EP0475428A1 - Gasturbine, für diese Gasturbine verwendete Schaufel und Verfahren zur Herstellung dieser Schaufel - Google Patents

Gasturbine, für diese Gasturbine verwendete Schaufel und Verfahren zur Herstellung dieser Schaufel Download PDF

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Publication number
EP0475428A1
EP0475428A1 EP91115542A EP91115542A EP0475428A1 EP 0475428 A1 EP0475428 A1 EP 0475428A1 EP 91115542 A EP91115542 A EP 91115542A EP 91115542 A EP91115542 A EP 91115542A EP 0475428 A1 EP0475428 A1 EP 0475428A1
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EP
European Patent Office
Prior art keywords
gas turbine
turbine blade
shank
less
wing
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.)
Granted
Application number
EP91115542A
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English (en)
French (fr)
Other versions
EP0475428B1 (de
Inventor
Akira Yoshinari
Tosiaki Saito
Katsumi Iijima
Tadami Ishida
Ryozo Hashida
Kimio Kano
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Tohoku Electric Power Co Inc
Hitachi Ltd
Original Assignee
Tohoku Electric Power Co Inc
Hitachi Ltd
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 Tohoku Electric Power Co Inc, Hitachi Ltd filed Critical Tohoku Electric Power Co Inc
Publication of EP0475428A1 publication Critical patent/EP0475428A1/de
Application granted granted Critical
Publication of EP0475428B1 publication Critical patent/EP0475428B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22D—CASTING OF METALS; CASTING OF OTHER SUBSTANCES BY THE SAME PROCESSES OR DEVICES
    • B22D3/00—Pig or like casting
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/607—Monocrystallinity

