US3373970A - Gas turbine blade - Google Patents

Gas turbine blade Download PDF

Info

Publication number
US3373970A
US3373970A US600540A US60054066A US3373970A US 3373970 A US3373970 A US 3373970A US 600540 A US600540 A US 600540A US 60054066 A US60054066 A US 60054066A US 3373970 A US3373970 A US 3373970A
Authority
US
United States
Prior art keywords
blade
insert body
blade member
cooling
cooling air
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.)
Expired - Lifetime
Application number
US600540A
Other languages
English (en)
Inventor
Brockmann Heinz
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.)
Daimler Benz AG
Original Assignee
Daimler Benz AG
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 Daimler Benz AG filed Critical Daimler Benz AG
Application granted granted Critical
Publication of US3373970A publication Critical patent/US3373970A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present disclosure relates to a blade for a gas turbine, particularly a turbine to be used in a flight propulsion unit.
  • the blade is constructed of inner and outer substantially identically contoured hollow blade members, which are connected together by cooling fins or ribs that extend substantially in the plane of propulsion gas flow.
  • a cooling fluid is introduced into the inner hollow blade and is discharged through radially elongated slots in the forward edge of the inner blade against the interior surface of the outer hollow blade forward edge.
  • the cooling fluid will travel from the forward edge between the inner and outer hollow blade members rearwardly where it is discharged through radially elongated slots in the rearward edge of the outer hollow member smoothly into the flow of propulsion gases so that it will not disturb the flow of propulsion gases and enhance the aerodynamic effect of the turbine blade.
  • the blade construction may be employed in either a rotating blade or a stationary guide vane.
  • the present invention avoids the above-mentioned disadvantages while obtaining a good heat transfer within the blade.
  • the blade of the present invention comprises inner and outer hollow substantially identically contoured blade members that are connected together by cooling ribs or fins extending in the plane of working fluid flow. Cooling gases are conducted from the interior of the inner hollow blade through conduits at its forward edge against the inner surface of the outer hollow member forward edge. From this point, the cooling gases extend along the fins rearwardly and pass through ducts in the rearward edge of the outer hollow blade member.
  • the cooling gases By discharging the cooling air through the rearward edge of the outer blade member, the cooling gases enter the working fluid without producing any eddies, turbulence or other friction loss producing disturbances. In fact, the flow pattern of the working gases is improved, because any vertexes that might form in the area immediately rearward of the rearward edge are eliminated by the discharge of fluid into this area.
  • the inner hollow blade or insert body may be formed in any way according to the broader aspects of the present invention, for example, by means of a small separate 3,373,970. Patented Mar. 19, 1968 ice tube inserted into the inside of the outer hollow blade.
  • a more narrow aspect of the present invention employs an inner hollow blade member or insert body that matches the internal contour of the outer hollow blade and is mounted on or supported by the cooling fins or ribs that are integral with the outer hollow blade member.
  • the strength of the overall blade structure is enhanced by rigidly connecting the cooling fins between and to the outer and inner hollow blade members.
  • Many prior art structures have required separate reinforcement struts, which are not required with the preferred construction of the present invention.
  • the cooling air is forced to flow in uniform narrow passages closely adjacent to the interior walls of the outer hollow blade member so that the heat exchange efliciency is improved and a relatively small quantity of cooling fluid is required.
  • the arrangement of the cooling fins according to the present invention not only increases the heat exchange surface area, but they also decrease the cross-section of the cooling air flow channel to further increase the heat exchange efficiency.
  • the present invention recognizes the fact that the forward edge of the outer blade member is heated to the greatest extent; by discharging the cooling air against the interior surface of the outer blade member forward edge, the cooling air first contacts the hottest portion of the outer blade member so that the blade member is more uniformly cooled. Even though this is a very considerable advantage over the preferred construction, the cooling air may be discharged from the insert body or inner hollow blade member at any arbitrary point according to the broader aspects of the present invention.
  • the apertures in the insert body and in the rear edge of the outer hollow member may be shaped in any way.
  • the conduits in the forward edge of the insert body extend the entire height or radial dimension of the insert body and that the ducts in the rearward edge of the outer hollow blade member also extend for the entire height or radial dimension of the outer blade.
  • the conduits and ducts may be relatively narrow to conform to the relatively small transverse dimension of the forward and rearward edges while providing a suflicient cross-sectional area to assure proper cooling fluid flow.
  • the flow cross-sectional area may be varied along the radial dimension to match the temperature distribution, that is, a greater cooling fluid flow may be provided at the areas of highest temperature while a smaller cooling fluid flow may be provided at an area of lower temperature that is radially spaced from the higher temperature area.
  • a very precise control of the radial fluid flow distribution may be obtained by a plurality of radially spaced transverse slots that each extend in the plane of the cooling fins or working fluid flow; the size of the slots may vary in the radial direction or their radial spacing may vary in the radial direction to obtain the proper fluid flow distribution and to match the temperature distribution. These slots may be easily and automatically produced by rotating cutters in a. separate insert body.
  • the manufacture of the blade may be accomplished in any manner.
  • each of the components may be assembled individually. However, this would be quite costly and complicated.
