EP1728969A2 - Gekühlte Turbomaschinenschaufel - Google Patents

Gekühlte Turbomaschinenschaufel Download PDF

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Publication number
EP1728969A2
EP1728969A2 EP06252658A EP06252658A EP1728969A2 EP 1728969 A2 EP1728969 A2 EP 1728969A2 EP 06252658 A EP06252658 A EP 06252658A EP 06252658 A EP06252658 A EP 06252658A EP 1728969 A2 EP1728969 A2 EP 1728969A2
Authority
EP
European Patent Office
Prior art keywords
leg
airfoil
centrifugal
intake
passage
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.)
Withdrawn
Application number
EP06252658A
Other languages
English (en)
French (fr)
Inventor
Atul Kohli
Edward F. Pietraszkiewicz
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.)
Raytheon Technologies Corp
Original Assignee
United Technologies Corp
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 United Technologies Corp filed Critical United Technologies Corp
Publication of EP1728969A2 publication Critical patent/EP1728969A2/de
Withdrawn legal-status Critical Current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
    • F01D5/14—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

Definitions

  • This invention relates to cooled airfoils of the type used in turbine engines and particularly to a cooled airfoil with reduced turning losses in an internal cooling passage of the airfoil.
  • Turbine engines include one or more turbines for extracting energy from a stream of hot working medium gases.
  • a typical turbine includes a rotatable hub with a set of circumferentially distributed blades projecting radially from the hub. Each blade includes an attachment for attaching the blade to the hub. Each blade also includes an airfoil that spans radially across a working medium flowpath from an airfoil root to an airfoil tip.
  • a typical turbine also includes one or more arrays of stationary vanes axially spaced from the blades. Each vane includes an airfoil that spans radially across the flowpath and a hook or other feature for securing the vane to a case.
  • the blades and vanes operate in a hot environment, it is common practice to provide internal coolant passages in at least the airfoils of the blades and vanes.
  • coolant flows through the internal passages to protect the airfoils from the intense heat of the combustion gases.
  • the coolant is usually relatively cool air that has been pressurized by a compressor powered by the turbine.
  • a multi-pass passage includes at least two spanwisely extending legs that are chordwisely adjacent to each other.
  • a spanwisely extending rib separates the legs from each other.
  • An elbow at the radially inner or outer ends of the legs wraps around one extremity of the rib to connect the legs in series.
  • a stream of coolant flows through one of the legs (the upstream leg), through the elbow and then through the other leg (the downstream leg).
  • the elbow reverses the direction of coolant flow, for example from radially outwardly in the upstream leg to radially inwardly in the downstream leg.
  • the coolant stream entering the downstream leg is susceptible to separation from the rib.
  • the region of the leg susceptible to fluid separation extends chordwisely a considerable distance across the downstream leg and is characterized by high aerodynamic losses. These losses can imperil the durability of the airfoil by restricting coolant flow and/or by reducing the pressure of the coolant downstream of the region of separation.
  • An engine designer can attempt to compensate for these effects by supplying higher pressure coolant to the passages. However such an approach may not be completely successful.
  • the turbine itself is the source of energy for pressurizing the coolant, the use of higher pressure coolant degrades engine efficiency.
  • an airfoil has an internal fluid passage that includes upstream and downstream legs, such as a co-centrifugal leg and a counter-centrifugal leg.
  • the legs are chordwisely separated from each other by a rib but are connected in series with each other.
  • the airfoil also includes a vent passage for venting fluid from the internal passage. The intake to the vent passage resides in the counter-centrifugal leg.
  • FIGS. 1 and 2 show a cooled turbine blade 10 for the turbine of a gas turbine engine.
  • the blade includes an attachment 12 for securing the blade to a hub, not shown.
  • the hub is rotatable about an engine centerline or axis 14.
  • the blade also includes a platform 16 and an airfoil 18.
  • the airfoil spans radially across a working medium flowpath 20 from an airfoil root 24 to an airfoil tip 26.
  • a notional chord line 28 (FIG. 2) extends from a leading edge 30 to a trailing edge 32 of the airfoil.
  • Internal coolant passages such as multi-pass serpentine passage 34, convey coolant 38 through the airfoil. The coolant protects the airfoil from the intense heat of combustion gases G flowing axially through the flowpath 20.
