WO2017105380A1 - Profil aérodynamique de turbine comprenant une structure de support de partie centrale et une enveloppe externe thermorésistante - Google Patents

Profil aérodynamique de turbine comprenant une structure de support de partie centrale et une enveloppe externe thermorésistante Download PDF

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Publication number
WO2017105380A1
WO2017105380A1 PCT/US2015/065424 US2015065424W WO2017105380A1 WO 2017105380 A1 WO2017105380 A1 WO 2017105380A1 US 2015065424 W US2015065424 W US 2015065424W WO 2017105380 A1 WO2017105380 A1 WO 2017105380A1
Authority
WO
WIPO (PCT)
Prior art keywords
airfoil
air
truss
heat resistant
support
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.)
Ceased
Application number
PCT/US2015/065424
Other languages
English (en)
Inventor
Zachary D DYER
Allister William James
Sachin R. Shinde
Jay A. Morrison
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.)
Siemens AG
Siemens Corp
Siemens Energy Inc
Original Assignee
Siemens AG
Siemens Corp
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 AG, Siemens Corp, Siemens Energy Inc filed Critical Siemens AG
Priority to PCT/US2015/065424 priority Critical patent/WO2017105380A1/fr
Publication of WO2017105380A1 publication Critical patent/WO2017105380A1/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • 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
    • 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
    • F01D5/282—Selecting composite materials, e.g. blades with reinforcing filaments
    • 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
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/50—Building or constructing in particular ways
    • F05D2230/51—Building or constructing in particular ways in a modular way, e.g. using several identical or complementary parts or features
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • 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
    • F05D2260/30—Retaining components in desired mutual position
    • F05D2260/37—Retaining components in desired mutual position by a press fit connection
    • 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
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/17—Alloys
    • F05D2300/175—Superalloys
    • 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
    • F05D2300/603—Composites; e.g. fibre-reinforced
    • F05D2300/6033—Ceramic matrix composites [CMC]

