EP2938824A1 - Rotor de compresseur pour moteur à turbine à gaz présentant une profonde encoche de fixation d'aube - Google Patents

Rotor de compresseur pour moteur à turbine à gaz présentant une profonde encoche de fixation d'aube

Info

Publication number
EP2938824A1
EP2938824A1 EP13867087.2A EP13867087A EP2938824A1 EP 2938824 A1 EP2938824 A1 EP 2938824A1 EP 13867087 A EP13867087 A EP 13867087A EP 2938824 A1 EP2938824 A1 EP 2938824A1
Authority
EP
European Patent Office
Prior art keywords
groove
tangential
sides
tangential sides
radial distance
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.)
Withdrawn
Application number
EP13867087.2A
Other languages
German (de)
English (en)
Other versions
EP2938824A4 (fr
Inventor
Nicholas Aiello
Uyen Phan
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.)
RTX Corp
Original Assignee
United Technologies 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 Technologies Corp filed Critical United Technologies Corp
Publication of EP2938824A1 publication Critical patent/EP2938824A1/fr
Publication of EP2938824A4 publication Critical patent/EP2938824A4/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • 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
    • 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
    • F01D5/3023—Fixing blades to rotors; Blade roots ; Blade spacers of radial insertion type, e.g. in individual recesses
    • F01D5/303—Fixing blades to rotors; Blade roots ; Blade spacers of radial insertion type, e.g. in individual recesses in a circumferential slot
    • F01D5/3038—Fixing blades to rotors; Blade roots ; Blade spacers of radial insertion type, e.g. in individual recesses in a circumferential slot the slot having inwardly directed abutment faces on both sides
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/29—Three-dimensional machined; miscellaneous
    • F05D2250/292—Three-dimensional machined; miscellaneous tapered
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/29—Three-dimensional machined; miscellaneous
    • F05D2250/294—Three-dimensional machined; miscellaneous grooved
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • F05D2260/941—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF] particularly aimed at mechanical or thermal stress reduction

Definitions

  • This application relates to a rotor for use in a compressor for a gas turbine engine, wherein a blade groove has tangential sides that are at a relatively deep location.
  • Gas turbine engines are known, and typically include a fan delivering air into a compressor. The air is compressed and delivered into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over the turbine rotors, driving them to rotate.
  • the compressors typically include a plurality of rotors, some of which include grooves that receive removable blades.
  • the blades extend from a root radially outwardly to an airfoil.
  • the grooves are provided with lock and load slots which are utilized to load the blades into the grooves.
  • lock and load slots which are utilized to load the blades into the grooves.
  • tangential sides to the grooves which provide a reaction surface for the sides of the blades. These tangential sides transmit mechanical tensile stress back into the rotor. These tensile stresses offset thermally induced compressive stresses, which are particularly concentrated at such slots as load or lock slots.
  • a compressor rotor has a rotor centered on an axis, and a groove with opposed side edges.
  • the groove receives a plurality of removable compressor blades.
  • the groove has tangential sides.
  • the blades have tangential side surfaces to be in contact with the tangential sides of the groove.
  • At least one slot is cut into the side edges.
  • a first radial distance is defined measured from a radially outer edge of the side edge to a radially outer beginning point of the tangential sides of the groove.
  • a second radial distance is defined radially between the radially outer beginning point of the tangential sides to a radially inner end of the tangential sides of the groove.
  • a ratio of the first radial distance to the second radial distance is between 1.1 and 5.0.
  • the tangential sides of the groove are defined at an angle. The angle is between 0 and 75 degrees.
  • the slots include lock slots and load slots to assist in loading blades into the grooves.
  • a bearing surface slot is formed within at least one of the tangential sides.
  • Figure 1 schematically shows a gas turbine engine.
  • Figure 2A shows a first feature of a compressor rotor.
  • Figure 2B shows another feature.
  • Figure 3A is a cross sectional view through a rotor.
  • Figure 3B is a view similar to Figure 3A, but with a blade removed.
  • a Figure 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flowpath C for compression and communication into the combustor section 26 then expansion through the turbine section 28.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath B in a bypass duct defined within
  • the engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
  • the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46.
  • the inner shaft 40 may be connected to the fan 42 through a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30.
  • the high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54.
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
  • a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46.
  • the mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
  • the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
  • the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path.
  • the turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • the engine 20 in one example is a high-bypass geared aircraft engine.
  • the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10)
  • the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5.
  • the engine 20 bypass ratio is greater than about ten (10:1)
  • the fan diameter is significantly larger than that of the low pressure compressor 44
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5:1.
  • Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
  • the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
  • the fan section 22 of the engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet.
  • the flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point.
  • 'TSFC' Thrust Specific Fuel Consumption
  • Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
  • the low fan pressure ratio as disclosed herein according to one non- limiting embodiment is less than about 1.45.
  • Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)] ° '5 .
  • the "Low corrected fan tip speed” as disclosed herein according to one non- limiting embodiment is less than about 1150 ft / second.
  • Figure 2A shows a portion of a rotor 80 which may be incorporated into a compressor section in a gas turbine engine such as gas turbine engine 20. As shown, there are tangential sides 82 radially within a groove 200 defined between rim edges 88.
  • a bearing surface slot 86 is shown within the side 82.
  • Figure 2B shows opposed rim edges 88, a load slot 92, and lock slots 90, all of which are cut into the facing edges 88.
  • edges of the blades contact the tangential sides 82, and transmit mechanical tensile stresses. As mentioned above, those mechanical tensile stresses can offset thermally induced compressive stresses, which are concentrated in the slots.
  • Figure 3A is a cross-sectional view through the rotor 80. As shown, sides 83 of a blade 94 contact the sides 82. An airfoil 95, partially illustrated, extends radially outwardly. As the rotor 80 is driven to rotate at high rates of speed, the blade is urged radially outwardly due to centrifugal forces, and the mechanical tensile stresses from sides 83 contacting the tangential sides 82 become high.
  • the present invention addresses this concern by making the sides 82 relatively radially deep compared to the prior art.
  • a first distance di can be defined between a radially outer point 89 of the side edge 88, and extending inwardly to a radially inner end 91 of the side edge 88. These distances are measured relative to a center axis A.
  • the tangential sides 82 begin at point 91 and extend radially inwardly to point 93. In one embodiment the tangential side extends at an angle A.
  • the radial distance between point 91 and 93 is d 2 . This is a radial distance and not the length along the surface of side 82.
  • a ratio of di to d 2 was between 1.1 and 5.0.
  • the angle A was 45 degrees in one embodiment. In embodiments, A may be between 0 and 75 degrees.

