US3846041A - Impingement cooled turbine blades and method of making same - Google Patents

Impingement cooled turbine blades and method of making same Download PDF

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Publication number
US3846041A
US3846041A US00302421A US30242172A US3846041A US 3846041 A US3846041 A US 3846041A US 00302421 A US00302421 A US 00302421A US 30242172 A US30242172 A US 30242172A US 3846041 A US3846041 A US 3846041A
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Prior art keywords
insert
chamber
tang
airfoil
passageway
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US00302421A
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P Albani
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Avco Corp
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Avco Corp
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Priority to US00302421A priority Critical patent/US3846041A/en
Priority to GB5024273A priority patent/GB1441969A/en
Priority to SE7314736A priority patent/SE389162B/en
Priority to DE19732354693 priority patent/DE2354693C3/en
Priority to FR7338953A priority patent/FR2205097A5/fr
Priority to JP48121871A priority patent/JPS5249085B2/ja
Priority to IT53459/73A priority patent/IT997587B/en
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
    • B23P15/00—Making specific metal objects by operations not covered by a single other subclass or a group in this subclass
    • B23P15/04—Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from several pieces
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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

  • Air is introduced, through the tang, into the plenum References Cited formed by the insert and then discharged through ori- UNITED STATES PATENTS fices to impinge against the chamber walls and cool 2.779565 1 1957 Bruckmann 416/96 the airfoil
  • the insert is assembled y introducing it 2,873,944 2/1959 Wiese et al. 416/92 through the tip end openin and an opening in the 3,446,480 5/1969 Emmerson et at... 416/90 lower end of the chamber.
  • the inner end of the insert 5 1/1972 G m t a 9 9 v 7 is brought into register with openings in the tang for Corsmerer et al.
  • a more recent innovation in air cooling of blades has been the so called impingement cooling of blades.
  • a central plenum is provided from which air is discharged from relatively small holes, or orifices, to impinge against selected portions of the interior of the blade and thus provide a highly effective cooling mechanism.
  • the blade is formed as a hollow shell and the plenum is formed by a separate sheet metal insert which communicates, at its inner end, with a source of pressurized cooling air.
  • the blades of a turbine are generally formed as separate elements which are provided with tangs at their inner ends for attachment to the turbine rotor. This not only facilitates manufacture but also facililates maintenance and overhaul.
  • the blades are mounted on the turbine rotor by means of a dovetail slot arrangement or pins so that they can be separately replaced.
  • turbine blade comprising an intergrally formed airfoil and tang.
  • the airfoil has thin walls which define a chamber extending from its tip end to its inner end.
  • An insert is disposed in this chamber and it extends into the blade tang with its inner end mechanically locked thereto.
  • the insert forms a plenum sealed from the airfoil chamber.
  • Passageway means extend from the exterior of the tang into the insert for connecting the plenum with a source of pressurized cooling air.
  • This cooling air is discharged from a plurality of orifices directed toward the chamber wall to impinge there against and maintain the metal temperature of the airfoil at an acceptable level. The cooling air is then discharged from the chamber.
  • the inner end of the insert is locked to the tang by means of a pin extending therethrough.
  • the inner end of the insert, through which the pin passes is thickened or solid.
  • the passageway means for introducing cooling air into the insert may include a passageway through the upper portion of the tang, extending from one end thereof, and an opening in the lower portion of the insert above its thickened portion.
  • a lug may be formed in the tang passageway and be slotted to receive the lower end of the insert and its retaining pin.
  • the pin may terminate at the face of the lug within the passageway and an opening may be provided in the opposite face or opposite side of the tang to facilitate riveting of the pin.
  • the insert may be formed as a thin walled shell, at its outer end, having an end cap with 'an outwardly projecting, peripheral flange.
  • This flange may be secured to the wall of the insert shell and then flared outwardly into an undercut peripheral groove around the tip end opening in the airfoil to positively retain the cap itself as well as to provide for sealing of the chamber from leakage through its tip end.
  • Another feature of the invention is found in the provision of integrally formed pins projecting from the airfoil into the chamber to engage and support the insert therein.
  • Aligned openings in the tip end of the airfoil and at the inner end of the chamber may be provided to permit insertion of the insert into the blade. After such insertion, the pin may be introduced through aligned openings in the tang and insert to mechanically lock the insert.
  • FIG. 1 a portion of a turbine rotor is shown.
  • a plurality of turbine blades 12 (one is shown) are mounted, in the usual configuration,
  • the blades are attached to the rotor by dovetail tangs 14 (FIG. 2) which are received by correspondingly shaped grooves in the rotor.
  • the blades are axially held in place on the rotor 10 by retaining rings, or discs, 16 and 18.
  • the disc 18 is spaced from the rotor 10 and defines apassageway 20 for pressurized cooling air which is introduced centrally from the engines compressor. This cooling air enters the blade through a tang passageway in a manner later described to provide cooling of the blade in accordance with the present invention.
  • the blade comprises an aerodynamically shaped airfoil section or portion 22 which projects outwardly from a platform 24.
  • the tang 14 then extends inwardly of the platform 24 which defines the inner bonds of the hot gas flowpath through the engine.
  • the tang 14, as previously referenced, provides for attachment of the blade to the turbine rotor 10.
  • the airfoil 22 is in the form of a hollow shell, defining a chamber 23 at its relatively thick and blunt leading edge portion.
