US3597109A - Gas turbine engine axial flow multistage compressor - Google Patents

Gas turbine engine axial flow multistage compressor Download PDF

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US3597109A
US3597109A US826803A US3597109DA US3597109A US 3597109 A US3597109 A US 3597109A US 826803 A US826803 A US 826803A US 3597109D A US3597109D A US 3597109DA US 3597109 A US3597109 A US 3597109A
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rotor
platform
compressor
members
blade
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US826803A
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James Alexander Petrie
Kenneth Edward George Bracey
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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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • 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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/06—Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections

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  • the invention concerns a gas turbine engine axial flow multistage compressor having a rotor member, and a plurality of platform elements each of which has rotor blade members mounted thereon and each of which is itself disposed radially outwardly of and spaced from the rotor member, the plurality of platform elements defining an annular platform member and the axially consecutive annular platform member of different stages of rotor blade members contacting each other to form a drumlike structure.
  • FIGS. 1 A first figure.
  • GAS TURBINE ENGINE AXIAL FLOW MULTISTAGE COMPRESSOR This invention concerns a gas turbine engine axial flow multistage compressor.
  • a gas turbine engine axial flow multistage compressor having rotor blade members each of which is mounted on a platform which is itself disposed radially outwardly of and spaced from a rotor member, the axially consecutive platforms of different stages of rotor blade members contacting each other to form a drumlike structure.
  • Each platform may carry a plurality of rotor blade members ofthe same stage.
  • each platform may carry a plurality of rotor blade members ofdifferent stages.
  • Each platform is preferably connected by connecting structure to a root portion which is mounted in a rotor disc, the connecting structure extending radially outwardly of the rotor disc.
  • the connecting structure may comprise a plurality of rib members, each platform may be integral with its blade or blade members, connecting structure and root portion.
  • Angularly consecutive platforms may be connected to root portions which contact each other and which are mounted in a common slot in a rotor disc.
  • FIG. 1 is a broken-away plan view of part of a gas turbine engine axial flow multistage compressor according to the present invention.
  • FIG. 2 is a perspective view of a blade assembly forming part ofthe compressor in FIG. 1,
  • FIG. 3 is a broken-away plan view of part of another embodiment of a gas turbine engine axial flow multistage compressor according to the present invention
  • FIGS. 4 and 5 are respectively an elevation and a plan view of a blade assembly forming part of the compressor shown in FIG. 3.
  • FIG. 6 is a broken-away sectional view through part of yet another embodiment of a gas turbine engine axial flow multistage compressor according to the present invention.
  • FIGS. 7 and 8 are sections taken respectively on the lines 7-7 and 88 of FIG. 6, and
  • FIG. 9 is a plan view looking in the direction of the arrow 9 of FIG. 6.
  • a gas turbine engine, axial flow, multistage, high-pressure compressor 10 comprises rotor discs l1, 12 which are carried by a shaft 13, each of the discs ll, 12 being provided with a plurality ofangularly spacedapart blade assemblies 14.
  • each blade assembly 14 comprises two adjacent angularly consecutive platforms l5, 16 which are mounted in contact with each other and which extend axially of the compressor.
  • the platforms 15, 16 are respectively integral, via a plurality ofrib members 17, with root portions 20, 21 which collectively form a fir-tree-shaped root and are mounted in a common slot 22 in the respective rotor disc ll, 12.
  • the platform I5 is provided with two rotor blade members 23, 24 which are mounted thereon and integral therewith wh le the platform 16 is similarly provided with two rotor blade members 25, 26.
  • the rotor blade members 23, 25 con stitute members of a common rotor stage, while the rotor blade members 24, 26 constitute members of a different, i.e. the adjacent. rotor stage.
  • each of the platforms [5, 16 supports rotor blade members of two adjacent rotor stages.
  • each of the platforms 15, which carries two consecutive stages of rotor blade members 23, 24 contacts the axially consecutive platform 15, which carries the next two consecutive stages of rotor blade members 23, 24, the axially consecutive platforms 15 contacting each other at their adjacent edges 27 to form a drumlike structure.
  • the axially consecutive platforms 16 con tact each other at their adjacent edges.
  • the compressor 10 is also provided with stator blades 28.
  • annular flange member extends between and is' sealed to adjacent rotor discs and forms with the radial extremities of the adjacent rotor discs an annular groove.
  • annular groove there extends an annular inner shroud structure of the respective stator row, this inner shroud structure having sealing members which cooperate with corresponding sealing members on the annular flange member. Since, however, in the case of the structure shown in FIGS.
  • the axially consecutive platforms 15, 16 contact each other to form a drumlike structure, the sealing losses which normally occur through the said annular groove, and the heating of the rotor discs due to air friction, are avoided, while the stator blades 28 do not need to be provided with the said inner shroud structure, so that their efficiency is improved.
  • the provision of the radially elongated rib members 17 also ensures that the air which is heated by being compressed in the compressor 10 is kept well away from the rotor discs 11, 12. This is of particular value in a high-pressure compressor having a high overall compression ratio since the temperature of the air at the downstream end of such a compressor would otherwise necessitate that the rotor discs should be formed of high-temperature-resistant materials which are heavy and expensive.
  • FIGS. 1 and 2 it is necessary to use somewhat heavier blade members and rib members than usual, it is also possible to employ lighter materials in the rotor discs than would otherwise be the case and thus to effect an overall saving in weight.
  • each root portion 20, 21 carries a plurality of rotor blade members, this space problem is solved.
  • each root portion may be cast integrally with its rib members, platform and blade members, the assembly of the compressor is simplified by reason of a reduction in the number of parts requiring to be assembled together.
  • a gas turbine en gine axial flow, multistage compressor 30 has rotor discs 31, 32, which are carried by a shaft 33 and each of which is pro vided with a plurality of angularly spaced-apart blade assemblies 34.
  • Each of the blade assemblies 34 has four rotor blade members 35 (see FIGS. 4 and 5) which are integral with and mounted on a common platform 36.
  • the rotor blade members 35 are all members of a common rotor stage.
  • the platform 36 of each of the blade assemblies 34 is connected to a root portion 40 by a plurality of rib members 41.
  • the rib members 41 by means of which each of the platforms 36 is disposed radially outwardly of and is spaced from the rotor discs 31, 32, extend radially outwardly of the latter throughout a radial length which exceeds that of the rotor blade members 35.
  • the compressor 30 is also provided with stator blades 42.
  • the axially consecutive plat forms 36 ofdifferent stages of rotor blade members 35 contact each other to form a drumlike structure.
  • a gas turbine engine axial flow, multistage compressor 43 (FIG. 6) has a plurality of rotor discs 44 each of which is provided with a plurality of angularly spaced-apart blade assemblies 45.
  • Each of the blade assemblies 45 has two rotor blade members 46 47, which form members of different rotor stages and which are integral with and mounted on a common platform 50.
  • the platform 50 is itself integral with rib members 51, 52. 53 which serve to connect the platform 50 to a root portion 54 by means ofwhich the blade assembly is mounted in the disc 44.
  • each ofthe platforms 50 is disposed radially outwardly of and is spaced from the rotor disc 44 extend outwardly of the latter throughout a radial length which exceeds that of the rotor blade members 46, 47.
  • each of the platforms 50 which carries two consecutive stages of rotor blade members 46, 47, contacts the axially consecutive platform 50, which carries the next two consecutive stages of rotor blade members 46, 47, the axially consecutive platforms 50 contacting each other at their adjacent edges 55 to form a drumlike structure.
  • a gas turbine engine axial flow multistage compressor comprising: a plurality of axially spaced rotor discs; each of said rotor discs supporting a plurality of angularly spaced blade assemblies, the angularly spaced blade assemblies of one of said rotor discs being axially spaced with respect to the angularly spaced blade assemblies of an adjacent one of said discs; each blade assembly of said plurality of angularly spaced blade assemblies being an integral structure including a root portion mounted in a slot in a respective rotor disc, a connecting structure extending from the root portion radially outwardly of the respective rotor disc, a platform element disposed on the outer end of said connecting structure, and a plurality of rotor blade members mounted on said platform element; and said platform elements of said angularly spaced blade assemblies for the respective rotor discs defining an annular platform member and consecutive annular platform members of axially spaced-apart rotor discs contacting one another to form a drumlike structure.
  • each platform element carries a plurality of rotor blade members of a same stage.
  • each platform element carries a plurality of rotor blade members of different stages.
  • each blade assembly comprises at least one rib member angularly spaced from a rib member of an adjacent blade assembly.
  • a compressor as claimed in claim I in which adjacent angularly spaced blade assemblies have root portions which contact each other and which are mounted in a common slot in the respective rotor disc.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

The invention concerns a gas turbine engine axial flow multistage compressor having a rotor member, and a plurality of platform elements each of which has rotor blade members mounted thereon and each of which is itself disposed radially outwardly of and spaced from the rotor member, the plurality of platform elements defining an annular platform member and the axially consecutive annular platform member of different stages of rotor blade members contacting each other to form a drumlike structure.

Description

United States Patent Inventors James Alexander Petrie Llttleover; Kenneth Edward George Bracey, Findern, both of Derby. England Appl. No 826,803 Filed May 22, 1969 Patented Aug. 3, 1971 Assignee Rolb-Royce Limited Derby, England Priority May 31, 1968 Great Britain 26169/68 GAS TURBINE ENGINE AXIAL FLOW MULTISI AGE COMPRESSOR 5 Claims, 9 Drawing Figs.
US. Cl 416/198, 416/193, 416/201, 416/210, 416/217, 416/244 Int. Cl F0ld 5/30 Field oiSearch 416/198,
20l,2l0,213,2l5,217,244
[56] References Cited UNITED STATES PATENTS 2,869,820 l/1959 Marchant et al 416/201 2,951,677 9/1960 Howald t 416/201 3,010,643 11/1961 Ricketts 416/201 X 3,249,293 5/1966 Koff 416/198 Primary Examiner- Everette A. Powell, Jr. AtlorneyCushman, Darby & Cushman ABSTRACT: The invention concerns a gas turbine engine axial flow multistage compressor having a rotor member, and a plurality of platform elements each of which has rotor blade members mounted thereon and each of which is itself disposed radially outwardly of and spaced from the rotor member, the plurality of platform elements defining an annular platform member and the axially consecutive annular platform member of different stages of rotor blade members contacting each other to form a drumlike structure.
Patented Aug. 3, 1971 3,597,109
5 Sheets-Sheet 1 ATTORNEYS Patented Aug. 3, 1971 5 Sheets-Sheet 2 Patented Aug. 3, 1971 5 Sheets-Sheet 5 ii 30 F/G.3.
M r W Patented Aug. 3, 1971 5 Sheets-Sheet 4 FIGS.
FIG]
FIGS;
7 M ATTORNEY h rfdda Patented Aug. 3, 1971 3,597,109
5 Sheets-Sheet 5 F/GB.
B ATTORNEYS:
GAS TURBINE ENGINE AXIAL FLOW MULTISTAGE COMPRESSOR This invention concerns a gas turbine engine axial flow multistage compressor.
According to the present invention, there is provided a gas turbine engine axial flow multistage compressor having rotor blade members each of which is mounted on a platform which is itself disposed radially outwardly of and spaced from a rotor member, the axially consecutive platforms of different stages of rotor blade members contacting each other to form a drumlike structure.
Each platform may carry a plurality of rotor blade members ofthe same stage.
Additionally, or alternatively, each platform may carry a plurality of rotor blade members ofdifferent stages.
Each platform is preferably connected by connecting structure to a root portion which is mounted in a rotor disc, the connecting structure extending radially outwardly of the rotor disc.
The connecting structure may comprise a plurality of rib members, each platform may be integral with its blade or blade members, connecting structure and root portion.
Angularly consecutive platforms may be connected to root portions which contact each other and which are mounted in a common slot in a rotor disc.
The invention is illustrated, merely by way of example, in the accompanying drawings, in which:
FIG. 1 is a broken-away plan view of part of a gas turbine engine axial flow multistage compressor according to the present invention.
FIG. 2 is a perspective view of a blade assembly forming part ofthe compressor in FIG. 1,
FIG. 3 is a broken-away plan view of part of another embodiment of a gas turbine engine axial flow multistage compressor according to the present invention,
FIGS. 4 and 5 are respectively an elevation and a plan view of a blade assembly forming part of the compressor shown in FIG. 3.
FIG. 6 is a broken-away sectional view through part of yet another embodiment of a gas turbine engine axial flow multistage compressor according to the present invention,
FIGS. 7 and 8 are sections taken respectively on the lines 7-7 and 88 of FIG. 6, and
FIG. 9 is a plan view looking in the direction of the arrow 9 of FIG. 6.
Referring first to the embodiment shown in FIGS. 1 and 2, a gas turbine engine, axial flow, multistage, high-pressure compressor 10 comprises rotor discs l1, 12 which are carried by a shaft 13, each of the discs ll, 12 being provided with a plurality ofangularly spacedapart blade assemblies 14.
As best seen in FIG. 2, each blade assembly 14 comprises two adjacent angularly consecutive platforms l5, 16 which are mounted in contact with each other and which extend axially of the compressor. The platforms 15, 16 are respectively integral, via a plurality ofrib members 17, with root portions 20, 21 which collectively form a fir-tree-shaped root and are mounted in a common slot 22 in the respective rotor disc ll, 12.
The platform I5 is provided with two rotor blade members 23, 24 which are mounted thereon and integral therewith wh le the platform 16 is similarly provided with two rotor blade members 25, 26. The rotor blade members 23, 25 con stitute members of a common rotor stage, while the rotor blade members 24, 26 constitute members of a different, i.e. the adjacent. rotor stage. Thus each of the platforms [5, 16 supports rotor blade members of two adjacent rotor stages.
The rib members 17, by means of which each of the platforms l5, I6 is disposed radially outwardly of and is spaced from the rotor disc 11, 12, extend radially outwardly of the rotor discs ll, 12 throughout a radial length 1, which is at least 50 percent of(and which may, indeed, be at least equal to) the radial length 1 of the rotor blade members 23-26.
As will be seen from FIG. 1, each of the platforms 15, which carries two consecutive stages of rotor blade members 23, 24 contacts the axially consecutive platform 15, which carries the next two consecutive stages of rotor blade members 23, 24, the axially consecutive platforms 15 contacting each other at their adjacent edges 27 to form a drumlike structure. Similarly, of course, the axially consecutive platforms 16 con tact each other at their adjacent edges.
The compressor 10 is also provided with stator blades 28.
The structure described above has a number of advantages.
In a conventional compressor, an annular flange member extends between and is' sealed to adjacent rotor discs and forms with the radial extremities of the adjacent rotor discs an annular groove. Into this annular groove there extends an annular inner shroud structure of the respective stator row, this inner shroud structure having sealing members which cooperate with corresponding sealing members on the annular flange member. Since, however, in the case of the structure shown in FIGS. 1 and 2, the axially consecutive platforms 15, 16 contact each other to form a drumlike structure, the sealing losses which normally occur through the said annular groove, and the heating of the rotor discs due to air friction, are avoided, while the stator blades 28 do not need to be provided with the said inner shroud structure, so that their efficiency is improved.
The provision of the radially elongated rib members 17 also ensures that the air which is heated by being compressed in the compressor 10 is kept well away from the rotor discs 11, 12. This is of particular value in a high-pressure compressor having a high overall compression ratio since the temperature of the air at the downstream end of such a compressor would otherwise necessitate that the rotor discs should be formed of high-temperature-resistant materials which are heavy and expensive. Although in the construction shown in FIGS. 1 and 2 it is necessary to use somewhat heavier blade members and rib members than usual, it is also possible to employ lighter materials in the rotor discs than would otherwise be the case and thus to effect an overall saving in weight. The provision of the radially elongated rib members would normally increase the volume of the above-mentioned annular groove and therefore increase the sealing losses to unacceptably high values. Since, however, in the construction of FIGS. 1 and 2, the said annular groove is avoided, the said construction is also such as to permit the use of the said radially elongated rib members.
The provision of these radially elongated rib members would, moreover, normally make it difficult to find sufficient circumferential space on the rim of the rotor disc to mount each rotor blade member individually therein. However, since each root portion 20, 21 carries a plurality of rotor blade members, this space problem is solved. At the same time, since each root portion may be cast integrally with its rib members, platform and blade members, the assembly of the compressor is simplified by reason of a reduction in the number of parts requiring to be assembled together.
In the embodiment shown in FIGS. 3 to 5, a gas turbine en gine axial flow, multistage compressor 30 has rotor discs 31, 32, which are carried by a shaft 33 and each of which is pro vided with a plurality of angularly spaced-apart blade assemblies 34. Each of the blade assemblies 34 has four rotor blade members 35 (see FIGS. 4 and 5) which are integral with and mounted on a common platform 36. The rotor blade members 35 are all members of a common rotor stage.
The platform 36 of each of the blade assemblies 34 is connected to a root portion 40 by a plurality of rib members 41. The rib members 41, by means of which each of the platforms 36 is disposed radially outwardly of and is spaced from the rotor discs 31, 32, extend radially outwardly of the latter throughout a radial length which exceeds that of the rotor blade members 35.
The compressor 30 is also provided with stator blades 42.
As will be seen from FIG. 3, the axially consecutive plat forms 36 ofdifferent stages of rotor blade members 35 contact each other to form a drumlike structure.
In the embodiment illustrated in FlGS. 6 to 9, a gas turbine engine axial flow, multistage compressor 43 (FIG. 6) has a plurality of rotor discs 44 each of which is provided with a plurality of angularly spaced-apart blade assemblies 45. Each of the blade assemblies 45 has two rotor blade members 46 47, which form members of different rotor stages and which are integral with and mounted on a common platform 50. The platform 50 is itself integral with rib members 51, 52. 53 which serve to connect the platform 50 to a root portion 54 by means ofwhich the blade assembly is mounted in the disc 44.
The rib members 51, S2, 53, by means of which each ofthe platforms 50 is disposed radially outwardly of and is spaced from the rotor disc 44 extend outwardly of the latter throughout a radial length which exceeds that of the rotor blade members 46, 47.
As will'be seen from FIG. 6, each of the platforms 50, which carries two consecutive stages of rotor blade members 46, 47, contacts the axially consecutive platform 50, which carries the next two consecutive stages of rotor blade members 46, 47, the axially consecutive platforms 50 contacting each other at their adjacent edges 55 to form a drumlike structure.
We claim:
1. A gas turbine engine axial flow multistage compressor comprising: a plurality of axially spaced rotor discs; each of said rotor discs supporting a plurality of angularly spaced blade assemblies, the angularly spaced blade assemblies of one of said rotor discs being axially spaced with respect to the angularly spaced blade assemblies of an adjacent one of said discs; each blade assembly of said plurality of angularly spaced blade assemblies being an integral structure including a root portion mounted in a slot in a respective rotor disc, a connecting structure extending from the root portion radially outwardly of the respective rotor disc, a platform element disposed on the outer end of said connecting structure, and a plurality of rotor blade members mounted on said platform element; and said platform elements of said angularly spaced blade assemblies for the respective rotor discs defining an annular platform member and consecutive annular platform members of axially spaced-apart rotor discs contacting one another to form a drumlike structure.
2. A compressor as claimed in claim 1 in which each platform element carries a plurality of rotor blade members of a same stage.
3. A compressor as claimed in claim 1 in which each platform element carries a plurality of rotor blade members of different stages.
4. A compressor as claimed in claim 1 in which said connecting structure of each blade assembly comprises at least one rib member angularly spaced from a rib member of an adjacent blade assembly.
5. A compressor as claimed in claim I in which adjacent angularly spaced blade assemblies have root portions which contact each other and which are mounted in a common slot in the respective rotor disc.

Claims (5)

1. A gas turbine engine axial flow multistage compressor comprising: a plurality of axially spaced rotor discs; each of said rotor discs supporting a plurality of angularly spaced blade assemblies, the angularly spaced blade assemblies of one of said rotor discs being axially spaced with respect to the angularly spaced blade assemblies of an adjacent one of said discs; each blade assembly of said plurality of angularly spaced blade assemblies being an integral structure including a root portion mounted in a slot in a respective rotor disc, a connecting structure extending from the root portion radially outwardly of the respective rotor disc, a platform element disposed on the outer end of said connecting structure, and a plurality of rotor blade members mounted on said platform element; and said platform elements of said angularly spaced blade assemblies for the respective rotor discs defining an annular platform member and consecutive annular platform members of axially spaced-apart rotor discs contacting one another to form a drumlike structure.
2. A compressor as claimed in claim 1 in which each platform element carries a plurality of rotor blade members of a same stage.
3. A compressor as claimed in claim 1 in which each platform element carries a plurality of rotor blade members of different stages.
4. A compressor as claimed in claim 1 in which said connecting structure of each blade assembly comprises at least one rib member angularly spaced from a rib member of an adjacent blade assembly.
5. A compressor as claimed iN claim 1 in which adjacent angularly spaced blade assemblies have root portions which contact each other and which are mounted in a common slot in the respective rotor disc.
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Cited By (16)

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Publication number Priority date Publication date Assignee Title
US4016636A (en) * 1974-07-23 1977-04-12 United Technologies Corporation Compressor construction
US4130379A (en) * 1977-04-07 1978-12-19 Westinghouse Electric Corp. Multiple side entry root for multiple blade group
US4265594A (en) * 1978-03-02 1981-05-05 Bbc Brown Boveri & Company Limited Turbine blade having heat localization segments
US4310286A (en) * 1979-05-17 1982-01-12 United Technologies Corporation Rotor assembly having a multistage disk
US4483659A (en) * 1983-09-29 1984-11-20 Armstrong Richard J Axial flow impeller
US5425622A (en) * 1993-12-23 1995-06-20 United Technologies Corporation Turbine blade attachment means
US5735673A (en) * 1996-12-04 1998-04-07 United Technologies Corporation Turbine engine rotor blade pair
US5743713A (en) * 1995-09-21 1998-04-28 Ngk Insulators, Ltd. Blade, turbine disc and hybrid type gas turbine blade
US6416276B1 (en) * 1999-03-29 2002-07-09 Alstom (Switzerland) Ltd Heat shield device in gas turbines
US20100111673A1 (en) * 2008-11-05 2010-05-06 General Electric Company Turbine with interrupted purge flow
ITTO20090522A1 (en) * 2009-07-13 2011-01-14 Avio Spa TURBOMACCHINA WITH IMPELLER WITH BALLED SEGMENTS
US20110200440A1 (en) * 2008-11-13 2011-08-18 Frank Stiehler Blade cluster having an offset axial mounting base
US20140174098A1 (en) * 2012-12-20 2014-06-26 United Technologies Corporation Turbine disc with reduced neck stress concentration
US20160108735A1 (en) * 2014-10-16 2016-04-21 United Technologies Corporation Tandem rotor blades
US9551353B2 (en) 2013-08-09 2017-01-24 General Electric Company Compressor blade mounting arrangement
US11549518B2 (en) * 2017-07-06 2023-01-10 Raytheon Technologies Corporation Tandem blade rotor disk

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DE1602951C3 (en) * 1967-03-08 1982-09-23 Thyssen Industrie Ag, 4300 Essen Multi-spindle drilling machine
DE102008034738A1 (en) * 2008-07-24 2010-01-28 Rolls-Royce Deutschland Ltd & Co Kg Compressor rotor for turbo-engine for use in aircraft industry, has hub, disk collar and shovel that is assembled to rotor blade carriers

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US2951677A (en) * 1956-03-12 1960-09-06 Curtiss Wright Corp Turbine rotor construction
US3010643A (en) * 1955-12-23 1961-11-28 Bristol Siddeley Engines Ltd Axial flow compressors
US3249293A (en) * 1964-01-23 1966-05-03 Gen Electric Ring-drum rotor

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US2869820A (en) * 1951-04-18 1959-01-20 Bristol Aero Engines Ltd Rotors for axial flow compressors or turbines
US3010643A (en) * 1955-12-23 1961-11-28 Bristol Siddeley Engines Ltd Axial flow compressors
US2951677A (en) * 1956-03-12 1960-09-06 Curtiss Wright Corp Turbine rotor construction
US3249293A (en) * 1964-01-23 1966-05-03 Gen Electric Ring-drum rotor

Cited By (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4016636A (en) * 1974-07-23 1977-04-12 United Technologies Corporation Compressor construction
US4130379A (en) * 1977-04-07 1978-12-19 Westinghouse Electric Corp. Multiple side entry root for multiple blade group
US4265594A (en) * 1978-03-02 1981-05-05 Bbc Brown Boveri & Company Limited Turbine blade having heat localization segments
US4310286A (en) * 1979-05-17 1982-01-12 United Technologies Corporation Rotor assembly having a multistage disk
US4483659A (en) * 1983-09-29 1984-11-20 Armstrong Richard J Axial flow impeller
US5425622A (en) * 1993-12-23 1995-06-20 United Technologies Corporation Turbine blade attachment means
US5743713A (en) * 1995-09-21 1998-04-28 Ngk Insulators, Ltd. Blade, turbine disc and hybrid type gas turbine blade
US5735673A (en) * 1996-12-04 1998-04-07 United Technologies Corporation Turbine engine rotor blade pair
EP0846845B1 (en) * 1996-12-04 2005-11-09 United Technologies Corporation Rotor blade pair and rotor comprising such a blade pair
US6416276B1 (en) * 1999-03-29 2002-07-09 Alstom (Switzerland) Ltd Heat shield device in gas turbines
US20100111673A1 (en) * 2008-11-05 2010-05-06 General Electric Company Turbine with interrupted purge flow
JP2010112376A (en) * 2008-11-05 2010-05-20 General Electric Co <Ge> Turbine with interrupted purge flow
US8137067B2 (en) 2008-11-05 2012-03-20 General Electric Company Turbine with interrupted purge flow
US20110200440A1 (en) * 2008-11-13 2011-08-18 Frank Stiehler Blade cluster having an offset axial mounting base
ITTO20090522A1 (en) * 2009-07-13 2011-01-14 Avio Spa TURBOMACCHINA WITH IMPELLER WITH BALLED SEGMENTS
US20140174098A1 (en) * 2012-12-20 2014-06-26 United Technologies Corporation Turbine disc with reduced neck stress concentration
US9551353B2 (en) 2013-08-09 2017-01-24 General Electric Company Compressor blade mounting arrangement
US20160108735A1 (en) * 2014-10-16 2016-04-21 United Technologies Corporation Tandem rotor blades
US10598024B2 (en) * 2014-10-16 2020-03-24 United Technologies Corporation Tandem rotor blades
US11852034B2 (en) 2014-10-16 2023-12-26 Rtx Corporation Tandem rotor blades
US11549518B2 (en) * 2017-07-06 2023-01-10 Raytheon Technologies Corporation Tandem blade rotor disk
US20230116394A1 (en) * 2017-07-06 2023-04-13 Raytheon Technologies Corporation Tandem blade rotor disk
US12049904B2 (en) * 2017-07-06 2024-07-30 Rtx Corporation Tandem blade rotor disk
US20240352942A1 (en) * 2017-07-06 2024-10-24 Rtx Corporation Tandem blade rotor disk

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GB1217275A (en) 1970-12-31
DE1927021A1 (en) 1970-04-02
FR2011882A1 (en) 1970-03-13

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