US3782852A - Gas turbine engine blades - Google Patents

Gas turbine engine blades Download PDF

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Publication number
US3782852A
US3782852A US00282779A US3782852DA US3782852A US 3782852 A US3782852 A US 3782852A US 00282779 A US00282779 A US 00282779A US 3782852D A US3782852D A US 3782852DA US 3782852 A US3782852 A US 3782852A
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blade
insert
flow
chambers
cooling fluid
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US00282779A
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A Moore
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Rolls Royce PLC
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Rolls Royce PLC
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    • 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

Definitions

  • a gas turbine blade has a hollow interior space in which is disposed a double-walled insert.
  • the insert forms chambers with the blade walls, and together with the interior of the insert these chambers form part of a path for the flow of cooling fluid through the blade.
  • the chambers provide separated in-flow and [52] US. [51] Int. [58] Field of Search..................
  • out-flow chamber(s) arranged so that cooling fluid S T N w 3 g e my 8 D E n N U supplied from one end of the blade passes through the in-flow chambers to the interior of the insert at the leading edge of the blade, and from there passes to the trailing edge of the blade where it leaves the blade by way of the out-flow chamber(s).
  • the present invention aims to provide a rotor to stator blade construction which enables improved cooling to be obtained.
  • the invention provides a gas-turbine blade with a hollow interior space divided into in-flow and outflow chambers for cooling fluid, having a doubled-walled insert of which the interior communicates with the in-flow chambers in the region of the leading edge of the insert and communicates with the out-flow chamber(s) in the region of the trailing edge thereof, the in-flow chambers lying on both sides of the insert and being arranged for supply with cooling fluid at one end of the blade.
  • the invention is applied to rotor blades which are fed from the root end.
  • FIG. 1 shows a sectional elevation of a gas-turbine rotor blade according to the invention, the section being taken on the line B--B in FIG. 2;
  • FIG. 2 shows a section on the line AA in FIG. 1 to a scale times as great.
  • the turbine blade has leading and trailing edges 10,
  • the blade is hollow and the space within it is divided into air flow chambers for cooling air.
  • a double-walled insert 14 is mounted within the blade, being brazed to the blade at its root end but otherwise unattached to allow for differential expansion.
  • the insert divides the interior of the blade effectively into two spaces lyingeither side of it which spaces are then subdivided by a fin 15 cast on the interior surfaces of each of the walls 17, 18 of the blade.
  • the insert 14 is located within the blade by these fins l5, and also by a large number of pimples 16 indicated in FIG. 1 by crosses.
  • the blade is strengthened by pedestals 19 which extend fully between the walls 17, 18 of the blade and which are also indicated in FIG. 1 by crosses.
  • the particular blade construction shown has an additional advantage in that the fins l5, pimples 16 and pedesdals 19 can all be cast with the blades, using a single piece core, which facilitates production of the blades.
  • the two walls on the insert 14 are held spaced apart by spacers 21.
  • the leading edge 22 of the insert is open along its length, while the trailing edge has a series of apertures 23.
  • the insert is narrower at the shroud end than at the root end, and the fins 15 which support it are correspondingly curved.
  • the leading edge of the blade has leading edge film cooling holes 24 and slots 25 are provided in the trailing edge.
  • a cooling air flow enters the blade at the root end 13 and passes into air inflow chambers 26 lying both sides of the insert over the full radial length of the blade.
  • the cross section of the inflow chambers 26 diminishes towards the .shroud end 12 to allow for the reduction in volume of airflow due to the cooling holes 24, and the flow of air which passes into the open leading edge of the insert 14.
  • Air which passes into the interior of the insert flows chordwise from the leading edge towards the trailing edge and emerges through the apertures 23 into an air outflow chamber 27, and from there the air escapes through the slots 25.
  • a certain amount of impingement cooling also takes place at the leading edge of the blade as the cooling air enters the passage 20 in the leading edge from the spaces between the first row of pimples 16.
  • the initial air flow is spanwise of the blade, and because of the small cross-sectional area between the insert and the blade wall, the air flow speed is substantially increased and the air at its coolest comes into direct contact with the parts of the blade to be cooled. Air entering this insert flows chordwise through a greater cross-sectional area and reaches the trailing edge with little pressure loss through the insert.
  • the apertures 24 are omitted; the apertures 25 are omitted and the cooling air is arranged to leave the blade at the srhoud end; a liquid or a gas other than air is used for cooling; the pimples 16 and/or fins 15 are other than integrally formed on the inside surfaces of the blade walls; the apertures 23 are replaced by a continuous slot along the trailing edge of the insert, and, conversely the leading edge of the insert is closed and apertures analogous to the apertures 23 are provided in the walls of the insert on either side of its leading edge.
  • Another possible variation of the described arrangement is the provision of one or more bleed holes in the fins 15 to allow a small amount of air at the root end of the blade to pass to the outflow chamber 27 directly, that is to say, without passing through the insert 14.
  • a gas turbine blade having opposed walls defining therebetween a hollow interior space, a double-walled insert within said space, and means defining two inflow chambers between said insert and said walls, one on each sideof the insert, and at least one out-flow chamber, the insert interior communicating with the in-flow chambers in the region of the leading edge of the insert and communicating with said out-flow chamber in the region of the trailing edge thereof, the in-flow chambers being arranged for supply with cooling fluid at one end of the blade.
  • a blade as claimed in claim 1 wherein location for the insert within the blade is provided by a plurality of projections integrally formed on the interior surfaces of each blade wall and abutting the insert on either side thereof.
  • a blade as claimed in claim 1 wherein the interior of the insert communicates along the leading edge of the insert with a chamber common to the in-flow chambers and separating the leading edges of the blade and the insert, the walls of the insert being maintained in spaced relation by the plurality of spacers.
  • a gas turbine rotor having a plurality of blades as claimed in claim 1 arranged for supply with cooling fluid at their root ends.
  • a blade as claimed in claim 1 which includes one or more bleed holes connecting the inflow chambers and the outflow chamber(s) directly together at the said one end of the blade.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A gas turbine blade has a hollow interior space in which is disposed a double-walled insert. The insert forms chambers with the blade walls, and together with the interior of the insert these chambers form part of a path for the flow of cooling fluid through the blade. The chambers provide separated in-flow and out-flow chamber(s), arranged so that cooling fluid supplied from one end of the blade passes through the in-flow chambers to the interior of the insert at the leading edge of the blade, and from there passes to the trailing edge of the blade where it leaves the blade by way of the out-flow chamber(s).