  • the airfoil 22 in cross section, tapers down to a relatively narrow trailing end section.
  • An insert 26 is disposed within the chamber 23 which extends from the tip end of the airfoil to below the platform 24.
  • the insert 26 comprises a thin walled shell which is closely spaced from the walls of chamber 23 at the leading and side surfaces of the airfoil.
  • the inner end of the insert 26 projects into a passageway 28 which extends from one end to the other of the upper portion of the tang l4 and connects with the cooling air passageway 20 of the rotor assembly.
  • the inner end of the insert is preferably formed with a thick walled or solid section 30 which is positioned within a slot 32 formed in a lug 34 which projects outwardly from the basic tang portion of the blade into the passageway 28.
  • the insert is held in place by a pin or rivet 36 which is inserted from one side of the tang.
  • the inner end of the rivet is then flared into locking relationship with the lug 34, as is particularly indicated in FIGS. 2 and 3, by inserting a swaging or flaring tool through an aligned hole 38 provided in the opposite side of the tang 14.
  • the end walls of the insert are cut away to provide for the flow of cooling air from the passageway 28 into the interior of the insert.
  • the insert sealingly engages a correspondingly shaped interior opening at the inner end of the chamber 23 so that the insert itself functions as a plenum for the pressurized cooling air.
  • This takes into account the fact that the end of the insert is sealed off by a cap 40.
  • the cap 40 has an outwardly projecting, peripheral flange 41 which may be welded or brazed to the marginal tip portions of the shell which forms the insert.
  • the tip end of the insert is positioned by the inner portion of a correspondingly shaped opening 42.
  • the opening 42 has an undercut groove 44 peripherally thereof.
  • the outer ends of the cap 40 and the insert 26 are flared outwardly into this groove. This provides further sealing of the airfoil chamber 23, and positive retention of the insert cap 40. Additionally, secondary retention of the insert itself is also provided.
  • the insert 26 is further positioned and supported by integrally formed pins 46 which project from the airfoil 22 into the chamber 23.
  • the majority of the air introduced into the plenum insert 26 is discharged therefrom through a series of relatively small, closely spaced orifices. or jets, 48 along its forward edge which impinge the air against the inner surface of the leading edge of the blade shell. This is one of the areas of greatest heat concentrations on the blade during its operation.
  • the impingement cooling effect helps to maintain the metal temperatures at an acceptable level.
  • a lesser number of cooling orifices 50 may be provided on the sides of the blade to impinge against the interior surfaces of the convex and concave surfaces of the airfoil section.
  • orifices 52 may be provided in the insert to impinge air against pins 54 which project into a cooling air discharge slot 56 running the length of the trailing edge portion of the blade.
  • the air impinging against the leading edge portion of the chamber 23 then flows around the insert and is discharge through the slot 56.
  • This flow of cooling air is made more effective by the pins 46 which additionally create turbulence and increase heat transfer as the air flows around the insert. Cooling effectiveness is further enhanced by the impingement air discharged from the lateral and trailing jets 50 and 52 with additional turbulence being provided by the pins 54.
  • the described blade is readily assembled by first fabricating the insert 26, with its end cap 40 in place, then introducing the insert through the opening 42 and positioning the thickened end portion 30 within the slot 32.
  • the rivet 36 may then be inserted and secured as previously described to positively retain the insert in place. Thereafter the tip end of the insert and the cap 40 may be flared outwardly to positively retain it in place for the advantages mentioned above.
  • a turbine blade comprising:
  • said airfoil having thin walls defining a chamber extending from its tip end to said inner end
  • said tang having a passageway from one exterior end face thereof through its upper portion to the insert, said insert having an opening registered with said passageway,
  • said insert having a plurality of orifices directed toward the chamber wall for impingement thereof of cooling air from said plenum
  • the pin extends through and terminates on the end face of the lug and is riveted there against and an opening is provided through the side face of said tang opposite said lug, said opening being aligned with said pin to facilitate riveting thereof.
  • said airfoil chamber extends beneath the level of said platform thereby providing for cooling thereof and the means for discharging cooling air from said chamber include passageway means extending from said chamber and discharging from the trailing edge of said airfoil.
  • said insert being insertable through said aligned openings and sealingly engagable therewith to seal said plenum from said chamber.
  • a turbine blade comprising:
  • said airfoil having thin walls defining a chamber extending from its tip end to its inner end
  • passageway menas extending from the exterior of said tang into said insert for connection of the plenum with a source of pressurized cooling air
  • said insert having a plurality of orifices directed toward the chamber wall for impingement thereof of cooling air from said plenums,
  • the tip end of the insert including a separate cap element having an outwardly projecting peripheral flange telescope within and secured to the sidewalls of the insert,
  • peripheral flange and outer wall portions of the insert being flared into said undercut groove.

Abstract

A turbine blade is described which comprises an integrally formed airfoil, platform and tang. An insert extends through a chamber formed in the airfoil with its inner end being pinned to the tang. The tip end of the insert is closed by a cap which is flared into a groove peripherally of an opening in the tip end of the airfoil. Air is introduced, through the tang, into the plenum formed by the insert and then discharged through orifices to impinge against the chamber walls and cool the airfoil. The insert is assembled by introducing it through the tip end opening and an opening in the lower end of the chamber. The inner end of the insert is brought into register with openings in the tang for insertion and riveting of the retaining pin.