  • the entire blade structure is manufactured in a single casting. With this preferred construction, the assembly time is greatly reduced and the blade strength is greatly increased. For casting the blade, the so-called wax melting process is especially suitable.
  • the insert body is integrally cast in the one piece structure. However, the insert body may preferably be made separately of thin sheet metal when very small blades are being manufactured.
  • FIGURE 1 is a cross-sectional view of a preferred embodiment of the present invention taken on a plane parallel to the flow of working fluid;
  • FIGURE 2 is a cross-sectional view taken along line IIII of FIGURE 1;
  • FIGURE 3 is a cross-sectional view taken along line III-III of FIGURE 1;
  • FIGURE 4 is an elevation view of the rear edge of the blade taken in the direction of arrow A in FIGURE 1'.
  • the preferred embodiment of the present invention includes an outer hollow blade member integrally cast in one piece with a plurality of reinforcing and cooling ribs or fins 11 that are also integrally connected at their inside portions with an insert body or inner hollow blade member 12.
  • the fins 11 extend in the plane of FIGURE 1, that is, in the plane of the working fluid and perpendicular to the radial axis of the blade.
  • the profile of the insert body 12 is substantially identically contoured with respect to the outer hollow blade 10.
  • the fins 11 integrally extend between the blade member 10 and the insert body 12 to reinforce them so that additional struts are unnecessary and a rigid strong blade structure is obtained.
  • the insert body 12 forms a plenum or distribution chamber for the cooling air within its hollow interior. As shown in FIGURE 2, the cooling air enters the blade head or radially outward portion of the insert body 12 through an opening 13.
  • the opening 13 would most advantageously be placed in the foot or radially inward portion 14 of the blade to take advantage of the centrifugal forces.
  • a stationary blade has been shown for purposes of illustration and the opening in the foot portion 14 of the insert body 12 is closed by means of a plate 15.
  • the insert body 12 is provided with two relatively narrow radially elongated slots 16, 17 in its forward edge for discharging cooling air against the inside surface 18 of the outer hollow blade 10 forward edge portion.
  • the slots 16 and 17 provide the only discharge opening in the insert body 12.
  • the cooling effect of the fresh cooling air is especially intensive at this highly thermally stressed area 18 of the outer hollow blade 10. From the area 13, the cooling air flows rearwardly around both sides of the insert body, FIGURE 1, along the inner wall portions of the blade member 10 between the cooling fins 11 to the rear edge 19 of the blade member 10.
  • the slots 16 and 17 constitute conduits for discharging the cooling air from the insert body 12.
  • the walls of the insert body 12 and blade member 10 form, along with the cooling fins 11, relatively small cross-sectional area passageways for conducting the cooling air rearwardly from area 18 to area 19. With these small passageways, a relatively small amount of cooling air will be conducted in efficient heat exchange relationship with a relatively large area of the blade member 10 and cooling fins 11. Thus, a very efficient heat exchange relationship is obtained.
  • the thus heated cooling air is discharged from the rear edge 19 in the direction of the working or propulsion gases flowing along the outer surface of the blade member 10 by means of three relatively narrow radially elongated slots or ducts 20, 21, 22, as shown in 4 FIGURES 1 and 4.
  • the cooling air that is discharged from the slots 20, 21, 22 mixes with the working gases without producing and disturbance and improves the flow of the working gases within the region of the rearward edge of the blade structure.
  • the cooling fins are integral with the outer blade member 10 and spaced from the forward edge of the insert body 12.
  • the blade member 10 and the insert body 12 have been shown with different cross-hatching to more clearly define their extent and shape; however, it appears to be understood that the preferred embodiment of the present invention relates to an integral one-piece single cast construction for the blade member 10, the cooling fins 11 and the insert body 12.
  • these elements may be separate according to the broader aspects of the present invention.
  • the cooling fins 11 are preferably rigidly attached to both the blade member 10 and the insert body 12 along their sides as was clearly shown in FIGURE 2 to increase the rigidity of the blade structure.
  • a turbine blade with an outer hollow blade member having an axially forward edge portion and an axially rearward edge portion, hollow insert body means within said blade member for feeding cooling air into the inside of said blade member; said blade member rearward portion having duct means for axially discharging the cooling air generally in the same direction as the driving gases passing said blade member, in combination with the improvement comprising: said insert body means having conduit means for feeding air into said blade member only at its forward edge portions; said blade member having a plurality of integral cooling fins on its inside surface that are generally within flow planes of the driving gases.
  • said insert body means generally matches the internal contours of said blade member and has an outer surface generally uniformly spaced from the adjacent inside surface of said blade member; said insert body means being mounted on and engaging said cooling fins.
  • said duct means consists of a plurality of relatively narrow radially elongated slots in the rearward edge of said blade member; and said conduit means include a plurality of relatively narrow radially elongated slots in the forward edge of said insert body means.
  • said duct means consists of a plurality of relatively narrow radially elongated slots in the rearward edge of said blade member; and said conduit means include a plurality of relatively narrow radially elongated slots in the forward edge of said insert body means.
  • said duct means consists of a plurality of relatively narrow radially elongated slots in the rearward edge of said blade member; and said conduit means include a plurality of relatively narrow radially elongated slots in the forward edge of said insert body means.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US600540A 1965-12-11 1966-12-09 Gas turbine blade Expired - Lifetime US3373970A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DED0048895 1965-12-11