  • the passage 34 includes a spanwisely extending upstream leg 40 and a spanwisely extending downstream leg 42 chordwisely separated from the upstream leg by a spanwisely extending rib 44.
  • the rib is truncated to accommodate an elbow 46 that connects the upstream and downstream legs in series flow relationship.
  • the upstream leg 40 is a co-centrifugal leg because the rotation of the blade about axis 14 assists the flow of coolant from root end of the leg toward the tip end of the leg.
  • the downstream leg 42 is a counter-centrifugal leg because the rotation of the blade about axis 14 resists the flow of coolant from tip end of the leg toward the root.
  • the downstream leg 42 has a chordwise width W, as measured from the rib 44 to a chordwisely neighboring rib 48 that defines the opposing sidewall of the leg.
  • a region 54 susceptible to fluid separation is present on the inside of the turn next to the rib 44.
  • the chordwise dimension of the separation susceptible region increases with increasing lengthwise distance along the passage, and exhibits its maximum chordwise dimension at spanwise location M.
  • the maximum chordwise dimension of the separation region is about 50% of the chordwise width W.
  • the region 54 then diminishes in chordwise dimension with additional lengthwise distance along the passage.
  • the overall spanwise dimension of the illustrated separation region 54 is about two and one half to three hydraulic diameters. This is a typical spanwise dimension for the separation region, however the spanwise dimension can vary depending on the geometry of the passage leg and the fluid properties of the coolant.
  • a vent passage 56 penetrates the suction sidewall 58 of the airfoil.
  • the vent passage has an intake 60 residing in the region susceptible to separation and an exit 62 on the suction wall.
  • the vent passage at least partially counteracts the separation potential of region 54 by allowing some of the coolant to vent from the passage leg 42.
  • the vent passage will be most effective if its inlet 60 resides immediately adjacent to rib 44 as seen best in FIG. 3.
  • a vent intake chordwisely spaced from the rib by as much as about half the maximum chordwise width of the unmoderated region 54 (FIG. 4) would also be quite effective. Because the widest portion of the unmoderated separation zone (seen in FIG.
  • vent intake occupies about 50% of the local passage width W, such a vent intake would be spaced from rib 44 by about 25% of the local passage width W.
  • the vent intake should also be within three and one half hydraulic diameters from the inlet, more typically about two and one half to three hydraulic diameters from the inlet in order that the vent intake is within the separation region.
  • the intake is at the lengthwise location M, where the chordwise dimension of the unmoderated separation susceptible region is widest.
  • the vent passage may be a single passage or it may be an array of passages, one example of which is the single linear array, seen in FIG. 3.
  • the above described principles for positioning the passage intakes apply equally to a single passage or to an array of passages.
  • the row is centered at spanwise location M where the unmoderated separation susceptible region has its maximum chordwise dimension.
  • vent passages may be installed by any suitable technique, for example laser drilling, electron beam drilling or electro-discharge machining. Since turbine blades are usually cast, the passages may also be provided for in the casting itself.
  • One possible advantage to cast passages is the relative ease with which they may be precisely and repeatably positioned in a group of serially produced airfoils.
  • vent passage is a film cooling hole that exhausts some of the coolant 38 to the surface of the suction wall 58 where it spreads out to form a thermally protective cooling film on the wall surface.
  • a film cooling hole that vents coolant to the surface of the pressure wall 64 would also be effective, provided the pressure difference across the passage is large enough to drive the coolant through the passage.
  • film cooling holes venting to both the suction and pressure sides would also be effective. Vent passages that do not also serve as film cooling holes will also suffice to moderate the region of separation.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP06252658A 2005-05-31 2006-05-22 Gekühlte Turbomaschinenschaufel Withdrawn EP1728969A2 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US11/140,851 US20070009358A1 (en) 2005-05-31 2005-05-31 Cooled airfoil with reduced internal turn losses

Publications (1)

Publication Number Publication Date
EP1728969A2 true EP1728969A2 (de) 2006-12-06

Family

ID=36675905

Family Applications (1)