Definitions

  • the present invention relates to a gas turbine, a heavy-duty gas turbine blade, which has horizontally extending protrusions, and a manufacturing method for the gas turbine blade.
  • Ni-base superalloys have hitherto been used as materials for the rotor blades of electricity generating gas turbines.
  • the temperature of gas has been increased year after year.
  • conventional casting blades having complicated cooling holes therein have been employed.
  • Single-crystal wings have already been used as rotor blades of aircraft jet engines. Alloys for casting the single-crystal wing are developed on the assumption that they do not have grain boundaries, and therefore they do not contain grain boundary strengthening elements such as B, Zr and Hf. For this reason, the grain boundaries of single-crystal alloys are weak. At least a portion of a casting must be single-crystallized before the casting can be used. In order to use the single-crystal wing as a gas turbine rotor blade, it is indispensable for the entire casting to be single-crystallized.
  • the rotor blade for the aircraft jet engine has a length of approximately 10 cm, and the cross-section area of a shaft is 10 cm 2 at the largest.
  • the size of a platform extending horizontally from the main body of the rotor blade is small. Because the entire rotor blade is such a small component, a single-crystal wing can be manufactured by solidifying a wing-shaped casting through the above unidirectional solidification process.
  • rotor blades in electricity generating gas turbines are larger than those in aircraft jet engines.
  • the former have a length of 14-16 cm at the shortest or more and shanks having a cross-section area of 15 cm 2 or more. It is therefore difficult to manufacture the former in a single-crystal structure.
  • the horizontally protruding portion Since the horizontally protruding portion has no relationship with the other portion of the casting, it will have crystal orientation different from that of the other portion. When this portion and the other portion of the casting are further solidified and the crystals of both come into contact with each other, the contacting surface is formed into a grain boundary, thus preventing a single crystal from growing.
  • An object of the present invention is to provide a large single-crystal turbine blade excellent in tensile and creep strength and in thermal fatigue performance at heat and stress. Another object of the invention is to provide a manufacturing method for such a turbine blade. A further object is to provide a heavy-duty gas turbine having high thermal efficiency.
  • this invention provides a gas turbine blade comprising: a dovetail serving as a portion secured to a disk; a shank which is connected to the dovetail and has one or more protrusions integrally formed on the side of the dovetail; and a wing connected to the shank; wherein the gas turbine blade is made of a Ni-base alloy in which a y' phase is precipitated substantially in a y phase which is formed in a single-crystal structure.
  • the protrusions provided in the shank of the turbine blade may be sealing portions, in a single stage or multi-stages, provided on both surfaces along a surface where the wing rotates. The edge of the sealing portion bends towards the wing.
  • the protrusion provided in the shank is one platform provided on both surfaces intersecting with the surface where the wing rotates.
  • the shank, in which the protrusions are provided has a cross-section area of not less than 15 cm 2.
  • the shank and the wing including the dovetail and the protrusions are made of the Ni-base alloy in which the y' phase is precipitated in a single-crystal base of the ⁇ phase.
  • the gas turbine blade has an overall length of not less than 180 mm in a longer direction thereof.
  • the wing weighs not more than 30%, particularly 20-30%, of the overall weight of the gas turbine blade.
  • This invention also provides a manufacturing method for a gas turbine blade including a dovetail serving as a portion secured to a disk; a shank which is connected to the dovetail and has protrusions integrally formed on the side of the dovetail; and a wing connected to the shank, the manufacturing method comprising the steps of: connecting a by-pass mold corresponding to the protrusions to a main mold corresponding to the dovetail, the shank and the wing; and casting a single-crystal structure by gradually solidifying at the same speed in one direction molten metal of Ni-base alloy in the main mold and the by-pass mold.
  • the invention further provides a gas turbine blade comprising: a dovetail serving as a portion secured to a disk; a shank which is connected to the dovetail and has one or more protrusions integrally formed on the side of the dovetail; and a wing connected to the shank; wherein the gas turbine blade is solidified from the edge of the wing to the dovetail by a unidirectional solidification process, a ⁇ phase being made of a single-crystal Ni-base alloy.
  • the invention provides a heavy-duty gas turbine comprising: a compressor; a combustion liner; a turbine blade, in a single stage or multi-stages, which has a dovetail secured to a turbine disk and has an overall length of not less than 180 mm, and which is made of a single-crystal Ni-base alloy whose ⁇ phase is a single crystal; and a turbine nozzle provided in correspondence to the turbine blade; wherein operating gas temperature is not less than 1400°C, and metal temperature of a first blade is not less than 10000 C under working stress.
  • the mold having the by-pass formed in the protrusion is employed separately from the other mold used for the dovetail, the shank and the wing.
  • the manufacturing method for the gas turbine blade, according to this invention is capable of manufacturing a large gas turbine blade having a complicated configuration and the single-crystal structure.
  • the turbine blade of the invention is a large blade having the protrusion formed where the cross-section area of the blade is 15 cm 2 or more, it has more strength than a blade made of a polycrystal having grain boundaries because it is made in the single-crystal structure.
