US3801222A - Platform for compressor or fan blade - Google Patents

Platform for compressor or fan blade Download PDF

Info

Publication number
US3801222A
US3801222A US00229872A US3801222DA US3801222A US 3801222 A US3801222 A US 3801222A US 00229872 A US00229872 A US 00229872A US 3801222D A US3801222D A US 3801222DA US 3801222 A US3801222 A US 3801222A
Authority
US
United States
Prior art keywords
platform
blade
compressor
disc
arms
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US00229872A
Inventor
J Violette
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Technologies Corp
Original Assignee
United Aircraft Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Aircraft Corp filed Critical United Aircraft Corp
Application granted granted Critical
Publication of US3801222A publication Critical patent/US3801222A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • F01D11/006—Sealing the gap between rotor blades or blades and rotor
    • F01D11/008—Sealing the gap between rotor blades or blades and rotor by spacer elements between the blades, e.g. independent interblade platforms
    • 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
    • F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/322—Blade mountings
    • 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

  • ABSTRACT 52 U.S. Cl. 416 220 416 248 ⁇ mi 1m. 01. 1 01a s/32
  • the a blade is 58 Field of Search 416/219-221, into two complementary Separate halves 416/248 215 214 193 adapted to surround the root of each blade to define together with a plurality thereof a smooth aerody- [56] References Cited namic surface and each half having depending arms at the ends adapted to fit into the dovetail of the blade UNITED STATES PATENTS disc transmitting the centrifugal load of the platform 2,656,146 Sollinger X independent of the blade loads 2,749,029 6/1956 Goetzel et al 416/248 X 8 Claims, 4 Drawing Figures PLATFORM FOR COMPRESSOR OR FAN BLADE BACKGROUND OF THE INVENTION
  • This invention relates to the platform of a compressor and moreparticularly to an improved platform that serves the multi-functions of providing an aerodynamically smooth surface, seal
  • the compressor of a jet engine typically includes a platform formed adjacent the disc and in proximity to the blade root.
  • the platform normally serves to define an aerodynamic smooth surface and seals the working medium at this junction point.
  • the platform is formed integrally with the blade and is not a load carrying member. In the integral platform-blade configuration care must be taken in the manufacturing thereof so that sufficient strength at the point where the platform radiates from the blade is provided since this is a high stress point, where breakage from fatigue is likely. Thus, in certain applications this area of the blade and platform must be beefed up. Additionally the sealing around the platform is generally accomplished by a seal extension machined integral with the disc which has proven to be a difficult and expensive operation.
  • the primary object of this invention is to provide a platform for a compressor that is fabricated in-separate halves that surround the compressor blade at its tab ending and carries depending arms adapted to fit into the dovetail slot of the disc in such a manner as to transmit thereto the certrifugal loads produced by the rotating platform.
  • a further object of this invention is to provide an improved compressor platform fabricated into separate complementary halves that reduces the tab ending stress, provides means for transmitting the centrifugal load to the compressor disc, eliminates the disc seal structure and is especially efficacious for compressor blades fabricated from composite materials.
  • FIG. 1 is an exploded perspective view of the blade and platform.
  • FIG. 2 is a partial view in perspective of the assembled blade and platform of FIG. 1.
  • FIG. 3 is a partial view partly'in elevation and partly in section, taken along line 3-3 of FIG. 4, showing the platform-blade mounted in thedovetail slot of the compressor disc.
  • FIG. 4 is a partial view in elevation taken endwise of FIG. 3.
  • FIGS. 1-4 illustrating the preferred embodiment of this invention as used on a compressor for a turbine type of power plant.
  • the blade 10 comprises an airfoil shaped shell 12 which preferably is fabricated from a composite material such as boron epoxy or boron aluminum as described andclaimed in a patent application entitled Composite Fan Blade and Method of Construction filed by E. Rothman on even date and assigned to the same assignee.
  • the spar 14, extends beyond shell 10 to form the root of the blade and defines a retention 15 that is adapted to fit into the dovetailslot 16 of compressor disc 18.
  • the compressor disc obviously carries a plurality of such slots for supporting a like number of blades to form the'compressor rotor and only one being shown for the sake of clarity and simplicity.
  • the separable platform 20 is fabricated into two separate halves, the camber side half 22 and the face side half 24.
  • the platform like the core is fabricated from a suitable metal such as titanium.
  • the inner edge of the platform surface 28 of face-side half 24 and the inner edge 30 of the platform 32 are shaped to complement the face and camber sides of shell 12.
  • Each platform half carries fore and aft depending arms 34,-34'- and 36, 36' the primed referenced numerals refer to like reference numerals being mating arms and each pair of arms being dimensioned to overlie the edges or end faces of retention 15 and extend downwardly to be substantially coextensive therewith as can best be seen in FIG. 2.
  • Each arm, 34, 34', 36 and 36' has defined thereon at its lower outer extremity a lip 38 each being identically dimensioned and each having a load transmitting face 40.
  • the load transmitting faces 40 mate with a complementary face formed in the dovetail slot. These mating faces serve to transmit the centrifugal loads which tends to force the platform outwardly in the direction of arrow A produced by the rotation of the platform.
  • Dowel pins 44 may be incorporated in arms 34' and 36 to mate complementary holes 42 formed in arms 34 and 36 to align the halves of the platform. It is apparent from FIG. 2 that the platform when assembled extends above the tab portion lower shell extremity portion of the blade. This provides a smooth transition of the blade shell and blade root when in the assembled position. It will also be apparent that no securing means are necessary to hold the halves in their relative positions.
  • abutting edges 46 those edges that abut the next adjacent platform in the disc 18 form an air seal when assembled so as to prevent flow of air between adjacent fan blades and surfaces 47 prevent flow of air to adjacent stages.
  • junction edges of the platform and shell in the assembled position may be treated with a suitable plastic material such as silicone rubberto seal off the root of the blade from the air working surfaces. Additionally, this treatment serves to permit slight movement of the blade relative to the platform occasioned by vibrations at high speed.
  • a compressor blade platform adapted to define a smooth aerodynamic surface when assembled in a compressor disc, said platform comprising for each blade supported by said compressor disc a pair of complementary elements each having a top platform surface which surfaces are contoured on one edge to closely accommodate the blade extending adjacent thereto, each blade having a root section having end faces extending into a slot formed in the compressor disc, each platform having depending members extending'from the ends thereof into said slot adjacent to and in overlapping relation with said end faces of said root section, each of said depending members having means formed thereon for transmitting the load imposed by the rotation of said platform onto said disc independently of the compressor blade.
  • a compressor blade platform as claimed in claim 4 including lip portions having force reaction surfaces formed on the lower end of said depending arm, a complementary reaction surface formed on said disc in said slot for receiving the load transmitted by said arms when said platform is rotating.
  • a compressor blade a compressor blade platform adapted to define a smooth'aerodynamic surface when assembled in a compressor disc, said compressor blade having a retention formed at its root and supported in a slot formed in the compressor disc, said platform comprising a pair of complementary elements for each blade each having a top substantially flat platform surface which surfaces are contoured on an inner edge to closely surround the blade supported in said compressor disc and extending adjacent thereto, each platform having spaced depending arms substantially dimensioned the length of said retention and extending from the end of said platform into said slot formed in the compressor disc, each of said arms having means formed thereon for transmitting the load imposed by the rotation of said platform onto said disc in-' dependently of the compressor blade and each of said arms overlying the side faces of said root of the blade.
  • each of said depending arms carries rearwardly relative to said blade facing lip portion, a force reaction surface formed on the upper edge of said lip, a complementary reaction face formed in said disc engaging said reaction surface formed on said lip preventing said platform from moving radially outward with respect to the rotating center line when said platform rotates.

Abstract

The platform of a compressor or fan blade is fabricated into two complementary separate halves adapted to surround the root of each blade to define together with a plurality thereof a smooth aerodynamic surface and each half having depending arms at the ends adapted to fit into the dovetail of the blade disc transmitting the centrifugal load of the platform independent of the blade loads.

Description

h v United States Patent 1191 [111 3,801,222 Violette Apr. *2, 1974 [54] PLATFORM FOR COMPRESSOR OR FAN 2,819,870 l/1958 Wayne 416/248 BLADE 2,825,530 3/1958 Schum et al..... 416/248 X 2,967,043 1/1961 Dennis 416/221 Inventor: J A- Violette, y, n 3,008,689 11/1961 Morley et al. 416 221 [731 Assignees: United Aircm corpomfiqn, Egslwh 3,294,364 12/1966 Stanley 416/221 UX I V Hartford, Conn. Primary ExaminerEverette A. Powell, Jr. [22] Filed 1972 Attorney, Agent, or FirmNorman Friedland [21] Appl. No.: 229,872
' [57] ABSTRACT 52 U.S. Cl. 416 220 416 248 {mi 1m. 01. 1 01a s/32 The a blade is 58 Field of Search 416/219-221, into two complementary Separate halves 416/248 215 214 193 adapted to surround the root of each blade to define together with a plurality thereof a smooth aerody- [56] References Cited namic surface and each half having depending arms at the ends adapted to fit into the dovetail of the blade UNITED STATES PATENTS disc transmitting the centrifugal load of the platform 2,656,146 Sollinger X independent of the blade loads 2,749,029 6/1956 Goetzel et al 416/248 X 8 Claims, 4 Drawing Figures PLATFORM FOR COMPRESSOR OR FAN BLADE BACKGROUND OF THE INVENTION This invention relates to the platform of a compressor and moreparticularly to an improved platform that serves the multi-functions of providing an aerodynamically smooth surface, seal and centrifugal load transmitting means.
As is well known in the art the compressor of a jet engine typically includes a platform formed adjacent the disc and in proximity to the blade root. The platform normally serves to define an aerodynamic smooth surface and seals the working medium at this junction point. Also, typically, the platform is formed integrally with the blade and is not a load carrying member. In the integral platform-blade configuration care must be taken in the manufacturing thereof so that sufficient strength at the point where the platform radiates from the blade is provided since this is a high stress point, where breakage from fatigue is likely. Thus, in certain applications this area of the blade and platform must be beefed up. Additionally the sealing around the platform is generally accomplished by a seal extension machined integral with the disc which has proven to be a difficult and expensive operation.
I have found that not only can I obviate the stress problem of heretofore known compressor blades but in the case of a core-shell composite blade I can achieve greater dimensional latitude in the tab ending area. This is by virtue of my invention which contemplates fabricating the platform in a pair of mating halves to surround the blade at its tab ending. Additionally by providing properly shaped depending arms on the platform that fit into the dovetail slot of the disc, the platform can also serve the additional function of transmitting the centrifugal loads produced thereby to the disc and relieve this load which in some applications amounts to over one ton force from the platform. Furthermore, in accordance with my invention the complex and costly machine operation of forming the sealing structure on the disc is entirely eliminated.
SUMMARY OF THE INVENTION The primary object of this invention is to provide a platform for a compressor that is fabricated in-separate halves that surround the compressor blade at its tab ending and carries depending arms adapted to fit into the dovetail slot of the disc in such a manner as to transmit thereto the certrifugal loads produced by the rotating platform.
A further object of this invention is to provide an improved compressor platform fabricated into separate complementary halves that reduces the tab ending stress, provides means for transmitting the centrifugal load to the compressor disc, eliminates the disc seal structure and is especially efficacious for compressor blades fabricated from composite materials.
Other features and advantages will be apparent from the specification and claims and from the accompanying drawing which illustrate an embodiment of the invention.
BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is an exploded perspective view of the blade and platform.
FIG. 2 is a partial view in perspective of the assembled blade and platform of FIG. 1.
FIG. 3 is a partial view partly'in elevation and partly in section, taken along line 3-3 of FIG. 4, showing the platform-blade mounted in thedovetail slot of the compressor disc.
FIG. 4 is a partial view in elevation taken endwise of FIG. 3.
DESCRIPTION OF THE PREFERRED EMBODIMENT Reference is now made to FIGS. 1-4 illustrating the preferred embodiment of this invention as used on a compressor for a turbine type of power plant. The blade 10 comprises an airfoil shaped shell 12 which preferably is fabricated from a composite material such as boron epoxy or boron aluminum as described andclaimed in a patent application entitled Composite Fan Blade and Method of Construction filed by E. Rothman on even date and assigned to the same assignee. The spar 14, extends beyond shell 10 to form the root of the blade and defines a retention 15 that is adapted to fit into the dovetailslot 16 of compressor disc 18. The compressor disc obviously carries a plurality of such slots for supporting a like number of blades to form the'compressor rotor and only one being shown for the sake of clarity and simplicity.
The separable platform 20 is fabricated into two separate halves, the camber side half 22 and the face side half 24. Preferably, the platform like the core is fabricated from a suitable metal such as titanium. The inner edge of the platform surface 28 of face-side half 24 and the inner edge 30 of the platform 32 are shaped to complement the face and camber sides of shell 12.
Each platform half carries fore and aft depending arms 34,-34'- and 36, 36' the primed referenced numerals refer to like reference numerals being mating arms and each pair of arms being dimensioned to overlie the edges or end faces of retention 15 and extend downwardly to be substantially coextensive therewith as can best be seen in FIG. 2.
Each arm, 34, 34', 36 and 36' has defined thereon at its lower outer extremity a lip 38 each being identically dimensioned and each having a load transmitting face 40.
As can best be seen in FIG. 4 when retention 15 is inserted in the dovetail slot 16 of the disc 18, the load transmitting faces 40 mate with a complementary face formed in the dovetail slot. These mating faces serve to transmit the centrifugal loads which tends to force the platform outwardly in the direction of arrow A produced by the rotation of the platform.
Dowel pins 44 may be incorporated in arms 34' and 36 to mate complementary holes 42 formed in arms 34 and 36 to align the halves of the platform. It is apparent from FIG. 2 that the platform when assembled extends above the tab portion lower shell extremity portion of the blade. This provides a smooth transition of the blade shell and blade root when in the assembled position. It will also be apparent that no securing means are necessary to hold the halves in their relative positions.
The abutting edges 46, those edges that abut the next adjacent platform in the disc 18 form an air seal when assembled so as to prevent flow of air between adjacent fan blades and surfaces 47 prevent flow of air to adjacent stages. v
If necessary, the junction edges of the platform and shell in the assembled position may be treated with a suitable plastic material such as silicone rubberto seal off the root of the blade from the air working surfaces. Additionally, this treatment serves to permit slight movement of the blade relative to the platform occasioned by vibrations at high speed.
What has been shown by this invention is a twopiece, separable platform with a connection directly to the disc dovetail slot, independent of the blade. This assembly not only reduces tab ending stresses but also alleviates the retention loading occasioned by the centrifugal loads by transmitting this load through the platform directly to the disc.
It should be understood that the invention is not limited to the particular embodiments shown and described herein, but that various changes and modifications may be made without departing from the spirit or scope of this novel concept as defined by the following claims.
I claim:
1. A compressor blade platform adapted to define a smooth aerodynamic surface when assembled in a compressor disc, said platform comprising for each blade supported by said compressor disc a pair of complementary elements each having a top platform surface which surfaces are contoured on one edge to closely accommodate the blade extending adjacent thereto, each blade having a root section having end faces extending into a slot formed in the compressor disc, each platform having depending members extending'from the ends thereof into said slot adjacent to and in overlapping relation with said end faces of said root section, each of said depending members having means formed thereon for transmitting the load imposed by the rotation of said platform onto said disc independently of the compressor blade.
2. A compressor blade platform as claimed in claim 1 wherein said means includes a lip portion having a reaction surface bearing against a complementary surface adjacent said end faces of said root section formed in said compressor disc adjacent the slot formed therein and the root section transmits the load of the each having a top substantially flat platform surface which surfaces are contoured on an inner edge to closely accommodate the blade extending adjacent thereto, each platform having depending arms extending from the end thereof into a slot formed in the compressor disc, each of said arms having means formed thereon for transmitting the load imposed by the rotation of said platform onto said disc independently of the compressor blade, said blade having a retention portion having opposing side end faces formed at its root being supported in said slot formed in said compressor disc and said arms beingdisposed in overlying relation with said end faces.
5. A compressor blade platform as claimed in claim 4 wherein said depending arms are adjacent to and substantially coextensive with said end faces.
6 A compressor blade platform as claimed in claim 4 including lip portions having force reaction surfaces formed on the lower end of said depending arm, a complementary reaction surface formed on said disc in said slot for receiving the load transmitted by said arms when said platform is rotating.
7. In combination, a compressor blade, a compressor blade platform adapted to define a smooth'aerodynamic surface when assembled in a compressor disc, said compressor blade having a retention formed at its root and supported in a slot formed in the compressor disc, said platform comprising a pair of complementary elements for each blade each having a top substantially flat platform surface which surfaces are contoured on an inner edge to closely surround the blade supported in said compressor disc and extending adjacent thereto, each platform having spaced depending arms substantially dimensioned the length of said retention and extending from the end of said platform into said slot formed in the compressor disc, each of said arms having means formed thereon for transmitting the load imposed by the rotation of said platform onto said disc in-' dependently of the compressor blade and each of said arms overlying the side faces of said root of the blade.
8. The combination as claimed in claim 7 wherein each of said depending arms carries rearwardly relative to said blade facing lip portion, a force reaction surface formed on the upper edge of said lip, a complementary reaction face formed in said disc engaging said reaction surface formed on said lip preventing said platform from moving radially outward with respect to the rotating center line when said platform rotates.

Claims (8)

1. A compressor blade platform adapted to define a smooth aerodynamic surface when assembled in a compressor disc, said platform comprising for each blade supported by said compressor disc a pair of complementary elements each having a top platform surface which surfaces are contoured on one edge to closely accommodate the blade extending adjacent thereto, each blade having a root section having end faces extending into a slot formed in the compressor disc, each platform having depending members extending from the ends thereof into said slot adjacent to and in overlapping relation with said end faces of said root section, each of said depending members having means formed thereon for transmitting the load imposed by the rotation of said platform onto said disc independently of the compressor blade.
2. A compressor blade platform as claimed in claim 1 wherein said means includes a lip portion having a reaction surface bearing against a complementary surface adjacent said end faces of said root section formed in said compressor disc adjacent the slot formed therein and the root section transmits the load of the blade and said means transmits the load of the platform.
3. A compressor blade platform as claimed in claim 1 wherein said platform includes a side edge outwardly of said contoured edge adapted to engage the next adjacent platform supported in said compressor disc.
4. A compressor blade platform adapted to define a smooth aerodynamic surface when assembled in a compressor disc, said platform comprising a pair of complementary elements for each blade supported by said disc each having a top substantially flat platform surface which surfaces are contoured on an inner edge to closely accommodate the blade extending adjacent thereto, each platform having depending arms extending from the end thereof into a slot formed in the compressor disc, each of said arms having means formed thereon for transmitting the load imposed by the rotation of said platform onto said disc independently of the compressor blade, said blade having a retention portion having opposing side end faces formed at its root being supported in said slot formed in said compressor disc and said arms being disposed in overlying relation with said end faces.
5. A compressor blade platform as claimed in claim 4 wherein said depending arms are adjacent to and substantially coextensive with said end faces.
6. A compressor blade platform as claimed in claim 4 including lip portions having force reaction surfaces formed on the lower end of said depending arm, a complementary reaction surface formed on said disc in said slot for receiving the load transmitted by said arms when said platform is rotating.
7. In combination, a compressor blade, a compressor blade platform adapted to define a smooth aerodynamic surface when assembled in a compressor disc, said compressor blade having a retention formed at its root and supported in a slot formed in the compressor disc, said platform comprising a pair of complementary elements for each blade each having a top substantially flat platform surface which surfaces are contoured on an inner edge to closely surround the blade supported in said compressor disc and extending adjacent thereto, each platform having spaced depending arms substantially dimensioned the length of said retention and extending from the end of said platform into said slot formed in the compressor disc, each of said arms having means formed thereon for transmitting the load imposed by the rotation of said platform onto said disc independently of the compressor blade and each of said arms overlying the side faces of said Root of the blade.
8. The combination as claimed in claim 7 wherein each of said depending arms carries rearwardly relative to said blade facing lip portion, a force reaction surface formed on the upper edge of said lip, a complementary reaction face formed in said disc engaging said reaction surface formed on said lip preventing said platform from moving radially outward with respect to the rotating center line when said platform rotates.
US00229872A 1972-02-28 1972-02-28 Platform for compressor or fan blade Expired - Lifetime US3801222A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US22987272A 1972-02-28 1972-02-28

Publications (1)

Publication Number Publication Date
US3801222A true US3801222A (en) 1974-04-02

Family

ID=22863003

Family Applications (1)

Application Number Title Priority Date Filing Date
US00229872A Expired - Lifetime US3801222A (en) 1972-02-28 1972-02-28 Platform for compressor or fan blade

Country Status (6)

Country Link
US (1) US3801222A (en)
DE (1) DE2307967C2 (en)
FR (1) FR2174577A5 (en)
GB (1) GB1386575A (en)
IT (1) IT979575B (en)
SE (1) SE389172B (en)

Cited By (50)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4019832A (en) * 1976-02-27 1977-04-26 General Electric Company Platform for a turbomachinery blade
US4045149A (en) * 1976-02-03 1977-08-30 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Platform for a swing root turbomachinery blade
US5030063A (en) * 1990-02-08 1991-07-09 General Motors Corporation Turbomachine rotor
US5183389A (en) * 1992-01-30 1993-02-02 General Electric Company Anti-rock blade tang
US5277548A (en) * 1991-12-31 1994-01-11 United Technologies Corporation Non-integral rotor blade platform
US5318406A (en) * 1992-11-02 1994-06-07 General Electric Company Multipart gas turbine blade
US6447250B1 (en) * 2000-11-27 2002-09-10 General Electric Company Non-integral fan platform
US20060251508A1 (en) * 2005-05-05 2006-11-09 United Technologies Corporation Seal arrangement for a fan-turbine rotor assembly
US20070189901A1 (en) * 2003-03-22 2007-08-16 Dundas Jason E Separable blade platform
US20080159866A1 (en) * 2006-07-21 2008-07-03 Rolls-Royce Plc Fan blade for a gas turbine engine
US20080226457A1 (en) * 2007-03-16 2008-09-18 Snecma Turbomachine rotor disk
US20080286109A1 (en) * 2007-05-15 2008-11-20 Sean Robert Keith Turbine rotor blade and method of fabricating the same
US20080286106A1 (en) * 2007-05-15 2008-11-20 Sean Robert Keith Turbine rotor blade assembly and method of fabricating the same
US20100061858A1 (en) * 2008-09-08 2010-03-11 Siemens Power Generation, Inc. Composite Blade and Method of Manufacture
US20100166551A1 (en) * 2008-12-29 2010-07-01 Morrison Adam J Hybrid turbomachinery component for a gas turbine engine
US20100189562A1 (en) * 2009-01-28 2010-07-29 Snecma Composite material turbomachine blade with a reinforced root
US20100290914A1 (en) * 2009-05-15 2010-11-18 Souers Philip F Blade Closing Key System for a Turbine Engine
US7874804B1 (en) * 2007-05-10 2011-01-25 Florida Turbine Technologies, Inc. Turbine blade with detached platform
US7972113B1 (en) * 2007-05-02 2011-07-05 Florida Turbine Technologies, Inc. Integral turbine blade and platform
US20110243746A1 (en) * 2010-04-06 2011-10-06 General Electric Company Composite turbine bucket assembly
EP2441917A1 (en) * 2010-10-18 2012-04-18 Siemens Aktiengesellschaft Root adapting device and method of attaching a blade in a recess of a rotatable shaft of a steam turbine
US20120156045A1 (en) * 2010-12-17 2012-06-21 General Electric Company Methods, systems and apparatus relating to root and platform configurations for turbine rotor blades
US20120163978A1 (en) * 2010-12-23 2012-06-28 General Electric Company Turbine airfoil components containing ceramic-based materials and processes therefor
US20120301317A1 (en) * 2011-05-26 2012-11-29 Ioannis Alvanos Hybrid rotor disk assembly with ceramic matrix composites platform for a gas turbine engine
US20130004327A1 (en) * 2011-06-30 2013-01-03 United Technologies Corporation Aluminum fan blade root
JP2013139765A (en) * 2012-01-03 2013-07-18 General Electric Co <Ge> Composite airfoil assembly
WO2013185044A1 (en) 2012-06-07 2013-12-12 United Technologies Corporation Fan blade platform
US8721290B2 (en) 2010-12-23 2014-05-13 General Electric Company Processes for producing components containing ceramic-based and metallic materials
US8777582B2 (en) 2010-12-27 2014-07-15 General Electric Company Components containing ceramic-based materials and coatings therefor
US8777583B2 (en) 2010-12-27 2014-07-15 General Electric Company Turbine airfoil components containing ceramic-based materials and processes therefor
US20140212292A1 (en) * 2012-12-18 2014-07-31 United Technologies Corporation Airfoil Member and Composite Platform Having Contoured Endwall
US8827651B2 (en) 2010-11-01 2014-09-09 Rolls-Royce Plc Annulus filler
US8834125B2 (en) 2011-05-26 2014-09-16 United Technologies Corporation Hybrid rotor disk assembly with a ceramic matrix composite airfoil for a gas turbine engine
EP2778347A1 (en) * 2013-03-11 2014-09-17 Siemens Aktiengesellschaft Rotor blade assembly, turbomachine comprising a rotor blade assembly and method of assembling a rotor blade assembly
CN104220700A (en) * 2012-03-29 2014-12-17 西门子公司 Modular turbine blade having platform
US20150017007A1 (en) * 2012-03-30 2015-01-15 Alstom Technology Ltd Turbine blade or vane
WO2015053848A2 (en) 2013-09-18 2015-04-16 United Technologies Corporation Fan platform with leading edge tab
US20150192027A1 (en) * 2012-06-30 2015-07-09 General Electric Company Turbine blade sealing structure
US9145784B2 (en) 2011-04-14 2015-09-29 Rolls-Royce Plc Annulus filler system
US9228444B2 (en) 2011-11-15 2016-01-05 Rolls-Royce Plc Annulus filler
CN106103898A (en) * 2013-11-25 2016-11-09 通用电器技术有限公司 Blade assembly based on modular construction for turbine
US20170058916A1 (en) * 2015-09-01 2017-03-02 United Technologies Corporation Gas turbine fan fairing platform and method of fairing a root leading edge of a fan blade of a gas turbine engine
CN107023514A (en) * 2016-02-02 2017-08-08 通用电气公司 Load absorption system and method
EP3287601A1 (en) * 2016-08-23 2018-02-28 United Technologies Corporation Multi-piece non-linear fan blade
US20180100516A1 (en) * 2016-10-12 2018-04-12 Safran Aircraft Engines Vane comprising an assembled platform and blade
US10202853B2 (en) 2013-09-11 2019-02-12 General Electric Company Ply architecture for integral platform and damper retaining features in CMC turbine blades
EP3511525A1 (en) * 2018-01-16 2019-07-17 Rolls-Royce plc Annulus filler for the fan of a gas turbine engine and corresponding gas turbine engine
US10557350B2 (en) 2017-03-30 2020-02-11 General Electric Company I beam blade platform
US11131203B2 (en) * 2018-09-26 2021-09-28 Rolls-Royce Corporation Turbine wheel assembly with offloaded platforms and ceramic matrix composite blades
US11261744B2 (en) 2019-06-14 2022-03-01 Raytheon Technologies Corporation Ceramic matrix composite rotor blade attachment

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2608674B1 (en) * 1986-12-17 1991-04-19 Snecma CERAMIC BLADE TURBINE WHEEL
DE19826897A1 (en) * 1998-06-17 1999-12-23 Abb Patent Gmbh Lock for blades of a turbine runner

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2656146A (en) * 1948-04-08 1953-10-20 Curtiss Wright Corp Turbine blade construction
US2749029A (en) * 1948-11-26 1956-06-05 Sintercast Corp America Compressor blade
US2819870A (en) * 1955-04-18 1958-01-14 Oleh A Wayne Sheet metal blade base
US2825530A (en) * 1955-05-13 1958-03-04 Eugene F Schum Air-cooled, strut supported turbine blade
US2967043A (en) * 1956-11-30 1961-01-03 Napier & Sons Ltd D Blades and blade mounting assemblies for turbines and axial flow compressors
US3008689A (en) * 1954-08-12 1961-11-14 Rolls Royce Axial-flow compressors and turbines
US3294364A (en) * 1962-01-02 1966-12-27 Gen Electric Rotor assembly

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3114533A (en) * 1961-04-27 1963-12-17 Gen Electric Rotor construction
GB1276106A (en) * 1969-12-19 1972-06-01 Rolls Royce FLUID FLOW MACHINE, e.g. GAS TURBINE ENGINE COMPRESSOR

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2656146A (en) * 1948-04-08 1953-10-20 Curtiss Wright Corp Turbine blade construction
US2749029A (en) * 1948-11-26 1956-06-05 Sintercast Corp America Compressor blade
US3008689A (en) * 1954-08-12 1961-11-14 Rolls Royce Axial-flow compressors and turbines
US2819870A (en) * 1955-04-18 1958-01-14 Oleh A Wayne Sheet metal blade base
US2825530A (en) * 1955-05-13 1958-03-04 Eugene F Schum Air-cooled, strut supported turbine blade
US2967043A (en) * 1956-11-30 1961-01-03 Napier & Sons Ltd D Blades and blade mounting assemblies for turbines and axial flow compressors
US3294364A (en) * 1962-01-02 1966-12-27 Gen Electric Rotor assembly

Cited By (82)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4045149A (en) * 1976-02-03 1977-08-30 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Platform for a swing root turbomachinery blade
US4019832A (en) * 1976-02-27 1977-04-26 General Electric Company Platform for a turbomachinery blade
US5030063A (en) * 1990-02-08 1991-07-09 General Motors Corporation Turbomachine rotor
US5277548A (en) * 1991-12-31 1994-01-11 United Technologies Corporation Non-integral rotor blade platform
US5183389A (en) * 1992-01-30 1993-02-02 General Electric Company Anti-rock blade tang
US5318406A (en) * 1992-11-02 1994-06-07 General Electric Company Multipart gas turbine blade
US6447250B1 (en) * 2000-11-27 2002-09-10 General Electric Company Non-integral fan platform
US20070189901A1 (en) * 2003-03-22 2007-08-16 Dundas Jason E Separable blade platform
US7284958B2 (en) * 2003-03-22 2007-10-23 Allison Advanced Development Company Separable blade platform
US20060251508A1 (en) * 2005-05-05 2006-11-09 United Technologies Corporation Seal arrangement for a fan-turbine rotor assembly
US7201558B2 (en) * 2005-05-05 2007-04-10 United Technologies Corporation Seal arrangement for a fan-turbine rotor assembly
US20080159866A1 (en) * 2006-07-21 2008-07-03 Rolls-Royce Plc Fan blade for a gas turbine engine
US8070438B2 (en) 2006-07-21 2011-12-06 Rolls-Royce Plc Fan blade for a gas turbine engine
US20080226457A1 (en) * 2007-03-16 2008-09-18 Snecma Turbomachine rotor disk
US8277188B2 (en) * 2007-03-16 2012-10-02 Snecma Turbomachine rotor disk
US7972113B1 (en) * 2007-05-02 2011-07-05 Florida Turbine Technologies, Inc. Integral turbine blade and platform
US7874804B1 (en) * 2007-05-10 2011-01-25 Florida Turbine Technologies, Inc. Turbine blade with detached platform
US20080286106A1 (en) * 2007-05-15 2008-11-20 Sean Robert Keith Turbine rotor blade assembly and method of fabricating the same
US7878763B2 (en) 2007-05-15 2011-02-01 General Electric Company Turbine rotor blade assembly and method of assembling the same
US7976281B2 (en) 2007-05-15 2011-07-12 General Electric Company Turbine rotor blade and method of assembling the same
US20080286109A1 (en) * 2007-05-15 2008-11-20 Sean Robert Keith Turbine rotor blade and method of fabricating the same
US20100061858A1 (en) * 2008-09-08 2010-03-11 Siemens Power Generation, Inc. Composite Blade and Method of Manufacture
US8075280B2 (en) * 2008-09-08 2011-12-13 Siemens Energy, Inc. Composite blade and method of manufacture
US20100166551A1 (en) * 2008-12-29 2010-07-01 Morrison Adam J Hybrid turbomachinery component for a gas turbine engine
US8435007B2 (en) 2008-12-29 2013-05-07 Rolls-Royce Corporation Hybrid turbomachinery component for a gas turbine engine
US20100189562A1 (en) * 2009-01-28 2010-07-29 Snecma Composite material turbomachine blade with a reinforced root
US20100290914A1 (en) * 2009-05-15 2010-11-18 Souers Philip F Blade Closing Key System for a Turbine Engine
US8215915B2 (en) 2009-05-15 2012-07-10 Siemens Energy, Inc. Blade closing key system for a turbine engine
CN102213108A (en) * 2010-04-06 2011-10-12 通用电气公司 Composite turbine bucket assembly
US8727730B2 (en) * 2010-04-06 2014-05-20 General Electric Company Composite turbine bucket assembly
US20110243746A1 (en) * 2010-04-06 2011-10-06 General Electric Company Composite turbine bucket assembly
EP2441917A1 (en) * 2010-10-18 2012-04-18 Siemens Aktiengesellschaft Root adapting device and method of attaching a blade in a recess of a rotatable shaft of a steam turbine
WO2012052358A1 (en) * 2010-10-18 2012-04-26 Siemens Aktiengesellschaft Root adapting device for and method of attaching a blade in a recess of a rotatable shaft of a steam turbine
US8827651B2 (en) 2010-11-01 2014-09-09 Rolls-Royce Plc Annulus filler
US20120156045A1 (en) * 2010-12-17 2012-06-21 General Electric Company Methods, systems and apparatus relating to root and platform configurations for turbine rotor blades
CN102536334A (en) * 2010-12-17 2012-07-04 通用电气公司 Methods, systems and apparatus relating to root and platform configurations for turbine rotor blades
US8721290B2 (en) 2010-12-23 2014-05-13 General Electric Company Processes for producing components containing ceramic-based and metallic materials
US9228445B2 (en) * 2010-12-23 2016-01-05 General Electric Company Turbine airfoil components containing ceramic-based materials and processes therefor
US20120163978A1 (en) * 2010-12-23 2012-06-28 General Electric Company Turbine airfoil components containing ceramic-based materials and processes therefor
US8777582B2 (en) 2010-12-27 2014-07-15 General Electric Company Components containing ceramic-based materials and coatings therefor
US8777583B2 (en) 2010-12-27 2014-07-15 General Electric Company Turbine airfoil components containing ceramic-based materials and processes therefor
US9145784B2 (en) 2011-04-14 2015-09-29 Rolls-Royce Plc Annulus filler system
US8936440B2 (en) * 2011-05-26 2015-01-20 United Technologies Corporation Hybrid rotor disk assembly with ceramic matrix composites platform for a gas turbine engine
US20120301317A1 (en) * 2011-05-26 2012-11-29 Ioannis Alvanos Hybrid rotor disk assembly with ceramic matrix composites platform for a gas turbine engine
US8834125B2 (en) 2011-05-26 2014-09-16 United Technologies Corporation Hybrid rotor disk assembly with a ceramic matrix composite airfoil for a gas turbine engine
EP2570609B1 (en) * 2011-05-26 2018-05-16 United Technologies Corporation Ceramic matrix composite component and corresponding rotor disk assembly
US20130004327A1 (en) * 2011-06-30 2013-01-03 United Technologies Corporation Aluminum fan blade root
US9228444B2 (en) 2011-11-15 2016-01-05 Rolls-Royce Plc Annulus filler
JP2013139765A (en) * 2012-01-03 2013-07-18 General Electric Co <Ge> Composite airfoil assembly
EP2612997B1 (en) 2012-01-03 2017-04-26 General Electric Company Composite blade assembly, corresponding turbine rotor wheel and assembly method
CN104220700A (en) * 2012-03-29 2014-12-17 西门子公司 Modular turbine blade having platform
US20150017007A1 (en) * 2012-03-30 2015-01-15 Alstom Technology Ltd Turbine blade or vane
US9920636B2 (en) * 2012-03-30 2018-03-20 Ansaldo Energia Ip Uk Limited Turbine blade or vane
EP2831376B1 (en) * 2012-03-30 2018-10-24 Ansaldo Energia IP UK Limited Turbine blade
EP3473811A1 (en) * 2012-06-07 2019-04-24 United Technologies Corporation Fan blade platform
WO2013185044A1 (en) 2012-06-07 2013-12-12 United Technologies Corporation Fan blade platform
EP2859188A4 (en) * 2012-06-07 2016-07-27 United Technologies Corp Fan blade platform
US10648352B2 (en) * 2012-06-30 2020-05-12 General Electric Company Turbine blade sealing structure
US20150192027A1 (en) * 2012-06-30 2015-07-09 General Electric Company Turbine blade sealing structure
US9598967B2 (en) * 2012-12-18 2017-03-21 United Technologies Corporation Airfoil member and composite platform having contoured endwall
US20140212292A1 (en) * 2012-12-18 2014-07-31 United Technologies Corporation Airfoil Member and Composite Platform Having Contoured Endwall
WO2014139714A1 (en) * 2013-03-11 2014-09-18 Siemens Aktiengesellschaft Rotor blade assembly, turbomachine comprising a rotor blade assembly and method of assembling a rotor blade assembly
EP2778347A1 (en) * 2013-03-11 2014-09-17 Siemens Aktiengesellschaft Rotor blade assembly, turbomachine comprising a rotor blade assembly and method of assembling a rotor blade assembly
US10202853B2 (en) 2013-09-11 2019-02-12 General Electric Company Ply architecture for integral platform and damper retaining features in CMC turbine blades
EP3047109A4 (en) * 2013-09-18 2016-10-05 United Technologies Corp Fan platform with leading edge tab
WO2015053848A2 (en) 2013-09-18 2015-04-16 United Technologies Corporation Fan platform with leading edge tab
US10227884B2 (en) 2013-09-18 2019-03-12 United Technologies Corporation Fan platform with leading edge tab
CN106103898B (en) * 2013-11-25 2019-02-22 安萨尔多能源英国知识产权有限公司 The blade assembly based on modular construction for turbine
CN106103898A (en) * 2013-11-25 2016-11-09 通用电器技术有限公司 Blade assembly based on modular construction for turbine
US20170058916A1 (en) * 2015-09-01 2017-03-02 United Technologies Corporation Gas turbine fan fairing platform and method of fairing a root leading edge of a fan blade of a gas turbine engine
US10371163B2 (en) 2016-02-02 2019-08-06 General Electric Company Load absorption systems and methods
CN107023514A (en) * 2016-02-02 2017-08-08 通用电气公司 Load absorption system and method
CN107023514B (en) * 2016-02-02 2020-03-24 通用电气公司 Load absorbing system and method
EP3287601A1 (en) * 2016-08-23 2018-02-28 United Technologies Corporation Multi-piece non-linear fan blade
US20180058469A1 (en) * 2016-08-23 2018-03-01 United Technologies Corporation Multi-piece non-linear airfoil
US10753368B2 (en) * 2016-08-23 2020-08-25 Raytheon Technologies Corporation Multi-piece non-linear airfoil
US20180100516A1 (en) * 2016-10-12 2018-04-12 Safran Aircraft Engines Vane comprising an assembled platform and blade
US10584720B2 (en) * 2016-10-12 2020-03-10 Safran Aircraft Engines Vane comprising an assembled platform and blade
US10557350B2 (en) 2017-03-30 2020-02-11 General Electric Company I beam blade platform
EP3511525A1 (en) * 2018-01-16 2019-07-17 Rolls-Royce plc Annulus filler for the fan of a gas turbine engine and corresponding gas turbine engine
US11131203B2 (en) * 2018-09-26 2021-09-28 Rolls-Royce Corporation Turbine wheel assembly with offloaded platforms and ceramic matrix composite blades
US11261744B2 (en) 2019-06-14 2022-03-01 Raytheon Technologies Corporation Ceramic matrix composite rotor blade attachment

Also Published As

Publication number Publication date
DE2307967A1 (en) 1973-09-06
GB1386575A (en) 1975-03-12
DE2307967C2 (en) 1983-08-25
FR2174577A5 (en) 1973-10-12
SE389172B (en) 1976-10-25
IT979575B (en) 1974-09-30

Similar Documents

Publication Publication Date Title
US3801222A (en) Platform for compressor or fan blade
US3728042A (en) Axial positioner and seal for cooled rotor blade
US2641440A (en) Turbine blade with cooling means and carrier therefor
US3752599A (en) Bucket vibration damping device
US4045149A (en) Platform for a swing root turbomachinery blade
JP3872830B2 (en) Vane passage hub structure for stator vane with cantilever and manufacturing method thereof
US3709631A (en) Turbine blade seal arrangement
EP1451446B1 (en) Turbine blade pocket shroud
EP0297120B1 (en) Interblade seal for turbomachine rotor
US6884028B2 (en) Turbomachinery blade retention system
US3814539A (en) Rotor sealing arrangement for an axial flow fluid turbine
US2714499A (en) Blading for turbomachines
CA1042348A (en) Platform for a turbomachinery blade
US5257909A (en) Dovetail sealing device for axial dovetail rotor blades
US3377050A (en) Shrouded rotor blades
US4460315A (en) Turbomachine rotor assembly
US3756745A (en) Composite blade root configuration
GB2353826A (en) Aerofoil to platform transition in gas turbine blade/vane
CA2041633C (en) Turbomachine blade fastening
US8956119B2 (en) Turbine wheel provided with an axial retention device that locks blades in relation to a disk
JPS6146644B2 (en)
US3378230A (en) Mounting of blades in turbomachine rotors
EP0900920A3 (en) Sealing device between a blade platform and two stator shrouds
US5183389A (en) Anti-rock blade tang
GB1355751A (en) Bladed rotor assemblies