Landscapes

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

Abstract

L'invention concerne un rotor de compresseur comprenant un rotor centré sur un axe et une encoche présentant deux bords latéraux se faisant face. L'encoche reçoit une pluralité d'aubes de compresseur amovibles et présente des côtés tangentiels. Les surfaces latérales tangentielles sont en contact avec lesdits côtés tangentiels de l'encoche et au moins une fente est découpée dans les bords latéraux. Une première distance radiale est mesurée depuis un bord radialement externe du bord latéral vers un point de départ radialement externe des côtés tangentiels de l'encoche. Une seconde distance radiale s'étend radialement entre le point de départ radialement externe des côtés tangentiels et un côté radialement interne des côtés tangentiels de l'encoche. Le rapport entre les première et seconde distances radiales est compris entre 1,1 et 5,0.
EP13867087.2A 2012-12-31 2013-12-19 Rotor de compresseur pour moteur à turbine à gaz présentant une profonde encoche de fixation d'aube Withdrawn EP2938824A4 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/731,147 US20140182293A1 (en) 2012-12-31 2012-12-31 Compressor Rotor for Gas Turbine Engine With Deep Blade Groove
PCT/US2013/076351 WO2014105593A1 (fr) 2012-12-31 2013-12-19 Rotor de compresseur pour moteur à turbine à gaz présentant une profonde encoche de fixation d'aube

Publications (2)

Publication Number Publication Date
EP2938824A1 true EP2938824A1 (fr) 2015-11-04
EP2938824A4 EP2938824A4 (fr) 2015-12-30

Family

ID=51015616

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13867087.2A Withdrawn EP2938824A4 (fr) 2012-12-31 2013-12-19 Rotor de compresseur pour moteur à turbine à gaz présentant une profonde encoche de fixation d'aube

Country Status (3)

Country Link
US (1) US20140182293A1 (fr)
EP (1) EP2938824A4 (fr)
WO (1) WO2014105593A1 (fr)

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6434780B2 (ja) * 2014-11-12 2018-12-05 三菱日立パワーシステムズ株式会社 タービン用ロータアセンブリ、タービン、及び、動翼
US11242761B2 (en) 2020-02-18 2022-02-08 Raytheon Technologies Corporation Tangential rotor blade slot spacer for a gas turbine engine

Family Cites Families (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2295012A (en) * 1941-03-08 1942-09-08 Westinghouse Electric & Mfg Co Turbine blading
US2931625A (en) * 1956-12-17 1960-04-05 Gen Electric Compressor rotor
GB1151937A (en) * 1966-08-26 1969-05-14 Mini Of Technology Bladed Rotors for Fluid Flow Machines
JPS4946102U (fr) * 1972-07-31 1974-04-23
US4482297A (en) * 1981-11-16 1984-11-13 Terry Corporation Bladed rotor assembly
US5141401A (en) * 1990-09-27 1992-08-25 General Electric Company Stress-relieved rotor blade attachment slot
US5522706A (en) * 1994-10-06 1996-06-04 General Electric Company Laser shock peened disks with loading and locking slots for turbomachinery
JPH10299407A (ja) * 1997-04-22 1998-11-10 Hitachi Ltd ガスタービンエンジンのロータ
GB2364554B (en) * 2000-07-07 2004-04-07 Alstom Power Nv Turbine disc
US6619030B1 (en) * 2002-03-01 2003-09-16 General Electric Company Aircraft engine with inter-turbine engine frame supported counter rotating low pressure turbine rotors
DE102004051116A1 (de) * 2004-10-20 2006-04-27 Mtu Aero Engines Gmbh Rotor einer Turbomaschine, insbesondere Gasturbinenrotor
US8206116B2 (en) * 2005-07-14 2012-06-26 United Technologies Corporation Method for loading and locking tangential rotor blades and blade design
US8251667B2 (en) * 2009-05-20 2012-08-28 General Electric Company Low stress circumferential dovetail attachment for rotor blades
US8414268B2 (en) * 2009-11-19 2013-04-09 United Technologies Corporation Rotor with one-sided load and lock slots
JP5730085B2 (ja) * 2011-03-17 2015-06-03 三菱日立パワーシステムズ株式会社 ロータ構造

Also Published As

Publication number Publication date
US20140182293A1 (en) 2014-07-03
WO2014105593A1 (fr) 2014-07-03
EP2938824A4 (fr) 2015-12-30

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