Description

Albani on 3,846,941 [451 Nov. 5, 1974 IMPINGEMENT COOLED TURBINE BLADES AND METHOD OF MAKING SAME Primary ExaminerEvcrette A. Powell. Jr. Attorney, Agent or Firm-Charles M. Hogan; lrwin P.
[75] Inventor: Peter E. Albani, Hamden, Conn. Glrfinklc [73] Assignee: Avco Corporation, Stratford, Conn.
[22] Filed: Oct. 31, 1972 [57] ABSTRACT [21] Appl. No.: 302,421 I I A turbine blade 15 described which comprises an mtcgrally formed airfoil, platform and tang. An insert ex- U-S. Cl. tends through a hamber formed in [he its [5 CL inner end being pinned to the The end of the Fleld of Search insert is closed a cap is flared into a groove 415/1 l6 peripherally of an opening in the tip end of the airfoil. Air is introduced, through the tang, into the plenum References Cited formed by the insert and then discharged through ori- UNITED STATES PATENTS fices to impinge against the chamber walls and cool 2.779565 1 1957 Bruckmann 416/96 the airfoil The insert is assembled y introducing it 2,873,944 2/1959 Wiese et al. 416/92 through the tip end openin and an opening in the 3,446,480 5/1969 Emmerson et at... 416/90 lower end of the chamber. The inner end of the insert 5 1/1972 G m t a 9 9 v 7 is brought into register with openings in the tang for Corsmerer et al. X insertion and riveting of the retaining FOREIGN PATENTS OR APPLICATIONS 7 C 5 D F 926.397 4/1955 Germany 416/92 rawmg r E 3 E 4e 5 l2 4B o V22 r: 3 3& 2 48 I 7 \g a E o I 24 E Z 2 2e 7 /3O e t 4 IMPINGEMENT COOLED TURBINE BLADES AND METHOD OF MAKING SAME to high temperatures of the hot gas stream are also sub-.
jected to extremely high stress loadings particularly because of the high rates of rotation of the turbine rotors and the consequent high centrifugal forces to which they are subjected.
It is long been recognized that operation of turbine blades in this adverse, high temperature environment, can be obtained through the use of a cooling mechanism for the blades. Most, if not all, practical cooling system utilize relativelycool compressor air which is flowed through the blades so that the actual metal temperature of the blades is maintained at a relatively low level and blade strength and life is thus greatly improved.
A more recent innovation in air cooling of blades, has been the so called impingement cooling of blades. In this system, a central plenum is provided from which air is discharged from relatively small holes, or orifices, to impinge against selected portions of the interior of the blade and thus provide a highly effective cooling mechanism. Usually the blade is formed as a hollow shell and the plenum is formed by a separate sheet metal insert which communicates, at its inner end, with a source of pressurized cooling air.
For many reasons well known to those skilled in the art, the blades of a turbine are generally formed as separate elements which are provided with tangs at their inner ends for attachment to the turbine rotor. This not only facilitates manufacture but also facililates maintenance and overhaul. Usually the blades are mounted on the turbine rotor by means of a dovetail slot arrangement or pins so that they can be separately replaced.
The advantages of an impingement cooling system have long been complicated, however, by the difficulties in providing a separate insert within a hollow blade. There are two primary areas of difficulty first in providing an attachment of the insert which does not unduly affect the structural integrity of the blade itself and secondly in assuring structural integrity of the insert itself so that it first remains secure in operation and further minimizes the possibility of any portion of the insert to escape from the blade. These problems are Another objectof the invention is to provide an im proved method for manufacturing turbine blades having inserts for the impingement cooling thereof.
The above ends are attained by turbine blade comprising an intergrally formed airfoil and tang. The airfoil has thin walls which define a chamber extending from its tip end to its inner end. An insert is disposed in this chamber and it extends into the blade tang with its inner end mechanically locked thereto. The insert forms a plenum sealed from the airfoil chamber. Passageway means extend from the exterior of the tang into the insert for connecting the plenum with a source of pressurized cooling air. This cooling air is discharged from a plurality of orifices directed toward the chamber wall to impinge there against and maintain the metal temperature of the airfoil at an acceptable level. The cooling air is then discharged from the chamber.
Preferably, the inner end of the insert is locked to the tang by means of a pin extending therethrough. Advantageously, the inner end of the insert, through which the pin passes is thickened or solid. Further, the passageway means for introducing cooling air into the insert may include a passageway through the upper portion of the tang, extending from one end thereof, and an opening in the lower portion of the insert above its thickened portion. A lug may be formed in the tang passageway and be slotted to receive the lower end of the insert and its retaining pin. The pin may terminate at the face of the lug within the passageway and an opening may be provided in the opposite face or opposite side of the tang to facilitate riveting of the pin.
The insert may be formed as a thin walled shell, at its outer end, having an end cap with 'an outwardly projecting, peripheral flange. This flange may be secured to the wall of the insert shell and then flared outwardly into an undercut peripheral groove around the tip end opening in the airfoil to positively retain the cap itself as well as to provide for sealing of the chamber from leakage through its tip end.
Another feature of the invention is found in the provision of integrally formed pins projecting from the airfoil into the chamber to engage and support the insert therein.
Aligned openings in the tip end of the airfoil and at the inner end of the chamber may be provided to permit insertion of the insert into the blade. After such insertion, the pin may be introduced through aligned openings in the tang and insert to mechanically lock the insert.
The above and other related objects and features of the invention will be apparent from a reading of the following description of the prefered embodiment found in the accompanying drawings and the novelty thereof pointed out in the accompanying claims.
In the drawings:
With reference first to FIG. 1, a portion of a turbine rotor is shown. A plurality of turbine blades 12 (one is shown) are mounted, in the usual configuration,
around the periphery of the rotor 10. The blades are attached to the rotor by dovetail tangs 14 (FIG. 2) which are received by correspondingly shaped grooves in the rotor. The blades are axially held in place on the rotor 10 by retaining rings, or discs, 16 and 18. The disc 18 is spaced from the rotor 10 and defines apassageway 20 for pressurized cooling air which is introduced centrally from the engines compressor. This cooling air enters the blade through a tang passageway in a manner later described to provide cooling of the blade in accordance with the present invention.
The blade comprises an aerodynamically shaped airfoil section or portion 22 which projects outwardly from a platform 24. The tang 14 then extends inwardly of the platform 24 which defines the inner bonds of the hot gas flowpath through the engine. The tang 14, as previously referenced, provides for attachment of the blade to the turbine rotor 10. These components of the blade are integrally formed as by casting or the like.
The airfoil 22 is in the form of a hollow shell, defining a chamber 23 at its relatively thick and blunt leading edge portion. The airfoil 22 in cross section, tapers down to a relatively narrow trailing end section.