Definitions

  • This invention relates to airfoils, such as vanes or blades, used in a gas turbine, and more particularly, to an airfoil having a high heat resistant outer shell portion and an internally cooled support structure.
  • an axial flow gas turbine 10 includes a multi-stage compressor section 12, a combustion section 14, a multi stage turbine section 16 and an exhaust system 18 arranged along a center axis 20.
  • Air at atmospheric pressure is drawn into the compressor section 12 generally in the direction of the flow arrows F along the axial length of the turbine 10.
  • the intake air is progressively compressed in the compressor section 12 by rows of rotating compressor blades, thereby increasing pressure, and directed by mating compressor vanes to the combustion section 14, where it is mixed with fuel, such as natural gas, and ignited to create a combustion gas.
  • the combustion gas which is under greater pressure, temperature and velocity than the original intake air, is directed to the turbine section 16.
  • the turbine section 16 includes a plurality of airfoil shaped turbine blades 22 arranged in a plurality of rows Ri, R 2 , etc. on a shaft 24 that rotates about the axis 20.
  • the combustion gas expands through the turbine section 16 where it is directed in a combustion flow direction F across the rows of blades 22 by associated rows of stationary vanes 24.
  • a row of blades 22 and associated row of vanes 24 form a stage.
  • the turbine section 16 may include four stages. As the combustion gas passes through the turbine section 16, the combustion gas causes the blades 22 and thus the shaft 24 to rotate about the axis 20, thereby extracting energy from the flow to produce mechanical work.
  • a method for increasing the efficiency of a turbine is to increase an operating temperature of the turbine. Operation of a turbine at higher temperatures frequently results in the use of specialized high heat resistant materials for turbine components such as vanes and/or blades. It is desirable to enhance the structural support provided for the high heat resistant materials used in turbine vanes and/or blades.
  • the airfoil for a turbine, wherein the airfoil includes leading and trailing edges and high and low pressure surfaces.
  • the airfoil includes a core support structure having first and second support sections that include first and second air inlets, respectively.
  • the airfoil also includes at least one first air channel that extends from the first air inlet and through the first support section to an associated air exit hole.
  • the airfoil includes at least one truss connected between the first and second support sections and at least one second air channel that extends from the second air inlet and through the second support section, truss and first support section to an associated air exit hole.
  • the airfoil includes a high heat resistant outer shell, wherein a portion of the core support structure is located within the outer shell.
  • the outer shell may include first and second high heat resistant members.
  • Fig. 1 is a partial cross sectional view of an axial flow gas turbine.
  • Fig. 2 is a perspective view of a hybrid airfoil in accordance with an embodiment of the present invention.
  • FIG. 3 is a perspective view of a core structure of the airfoil.
  • Fig. 4 is a partial cross sectional side view of the airfoil and an exemplary first member.
  • Fig. 5 is a view of an exemplary first modular support structure.
  • FIGs. 6A and 6B are views of exemplary first and second members, respectively.
  • Fig. 7 is an exploded view of exemplary first and second members and the first modular support structure.
  • FIG. 8 A depicts an alternate embodiment of the present invention for a modular support structure.
  • Fig. 8B is a cross sectional view of a truss along view line 8B-8B of Fig. 8 A.
  • FIG. 9 depicts an alternate embodiment of the present invention for a member that mates with the modular support structure shown in Figs. 8 A and 8B.
  • FIG. 2 a perspective view of a hybrid airfoil 30 in accordance with an embodiment of the present invention is shown.
  • the airfoil 30 may be either a vane or blade used in a gas turbine 10.
  • the airfoil 30 includes a high heat resistant outer shell portion 32 and a core structure 34, This arrangement substantially separates and isolates thermal loading from mechanical loading that occurs in the airfoil 30 during operation of the gas turbine 10,
  • the outer shell 32 is subjected to a substantial portion of the thermal loading whereas the core structure 34 is subjected a substantial portion of the mechanical loading.
  • the airfoil 30 includes leading 36 and trailing 38 edges and a first airfoil surface having a substantially concave shaped profile to provide a high-pressure surface 40 and a second airfoil surface having a substantially convex shaped profile to provide a low-pressure surface 42.
  • the outer shell 32 includes at least one high heat resistant first member 44 and at least one high heat resistant second member 45 (see Figs. 6A and 6B) arranged in a stacked configuration around the core structure 34 to form a layered configuration.
  • the first 44 and second 45 members are each fabricated from a high heat resistant material suitable for use in a vane or blade used in a gas turbine 10.
  • the material may be a known ceramic matrix composite (CMC) material.
  • At least one first member 44 or second member 45 may be fabricated from one type of high heat resistant material whereas the remaining members 44, 45 may be fabricated from another type of high heat resistant material.
  • the first 44 and second 45 members and core structure 34 are shaped such that a suitable airfoil shape is formed when the first 44 and second 45 members are stacked or assembled around the core support structure.
  • the first 44 and second 45 members may have equal thicknesses.
  • at least one first member 44 or second member 45 may have a thickness that is greater or less than the thickness of the remaining members 44, 45.
  • a thickness of a first member 44 or second member 45 fabricated from CMC material may be approximately 3 mm.
  • the core structure 34 may be fabricated from an alloy suitable for use in a turbine vane or blade such as INCO EL® alloy available from Special Metals Corporation of New Hartford, NY, US.
  • the core structure 34 includes a first support section 46 having an internal portion 48 located within the outer shell 32.
  • the internal portion 48 includes a first air inlet 50 that extends through the internal portion 48.