Description

[ Jan. 1, 1974 3,560,107 2/1971 Helms..t.............. 3,635,587 l/l972 Giesmanetal...
[ GAS TURBINE ENGINE BLADES [75] Inventor: Alan Moore, Bristol, England [73] Assignee: Rolls-Royce (1971) Limited, Primary Examiner-Everette A. Powell, Jr.
London, England Attorney-Richard K. Stevens et a1. 1
Aug. 22, 1972 Appl. No.: 282,779
ABSTRACT [22] Filed:
A gas turbine blade has a hollow interior space in which is disposed a double-walled insert. The insert forms chambers with the blade walls, and together with the interior of the insert these chambers form part of a path for the flow of cooling fluid through the blade. The chambers provide separated in-flow and [52] US. [51] Int. [58] Field of Search..................
out-flow chamber(s), arranged so that cooling fluid S T N w 3 g e my 8 D E n N U supplied from one end of the blade passes through the in-flow chambers to the interior of the insert at the leading edge of the blade, and from there passes to the trailing edge of the blade where it leaves the blade by way of the out-flow chamber(s).
66 ma 66% ii 446 Meyer et al...... Banthin et a1. Kercher 10 Claims, 2 Drawing Figures GAS TURBINE ENGINE BLADES This invention relates to the blading of gas turbine engines, and although not so restricted, has particular reference to rotor blades for such engines.
As is known, it is advantageous to provide for the cooling of gas turbine engine blades, and many different ways of doing this are known. In particular, it is known to provide a cooling air flow through the interior of the blade, and the cooling is the more effective, the greater the volume of air flow and the greater the velocity of the air as it passes through the blade.
In very small rotor blades, e.g., less than 1 Va inch long considerable difficulties are encountered in providing flow channels by means of which satisfactory cooling can be achieved. The present invention aims to provide a rotor to stator blade construction which enables improved cooling to be obtained.
Accordingly the invention provides a gas-turbine blade with a hollow interior space divided into in-flow and outflow chambers for cooling fluid, having a doubled-walled insert of which the interior communicates with the in-flow chambers in the region of the leading edge of the insert and communicates with the out-flow chamber(s) in the region of the trailing edge thereof, the in-flow chambers lying on both sides of the insert and being arranged for supply with cooling fluid at one end of the blade.
In a preferred form, the invention is applied to rotor blades which are fed from the root end.
The invention will now be described with reference to the accompanying'drawing, in which:
FIG. 1 shows a sectional elevation of a gas-turbine rotor blade according to the invention, the section being taken on the line B--B in FIG. 2; and
FIG. 2 shows a section on the line AA in FIG. 1 to a scale times as great.
The turbine blade has leading and trailing edges 10,
11 a shroud end 12 and a root end 13. The bladeis hollow and the space within it is divided into air flow chambers for cooling air.
A double-walled insert 14 is mounted within the blade, being brazed to the blade at its root end but otherwise unattached to allow for differential expansion. The insert divides the interior of the blade effectively into two spaces lyingeither side of it which spaces are then subdivided by a fin 15 cast on the interior surfaces of each of the walls 17, 18 of the blade. The insert 14 is located within the blade by these fins l5, and also by a large number of pimples 16 indicated in FIG. 1 by crosses. Towards the trailing edge 11 of the blade, at points to which the insert 14 does not extend, the blade is strengthened by pedestals 19 which extend fully between the walls 17, 18 of the blade and which are also indicated in FIG. 1 by crosses. The particular blade construction shown has an additional advantage in that the fins l5, pimples 16 and pedesdals 19 can all be cast with the blades, using a single piece core, which facilitates production of the blades.
The two walls on the insert 14 are held spaced apart by spacers 21. The leading edge 22 of the insert is open along its length, while the trailing edge has a series of apertures 23. The insert is narrower at the shroud end than at the root end, and the fins 15 which support it are correspondingly curved. The leading edge of the blade has leading edge film cooling holes 24 and slots 25 are provided in the trailing edge.