Application Number Title Priority Date Filing Date
EP06252658A Withdrawn EP1728969A2 (de) 2005-05-31 2006-05-22 Gekühlte Turbomaschinenschaufel

Country Status (5)

Country Link
US (1) US20070009358A1 (de)
EP (1) EP1728969A2 (de)
JP (1) JP2006336651A (de)
AU (1) AU2006202304A1 (de)
RU (1) RU2006118026A (de)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20140219813A1 (en) * 2012-09-14 2014-08-07 Rafael A. Perez Gas turbine engine serpentine cooling passage
US9797258B2 (en) * 2013-10-23 2017-10-24 General Electric Company Turbine bucket including cooling passage with turn
US9726023B2 (en) * 2015-01-26 2017-08-08 United Technologies Corporation Airfoil support and cooling scheme

Family Cites Families (25)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3628885A (en) * 1969-10-01 1971-12-21 Gen Electric Fluid-cooled airfoil
US4162136A (en) * 1974-04-05 1979-07-24 Rolls-Royce Limited Cooled blade for a gas turbine engine
US4353679A (en) * 1976-07-29 1982-10-12 General Electric Company Fluid-cooled element
FR2468727A1 (fr) * 1979-10-26 1981-05-08 Snecma Perfectionnement aux aubes de turbine refroidies
US4775296A (en) * 1981-12-28 1988-10-04 United Technologies Corporation Coolable airfoil for a rotary machine
JPS62228603A (ja) * 1986-03-31 1987-10-07 Toshiba Corp ガスタ−ビンの翼
US4753575A (en) * 1987-08-06 1988-06-28 United Technologies Corporation Airfoil with nested cooling channels
US4767268A (en) * 1987-08-06 1988-08-30 United Technologies Corporation Triple pass cooled airfoil
US4930980A (en) * 1989-02-15 1990-06-05 Westinghouse Electric Corp. Cooled turbine vane
JP3666602B2 (ja) * 1992-11-24 2005-06-29 ユナイテッド・テクノロジーズ・コーポレイション 冷却可能なエアフォイル構造
US5403159A (en) * 1992-11-30 1995-04-04 United Technoligies Corporation Coolable airfoil structure
WO1995014848A1 (en) * 1993-11-24 1995-06-01 United Technologies Corporation Cooled turbine airfoil
US5387085A (en) * 1994-01-07 1995-02-07 General Electric Company Turbine blade composite cooling circuit
US5498126A (en) * 1994-04-28 1996-03-12 United Technologies Corporation Airfoil with dual source cooling
US5538393A (en) * 1995-01-31 1996-07-23 United Technologies Corporation Turbine shroud segment with serpentine cooling channels having a bend passage
US5669759A (en) * 1995-02-03 1997-09-23 United Technologies Corporation Turbine airfoil with enhanced cooling
US5645397A (en) * 1995-10-10 1997-07-08 United Technologies Corporation Turbine vane assembly with multiple passage cooled vanes
JP3411775B2 (ja) * 1997-03-10 2003-06-03 三菱重工業株式会社 ガスタービン動翼
US5827043A (en) * 1997-06-27 1998-10-27 United Technologies Corporation Coolable airfoil
US5902093A (en) * 1997-08-22 1999-05-11 General Electric Company Crack arresting rotor blade
US6004100A (en) * 1997-11-13 1999-12-21 United Technologies Corporation Trailing edge cooling apparatus for a gas turbine airfoil
US6474947B1 (en) * 1998-03-13 2002-11-05 Mitsubishi Heavy Industries, Ltd. Film cooling hole construction in gas turbine moving-vanes
DE19921644B4 (de) * 1999-05-10 2012-01-05 Alstom Kühlbare Schaufel für eine Gasturbine
US6241467B1 (en) * 1999-08-02 2001-06-05 United Technologies Corporation Stator vane for a rotary machine
US6595748B2 (en) * 2001-08-02 2003-07-22 General Electric Company Trichannel airfoil leading edge cooling

Also Published As

Publication number Publication date
RU2006118026A (ru) 2007-12-10
JP2006336651A (ja) 2006-12-14
US20070009358A1 (en) 2007-01-11
AU2006202304A1 (en) 2006-12-14

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