  • Ni-base alloys should be used for the turbine blade in this invention, each alloy containing by weight 0.15% or less C or preferably 0.02% as an impurity; 0.03% or less Si; more preferably an impurity; 2.0% or less Mn; 5-14% Cr; 1-7% AR; 1-5% Ti; 2.0% or less Nb; 2-15% W; 5% or less Mo; 12% or less Ta, more preferably 2-10%; 10% or less Co; 0.2% or less Hf; 3.0% or less Re; and 0.02% or less B.
  • Table 1 shows the above Ni-base alloys, indicating weight percent of the elements in the alloys.
  • Co-based alloys may be used in this invention, each alloy containing by weight 0.2-0.6% C; 0.5% or less Si; 2% or less Mn; 20-30% Cr; 20% or less Ni; 5% or less Mo; 2-15% W; 5% or less Nb; 0.5% or less Ti; 0.5% or less At ; 5% or less Fe; 0.02% or less B; 0.5% or less Zr; 5% or less Ta; and the remaining weight percent constitutes Co.
  • Table 2 shows the above Co-based alloys, used for a turbine nozzle serving as a stator blade, indicating weight percent of the elements in the alloys.
  • the gas turbine of this invention is more efficient because it is large and permits an operating gas temperature to increase to 1400 °C or more at an early stage of the operation.
  • Crystal orientation in the horizontally protruding portion with respect to the direction in which solidification advances is oriented so that it may be in the same crystal orientation as the casting. It is thus possible to efficiently manufacture the large single-crystal rotor blade.
  • the characteristics of the single-crystal rotor blade of the invention are excellent at high temperatures, the service life of the blade is extended, and the thermal efficiency of the gas turbine caused by an increase in the fuel gas temperature is increased to 34%.
  • Fig. 1 is a perspective view of the rotor blade, according to the present invention, of an electricity generating gas turbine.
  • Fig. 2 is a vertical cross-sectional view showing a manufacturing method for the rotor blade. This method employs a mold of this invention to manufacture the rotor blade.
  • a shell mold 2 made of alumina, is secured to a water-cooled chill 1, and is placed in a mold heating heater 3 in which it it heated to not less than the melting temperature of a Ni-base alloy.
  • a dissolved alloy is poured into the mold 2, and then the water-cooled chill 1 is withdrawn downwardly to solidify the alloy by a unidirectional solidification process.
  • many crystals are first formed in a starter 4 at the lower end of the mold 2, and are then formed into one single crystal in a selector 5, capable of rotating 3600, while the alloy is still being solidified.
  • the single crystal becomes larger in an enlarged section 6.
  • the alloy is solidified and formed into a casting 7, which is composed of a wing 8 having cooling holes formed therein, a shank 9 on the wing 8, and a Christmas tree- shaped dovetail 10 on the shank 9. (These components 8, 9 and 10 are shown upside down in Fig. 1.) Sealing portions or protrusions 11, the end of each bending toward the wing 8, protrude from the dovetail 10. As shown in Fig. 2, the turbine blade is cast from the wing 8 of the turbine rotor blade to the shank 9 and the dovetail 10 shown in Fig. 1.
  • a by-pass mold 12 different from the casing 7 is provided from the point of section enlargement 6 to the sealing portions or protrusions 11.
  • the provision of the by-pass mold 12 permits the entire rotor blade of the turbine to be single-crystallized.
  • the turbine rotor blade shown in Fig. 1 measures approximately 180 mm high x 40 mm wide x 100 mm long, as denoted by numerals 13, 14 and 15, respectively.
  • the wing 8 measures approximately 90 mm high, and weighs approximately 30% of the weight of the entire turbine rotor blade.
  • the cross-section area of the shank 9, where the sealing portions or protrusions 11 are formed, is 40 cm 2.
  • the sealing portions 11 each extend approximately 15 mm.
  • the casting heater 3 is maintained at high temperatures until the casting 7 is withdrawn and solidified completely.
  • the casting process mentioned above is performed in a vacuum.
  • the turbine rotor blade made from the single crystal After the turbine rotor blade made from the single crystal has been cast, it is subjected to a solution heat treatment in a vacuum at temperatures of 1300-1350 ° C for 2-10 hours.
  • a eutectic y' phase formed by solidifying the alloy is changed into a ⁇ phase.
  • the turbine rotor blade is then subjected to an aging treatment at temperatures of 980-10800 C for 4-15 hours and at temperatures of 800-900 ° for 10-25 hours.
  • Horn-shaped ⁇ ' phases, each having an average size of 3-5 ⁇ m, are precipitated in the ⁇ phase.
  • Table 3 shows conditions for casting the single-crystal wing.
  • Table 4 shows the comparison between the yield of single-crystal wings manufactured by the method of this invention and the yield of such wings manufactured by the conventional method.
  • the turbine rotor blade is shrunk at the upper portion of a platform, and the secondary growth of a long, thin dendrite is found at the lower portion of the platform.
  • this invention makes it possible to manufacture a large single-crystal wing which cannot be manufactured by the conventional method.
  • the time during which the rotor blade is in contact with the molten mold is shortened. It is possible to obtain a turbine rotor blade made of an alloy containing elements which vary little and have few defects; as a result, a turbine rotor blade having the required characteristics can be manufactured. It takes approximately one hour for the wing to solidify, and approximately two hours for the other components and the dovetail to solidify finally.
  • the elements in an alloy vary, and particularly Cr varies greatly.
  • the by-pass mold 12 different from the mold used for forming the turbine rotor blade, may be provided in a position which is above the selector 5 in a selector method or above a seed in a seed method, and which is anywhere below the sealing portions or protrusions 11.
  • the by-pass mold 12 must be removed; therefore desirably, the by-pass mold 12 should be provided in the enlarged section 6, shown in Fig. 2, which is above the selector 5 or the seed and is below the wing 8.