In operation, a cooling air flow enters the blade at the root end 13 and passes into air inflow chambers 26 lying both sides of the insert over the full radial length of the blade. The cross section of the inflow chambers 26 diminishes towards the .shroud end 12 to allow for the reduction in volume of airflow due to the cooling holes 24, and the flow of air which passes into the open leading edge of the insert 14. Air which passes into the interior of the insert flows chordwise from the leading edge towards the trailing edge and emerges through the apertures 23 into an air outflow chamber 27, and from there the air escapes through the slots 25.
A certain amount of impingement cooling also takes place at the leading edge of the blade as the cooling air enters the passage 20 in the leading edge from the spaces between the first row of pimples 16.
With this arrangement, the initial air flow is spanwise of the blade, and because of the small cross-sectional area between the insert and the blade wall, the air flow speed is substantially increased and the air at its coolest comes into direct contact with the parts of the blade to be cooled. Air entering this insert flows chordwise through a greater cross-sectional area and reaches the trailing edge with little pressure loss through the insert.
Although pimples 18 of any desired shape may be used for the location of the insert, diagonal fins could be used in addition, or in their place, which would also serve to guide cooling airfrom the root end toward the shroud end of the blade.
Amongst other possible variations of the described arrangement are: the apertures 24 are omitted; the apertures 25 are omitted and the cooling air is arranged to leave the blade at the srhoud end; a liquid or a gas other than air is used for cooling; the pimples 16 and/or fins 15 are other than integrally formed on the inside surfaces of the blade walls; the apertures 23 are replaced by a continuous slot along the trailing edge of the insert, and, conversely the leading edge of the insert is closed and apertures analogous to the apertures 23 are provided in the walls of the insert on either side of its leading edge.
Another possible variation of the described arrangement is the provision of one or more bleed holes in the fins 15 to allow a small amount of air at the root end of the blade to pass to the outflow chamber 27 directly, that is to say, without passing through the insert 14.
What is claimed is:
l. A gas turbine blade having opposed walls defining therebetween a hollow interior space, a double-walled insert within said space, and means defining two inflow chambers between said insert and said walls, one on each sideof the insert, and at least one out-flow chamber, the insert interior communicating with the in-flow chambers in the region of the leading edge of the insert and communicating with said out-flow chamber in the region of the trailing edge thereof, the in-flow chambers being arranged for supply with cooling fluid at one end of the blade. I
2. A blade as claimed in claim 1, where the in-flow and out-flow chambers are separated from one another by fins integrally formed on the interior surfaces of the blade walls and abutting the insert on either side thereof.
3. A blade as claimed in claim 1 wherein location for the insert within the blade is provided by a plurality of projections integrally formed on the interior surfaces of each blade wall and abutting the insert on either side thereof.
4. A blade as claimed in claim 3, wherein at least some of the projections are in the form of fins arranged to guide cooling fluid from the said one end of the blade towards the other end thereof.
5. A blade as claimed in claim 1, wherein the interior of the insert communicates along the leading edge of the insert with a chamber common to the in-flow chambers and separating the leading edges of the blade and the insert, the walls of the insert being maintained in spaced relation by the plurality of spacers.
6. A blade as claimed in claim 1, wherein a plurality of apertures are formed in the region of the leading edge of the blade for allowing cooling fluid to pass to the exterior of the blade from the inflow chambers.
7. A blade as claimed in claim 1, wherein a plurality of apertures are formed in the region of the trailing edge of the blade for allowing cooling fluid to pass to the exterior of the blade from the out-flow chamber(s).
8. A gas turbine rotor having a plurality of blades as claimed in claim 1 arranged for supply with cooling fluid at their root ends.
9. A blade as claimed in claim 1, which includes one or more bleed holes connecting the inflow chambers and the outflow chamber(s) directly together at the said one end of the blade.
10. A blade as claimed in claim 1, wherein the insert is mounted within the blade by brazing at one end thereof, being otherwise unattached to allow for differential expansion.