An insert 26 is disposed within the chamber 23 which extends from the tip end of the airfoil to below the platform 24. The insert 26 comprises a thin walled shell which is closely spaced from the walls of chamber 23 at the leading and side surfaces of the airfoil. The inner end of the insert 26 projects into a passageway 28 which extends from one end to the other of the upper portion of the tang l4 and connects with the cooling air passageway 20 of the rotor assembly.
The inner end of the insert is preferably formed with a thick walled or solid section 30 which is positioned within a slot 32 formed in a lug 34 which projects outwardly from the basic tang portion of the blade into the passageway 28. The insert is held in place by a pin or rivet 36 which is inserted from one side of the tang. The inner end of the rivet is then flared into locking relationship with the lug 34, as is particularly indicated in FIGS. 2 and 3, by inserting a swaging or flaring tool through an aligned hole 38 provided in the opposite side of the tang 14.
Above the solid portion 30, the end walls of the insert are cut away to provide for the flow of cooling air from the passageway 28 into the interior of the insert. The insert sealingly engages a correspondingly shaped interior opening at the inner end of the chamber 23 so that the insert itself functions as a plenum for the pressurized cooling air. This, of course, takes into account the fact that the end of the insert is sealed off by a cap 40. The cap 40 has an outwardly projecting, peripheral flange 41 which may be welded or brazed to the marginal tip portions of the shell which forms the insert.
. The tip end of the insert is positioned by the inner portion of a correspondingly shaped opening 42. The opening 42 has an undercut groove 44 peripherally thereof. The outer ends of the cap 40 and the insert 26 are flared outwardly into this groove. This provides further sealing of the airfoil chamber 23, and positive retention of the insert cap 40. Additionally, secondary retention of the insert itself is also provided. The insert 26 is further positioned and supported by integrally formed pins 46 which project from the airfoil 22 into the chamber 23.
The majority of the air introduced into the plenum insert 26 is discharged therefrom through a series of relatively small, closely spaced orifices. or jets, 48 along its forward edge which impinge the air against the inner surface of the leading edge of the blade shell. This is one of the areas of greatest heat concentrations on the blade during its operation. The impingement cooling effect, thus provided, helps to maintain the metal temperatures at an acceptable level. A lesser number of cooling orifices 50 may be provided on the sides of the blade to impinge against the interior surfaces of the convex and concave surfaces of the airfoil section. Similarly, orifices 52 may be provided in the insert to impinge air against pins 54 which project into a cooling air discharge slot 56 running the length of the trailing edge portion of the blade. The air impinging against the leading edge portion of the chamber 23 then flows around the insert and is discharge through the slot 56. This flow of cooling air is made more effective by the pins 46 which additionally create turbulence and increase heat transfer as the air flows around the insert. Cooling effectiveness is further enhanced by the impingement air discharged from the lateral and trailing jets 50 and 52 with additional turbulence being provided by the pins 54.
Another feature to be noted is that the described blade is readily assembled by first fabricating the insert 26, with its end cap 40 in place, then introducing the insert through the opening 42 and positioning the thickened end portion 30 within the slot 32. The rivet 36 may then be inserted and secured as previously described to positively retain the insert in place. Thereafter the tip end of the insert and the cap 40 may be flared outwardly to positively retain it in place for the advantages mentioned above.
While a prefered embodiment of the invention has been described, variations thereof will occur to those skilled in the art within the broader aspects of the invention. The spirit and scope of the invention is thus to be derived solely from the appended claims.
Having thus described the invention, what is claimed as novel and desired to be secured by Letters Patent of the United States is:
l. A turbine blade comprising:
an integral airfoil and tang at the inner end thereof,
said airfoil having thin walls defining a chamber extending from its tip end to said inner end,
an insert disposed in said chamber in spaced relation to the walls thereof and extending into said tang, said insert forming a plenum sealed from the airfoil chamber and having a substantial thickness in the innermost end thereof,
said tang having a passageway from one exterior end face thereof through its upper portion to the insert, said insert having an opening registered with said passageway,
a pin extending through said tang and the inner end of said insert in a direction generally normal to the side faces of the tang for mechanically locking the inner end of said insert to said tang,
said insert having a plurality of orifices directed toward the chamber wall for impingement thereof of cooling air from said plenum, and
means for discharging cooling air from said chamber.
2. A turbine blade as in claim 1 wherein at least the majority of said orifices are formed in spaced relation along the portion of the insert opposed to the leading edge portion of the arifoil.
3. A turbine blade as in claim 2 wherein a slotted lug is formed in said tang passageway with an end'face exposed therein,
the pin extends through and terminates on the end face of the lug and is riveted there against and an opening is provided through the side face of said tang opposite said lug, said opening being aligned with said pin to facilitate riveting thereof.
4. A turbine blade as in claim 2 wherein a platform is integrally formed intermediate said airfoil and tang, at the base of said airfoil,
said airfoil chamber extends beneath the level of said platform thereby providing for cooling thereof and the means for discharging cooling air from said chamber include passageway means extending from said chamber and discharging from the trailing edge of said airfoil.
5. A turbine blade as in claim 3 wherein aligned openings are provided between the tip end of the blade and said chamber and between the inner end of the chamber and the tang passageway,
said insert being insertable through said aligned openings and sealingly engagable therewith to seal said plenum from said chamber.
6. A turbine blade comprising:
an integral airfoil and tang,
said airfoil having thin walls defining a chamber extending from its tip end to its inner end,
an insert disposed in said chamber in spaced relation to the walls thereof and extending into said tang, said insert forming a plenum sealed from the airfoil chamber,
mechanical means for locking the inner end of said insert to said tang,
passageway menas extending from the exterior of said tang into said insert for connection of the plenum with a source of pressurized cooling air,
said insert having a plurality of orifices directed toward the chamber wall for impingement thereof of cooling air from said plenums,
means for discharging cooling air from said chamber,
the tip end of the insert including a separate cap element having an outwardly projecting peripheral flange telescope within and secured to the sidewalls of the insert,
an undercut peripheral groove being formed around said tip end opening, and
the peripheral flange and outer wall portions of the insert being flared into said undercut groove.
7. A turbine blade as in claim 6 wherein said airfoil has integral pins projecting into said chamber and engaging and supporting said insert and said airfoil passageway extends substantially along the height of said trailing edge and pins project into said passageway to provide for increased heat turbulence and heat transfer.