  • the first support section 46 also includes an external portion 52 located outside of the outer shell 32 that includes the trailing edge 38.
  • the core structure 34 includes a second support section 54 located within the outer shell 32.
  • the second support section 54 includes a second air inlet 56 that extends through the support section 54.
  • the first 46 and second 54 support sections are connected by at least one truss element 58.
  • at least one truss element 58 For purposes of illustration, four trusses 58 are shown in Fig. 3.
  • an airfoil 30 is subjected to various loads including centrifugal force due to rotation of the airfoil 30 and a gas bending load.
  • use of at least one truss element 58 increases stiffness and load carrying capability of the core structure 34 and thus the airfoil 30.
  • a plurality of air exit holes 60 may be formed on the high-pressure surface 40 of the airfoil 30 near the trailing edge 38. Alternatively, the air exit holes 60 may be formed on the low-pressure surface 42 of the airfoil 30.
  • a plurality of first air channels extend through the internal 48 and external 52 portions between the first air inlet 50 and associated air exit holes 60. Referring to Fig. 4, a partial cross sectional side view of the airfoil 30 and exemplary first member 44 is shown. Fig. 4 depicts an exemplary enclosed first air channel 62 that extends through the internal 48 and external 52 portions between the first air inlet 50 and an associated air exit hole 60.
  • the first air channel 62 is in fluid communication with the first air inlet 50 and the associated air exit hole 60 to form an exemplary first air passageway 64.
  • Each truss 58 includes at least one enclosed second air channel 66 that extends through the truss 58 between the second air inlet 56 and an associated air exit hole 60.
  • the second air channel 66 is in fluid communication with the second air inlet 56 and an associated air exit hole 60 to form a second air passageway 68.
  • the second air channel 66 may feed into the first air channel 62 or the first air inlet 50.
  • cooling air from the gas turbine 10 is received by the first 50 and second 56 air inlets and flows through the first 62 and second 66 air channels, respectively, and is then ultimately emitted from an associated air exit hole 60 in order to cool the airfoil 30.
  • all airflow through the airfoil 30 occurs within the core structure 34 and thus only the core structure 34 is subjected to an internal pressure load due to cooling airflow and substantially little or no load is placed on the outer shell 32 due to cooling airflow.
  • a fastener 47 such as a tie bolt may be used to affix or compress the first 44 and second 45 members together.
  • the fastener 47 may be inserted through the second air inlet 56 and spans the length of the second support section 54.
  • the fastener 47 may block middle portions of the second air inlet 56.
  • an interior wall 49 of the second air inlet 56 includes a plurality of cutout portions 51 located beyond edges 53 of the fastener 47 such that the cutout portions 51 are not blocked by the fastener 47, thus enabling air flow through the second air inlet 56.
  • the first member 44 includes first 70 and second 72 member sections that are connected by an end member section 74.
  • An inner surface 76 of the first 70, second 72 and end 74 member sections contacts the internal portion 48, truss 58 and second support section 54 such that top 78 and bottom 80 sections of the external portion 52 are exposed.
  • the first 70 and second 72 member sections include first 82 and second 84 mating surfaces, respectively, that contact corresponding first 86 and second 88 mating surfaces of the top 78 and bottom 80 sections, respectively.
  • the first mating surfaces 82, 86 taper away from the trailing edge 38 to form a first projection 90 at the inner surface 76 that extends into the top section 78 to form a first mating joint 92 having a first dovetail configuration that captures the first member section 70.
  • the second mating surfaces 84, 88 also taper away from the trailing edge 38 to form a second projection 94 at the inner surface 76 that extends into the bottom section 80 to form a second mating joint 95 having a second dovetail configuration that captures the second member section 72.
  • the first and second dovetail configurations serve to inhibit movement or separation of the first 70 and second 72 member sections away from the core structure 34.
  • the first member section 70 and top section 78 have substantially convex shapes that together form the low-pressure surface 42 and the second member section 72 and bottom section 80 have substantially concave shapes that together form the high-pressure surface 40.
  • the core structure 34 may be divided into a plurality of modular support structures.
  • an exemplary first modular support structure 96 is shown.
  • the first modular structure 96 includes first 98 and second 100 modular support sections that correspond to the first 46 and second 54 support sections, respectively, and at least one truss 58 for connecting the first 98 and second 100 modular support sections.
  • the first 98 and second 100 modular support sections include first 102 and second 104 modular air inlets that form part of the first 50 and second 56 air inlets, respectively, shown in Fig. 3.
  • the first 50 and second 56 modular air inlets are in fluid communication with air exit holes 60 via enclosed first and second air channels as previously described.
  • First 106 and second 108 edge portions extend from a top portion 110 of the first 98 and second 100 modular support sections to form first 112 and second 114 recesses, respectively.
  • First 116 and second 118 lip portions extend from a bottom portion 120 of the first 98 and second 100 modular support sections having a shape corresponding to the first 112 and second 114 recesses, respectively.
  • the first 116 and second 118 lip portions of the first modular structure 96 are inserted inside the first 112 and second 114 recesses, respectively, of the second modular structure 122.
  • the first 116 and second 118 lip portions and first 112 and second 114 recesses are sized such that a press-fit is formed between the lip portions 116, 118 and recesses 112, 114 to affix the second modular structure 122 to the first modular structure 96.
  • the first 96 and second 122 modular structures may be affixed to each other by brazing.
  • shrink fitting could be used.
  • the core structure 34 may include four modular structures although it is understood that additional or fewer modular structures may be used.