Claims (10)

1. A gas turbine blade having opposed walls defining therebetween a hollow interior space, a double-walled insert within said space, and means defining two inflow chambers between said insert and said walls, one on each side of the insert, and at least one out-flow chamber, the insert interior communicating with the in-flow chambers in the region of the leading edge of the insert and communicating with said out-flow chamber in the region of the trailing edge thereof, the in-flow chambers being arranged for supply with cooling fluid at one end of the blade.
2. A blade as claimed in claim 1, where the in-flow and out-flow chambers are separated from one another by fins integrally formed on the interior surfaces of the blade walls and abutting the insert on either side thereof.
3. A blade as claimed in claim 1 wherein location for the insert within the blade is provided by a plurality of projections integrally forMed on the interior surfaces of each blade wall and abutting the insert on either side thereof.
4. A blade as claimed in claim 3, wherein at least some of the projections are in the form of fins arranged to guide cooling fluid from the said one end of the blade towards the other end thereof.
5. A blade as claimed in claim 1, wherein the interior of the insert communicates along the leading edge of the insert with a chamber common to the in-flow chambers and separating the leading edges of the blade and the insert, the walls of the insert being maintained in spaced relation by the plurality of spacers.
6. A blade as claimed in claim 1, wherein a plurality of apertures are formed in the region of the leading edge of the blade for allowing cooling fluid to pass to the exterior of the blade from the inflow chambers.
7. A blade as claimed in claim 1, wherein a plurality of apertures are formed in the region of the trailing edge of the blade for allowing cooling fluid to pass to the exterior of the blade from the out-flow chamber(s).
8. A gas turbine rotor having a plurality of blades as claimed in claim 1 arranged for supply with cooling fluid at their root ends.
9. A blade as claimed in claim 1, which includes one or more bleed holes connecting the inflow chambers and the outflow chamber(s) directly together at the said one end of the blade.
10. A blade as claimed in claim 1, wherein the insert is mounted within the blade by brazing at one end thereof, being otherwise unattached to allow for differential expansion.
US00282779A 1971-08-25 1972-08-22 Gas turbine engine blades Expired - Lifetime US3782852A (en)

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Cited By (15)

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US3902820A (en) * 1973-07-02 1975-09-02 Westinghouse Electric Corp Fluid cooled turbine rotor blade
US3930748A (en) * 1972-08-02 1976-01-06 Rolls-Royce (1971) Limited Hollow cooled vane or blade for a gas turbine engine
US4153386A (en) * 1974-12-11 1979-05-08 United Technologies Corporation Air cooled turbine vanes
DE3040192A1 (en) * 1979-10-26 1981-05-07 Société Nationale d'Etude et de Construction de Moteurs d'Aviation (S.N.E.C.M.A.), 75015 Paris COOLED TURBINE SHOVEL
US4278400A (en) * 1978-09-05 1981-07-14 United Technologies Corporation Coolable rotor blade
US4403917A (en) * 1980-01-10 1983-09-13 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Turbine distributor vane
US4767261A (en) * 1986-04-25 1988-08-30 Rolls-Royce Plc Cooled vane
US5022817A (en) * 1989-09-12 1991-06-11 Allied-Signal Inc. Thermostatic control of turbine cooling air
US5232343A (en) * 1984-05-24 1993-08-03 General Electric Company Turbine blade
US5279111A (en) * 1992-08-27 1994-01-18 Inco Limited Gas turbine cooling
US20080226461A1 (en) * 2007-03-13 2008-09-18 Siemens Power Generation, Inc. Intensively cooled trailing edge of thin airfoils for turbine engines
US7497655B1 (en) * 2006-08-21 2009-03-03 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall impingement and vortex cooling
US20140093391A1 (en) * 2012-09-28 2014-04-03 Solar Turbines Incorporated Cooled turbine blade with trailing edge flow metering
US20170234145A1 (en) * 2016-02-15 2017-08-17 General Electric Company Accelerator insert for a gas turbine engine airfoil
US10641103B2 (en) 2017-01-19 2020-05-05 United Technologies Corporation Trailing edge configuration with cast slots and drilled filmholes

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NL7802688A (en) * 1978-03-13 1979-09-17 Philips Nv DEVICE FOR CONVERSION FROM ACOUSTIC TO ELECTRICAL VIBRATIONS AND VERSIONS, EQUIPPED WITH AT LEAST ONE CONDENSER ELECTRICAL ELEMENT CONNECTED TO AN ELECTRONIC CIRCUIT.
US4515526A (en) * 1981-12-28 1985-05-07 United Technologies Corporation Coolable airfoil for a rotary machine
US5097660A (en) * 1988-12-28 1992-03-24 Sundstrand Corporation Coanda effect turbine nozzle vane cooling
DE19634238A1 (en) * 1996-08-23 1998-02-26 Asea Brown Boveri Coolable shovel
JPH10266803A (en) * 1997-03-25 1998-10-06 Mitsubishi Heavy Ind Ltd Gas turbine cooling moving blade
DE59709275D1 (en) * 1997-07-14 2003-03-13 Alstom Switzerland Ltd Cooling system for the trailing edge area of a hollow gas turbine blade
SE512384C2 (en) * 1998-05-25 2000-03-06 Abb Ab Component for a gas turbine
EP1000698B1 (en) * 1998-11-09 2003-05-21 ALSTOM (Switzerland) Ltd Cooled components with conical cooling passages
DE19860787B4 (en) * 1998-12-30 2007-02-22 Alstom Turbine blade with cooling channels
US6955525B2 (en) * 2003-08-08 2005-10-18 Siemens Westinghouse Power Corporation Cooling system for an outer wall of a turbine blade
SE526847C2 (en) * 2004-02-27 2005-11-08 Demag Delaval Ind Turbomachine A component comprising a guide rail or a rotor blade for a gas turbine
EP1847684A1 (en) * 2006-04-21 2007-10-24 Siemens Aktiengesellschaft Turbine blade
US7563072B1 (en) 2006-09-25 2009-07-21 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall spiral flow cooling circuit
US8636463B2 (en) * 2010-03-31 2014-01-28 General Electric Company Interior cooling channels
EP2378073A1 (en) * 2010-04-14 2011-10-19 Siemens Aktiengesellschaft Blade or vane for a turbomachine
US9314838B2 (en) 2012-09-28 2016-04-19 Solar Turbines Incorporated Method of manufacturing a cooled turbine blade with dense cooling fin array
US9228439B2 (en) 2012-09-28 2016-01-05 Solar Turbines Incorporated Cooled turbine blade with leading edge flow redirection and diffusion
US10822963B2 (en) * 2018-12-05 2020-11-03 Raytheon Technologies Corporation Axial flow cooling scheme with castable structural rib for a gas turbine engine