Publications (1)

Publication Number Publication Date
US3373970A true US3373970A (en) 1968-03-19

Family

ID=7051492

Family Applications (1)

Application Number Title Priority Date Filing Date
US600540A Expired - Lifetime US3373970A (en) 1965-12-11 1966-12-09 Gas turbine blade

Country Status (3)

Country Link
US (1) US3373970A (fr)
FR (1) FR1503348A (fr)
GB (1) GB1097300A (fr)

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3574481A (en) * 1968-05-09 1971-04-13 James A Pyne Jr Variable area cooled airfoil construction for gas turbines
US3628880A (en) * 1969-12-01 1971-12-21 Gen Electric Vane assembly and temperature control arrangement
US3809494A (en) * 1971-06-30 1974-05-07 Rolls Royce 1971 Ltd Vane or blade for a gas turbine engine
US4019831A (en) * 1974-09-05 1977-04-26 Brown Boveri Sulzer Turbomachinery Ltd. Cooled rotor blade for a gas turbine
EP0154893A1 (fr) * 1984-03-13 1985-09-18 Kabushiki Kaisha Toshiba Aube pour une turbine à gaz
US5203873A (en) * 1991-08-29 1993-04-20 General Electric Company Turbine blade impingement baffle
EP1277918A1 (fr) * 2001-07-18 2003-01-22 FIATAVIO S.p.A. Aube de guidage à doubles parois pour une turbine à gaz
US20040096313A1 (en) * 2002-11-12 2004-05-20 Harvey Neil W. Turbine components
US20050220626A1 (en) * 2002-03-27 2005-10-06 Christopher Gray Impingement cooling of gas turbine blades or vanes
US20060120870A1 (en) * 2004-12-02 2006-06-08 Ricardo Trindade Internally cooled airfoil for a gas turbine engine and method
US20120163994A1 (en) * 2010-12-28 2012-06-28 Okey Kwon Gas turbine engine and airfoil
WO2015181497A1 (fr) * 2014-05-28 2015-12-03 Snecma Aube de turbine a refroidissement optimise
WO2015181482A1 (fr) * 2014-05-28 2015-12-03 Snecma Aube de turbine comprenant un conduit central de refroidissement et deux cavites laterales jointives en aval du conduit central