Claims (7)

1. A turbine blade comprising: an integral airfoil and tang at the inner end thereof, said airfoil having thin walls defining a chamber extending from its tip end to said inner end, an insert disposed in said chamber in spaced relation to the walls thereof and extending into said tang, said insert forming a plenum sealed from the airfoil chamber and having a substantial thickness in the innermost end thereof, said tang having a passageway from one exterior end face thereof through its upper portion to the insert, said insert having an opening registered with said passageway, a pin extending through said tang and the inner end of said insert in a direction generally normal to the side faces of the tang for mechanically locking the inner end of said insert to said tang, said insert having a plurality of orifices directed toward the chamber wall for impingement thereof of cooling air from said plenum, and means for discharging cooling air from said chamber.
2. A turbine blade as in claim 1 wherein at least the majority of said orifices are formed in spaced relation along the portion of the insert opposed to the leading edge portion of the arifoil.
3. A turbine blade as in claim 2 wherein a slotted lug is formed in said tang passageway with an end face exposed therein, the pin extends through and terminates on the end face of the lug and is riveted there against and an opening is provided through the side face of said tang opposite said lug, said opening being aligned with said pin to facilitate riveting thereof.
4. A turbine blade as in claim 2 wherein a platform is integrally formed intermediate said airfoil and tang, at the base of said airfoil, said airfoil chamber extends beneath the level of said platform thereby providing for cooling thereof and the means for discharging cooling air from said chamber include passageway means extending from said chamber and discharging from the trailing edge of said airfoil.
5. A turbine blade as in claim 3 wherein aligned openings are provided between the tip end of the blade and said chamber and between the inner end of the chamber and the tang passageway, said inSert being insertable through said aligned openings and sealingly engagable therewith to seal said plenum from said chamber.
6. A turbine blade comprising: an integral airfoil and tang, said airfoil having thin walls defining a chamber extending from its tip end to its inner end, an insert disposed in said chamber in spaced relation to the walls thereof and extending into said tang, said insert forming a plenum sealed from the airfoil chamber, mechanical means for locking the inner end of said insert to said tang, passageway menas extending from the exterior of said tang into said insert for connection of the plenum with a source of pressurized cooling air, said insert having a plurality of orifices directed toward the chamber wall for impingement thereof of cooling air from said plenums, means for discharging cooling air from said chamber, the tip end of the insert including a separate cap element having an outwardly projecting peripheral flange telescope within and secured to the sidewalls of the insert, an undercut peripheral groove being formed around said tip end opening, and the peripheral flange and outer wall portions of the insert being flared into said undercut groove.
7. A turbine blade as in claim 6 wherein said airfoil has integral pins projecting into said chamber and engaging and supporting said insert and said airfoil passageway extends substantially along the height of said trailing edge and pins project into said passageway to provide for increased heat turbulence and heat transfer.
US00302421A 1972-10-31 1972-10-31 Impingement cooled turbine blades and method of making same Expired - Lifetime US3846041A (en)