  • the rotor blade is solidified from the wing 8 to the dovetail 10 for the following reasons.
  • the wing 8 of the gas turbine rotor blade is the essential part of the rotor blade, and is subjected to high temperatures and stress. It therefore must possess fewer defects and be of a higher-quality than any other components.
  • the wing 8 is first solidified, so that the time during which it is held at high temperatures is shortened. In order to make the elements vary little, such casting is suitable for manufacturing the rotor blade of the gas turbine.
  • a plurality of cooling holes are provided leading from the wing 8 to the dovetail 10, and are used for cooling the components by a refrigerant.
  • a core for the cooling holes is used as the mold.
  • the speed at which the alloy is solidified varies from 1 to 50 cm/h according to the size of the casting to be solidified.
  • the wing 8 can be solidified faster than the shank 9 and the dovetail 10.
  • a rotor blade having substantially the same configuration as that of the rotor blade in the first embodiment is cast using the alloy of No. 2.
  • the same casting conditions and the unidirectional solidification process as those in the first embodiment are employed in the second embodiment.
  • the blade measures 160 mm high; a wing measures 70 mm high; and a shank and a dovetail each measure 90 mm high.
  • Fig. 3 shows the front view of this rotor blade. Because the rotor blade has a wide platform 17, when it is solidified by the unidirectional solidification process, a new crystal is formed at the platform 17, thus preventing a single crystal from growing. To solve this problem, the present invention is applied to the method of manufacturing the rotor blade. As shown in Fig. 4, a portion near the edge of the platform 17 is connected to a portion above a selector 5 by means of a by-pass mold 12, different from the mold for forming a casting 7. Such connection makes it possible for a single crystal to grow.
  • the by-pass mold 12 has a thickness of 1 mm and a width of 20 mm.
  • Fig. 4 shows the cross-sectional configuration of the rotor blade; Fig.
  • Fig. 5 shows how the new crystal grows in the conventional method, as seen from the upper portion of the wing 8; and Fig. 6 shows how the new crystal does not grow in this invention, as seen also from the upper portion of the wing 8.
  • numeral 18 denotes a grain boundary
  • numeral 19 denotes the new crystal. This invention makes it possible for the single crystal to grow, instead of a new crystal growing.
  • Fig. 7 is a partial cross-sectional view showing the rotary portion of a gas turbine.
  • the Ni-base alloy of No. 2 made of the single crystal, obtained in the first embodiment of this invention is used for a first turbine blade 20.
  • a turbine disk 21 has two stages. The first stage is disposed upstream of a gas flow, whereas the second stage, having a central hole 22 formed therein, is disposed downstream of the gas flow.
  • a martensitic heat resisting steel containing 12% Cr is used for the final stage of a compressor disk 23, a distant piece 24, a turbine spacer 25, a turbine stacking bolt 26 and a compressor stacking bolt 27.
  • the turbine blade 20 in a second stage, a turbine nozzle 28, the liner 30 of a combustor 29, a compressor blade 31, a compressor nozzle 32, diaphragm 33 and a shroud 34 are made of alloys. The elements contained in these alloys are shown in Table 5.
  • the turbine nozzle 28 in a first stage and the turbine blade 20 are made of a single-crystal casting.
  • the turbine nozzle 28 in the first stage is made of an alloy of No. 13, and is composed of one segment for each wing in the same manner as in the turbine blade.
  • the turbine nozzle 28 is disposed on a circumference, and has a diaphragm and a length which is substantially equal to the wing of the blade.
  • Numeral 35 denotes a turbine stab shaft
  • numeral 36 denotes a compressor stab shaft.
  • a compressor used in this embodiment has 17 stages.
  • the turbine blade, the turbine nozzle, a shroud segment (1) and the diaphragm, all shown in Table 5, are used in the first stage upstream of the gas flow, whereas a shroud segment (2) is used in the second stage.
  • a layer made of a highly-concentrated alloy containing AR, Cr and other elements, or made of a mixture containing oxides may be used as a coating layer which is resistant to oxidation and corrosion at temperatures higher than those at which an alloy serving as a base material is resistant to oxidation and corrosion.
  • the crystal may be formed so that its orientation becomes [001 ] in the direction in which a centrifugal force is applied.
  • a blade having high strength is obtainable by forming the crystal in this way.
  • the gas temperature at the entrance of the turbine nozzle in the first stage is capable of rising as high as 1500°C
  • the metal temperature at the blade in the first stage is capable of rising as high as 1000 C.
  • Thirty four percent thermal efficiency is obtainable.
  • the heat resisting steel having higher creep rupture strength and fewer defects caused by heat is used for the turbine disk, the distant piece, the spacer, the final stage of the compressor disk, and the stacking bolt.
  • the alloy having strength at high temperatures is used for the turbine blade; the alloy having strength and ductility at high temperatures is used for the turbine nozzle; and the alloy having high fatigue performance and strength at high temperatures is used for the liner of the combustor. It is thus possible to obtain a gas turbine which is more reliable in various aspects than the conventional art.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Crystals, And After-Treatments Of Crystals (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP91115542A 1990-09-14 1991-09-13 Gasturbine, für diese Gasturbine verwendete Schaufel und Verfahren zur Herstellung dieser Schaufel Expired - Lifetime EP0475428B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
JP245210/90 1990-09-14
JP2245210A JP2729531B2 (ja) 1990-09-14 1990-09-14 ガスタービンブレード及びその製造方法並びにガスタービン

