamique et procédé associé pour diriger un flux de refroidissement
CN112031878A (zh) * 2020-11-05 2020-12-04 中国航发沈阳黎明航空发动机有限责任公司 一种涡轮转子叶片叶尖双层壁结构
CN112240229A (zh) * 2020-10-20 2021-01-19 西北工业大学 一种用于涡轮动力叶片顶部的高效冷却结构
US11118462B2 (en) 2019-01-24 2021-09-14 Pratt & Whitney Canada Corp. Blade tip pocket rib
EP3896258A1 (fr) * 2020-04-16 2021-10-20 General Electric Company Aube rotorique et moteur à turbine associé
WO2021236073A1 (fr) * 2020-05-20 2021-11-25 Siemens Aktiengesellschaft Aube de turbine
US11371359B2 (en) 2020-11-26 2022-06-28 Pratt & Whitney Canada Corp. Turbine blade for a gas turbine engine

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1059419A1 (fr) * 1999-06-09 2000-12-13 General Electric Company Aube avec trois nervures sur l'extrémité de l'aube

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1059419A1 (fr) * 1999-06-09 2000-12-13 General Electric Company Aube avec trois nervures sur l'extrémité de l'aube

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10815788B2 (en) 2017-01-30 2020-10-27 Raytheon Technologies Corporation Turbine blade with slot film cooling
EP3354853A1 (fr) * 2017-01-30 2018-08-01 United Technologies Corporation Aube de turbine à refroidissement pelliculaire à fente
US11028703B2 (en) 2018-06-07 2021-06-08 Raytheon Technologies Corporation Gas turbine engine airfoil with tip leading edge shelf discourager
US20190376395A1 (en) * 2018-06-07 2019-12-12 United Technologies Corporation Gas turbine engine airfoil with tip leading edge shelf discourager
EP3578759A1 (fr) * 2018-06-07 2019-12-11 United Technologies Corporation Profil aérodynamique et procédé associé pour diriger un flux de refroidissement
US11118462B2 (en) 2019-01-24 2021-09-14 Pratt & Whitney Canada Corp. Blade tip pocket rib
EP3896258A1 (fr) * 2020-04-16 2021-10-20 General Electric Company Aube rotorique et moteur à turbine associé
WO2021236073A1 (fr) * 2020-05-20 2021-11-25 Siemens Aktiengesellschaft Aube de turbine
CN115605668A (zh) * 2020-05-20 2023-01-13 西门子能源全球两合公司（De） 涡轮机叶片
US11761339B2 (en) 2020-05-20 2023-09-19 Siemens Energy Global GmbH & Co. KG Turbine blade
CN112240229A (zh) * 2020-10-20 2021-01-19 西北工业大学 一种用于涡轮动力叶片顶部的高效冷却结构
CN112031878A (zh) * 2020-11-05 2020-12-04 中国航发沈阳黎明航空发动机有限责任公司 一种涡轮转子叶片叶尖双层壁结构
US11371359B2 (en) 2020-11-26 2022-06-28 Pratt & Whitney Canada Corp. Turbine blade for a gas turbine engine

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