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US00302421A US3846041A (en) 1972-10-31 1972-10-31 Impingement cooled turbine blades and method of making same
GB5024273A GB1441969A (en) 1972-10-31 1973-10-29 Impingement cooled turbine blades and method of making same
SE7314736A SE389162B (en) 1972-10-31 1973-10-30 TURBINE BLOWER AND KIT FOR MANUFACTURE OF A TURBINE BLOWER
DE19732354693 DE2354693C3 (en) 1972-10-31 1973-10-30 Air-cooled turbine blade
FR7338953A FR2205097A5 (en) 1972-10-31 1973-10-31
JP48121871A JPS5249085B2 (en) 1972-10-31 1973-10-31
IT53459/73A IT997587B (en) 1972-10-31 1973-10-31 IMPACT-COOLED TURBINE BLADES AND METHOD FOR PRODUCING

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JP (1) JPS5249085B2 (en)
FR (1) FR2205097A5 (en)
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SE (1) SE389162B (en)

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3994622A (en) * 1975-11-24 1976-11-30 United Technologies Corporation Coolable turbine blade
US4019831A (en) * 1974-09-05 1977-04-26 Brown Boveri Sulzer Turbomachinery Ltd. Cooled rotor blade for a gas turbine
US4105364A (en) * 1975-12-20 1978-08-08 Rolls-Royce Limited Vane for a gas turbine engine having means for impingement cooling thereof
US4171184A (en) * 1977-05-05 1979-10-16 Rolls-Royce Limited Rotor blade for a gas turbine engine
US4177010A (en) * 1977-01-04 1979-12-04 Rolls-Royce Limited Cooled rotor blade for a gas turbine engine
US4297077A (en) * 1979-07-09 1981-10-27 Westinghouse Electric Corp. Cooled turbine vane
US4314794A (en) * 1979-10-25 1982-02-09 Westinghouse Electric Corp. Transpiration cooled blade for a gas turbine engine
US4400137A (en) * 1980-12-29 1983-08-23 Elliott Turbomachinery Co., Inc. Rotor assembly and methods for securing a rotor blade therewithin and removing a rotor blade therefrom
US4407632A (en) * 1981-06-26 1983-10-04 United Technologies Corporation Airfoil pedestaled trailing edge region cooling configuration
US4451959A (en) * 1980-12-29 1984-06-05 Elliott Turbomachinery Company, Inc. Methods for securing a rotor blade within a rotor assembly and removing a rotor blade therefrom
DE4430302A1 (en) * 1994-08-26 1996-02-29 Abb Management Ag Impact-cooled wall part
US5741117A (en) * 1996-10-22 1998-04-21 United Technologies Corporation Method for cooling a gas turbine stator vane
US6193465B1 (en) * 1998-09-28 2001-02-27 General Electric Company Trapped insert turbine airfoil
EP1149982A2 (en) * 2000-04-11 2001-10-31 General Electric Company A method of joining a vane cavity insert to a nozzle segment of a gas turbine
US20040022633A1 (en) * 2002-07-31 2004-02-05 Kraft Robert J. Insulated cooling passageway for cooling a shroud of a turbine blade
EP1411208A1 (en) * 2002-10-18 2004-04-21 General Electric Company Compressor blade with unloaded leading edge and method of unloading the leading edge of a compressor blade
US6824352B1 (en) * 2003-09-29 2004-11-30 Power Systems Mfg, Llc Vane enhanced trailing edge cooling design
US20050281667A1 (en) * 2004-06-17 2005-12-22 Siemens Westinghouse Power Corporation Cooled gas turbine vane
US20080219855A1 (en) * 2007-03-09 2008-09-11 Richard Whitton Turbine blade with micro-turbine nozzle provided in the blade root
US20090246006A1 (en) * 2008-03-26 2009-10-01 Siemens Power Generation, Inc. Mechanically Affixed Turbine Shroud Plug
US20100080687A1 (en) * 2008-09-26 2010-04-01 Siemens Power Generation, Inc. Multiple Piece Turbine Engine Airfoil with a Structural Spar
US7713029B1 (en) 2007-03-28 2010-05-11 Florida Turbine Technologies, Inc. Turbine blade with spar and shell construction
US7967565B1 (en) * 2009-03-20 2011-06-28 Florida Turbine Technologies, Inc. Low cooling flow turbine blade
US7972113B1 (en) * 2007-05-02 2011-07-05 Florida Turbine Technologies, Inc. Integral turbine blade and platform
US20120114495A1 (en) * 2010-11-10 2012-05-10 Richard Lex Seneff Gas turbine engine and blade for gas turbine engine
US8449249B2 (en) 2010-04-09 2013-05-28 Williams International Co., L.L.C. Turbine nozzle apparatus and associated method of manufacture
US20140271101A1 (en) * 2012-09-28 2014-09-18 United Technologies Corporation Modulated turbine vane cooling
WO2015109040A1 (en) * 2014-01-15 2015-07-23 Siemens Aktiengesellschaft Internal cooling system with corrugated insert forming nearwall cooling channels for gas turbine airfoil
EP2890880A4 (en) * 2012-08-30 2015-12-02 United Technologies Corp Gas turbine engine airfoil cooling circuit arrangement
WO2016036366A1 (en) * 2014-09-04 2016-03-10 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of a gas turbine airfoil
WO2016036367A1 (en) * 2014-09-04 2016-03-10 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in midchord cooling cavities of a gas turbine airfoil
US20170234145A1 (en) * 2016-02-15 2017-08-17 General Electric Company Accelerator insert for a gas turbine engine airfoil
DE102016216858A1 (en) 2016-09-06 2018-03-08 Rolls-Royce Deutschland Ltd & Co Kg Blade for a turbomachine and method for assembling a blade for a turbomachine
US20180355730A1 (en) * 2017-06-12 2018-12-13 General Electric Company Turbomachine rotor blade
CN110546348A (en) * 2017-04-10 2019-12-06 赛峰集团 Turbine blade with improved structure
US10626740B2 (en) 2016-12-08 2020-04-21 General Electric Company Airfoil trailing edge segment
US11904420B2 (en) * 2020-10-06 2024-02-20 Safran Aircraft Engines Method for manufacturing a turbomachine compressor blade by compacting