Publications (2)

Publication Number Publication Date
EP0475428A1 true EP0475428A1 (de) 1992-03-18
EP0475428B1 EP0475428B1 (de) 1998-01-07

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EP91115542A Expired - Lifetime EP0475428B1 (de) 1990-09-14 1991-09-13 Gasturbine, für diese Gasturbine verwendete Schaufel und Verfahren zur Herstellung dieser Schaufel

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Country Link
EP (1) EP0475428B1 (de)
JP (1) JP2729531B2 (de)
KR (1) KR0185206B1 (de)
CN (1) CN1034828C (de)
CA (1) CA2051133C (de)
DE (1) DE69128580T2 (de)

Cited By (5)

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EP0637476A1 (de) * 1993-08-06 1995-02-08 Hitachi, Ltd. Gasturbinenschaufel, Verfahren zur Herstellung sowie Gasturbine mit dieser Schaufel
EP3085902A1 (de) * 2015-04-24 2016-10-26 United Technologies Corporation Einzelkristallkornstrukturversiegelungen und verfahren zur formung
US10202853B2 (en) 2013-09-11 2019-02-12 General Electric Company Ply architecture for integral platform and damper retaining features in CMC turbine blades
US10287897B2 (en) 2011-09-08 2019-05-14 General Electric Company Turbine rotor blade assembly and method of assembling same
US10648352B2 (en) 2012-06-30 2020-05-12 General Electric Company Turbine blade sealing structure