  • Use of a modular support structure in accordance with aspects of the present invention enables use of a three dimensional (3D) printing process or a casting process to manufacture the structure, thus minimizing complexity and cost of the airfoil 30 while increasing manufacturing flexibility.
  • an exemplary first member 44 is shown.
  • the first member 44 includes the first 70 and second 72 member sections that are connected by the end section 74.
  • the first 70 and second 72 member sections are spaced apart to form an aperture 124 for receiving the first 46 and second 54 support sections and a gap 126 for receiving a truss 58.
  • an exemplary second member 45 is shown.
  • the second member 45 further includes a bridge member 128 that connects the first 70 and second 72 member sections at a location on the second member 45 corresponding to the location of the truss 58 (see Fig. 3) to form a structure that is more rigid than the first member 44.
  • a plurality of second members 45 is stacked in areas of the core structure 34 that do not include a truss 58. Further, a plurality of first members 44 is stacked in areas of the core structure 34 that include a truss 58. Referring back to Fig. 3, a plurality of second members 45 is stacked in first locations 130A-130E locations of the core structure 34 (i.e. locations that do not include a truss 58). In addition, a plurality of first members 44 is stacked in second locations 132A-132D of the core structure 34 (i.e. areas that include a truss 58).
  • FIG. 7 an exploded view of exemplary first 44 and second 45 members and a first modular support structure 96 is shown. In order to assemble the first
  • a plurality of second members 45 is installed and stacked in a first direction 134 onto the first modular support structure 96 in the first location 130A as previously described.
  • a plurality of first members 44 is then stacked in a second direction 136 opposite the first direction 134 in the second location 132A as previously described.
  • FIG. 8A an alternate embodiment of the present invention for a modular support structure 140 is shown.
  • the modular structure 140 includes the first 98 and second 100 modular support sections as previously described that are connected by a truss 142.
  • FIG. 8B a cross sectional view of the truss 142 along view line 8B-8B is shown.
  • the truss 142 includes first 144 and second 146 raised top edge portions that form a top channel 148 and first 150 and second 152 raised bottom edge portions that form a bottom channel 154 resulting in a substantially I-beam configuration.
  • FIG. 9 an alternate embodiment of the present invention for a member 156 that mates with the modular support structure 140 is shown.
  • the member 156 includes first 158 and second 160 grooves located in bridge member 128 that form a raised surface 162 located between the first 158 and second 160 grooves.
  • first 150 and second 152 raised bottom edge portions are received by the first 158 and second 160 grooves, respectively, and the raised surface 162 is received by the bottom channel 154 to form an interlocking arrangement that provides additional structural support for the modular support structure 140.
  • the member 156 also includes first 164 and second 166 raised surfaces located adjacent the first 158 and second 160 grooves that are of sufficient height so as to cover the truss 142 when the member 156 is assembled to the modular support structure 140.
  • aspects of the current invention result in a substantial reduction in the temperature of the outer shell 32.
  • hot spot temperatures are reduced to approximately 1340 degrees C or less. This enables a substantial reduction in the amount of cooling air needed to cool the airfoil 30.
  • aspects of the current invention also enable operation of a gas turbine at a higher operating temperature, thus increasing turbine efficiency. Further, a thickness of a known thermal barrier coating used on the airfoil 30 may be minimized.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Composite Materials (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne un profil aérodynamique (30) pour une turbine (10), le profil aérodynamique (30) comprend un bord d'attaque (36) et un bord de fuite (38) ainsi que des surfaces haute (40) et basse (42) pression. Le profil aérodynamique (30) comprend une structure de support de partie centrale (34) ayant des première (46) et seconde (54) sections de support qui comportent des premier (50) et second (56) orifices d'entrée d'air, respectivement. Le profil aérodynamique (30) comprend également au moins un premier canal d'air (62) qui s'étend depuis le premier orifice d'entrée d'air (50) et à travers la première section de support (46) jusqu'à un orifice de sortie d'air associé (60). En outre, le profil aérodynamique (30) comprend au moins un treillis (58) raccordé entre les première (46) et seconde (54) sections de support et au moins un second canal d'air (66) qui s'étend depuis le second orifice d'entrée d'air (56) et à travers la seconde section de support (54), le treillis (58) et la première section de support (46) jusqu'à un orifice de sortie d'air associé (60). En outre, le profil aérodynamique (30) comprend une enveloppe externe à haute résistance à la chaleur (32), une partie (48) de la structure de support de partie centrale (30) étant située dans l'enveloppe externe (32). L'enveloppe externe peut comprendre des premier et second éléments à haute résistance à la chaleur.
PCT/US2015/065424 2015-12-14 2015-12-14 Profil aérodynamique de turbine comprenant une structure de support de partie centrale et une enveloppe externe thermorésistante Ceased WO2017105380A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PCT/US2015/065424 WO2017105380A1 (fr) 2015-12-14 2015-12-14 Profil aérodynamique de turbine comprenant une structure de support de partie centrale et une enveloppe externe thermorésistante

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2015/065424 WO2017105380A1 (fr) 2015-12-14 2015-12-14 Profil aérodynamique de turbine comprenant une structure de support de partie centrale et une enveloppe externe thermorésistante

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Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20070128043A1 (en) * 2005-01-21 2007-06-07 Siemens Westinghouse Power Corporation Cmc component and method of fabrication

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20070128043A1 (en) * 2005-01-21 2007-06-07 Siemens Westinghouse Power Corporation Cmc component and method of fabrication

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