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5234111A (en) * 1975-09-10 1977-03-15 Toshiba Corp Turbine moving blade
US5352091A (en) * 1994-01-05 1994-10-04 United Technologies Corporation Gas turbine airfoil
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Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE926397C (en) * 1952-09-25 1955-04-18 Ernst Heinkel Dr Dr Hollow blade, preferably for axially flowed turbines
US2779565A (en) * 1948-01-05 1957-01-29 Bruno W Bruckmann Air cooling of turbine blades
US2873944A (en) * 1952-09-10 1959-02-17 Gen Motors Corp Turbine blade cooling
US3446480A (en) * 1966-12-19 1969-05-27 Gen Motors Corp Turbine rotor
US3635587A (en) * 1970-06-02 1972-01-18 Gen Motors Corp Blade cooling liner
US3700348A (en) * 1968-08-13 1972-10-24 Gen Electric Turbomachinery blade structure

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2779565A (en) * 1948-01-05 1957-01-29 Bruno W Bruckmann Air cooling of turbine blades
US2873944A (en) * 1952-09-10 1959-02-17 Gen Motors Corp Turbine blade cooling
DE926397C (en) * 1952-09-25 1955-04-18 Ernst Heinkel Dr Dr Hollow blade, preferably for axially flowed turbines
US3446480A (en) * 1966-12-19 1969-05-27 Gen Motors Corp Turbine rotor
US3700348A (en) * 1968-08-13 1972-10-24 Gen Electric Turbomachinery blade structure
US3635587A (en) * 1970-06-02 1972-01-18 Gen Motors Corp Blade cooling liner