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2476207A1 (fr) * 1980-02-19 1981-08-21 Snecma Perfectionnement aux aubes de turbines refroidies
US4623087A (en) * 1983-05-26 1986-11-18 Rolls-Royce Limited Application of coatings to articles
JP2609805B2 (ja) * 1994-02-07 1997-05-14 株式会社東芝 ガスタービン翼

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2514105A (en) * 1945-12-07 1950-07-04 Thomas Wilfred Airfoil conditioning means
DE767546C (de) * 1938-09-12 1952-11-04 Bmw Flugmotorenbau G M B H Innengekuehlte Turbinenschaufel
US2858100A (en) * 1952-02-01 1958-10-28 Stalker Dev Company Blade structure for turbines and the like
US2879028A (en) * 1954-03-31 1959-03-24 Edward A Stalker Cooled turbine blades
US2920866A (en) * 1954-12-20 1960-01-12 A V Roe Canada Ltd Hollow air cooled sheet metal turbine blade
US3032314A (en) * 1957-05-28 1962-05-01 Snecma Method of and device for cooling the component elements of machines

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE767546C (de) * 1938-09-12 1952-11-04 Bmw Flugmotorenbau G M B H Innengekuehlte Turbinenschaufel
US2514105A (en) * 1945-12-07 1950-07-04 Thomas Wilfred Airfoil conditioning means
US2858100A (en) * 1952-02-01 1958-10-28 Stalker Dev Company Blade structure for turbines and the like
US2879028A (en) * 1954-03-31 1959-03-24 Edward A Stalker Cooled turbine blades
US2920866A (en) * 1954-12-20 1960-01-12 A V Roe Canada Ltd Hollow air cooled sheet metal turbine blade
US3032314A (en) * 1957-05-28 1962-05-01 Snecma Method of and device for cooling the component elements of machines

Cited By (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3574481A (en) * 1968-05-09 1971-04-13 James A Pyne Jr Variable area cooled airfoil construction for gas turbines
US3628880A (en) * 1969-12-01 1971-12-21 Gen Electric Vane assembly and temperature control arrangement
US3809494A (en) * 1971-06-30 1974-05-07 Rolls Royce 1971 Ltd Vane or blade for a gas turbine engine
US4019831A (en) * 1974-09-05 1977-04-26 Brown Boveri Sulzer Turbomachinery Ltd. Cooled rotor blade for a gas turbine
EP0154893A1 (fr) * 1984-03-13 1985-09-18 Kabushiki Kaisha Toshiba Aube pour une turbine à gaz
US5203873A (en) * 1991-08-29 1993-04-20 General Electric Company Turbine blade impingement baffle
EP1277918A1 (fr) * 2001-07-18 2003-01-22 FIATAVIO S.p.A. Aube de guidage à doubles parois pour une turbine à gaz
US20030017051A1 (en) * 2001-07-18 2003-01-23 Fiatavio S.P.A. Double-wall blade for a turbine, particularly for aeronautical applications
US7056083B2 (en) * 2002-03-27 2006-06-06 Alstom (Switzerland) Ltd Impingement cooling of gas turbine blades or vanes
US20050220626A1 (en) * 2002-03-27 2005-10-06 Christopher Gray Impingement cooling of gas turbine blades or vanes
US7137781B2 (en) * 2002-11-12 2006-11-21 Rolls-Royce Plc Turbine components
US20040096313A1 (en) * 2002-11-12 2004-05-20 Harvey Neil W. Turbine components
US20060120870A1 (en) * 2004-12-02 2006-06-08 Ricardo Trindade Internally cooled airfoil for a gas turbine engine and method
US7258528B2 (en) * 2004-12-02 2007-08-21 Pratt & Whitney Canada Corp. Internally cooled airfoil for a gas turbine engine and method
EP2472062B1 (fr) 2010-12-28 2017-02-15 Rolls-Royce North American Technologies, Inc. Moteur de turbine à gaz et l'aube
US8961133B2 (en) * 2010-12-28 2015-02-24 Rolls-Royce North American Technologies, Inc. Gas turbine engine and cooled airfoil
US20120163994A1 (en) * 2010-12-28 2012-06-28 Okey Kwon Gas turbine engine and airfoil
WO2015181497A1 (fr) * 2014-05-28 2015-12-03 Snecma Aube de turbine a refroidissement optimise
WO2015181482A1 (fr) * 2014-05-28 2015-12-03 Snecma Aube de turbine comprenant un conduit central de refroidissement et deux cavites laterales jointives en aval du conduit central
FR3021698A1 (fr) * 2014-05-28 2015-12-04 Snecma Aube de turbine, comprenant un conduit central de refroidissement isole thermiquement de parois de l'aube par deux cavites laterales jointives en aval du conduit central
FR3021697A1 (fr) * 2014-05-28 2015-12-04 Snecma Aube de turbine a refroidissement optimise
US20170183969A1 (en) * 2014-05-28 2017-06-29 Safran Aircraft Engines Turbine blade with optimised cooling
US10337333B2 (en) 2014-05-28 2019-07-02 Safran Aircraft Engines Turbine blade comprising a central cooling duct and two side cavities connected downstream from the central duct
US10689985B2 (en) * 2014-05-28 2020-06-23 Safran Aircraft Engines Turbine blade with optimised cooling

Also Published As

Publication number Publication date
GB1097300A (en) 1968-01-03
FR1503348A (fr) 1967-11-24

Similar Documents

Publication Publication Date Title
US3373970A (en) Gas turbine blade
US3574481A (en) Variable area cooled airfoil construction for gas turbines
JP3330503B2 (ja) タービンの分配器用羽根
US4021139A (en) Gas turbine guide vane
US3540810A (en) Slanted partition for hollow airfoil vane insert
US3806276A (en) Cooled turbine blade
US3782852A (en) Gas turbine engine blades
RU2697211C2 (ru) Лопатка турбины с оптимизированным охлаждением
US5752801A (en) Apparatus for cooling a gas turbine airfoil and method of making same
JP2862536B2 (ja) ガスタービンの翼
JP5349503B2 (ja) 端部の冷却を備えているタービン翼ならびに関連のタービンおよびターボエンジン
EP1444418B1 (fr) Aube ou pale de turbine a gaz refroidi interne
US5288207A (en) Internally cooled turbine airfoil
US5720431A (en) Cooled blades for a gas turbine engine
US3301526A (en) Stacked-wafer turbine vane or blade
JP6132546B2 (ja) タービンロータブレードのプラットフォームの冷却
US3885609A (en) Cooled rotor blade for a gas turbine
RU2563046C2 (ru) Модульная лопатка или лопасть для газовой турбины и газовая турбина с такой лопаткой или лопастью
US8870537B2 (en) Near-wall serpentine cooled turbine airfoil
JPH0681675A (ja) ガスタービン及びガスタービンの段落装置
US9388699B2 (en) Crossover cooled airfoil trailing edge
US5813827A (en) Apparatus for cooling a gas turbine airfoil
JP2002004803A (ja) 翼形部の後部空洞用の蒸気出口流設計
US6146098A (en) Tip shroud for cooled blade of gas turbine
JP2010502872A (ja) 冷却形タービン動翼