US3706512A - Compressor blades - Google Patents

Compressor blades Download PDF

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Publication number
US3706512A
US3706512A US89640A US3706512DA US3706512A US 3706512 A US3706512 A US 3706512A US 89640 A US89640 A US 89640A US 3706512D A US3706512D A US 3706512DA US 3706512 A US3706512 A US 3706512A
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United States
Prior art keywords
blade
rib
edge
ribs
tip end
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
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US89640A
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English (en)
Inventor
Arthur D Strelshik
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.)
Pratt and Whitney Canada Corp
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United Aircraft of Canada Ltd
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Filing date
Publication date
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    • 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
    • F04D29/324—Blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D21/00—Pump involving supersonic speed of pumped fluids
    • 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
    • F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
    • F01D21/04—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position
    • F01D21/045—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position special arrangements in stators or in rotors dealing with breaking-off of part of rotor
    • 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
    • F05D2250/70—Shape
    • 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 An improved thin, plate-like, compressor blade having stiffening ribs on its airfoil surfaces extending from or adjacent the leading edge of the blade toward the trailing edge and being positioned on the blade to reduce the tendency of the blade to fail by buckling whenvthe leading edge of the blade is struck by foreign objects.
  • the invention is particularly directed to improvements in compressor blades, and compressor rotors incorporating the improved blades, used primarily in transonic and supersonic flow conditions.
  • Blades used under these conditions normally are relatively thin with low camber areas and have a substantial width in comparison to their length.
  • the blades have a dimensional configuration such that they act structurally, at least for a major portion of their length where the camber is low, as a plate member rather than as a beam member when impacted on an edge.
  • Such blades when used in the compressor rotor stage of a jet turbine engine, can be damaged by foreign objects, such as hail or birds, being ingested into the intake of the engine. The objects strike the leading edges of the blades and can cause the blades to fail by buckling and thus also cause possible failure of the compressor rotor itself.
  • the improvement resides in providing at least one stiffening rib on each airfoil surface of the blade in the area of low camber. Each rib is positioned to traverse a theoretical line on the blade in the area of low camber along which the blade, without a rib, is most likely to buckle when an object impacts the leading edge of the blade.
  • each rib is located to extend from or adjacent the leading edge of the blade toward the trailing edge, traversing the theoretical buckling line of the blade, and is located closer to the unsupported tip end of the blade than to the root end in the area of low camber.
  • each rib is preferably dimensioned, at least in the region it traverses the buckling line, to have a width no greater than twice the maximum thickness of the blade and to have a thickness no greater than' the maximum thickness of the blade.
  • Blades as described above have been found, in testing, to provide a significant increased resistance to failure by buckling when impacted with foreign objects.
  • the use of the stiffening ribs provides further advantages in that they can be used, by removing material from one rib or the other, to assist in the dynamic balancing of a rotor incorporating the blades, to assist in tuning the natural frequencies of the blade, and'to provide an additional cutting action, over and above that provided by the blades, to foreign objects, such as birds, thus reducing the size of matter that is passed further downstream into the engine.
  • FIG. 1 illustrates a portion of a compressor rotor having blades which incorporate the stiffening ribs
  • FIG. 2 is a cross-section along line 2-2 of FIG. 1 showing one blade in detail;
  • FIG. 3 is a cross-section taken along line 3 -3 of FIG. 2 showing a detail of the ribs
  • FIG. 4 is a cross-section taken along line 4-4 of FIG. 2 showing a further detail of the ribs.
  • the invention is particularly directed toward the blades used in the compression stage rotor of a jet turbine engine, such as a turbofan jet engine, for example, and a rotor incorporating the blades, handling transonic or supersonic flow.
  • the blades 1, as shown in FIG. 1 are connected at their root end 3 by suitable connecting means 2, such as, for example, a cooperating fir-tree member and slot, about the periphery of the compressor rotor 5.
  • suitable connecting means 2 such as, for example, a cooperating fir-tree member and slot
  • the tip end 7 of the .blade is unsupported.
  • Each blade 1 has two major airfoil surfaces 9, 11, a leading edge 13, facing generally in the direction of airflow shown by arrow A as shown in FIG. 2, and a trailing edge 15.
  • the blade 1 preferably is of the type having a dimensional configuration over at least a major portion of its length to act structurally, in theory, as a plate member rather than as a beam member.
  • a blade, to act structurally as a plate member can be defined as one which, when subjected to external and/or inertial loading, will deflect due to bending moments in the chord-wise direction. In other words, unlike a beam, the blade will deflect non-uniformly along the chord of the blade.
  • Such blades are relatively thin and have, over a major portion of the length of the blade, spaced from the root end 3, a low camber angle no greater than 30.
  • the camber angle is definedby the complement angle at of the internal angle 0 formed by the intersection of tangent lines A, B drawn from the mean line .C of the blade at the leading and trailing edges l3, 15 respectively, as shown in FIG. 3.
  • the area of minimum camber is between two-thirds and seven-eighths of the length L of the blade from the root end.
  • the compressor blades 1 used in transonic or supersonic flow conditions have a high camber at the root end of about 60.
  • the camber is reduced to about 30 at a distance about one-fourth of the length L of the blade toward the tip end 7 at line E-E.
  • the camber reduces to about 25 at a distance about one-half of the length of the blade toward the tip end at line F- -F.
  • the camber may then be reduced to a minimum of about at a distance about three-fourths of the blade length from the root end at line G-G from where it increases slightly to about 10 at the tip end 7.
  • blades can fail by'buckling if an object is ingested into the intake of a jet engine and strikesthe leading edge 13 of the blade, particularly along the leading edge of the blade in the area where the camber is 30 or less. Buckling will occur in these areas across a line Y-Y as shown in FIG. 2, the location of which can, for different blades, be determined analytically or experimentally.
  • at least one rib 17,17 is provided, one on each major airfoil surface 9, ll of the blade as shown in FIG. 1.
  • Each rib 17, 17 extends from or adjacent the leading edge 13 of the blade'toward the trailing edge 15 and is generally positioned to extend in a direction substantially parallelto the direction of the airflow past the blade, as shown by arrow A, to minimize any reduction in aerodynamic efficiency of the blade.
  • the ribs 17, 17' on the airfoil surfaces are opposite to one another to minimize unbalancing of the blade.
  • the ribs 17, 17 traverse the buckling line Y-Y and are centrally located with respect to the buckling line to extend substantially the same distance to either side of the buckling line.
  • each rib 17, 17 is located closer to the tip end of the blade rather than to the root end edge, since buckling is more likely to occur in the upper half of the blade where the camber is at a minimum, but are spaced from the tip end so as to be positioned approximately midway of the buckling line Y-Y length.
  • the ribs 17, 17' are preferably located in the area of minimum camber.
  • the ribs preferably should have a width w no greater than twice the maximum thickness T of the blade and a thickness 1 no greater than the maximum thickness T of the blade.
  • the ribs 17, 17' can be of any suitable shape in crosssection, such as, for example, rectangular or trapezoidal, and are integrally formed with the blade when the blade is being cast or otherwise manufactured.
  • the front 19, 19' and back 21, 21' edges of the ribs are preferably tapered to merge smoothly into the airfoil surfaces of the blade and reduce airflow resistance as shown in FIG. 3.
  • an additional pair of opposed ribs 23, 23' can be provided on the airfoil surfaces 9, 11 of the blade to improve the resistance of the blade to buckling. If additional ribs 23, 23' are used, they are generally spaced midway between a secondary line of buckling X -X which is developed after using a first pair'of ribs 17, 17.
  • the ribs -23, 23' have substantially the same dimensions as the ribs 17, 17.
  • a thin, plate-like, compressor blade said blade having a leading edge, a trailing edge, a tip end edge, a root end edge having means permitting connection of the blade to a rotor, and two opposed airfoil surfaces bounded by said edges, the blade having a decreasing camber toward the tip and becoming substantially flat adjacent the tip, said blade having a buckling line in the substantially flat portion adjacent the tip, at least one pair of elongate stiffening ribs positioned on the platelike blade in a low camber region of the blade to strengthen it against buckling by impact of foreign objects on the leading edge, the ribs being located on the opposite airfoil surfaces in directly opposite relation to one another, each rib extending from a point adjacent the leading edge of the blade and extending toward the trailing edge, each rib being located in the tip end edge area of the blade but spaced from the tip end edge, and said region of the blade has a camber of less than 30, this area of the blade including the location of the buckling line, the ribs
  • each stiffening rib projects from its airfoil surface a distance no greater than twice the maximum thickness of the blade.
  • each stiffening rib has a maximum thickness no greater than the maximum thickness of the blade.
  • a compressor blade as claimed in claim 1 including an additional pair of stiffening ribs on the blade, one rib of the additional pair being located on one of the airfoil surfaces, and the other rib of the additional pair being located opposite to said one ribof the additional pair on the other airfoil surface, said additional pair of ribs being located closer to the tip end of the blade than the root end, each rib of said additional pair extending from or adjacent the leading edge of the blade toward the trailing edge, said additional pair of ribs being spaced from the first pair of ribs toward the tip end.
  • a compressor rotor a plurality of thin, plate-like, blades mounted about the periphery of the rotor, each blade having a leading edge, a trailing edge, a tip end edge, a root end edge, and two airfoil surfaces bounded .by said edges, each blade being relatively thin and decreasing in camber toward the blade tip to a substantially flat configuration forming a region subject to buckling, cooperating means on the root end edge of each blade and the periphery of the rotor permitting connection of each blade to the periphery of the rotor, at least one pair of stiffening ribs positioned on each plate-like blade to strengthen it against buckling due to impact of foreign objects on its leading edge, one rib being located on one airfoil surface and the other rib being located on the other airfoil surface opposite to said one rib, each rib extending only from a point adjacent the leading edge of the blade toward the trailing edge, each rib being located in the tip end area of the blade but spaced

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US89640A 1970-11-16 1970-11-16 Compressor blades Expired - Lifetime US3706512A (en)

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US8964070A 1970-11-16 1970-11-16

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GB (1) GB1366924A (Direct)

Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19957718C1 (de) * 1999-11-30 2001-06-13 Mtu Muenchen Gmbh Schaufel mit optimiertem Schwingungsverhalten
JP2002540334A (ja) * 1999-03-24 2002-11-26 アーベーベー・ターボ・ジステムス・アクチエンゲゼルシヤフト タービン翼
US20030118447A1 (en) * 2001-11-16 2003-06-26 Fiatavio S.P.A. Bladed member, in particular for an axial turbine of an aircraft engine
JP2005076634A (ja) * 2003-08-28 2005-03-24 General Electric Co <Ge> 圧縮機翼形部に生じる振動を低減するための方法及び装置
FR2867506A1 (fr) * 2004-03-11 2005-09-16 Snecma Moteurs Aube de redresseur nervuree
EP1471209A3 (en) * 2003-04-23 2006-07-12 General Electric Company Apparatus to reduce the vibrations of gas turbine rotor blades
US20070201983A1 (en) * 2006-02-27 2007-08-30 Paolo Arinci Rotor blade for a ninth phase of a compressor
US7270519B2 (en) * 2002-11-12 2007-09-18 General Electric Company Methods and apparatus for reducing flow across compressor airfoil tips
US20140245753A1 (en) * 2013-01-08 2014-09-04 United Technologies Corporation Gas turbine engine rotor blade
US20150361808A1 (en) * 2014-06-17 2015-12-17 Snecma Turbomachine vane including an antivortex fin
EP2990604A1 (en) * 2014-08-27 2016-03-02 Pratt & Whitney Canada Corp. Rotary airfoil and method of forming a rotary blade
US20160123345A1 (en) * 2013-06-13 2016-05-05 Nuovo Pignone Srl Compressor impellers
CN110637151A (zh) * 2017-10-31 2019-12-31 三菱重工发动机和增压器株式会社 涡轮动叶片、涡轮增压器以及涡轮动叶片的制造方法
US10539157B2 (en) 2015-04-08 2020-01-21 Horton, Inc. Fan blade surface features
CN110873075A (zh) * 2018-08-31 2020-03-10 赛峰航空助推器股份有限公司 用于涡轮机的压缩机的具有突起的叶片
US10605087B2 (en) * 2017-12-14 2020-03-31 United Technologies Corporation CMC component with flowpath surface ribs
US20240254882A1 (en) * 2021-02-02 2024-08-01 Ge Avio S.R.L. Turbine engine with reduced cross flow airfoils

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2616486B1 (fr) * 1987-06-10 1993-05-28 Snecma Procede de fabrication d'une aube mobile de compresseur a talons intermediaires rapportes
GB2373548B (en) 2001-03-21 2004-06-09 Rolls Royce Plc Gas turbine engine aerofoils
US10641107B2 (en) 2012-10-26 2020-05-05 Rolls-Royce Plc Turbine blade with tip overhang along suction side
FR3081913B1 (fr) * 2018-06-04 2021-01-08 Safran Aircraft Engines Aube de turbomachine comportant une ailette anti-tourbillons
FR3087828B1 (fr) * 2018-10-26 2021-01-08 Safran Helicopter Engines Aubage mobile de turbomachine

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB840543A (en) * 1956-01-16 1960-07-06 Vickers Electrical Co Ltd Improvements in turbine blading
CA606617A (en) * 1960-10-11 M. Ganger Karl Ventilating fan with reversible motor
US2965180A (en) * 1954-12-20 1960-12-20 American Radiator & Standard Propeller fan wheel
US3012709A (en) * 1955-05-18 1961-12-12 Daimler Benz Ag Blade for axial compressors

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CA606617A (en) * 1960-10-11 M. Ganger Karl Ventilating fan with reversible motor
US2965180A (en) * 1954-12-20 1960-12-20 American Radiator & Standard Propeller fan wheel
US3012709A (en) * 1955-05-18 1961-12-12 Daimler Benz Ag Blade for axial compressors
GB840543A (en) * 1956-01-16 1960-07-06 Vickers Electrical Co Ltd Improvements in turbine blading

Cited By (31)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2002540334A (ja) * 1999-03-24 2002-11-26 アーベーベー・ターボ・ジステムス・アクチエンゲゼルシヤフト タービン翼
US6565324B1 (en) * 1999-03-24 2003-05-20 Abb Turbo Systems Ag Turbine blade with bracket in tip region
US6503053B2 (en) 1999-11-30 2003-01-07 MTU Motoren-und Turbinen München GmbH Blade with optimized vibration behavior
DE19957718C1 (de) * 1999-11-30 2001-06-13 Mtu Muenchen Gmbh Schaufel mit optimiertem Schwingungsverhalten
US20030118447A1 (en) * 2001-11-16 2003-06-26 Fiatavio S.P.A. Bladed member, in particular for an axial turbine of an aircraft engine
US7270519B2 (en) * 2002-11-12 2007-09-18 General Electric Company Methods and apparatus for reducing flow across compressor airfoil tips
EP1471209A3 (en) * 2003-04-23 2006-07-12 General Electric Company Apparatus to reduce the vibrations of gas turbine rotor blades
JP2005076634A (ja) * 2003-08-28 2005-03-24 General Electric Co <Ge> 圧縮機翼形部に生じる振動を低減するための方法及び装置
US6905309B2 (en) * 2003-08-28 2005-06-14 General Electric Company Methods and apparatus for reducing vibrations induced to compressor airfoils
CN1598248B (zh) * 2003-08-28 2010-12-08 通用电气公司 减小诱发压缩机翼面振动的装置
FR2867506A1 (fr) * 2004-03-11 2005-09-16 Snecma Moteurs Aube de redresseur nervuree
US7766624B2 (en) * 2006-02-27 2010-08-03 Nuovo Pignone S.P.A. Rotor blade for a ninth phase of a compressor
US7785074B2 (en) * 2006-02-27 2010-08-31 General Electric Company Rotor blade for a second stage of a compressor
US20070201983A1 (en) * 2006-02-27 2007-08-30 Paolo Arinci Rotor blade for a ninth phase of a compressor
US20080044288A1 (en) * 2006-02-27 2008-02-21 Alessio Novori Rotor blade for a second phase of a compressor
US20140245753A1 (en) * 2013-01-08 2014-09-04 United Technologies Corporation Gas turbine engine rotor blade
US9845683B2 (en) * 2013-01-08 2017-12-19 United Technology Corporation Gas turbine engine rotor blade
US20160123345A1 (en) * 2013-06-13 2016-05-05 Nuovo Pignone Srl Compressor impellers
US10260361B2 (en) * 2014-06-17 2019-04-16 Safran Aircraft Engines Turbomachine vane including an antivortex fin
US20150361808A1 (en) * 2014-06-17 2015-12-17 Snecma Turbomachine vane including an antivortex fin
US10443390B2 (en) 2014-08-27 2019-10-15 Pratt & Whitney Canada Corp. Rotary airfoil
EP2990604A1 (en) * 2014-08-27 2016-03-02 Pratt & