US2931624A - Gas turbine blade - Google Patents

Gas turbine blade Download PDF

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Publication number
US2931624A
US2931624A US657849A US65784957A US2931624A US 2931624 A US2931624 A US 2931624A US 657849 A US657849 A US 657849A US 65784957 A US65784957 A US 65784957A US 2931624 A US2931624 A US 2931624A
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United States
Prior art keywords
blade
root
portion
adjacent
ridges
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Expired - Lifetime
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US657849A
Inventor
Hyde John Alan Courtney
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Orenda Engines Ltd
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Orenda Engines Ltd
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Publication date
Application filed by Orenda Engines Ltd filed Critical Orenda Engines Ltd
Priority to US657849A priority Critical patent/US2931624A/en
Priority claimed from FR792741A external-priority patent/FR1220732A/en
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Publication of US2931624A publication Critical patent/US2931624A/en
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Expired - Lifetime legal-status Critical Current

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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02Blade-carrying members, e.g. rotors
    • F01D5/08Heating, heat-insulating or cooling means
    • F01D5/085Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor
    • F01D5/087Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor in the radial passages of the rotor disc
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/14Form or construction
    • F01D5/18Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187Convection cooling

Description

April 1960 J. A. c. HYDE 2,931,624

GAS TURBINE BLADE 2 Sheets-Sheet I- Filed Ilay 8, 1957 mvawrm J. A. c. HYDE TTORNE'YS 2 Sheets-s 2 INVENT0 ATTOPNEyg- April 5, 1960 J; A. c. HYDE GAS TURBINE. BLADE Filed May 8. 1957 l atented Apr. 5, 1960 GAS TURBINE BLADE John Alan Courtney Hyde, Georgetown, Ontario, Canada, assignor to Orenda Engines Limited, Malton, Ontario, Canada Application May 8, 1957, Serial No. 657,849

4 Claims. (Cl. 253-39.15)

It is an object of the present invention to provide a gas turbine blade embodying means to enable the blade to be cooled.

It is a further object to provide a gas turbine blade of this type in which the cooling means are entirely external of the solid blade member and, hence, do not weaken the blade asva whole.

According to the present invention a gas turbine blade comprises a root portion and a blade portion, a root platform extending laterally from one side of' the blade, the side of the blade portion remote from the root plat,- form being provided with ridges longitudinally of the blade to define grooves therebetween, a sheet metal skin overlying the grooves, the grooves defined by the ridges being open at the tip of the blade, the skin adjacent the root portion of the blade being bent away from the blade to define a flange lying in substantially the same plane as the plane of the root platform, the ridges beneath the skin extending towards the root portion of the blade to maintain the skin in spaced relationship to the blade so that the grooves defined thereby are open to the area adjacent the root and beneath the flange.

A single preferred embodiment of the invention Will'be described with reference to the accompanying drawings in which Figure 1 is a partly broken away perspective view of a segment of a blade mounting structure having two gas turbine blades made in accordance with the invention positioned therein, and Figure 2 is a sectional view taken near the central radial plane of Figure 1.

Referring now to the drawings it will be seen that the invention is being described with reference to rotor blades secured to a rotor disc of a type commonly used in gas turbine engines.

It will be immediately apparent, however, that the surface of the sheet metal skin 16 forms a smooth and flush continuation of the blade surface.

At the lower end of the sheet metal skin 16 (the radially inner end as seen in the drawing) the skin is bent outwardly from the blade at 19 to define a flange 20 which lies parallel to the plane of the upper surface of the root platform 13.

The gas turbine blades are mounted in an annular mounting member 21 which, in the embodiment illustrated, is a gas turbine rotor disc of the split-disc type. It will be observed that two disc-like members (only segments of which are shown) 23 and 24 are placed together with their radial faces 23a and 24a in tight abutment. Conventional blade mounting slots 25 are provided, each slot passing completely through the compositedisc as can be seen in the drawing. The periphery of thecomposite disc 21 is provided with a peripheral chan nel 26 which is formed by a pair of side walls 27 and 28,, one side wall lying on either side of the composite disc' 21 and a bottom wall 26a, one half of which is carried: by each disclike member 23 and 24. The blade mounting slots 25. are of a configuration to slidably receive theprinciples of the invention are equally applicable to stator blade arrangements and the text of this description should be read with such a construction in mind.

The turbine blade shown generally at 10 may be seen to comprise a root portion 11, and a blade portion 12 which are integrally formed from a single piece of material by any of the well known methods. r A root platform l3'extends laterally from only on side of the blade portion 12 adjacent the root portion 11.

The side of the blade portion 12 opposite the root platform 13 is provided with a series of longitudinally extending ridges 14 which define grooves 15 therebetween. A sheet metal skin 16 is applied over the ridged side of the blade portion 12 by means of welding, brazing or other suitable means to enclose the grooves 15 to define, in conjunction with the ridges 14, longitudinal passages for the flow of cooling fluid within the blade portion 12. It ,will be seen in the drawings that the edges 17 and 18 of the sheet metal skin 16 are recessed within the outer surface of the ridged side of the blade so that the outer root portion 11 of the blade 10. It will be seen from the drawings that the adjacent surfaces 29 and 30 of the root platform and the side walls defining the peripheral channel respectively are in tight sealing engagement with one another when the blades 10 are in position since the root platform overlies and closes a portion of the peripheral channel and the root portions hold the root platforms snugly against the outer peripheral surfaces of. the side walls 27 and 28.

Between each pair of blade mounting slots 25 is a port 31 which is in communication, at one end, with a source of cooling fluid 32 and, at the other end, with the peripheral channel 26.

When the baldes are mounted in position in the mounting member 21 the free edge 20a of the flange 20 of the sheet metal skin 16 of one blade is in tight sealing engagement with the adjacent free edge 33 of the root platform 13 of an adjacent blade. It will be seen that the free edge 33 of the root platform 13 adjacent the flange 20 of the sheet metal skin 16 of the adjacent blade is provided with a recess 34 into which the free edge 20a of the flange 20 of the sheet metal skin16 may fit. The adjacent surfaces 35 and 30 of the flange 20 and the side walls 27 and 28 defining the peripheral channel 26 are, accordingly, maintained in tight sealing engagement with one another, and as a result, the peripheral channel 26, in conjunction with the side walls 27 and 28, the root platforms 13 and the flanges 20 define a plenum chamber between each pair of blades which,

in. turn, is in communication with the gallery v32 by means of ports 31 which lie between each pair of blade mounting slots 25. l

As can be seen from the drawings the grooves 15 defined by the longitudinally extending ridges 14 on the face of each blade remote from the root platform 13 extend completely to the tip of the blade and are, at this point, open to atmosphere so that any fluid within the blades may pass out through the openings at the tip of the blade.

At their inner ends, the ridges 14 extend towards the root portions of the blade a distance sufficient to maintain the'skin 16 in spaced relationship to the blade so F that the grooves 15' are open to the space 36 beneath the flange 20 and adjacent the root portion 11. This area comprises the'plenurn chamber '36 and, accordingly, cooling fluid from the gallery 32 may enter the plenum chamber 36 through the ports 31 and, from there, may pass radially outwardly through the longitudinal grooves 15 defined by the ridges 14 and the skin 16 of the turbine blades to -eflect the cooling of the blade when the engine is in operation.

In the event that it is desired to employe the substance of this invention in conjunction with a turbine "stator assembly which, in the usual form, comprises a mounting member-in which the blades extendtradially :inwardly thereof rather than radially outwardly of the mounting member as shown in the drawings.

As is common practice in the gas turbine industry some means must be provided for locking the blades 10 in their position within the blade mounting slots 25.

, Since, however, this is common practice .in the industry no attempt has been made to describe any particular embodiment. Similarly, gas turbine manufacturers adopt various methods of providing mounting members and, accordingly, the split disc type of turbine wheel has been illustrated by way of example only.

In the following claims certain terms are used definitively and for the purpose of clarity the following definiitions are given herein.

The'blade portion of the 'gas turbine blade is deemed to be that portion on the side of the root platform remote from the root mounting means. The root platform is that portion of the blade lying between the rootmounting means and the blade portion which extends in a direction laterally of the outer surface of the blade. The root portion is that portion of the blade lying on the side of the root platform remote from the blade portion. Accordingly, when in the specification or in the claims, the words blade or turbine blade are used it is intended to mean the entire gas turbine blade including the root, the root platform and the blade portion. When, however, the word portion'follows the word blade it is then intended to mean only that portion of the blade on the side of the root platform remote from'the root portion.

The foregoing specific description has referred to only one preferred embodiment of the present invention and it is to be appreciated that minor modifications in the construction and arrangement of 'parts may be made without departing from the spirit of the invention or the scope of the appended claims.

What I claim as my invention is:

1. A gas turbine blading assembly comprising an annular mounting member, an annular peripheral channel in the mounting member defined by radially extending and axially spaced side walls on either side of the mounting member and a bottom wall connecting said side walls, a series of circumferentially spaced blade mounting slots formed in the mounting member, each slot being transverse to the channel and passing through the side walls and the bottom wall, a port for cooling fluid in the mounting member between each pair of blade mounting slots and opening through said bottom wall, the ports communicating with a source of cooling fluid, a plurality of turbine blades each with its 'root portion engaged in one of said blade mounting slots and mesa ing across and interrupting said channel, each blade having a root platform having a free edge extending laterally from one side only towards an adjacent blade and overlying and closing one portion of said peripheral channel, the side of the blade opposite the root platform being provided with ridges extending longitudinally of the blade to define grooves therebetween, a sheet metal skin overlying the grooves and secured to the ridges, the grooves defined by the ridges and the skin being open at the tip of the blade, the skin adjacent the root portion of the blade being bent away from the blade to define a flange overlying and closing another portion of said peripheral channel adjacent said one portion and having a free edge parallel to the free edge of the root platform, the free edge of the flange of one blade being in sealing engagement with the free edge of the root platform of an adjacent blade, the side walls and. bottom wall of the. channel in the mounting member in conjunction with the said root platform, adjacent root portions and the said flange defining a plenum chamber in communication with the grooves defined by the sheet metal skin and the ridges on the blade.

2. A gas turbine blading assembly comprising an annular mounting member, an annular peripheral channel in the mounting member defined by radially extending and axially spaced side walls on either side of the mounting member and a bottom wall connecting said side wall, a series of circumferentially spaced blade mountingjslots formed in the mounting member, each slot being transverse to the channel and passing through the side walls and the bottom wall, a port for cooling fluid in the mounting member between each pair of blade mounting slots, the ports communicating with a source of cooling fluid and opening through the bottom wall, a plurality of turbine blades each with its root portion engaged in one of said slots and extending across and interruptingsaid channel, each blade having a root platform having a free edge extending laterally from only one side thereof towards an adjacent blade and overlying and closing one portion of said peripheral channel, the adjacent surfaces of the root platform and the side walls lying in sealing engagement with one another, the side of the blade opposite the root platform being provided with ridges extending longitudinally of the blade to define grooves therebetween, a sheet metal skin overlying the grooves and secured to the ridges, the grooves defined by the ridges and the skin being open at the tip "of the blade, the skin adjacent the root portion of the blade being bent'away from the blade to define a flangeoverlying and closing another portion of said peripheral channel adjacent said one portion and having a free edge parallel to the free edge of the root platform, the free edge of the flange of one blade being in sealing engagement with the free edge of the root platform of an adjacent blade, the side walls and bottom wall of the channel in the mounting member, 'in conjunction with thesaid root platform, adjacent root portions and the saidflange, defining a plenum chamber in communication with the grooves defined by the sheet metal skin and the ridges on the blade.

3. A gas turbine blading assembly comprising an annular mounting member, an annular peripheral channel in the mounting member defined by radially extending and axially spaced side walls on either side of the mounting member and a bottom wall connecting said side walls, a series of circumferentially spaced blade mounting slots formed in the mounting member, each slot being transverse to the channel and passing through the side walls and the bottom wall, a port for cooling .fiuid in the mounting member between each pair of blade mounting slots, the ports communicating with a source of cooling fluid and opening through the bottom wall, a plurality of turbine blades each with its root portion engaged in one of said slots and extending across and interrupting said channel, each blade having aroot platform having a free edge extending laterally from only one side towards an adjacent blade and overlying and closing one portion of said channel, the adjacent surfaces of the root platform and the side Walls lying in sealing engagement with one another, the side of the blade opposite the root platform being provided with ridges extending longitudinally of the blade to define grooves therebetween, a sheet metal skin overlying the grooves and secured to the ridges, the outer surface of the skin forming a smooth continuation of the blade surface overthe ridges, the grooves defined by the ridges and the skin being open at the tip'of the blade, the skin adjacent the root portion of the blade being bent away from the blade to define a flange having a free edge parallel to the free edge of the root platform and overlying and closing another portion of said channel adjacent said one portion, the free edge of the flange of one blade being in sealingen'gagement with the free edge of the root platform of an adjacent blade, the side walls and bottom wall of the channel in the mounting member,

in conjunction with the said root platform, adjacent root portions and the said flange defining a plenum chamber in communication with the grooves defined by the sheet metal skin and the ridges on the blade.

4. A gas turbine blading assembly comprising an annular mounting member, an annular peripheral channel in the mounting member defined by radially extending and axially spaced side walls on either side of the mounting member and a bottom wall connecting said side walls, a series of circumferentially spaced blade mounting slots formed in the mounting member each slot passing through the side walls and the bottom wall of the channel, a port for cooling fluid in the mounting'member between each pair of blade mounting slots, the port communicating with a source of cooling fluid and opening through the bottom wall, a plurality of turbine blades each with its root portion engaged in one of said slots'and extending across and interrupting said channel, each blade having a root platform having a free edgeextending laterally from only one side towards an adjacent blade and overlying and closing one portion of said peripheral channel, the adjacent surfaces of the root platform and the side walls lying in sealing engagement with one another, the side of the blade oposite the root platform being provided with ridges the ridges, the grooves defined by the ridges and the skin being open at the tip of the blade, the skin adjacent the root portion of the blade being bent away from the blade to define a flange lying in substantially the same plane as the plane of the root platform and overlying and closing another portion of said peripheral channel adjacent said one portion and having a free'edge. parallel to the free edge of the root platform, the free edge of the flange of one blade being in sealing engagement with the free edge of the root platform of an adjacent blade, the side walls and bottom walls of the channel in the mounting member in conjunctionwith the root platform, adjacent said root portions and the said flange defining a plenum chamber in communication with the grooves defined by the sheet metal skin and the ridges on the blade.

References Cited in the file of this patent UNITED STATES PATENTS 2,641,439 Williams June 9, 1953 2,641,440 Williams June 9, 1953 2,656,147 Brownhill Oct. 20, 1953 2,700,530 Williams Jan. 25, 1955 2,779,565 Bruckmann Jan. 29, 1957 2,786,646 Grantham Mar. 26, 1957 FOREIGN PATENTS 161,997 Australia Mar. 17, 1955 737,479 Germany July 15. 1943 924,248 Germany Feb. 28, 1955

US657849A 1957-05-08 1957-05-08 Gas turbine blade Expired - Lifetime US2931624A (en)

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US657849A US2931624A (en) 1957-05-08 1957-05-08 Gas turbine blade
GB693758A GB854812A (en) 1957-05-08 1958-03-04 Gas turbine blade
FR792741A FR1220732A (en) 1957-05-08 1959-04-21 Blade of gas turbine

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Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3112914A (en) * 1960-08-01 1963-12-03 Gen Motors Corp Turbine rotor
US3761200A (en) * 1970-12-05 1973-09-25 Secr Defence Bladed rotors
FR2471474A1 (en) * 1979-12-17 1981-06-19 United Technologies Corp Rotor disc
US4451204A (en) * 1981-03-25 1984-05-29 Rolls-Royce Limited Aerofoil blade mounting
US4606701A (en) * 1981-09-02 1986-08-19 Westinghouse Electric Corp. Tip structure for a cooled turbine rotor blade
EP0859127A1 (en) * 1997-02-13 1998-08-19 BMW Rolls-Royce GmbH Channeling of cooling air in a turbine rotor disc
EP2233692A1 (en) * 2009-03-27 2010-09-29 Siemens Aktiengesellschaft Axial turboengine rotor with rotor cooling
EP2348191A2 (en) 2010-01-22 2011-07-27 Rolls-Royce plc A Rotor Disc
US20110182738A1 (en) * 2010-01-27 2011-07-28 Herbert Chidsey Roberts Method and apparatus for a segmented turbine bucket assembly
CH704716A1 (en) * 2011-03-22 2012-09-28 Alstom Technology Ltd Rotor disk for a turbine rotor and turbine as well as with such a rotor disk.

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE737479C (en) * 1939-01-26 1943-07-15 Versuchsanstalt Fuer Luftfahrt Gas turbine blade
US2641440A (en) * 1947-11-18 1953-06-09 Chrysler Corp Turbine blade with cooling means and carrier therefor
US2641439A (en) * 1947-10-01 1953-06-09 Chrysler Corp Cooled turbine or compressor blade
US2656147A (en) * 1946-10-09 1953-10-20 English Electric Co Ltd Cooling of gas turbine rotors
US2700530A (en) * 1948-08-27 1955-01-25 Chrysler Corp High temperature elastic fluid apparatus
DE924248C (en) * 1941-03-28 1955-02-28 Daimler Benz Ag Huelsenhohlschaufel for gas or exhaust gas turbines
US2779565A (en) * 1948-01-05 1957-01-29 Bruno W Bruckmann Air cooling of turbine blades
US2786646A (en) * 1949-08-10 1957-03-26 Power Jets Res & Dev Ltd Bladed rotors for axial flow turbines and similarly bladed fluid flow machines

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE737479C (en) * 1939-01-26 1943-07-15 Versuchsanstalt Fuer Luftfahrt Gas turbine blade
DE924248C (en) * 1941-03-28 1955-02-28 Daimler Benz Ag Huelsenhohlschaufel for gas or exhaust gas turbines
US2656147A (en) * 1946-10-09 1953-10-20 English Electric Co Ltd Cooling of gas turbine rotors
US2641439A (en) * 1947-10-01 1953-06-09 Chrysler Corp Cooled turbine or compressor blade
US2641440A (en) * 1947-11-18 1953-06-09 Chrysler Corp Turbine blade with cooling means and carrier therefor
US2779565A (en) * 1948-01-05 1957-01-29 Bruno W Bruckmann Air cooling of turbine blades
US2700530A (en) * 1948-08-27 1955-01-25 Chrysler Corp High temperature elastic fluid apparatus
US2786646A (en) * 1949-08-10 1957-03-26 Power Jets Res & Dev Ltd Bladed rotors for axial flow turbines and similarly bladed fluid flow machines

Cited By (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3112914A (en) * 1960-08-01 1963-12-03 Gen Motors Corp Turbine rotor
US3761200A (en) * 1970-12-05 1973-09-25 Secr Defence Bladed rotors
FR2471474A1 (en) * 1979-12-17 1981-06-19 United Technologies Corp Rotor disc
US4451204A (en) * 1981-03-25 1984-05-29 Rolls-Royce Limited Aerofoil blade mounting
US4606701A (en) * 1981-09-02 1986-08-19 Westinghouse Electric Corp. Tip structure for a cooled turbine rotor blade
US5957660A (en) * 1997-02-13 1999-09-28 Bmw Rolls-Royce Gmbh Turbine rotor disk
EP0859127A1 (en) * 1997-02-13 1998-08-19 BMW Rolls-Royce GmbH Channeling of cooling air in a turbine rotor disc
CN102365423A (en) * 2009-03-27 2012-02-29 西门子公司 Axial turbomachine rotor having blade cooling
EP2233692A1 (en) * 2009-03-27 2010-09-29 Siemens Aktiengesellschaft Axial turboengine rotor with rotor cooling
JP2012522160A (en) * 2009-03-27 2012-09-20 シーメンス アクティエンゲゼルシャフト Rotor of axial-flow turbomachine that cools blades
WO2010108972A1 (en) * 2009-03-27 2010-09-30 Siemens Aktiengesellschaft Axial turbomachine rotor having blade cooling
US20110182751A1 (en) * 2010-01-22 2011-07-28 Rolls-Royce Plc Rotor disc
EP2348191A2 (en) 2010-01-22 2011-07-27 Rolls-Royce plc A Rotor Disc
US8708657B2 (en) 2010-01-22 2014-04-29 Rolls-Royce Plc Rotor Disc
US20110182738A1 (en) * 2010-01-27 2011-07-28 Herbert Chidsey Roberts Method and apparatus for a segmented turbine bucket assembly
CH704716A1 (en) * 2011-03-22 2012-09-28 Alstom Technology Ltd Rotor disk for a turbine rotor and turbine as well as with such a rotor disk.

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