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DE69316251T2 (de) * 1992-03-09 1998-05-20 Hitachi Ltd Hochgradig heisskorrosionsbeständige und hochfeste Superlegierung, hochgradig heisskorrosionsbeständiges und hochfestes Gussstück mit Einkristallgefüge, Gasturbine und kombiniertes Kreislaufenergieerzeugungssystem
JPH0959747A (ja) * 1995-08-25 1997-03-04 Hitachi Ltd 高強度耐熱鋳鋼，蒸気タービンケーシング，蒸気タービン発電プラント及び蒸気タービン
JP3209902B2 (ja) * 1995-11-06 2001-09-17 キャノン・マスキーガン・コーポレーション 高温腐食抵抗性の単結晶ニッケル系スーパーアロイ
KR100530759B1 (ko) * 1999-02-18 2005-11-23 삼성테크윈 주식회사 항공기용 가스터어빈 엔진
DE10100790C2 (de) * 2001-01-10 2003-07-03 Mtu Aero Engines Gmbh Nickel-Basislegierung für die gießtechnische Herstellung einkristallin erstarrter Bauteile
US7526965B2 (en) * 2006-12-30 2009-05-05 General Electric Company Method for evaluating burnishing element condition
JP5232492B2 (ja) 2008-02-13 2013-07-10 株式会社日本製鋼所 偏析性に優れたＮｉ基超合金
US20100071812A1 (en) * 2008-09-25 2010-03-25 General Electric Company Unidirectionally-solidification process and castings formed thereby
JP5063550B2 (ja) * 2008-09-30 2012-10-31 株式会社日立製作所 ニッケル基合金及びそれを用いたガスタービン翼
KR101023783B1 (ko) * 2009-08-05 2011-03-21 한국전력공사 가스터빈 압축기의 고정익 고정구조
US8226886B2 (en) * 2009-08-31 2012-07-24 General Electric Company Nickel-based superalloys and articles
US9039375B2 (en) * 2009-09-01 2015-05-26 General Electric Company Non-axisymmetric airfoil platform shaping
US8641381B2 (en) * 2010-04-14 2014-02-04 General Electric Company System and method for reducing grain boundaries in shrouded airfoils
JP5396445B2 (ja) * 2011-08-29 2014-01-22 株式会社日立製作所 ガスタービン
KR101427801B1 (ko) * 2011-12-30 2014-09-25 두산중공업 주식회사 가스터빈 압축기의 블레이드 및 그의 제조방법
US20130177438A1 (en) * 2012-01-06 2013-07-11 General Electric Company Sectioned rotor, a steam turbine having a sectioned rotor and a method for producing a sectioned rotor
US9097128B2 (en) * 2012-02-28 2015-08-04 General Electric Company Seals for rotary devices and methods of producing the same
CN105108061A (zh) * 2015-09-30 2015-12-02 东方电气集团东方汽轮机有限公司 消除单晶叶片中杂晶缺陷的方法
EP3677697A1 (de) * 2019-01-07 2020-07-08 Siemens Aktiengesellschaft Co-legierung zur generativen fertigung und verfahren
FR3105048B1 (fr) * 2019-12-20 2022-08-05 Safran Solution de fabrication d'un disque aubage monobloc
CN114872909B (zh) * 2022-05-06 2023-03-24 中国航发四川燃气涡轮研究院 飞机型涡轮叶片冷却通道换热结构

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Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0637476A1 (de) * 1993-08-06 1995-02-08 Hitachi, Ltd. Gasturbinenschaufel, Verfahren zur Herstellung sowie Gasturbine mit dieser Schaufel
US5611670A (en) * 1993-08-06 1997-03-18 Hitachi, Ltd. Blade for gas turbine
US10287897B2 (en) 2011-09-08 2019-05-14 General Electric Company Turbine rotor blade assembly and method of assembling same
US10648352B2 (en) 2012-06-30 2020-05-12 General Electric Company Turbine blade sealing structure
US10202853B2 (en) 2013-09-11 2019-02-12 General Electric Company Ply architecture for integral platform and damper retaining features in CMC turbine blades
EP3085902A1 (de) * 2015-04-24 2016-10-26 United Technologies Corporation Einzelkristallkornstrukturversiegelungen und verfahren zur formung
US10830357B2 (en) 2015-04-24 2020-11-10 Raytheon Technologies Corporation Single crystal grain structure seals and method of forming

Also Published As

Publication number Publication date
DE69128580T2 (de) 1998-04-30
CN1034828C (zh) 1997-05-07
CA2051133A1 (en) 1992-03-15
CN1060890A (zh) 1992-05-06
JPH04124237A (ja) 1992-04-24
KR0185206B1 (ko) 1999-04-01
EP0475428B1 (de) 1998-01-07
DE69128580D1 (de) 1998-02-12
CA2051133C (en) 2000-08-29
KR920006057A (ko) 1992-04-27
JP2729531B2 (ja) 1998-03-18

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