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US3370829A (en) * 1965-12-20 1968-02-27 Avco Corp Gas turbine blade construction
US3540810A (en) * 1966-03-17 1970-11-17 Gen Electric Slanted partition for hollow airfoil vane insert
US3560107A (en) * 1968-09-25 1971-02-02 Gen Motors Corp Cooled airfoil
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Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3930748A (en) * 1972-08-02 1976-01-06 Rolls-Royce (1971) Limited Hollow cooled vane or blade for a gas turbine engine
US3902820A (en) * 1973-07-02 1975-09-02 Westinghouse Electric Corp Fluid cooled turbine rotor blade
US4153386A (en) * 1974-12-11 1979-05-08 United Technologies Corporation Air cooled turbine vanes
US4278400A (en) * 1978-09-05 1981-07-14 United Technologies Corporation Coolable rotor blade
DE3040192A1 (en) * 1979-10-26 1981-05-07 Société Nationale d'Etude et de Construction de Moteurs d'Aviation (S.N.E.C.M.A.), 75015 Paris COOLED TURBINE SHOVEL
FR2468727A1 (en) * 1979-10-26 1981-05-08 Snecma IMPROVEMENT TO COOLED TURBINE AUBES
US4456428A (en) * 1979-10-26 1984-06-26 S.N.E.C.M.A. Apparatus for cooling turbine blades
US4403917A (en) * 1980-01-10 1983-09-13 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Turbine distributor vane
US5232343A (en) * 1984-05-24 1993-08-03 General Electric Company Turbine blade
US4767261A (en) * 1986-04-25 1988-08-30 Rolls-Royce Plc Cooled vane
US5022817A (en) * 1989-09-12 1991-06-11 Allied-Signal Inc. Thermostatic control of turbine cooling air
US5279111A (en) * 1992-08-27 1994-01-18 Inco Limited Gas turbine cooling
US7497655B1 (en) * 2006-08-21 2009-03-03 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall impingement and vortex cooling
US20080226461A1 (en) * 2007-03-13 2008-09-18 Siemens Power Generation, Inc. Intensively cooled trailing edge of thin airfoils for turbine engines
US7722326B2 (en) * 2007-03-13 2010-05-25 Siemens Energy, Inc. Intensively cooled trailing edge of thin airfoils for turbine engines
US20140093391A1 (en) * 2012-09-28 2014-04-03 Solar Turbines Incorporated Cooled turbine blade with trailing edge flow metering
US9206695B2 (en) * 2012-09-28 2015-12-08 Solar Turbines Incorporated Cooled turbine blade with trailing edge flow metering
US20170234145A1 (en) * 2016-02-15 2017-08-17 General Electric Company Accelerator insert for a gas turbine engine airfoil
US10443407B2 (en) * 2016-02-15 2019-10-15 General Electric Company Accelerator insert for a gas turbine engine airfoil
US10641103B2 (en) 2017-01-19 2020-05-05 United Technologies Corporation Trailing edge configuration with cast slots and drilled filmholes

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DE2241192B2 (en) 1977-07-21
US3806274A (en) 1974-04-23
GB1361256A (en) 1974-07-24
DE2241192C3 (en) 1978-03-09
AU4587272A (en) 1974-03-07
FR2150476B1 (en) 1979-04-06
DE2241192A1 (en) 1973-03-08
AU467301B2 (en) 1975-11-27
SE378645B (en) 1975-09-08
FR2150476A1 (en) 1973-04-06

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