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4019831A (en) * 1974-09-05 1977-04-26 Brown Boveri Sulzer Turbomachinery Ltd. Cooled rotor blade for a gas turbine
US3994622A (en) * 1975-11-24 1976-11-30 United Technologies Corporation Coolable turbine blade
US4105364A (en) * 1975-12-20 1978-08-08 Rolls-Royce Limited Vane for a gas turbine engine having means for impingement cooling thereof
US4177010A (en) * 1977-01-04 1979-12-04 Rolls-Royce Limited Cooled rotor blade for a gas turbine engine
US4171184A (en) * 1977-05-05 1979-10-16 Rolls-Royce Limited Rotor blade for a gas turbine engine
US4297077A (en) * 1979-07-09 1981-10-27 Westinghouse Electric Corp. Cooled turbine vane
US4314794A (en) * 1979-10-25 1982-02-09 Westinghouse Electric Corp. Transpiration cooled blade for a gas turbine engine
US4400137A (en) * 1980-12-29 1983-08-23 Elliott Turbomachinery Co., Inc. Rotor assembly and methods for securing a rotor blade therewithin and removing a rotor blade therefrom
US4451959A (en) * 1980-12-29 1984-06-05 Elliott Turbomachinery Company, Inc. Methods for securing a rotor blade within a rotor assembly and removing a rotor blade therefrom
US4407632A (en) * 1981-06-26 1983-10-04 United Technologies Corporation Airfoil pedestaled trailing edge region cooling configuration
DE4430302A1 (en) * 1994-08-26 1996-02-29 Abb Management Ag Impact-cooled wall part
US5741117A (en) * 1996-10-22 1998-04-21 United Technologies Corporation Method for cooling a gas turbine stator vane
US6193465B1 (en) * 1998-09-28 2001-02-27 General Electric Company Trapped insert turbine airfoil
EP1149982A2 (en) * 2000-04-11 2001-10-31 General Electric Company A method of joining a vane cavity insert to a nozzle segment of a gas turbine
EP1149982A3 (en) * 2000-04-11 2004-05-26 General Electric Company A method of joining a vane cavity insert to a nozzle segment of a gas turbine
US20040022633A1 (en) * 2002-07-31 2004-02-05 Kraft Robert J. Insulated cooling passageway for cooling a shroud of a turbine blade
US6811378B2 (en) * 2002-07-31 2004-11-02 Power Systems Mfg, Llc Insulated cooling passageway for cooling a shroud of a turbine blade
EP1411208A1 (en) * 2002-10-18 2004-04-21 General Electric Company Compressor blade with unloaded leading edge and method of unloading the leading edge of a compressor blade
US6824352B1 (en) * 2003-09-29 2004-11-30 Power Systems Mfg, Llc Vane enhanced trailing edge cooling design
US20050281667A1 (en) * 2004-06-17 2005-12-22 Siemens Westinghouse Power Corporation Cooled gas turbine vane
US7118326B2 (en) * 2004-06-17 2006-10-10 Siemens Power Generation, Inc. Cooled gas turbine vane
US20080219855A1 (en) * 2007-03-09 2008-09-11 Richard Whitton Turbine blade with micro-turbine nozzle provided in the blade root
US7713029B1 (en) 2007-03-28 2010-05-11 Florida Turbine Technologies, Inc. Turbine blade with spar and shell construction
US7972113B1 (en) * 2007-05-02 2011-07-05 Florida Turbine Technologies, Inc. Integral turbine blade and platform
US20090246006A1 (en) * 2008-03-26 2009-10-01 Siemens Power Generation, Inc. Mechanically Affixed Turbine Shroud Plug
US8070421B2 (en) * 2008-03-26 2011-12-06 Siemens Energy, Inc. Mechanically affixed turbine shroud plug
US8033790B2 (en) * 2008-09-26 2011-10-11 Siemens Energy, Inc. Multiple piece turbine engine airfoil with a structural spar
US20100080687A1 (en) * 2008-09-26 2010-04-01 Siemens Power Generation, Inc. Multiple Piece Turbine Engine Airfoil with a Structural Spar
US7967565B1 (en) * 2009-03-20 2011-06-28 Florida Turbine Technologies, Inc. Low cooling flow turbine blade
US8449249B2 (en) 2010-04-09 2013-05-28 Williams International Co., L.L.C. Turbine nozzle apparatus and associated method of manufacture
US20120114495A1 (en) * 2010-11-10 2012-05-10 Richard Lex Seneff Gas turbine engine and blade for gas turbine engine
US8888455B2 (en) * 2010-11-10 2014-11-18 Rolls-Royce Corporation Gas turbine engine and blade for gas turbine engine
US9759072B2 (en) 2012-08-30 2017-09-12 United Technologies Corporation Gas turbine engine airfoil cooling circuit arrangement
EP2890880A4 (en) * 2012-08-30 2015-12-02 United Technologies Corp Gas turbine engine airfoil cooling circuit arrangement
US11377965B2 (en) 2012-08-30 2022-07-05 Raytheon Technologies Corporation Gas turbine engine airfoil cooling circuit arrangement
US20140271101A1 (en) * 2012-09-28 2014-09-18 United Technologies Corporation Modulated turbine vane cooling
US10519802B2 (en) 2012-09-28 2019-12-31 United Technologies Corporation Modulated turbine vane cooling
US9670797B2 (en) * 2012-09-28 2017-06-06 United Technologies Corporation Modulated turbine vane cooling
WO2015109040A1 (en) * 2014-01-15 2015-07-23 Siemens Aktiengesellschaft Internal cooling system with corrugated insert forming nearwall cooling channels for gas turbine airfoil
US9840930B2 (en) 2014-09-04 2017-12-12 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in midchord cooling cavities of a gas turbine airfoil
US9863256B2 (en) 2014-09-04 2018-01-09 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of an airfoil usable in a gas turbine engine
WO2016036366A1 (en) * 2014-09-04 2016-03-10 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of a gas turbine airfoil
WO2016036367A1 (en) * 2014-09-04 2016-03-10 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in midchord cooling cavities of a gas turbine airfoil
US10443407B2 (en) * 2016-02-15 2019-10-15 General Electric Company Accelerator insert for a gas turbine engine airfoil
US20170234145A1 (en) * 2016-02-15 2017-08-17 General Electric Company Accelerator insert for a gas turbine engine airfoil
DE102016216858A1 (en) 2016-09-06 2018-03-08 Rolls-Royce Deutschland Ltd & Co Kg Blade for a turbomachine and method for assembling a blade for a turbomachine
US10781699B2 (en) * 2016-09-06 2020-09-22 Rolls-Royce Deutschland Ltd & Co Kg Rotor blade for a turbomachine and method for the assembly of a rotor blade for a turbomachine
US20180066526A1 (en) * 2016-09-06 2018-03-08 Rolls-Royce Deutschland Ltd & Co Kg Rotor blade for a turbomachine and method for the assembly of a rotor blade for a turbomachine
US10626740B2 (en) 2016-12-08 2020-04-21 General Electric Company Airfoil trailing edge segment
CN110546348A (en) * 2017-04-10 2019-12-06 赛峰集团 Turbine blade with improved structure
US11248468B2 (en) 2017-04-10 2022-02-15 Safran Turbine blade having an improved structure
US20180355730A1 (en) * 2017-06-12 2018-12-13 General Electric Company Turbomachine rotor blade
US10851663B2 (en) * 2017-06-12 2020-12-01 General Electric Company Turbomachine rotor blade
US11904420B2 (en) * 2020-10-06 2024-02-20 Safran Aircraft Engines Method for manufacturing a turbomachine compressor blade by compacting

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JPS5249085B2 (en) 1977-12-14
DE2354693A1 (en) 1974-05-09
JPS4995011A (en) 1974-09-10
SE389162B (en) 1976-10-25
DE2354693B2 (en) 1975-07-03
FR2205097A5 (en) 1974-05-24
GB1441969A (en) 1976-07-07
IT997587B (en) 1975-12-30

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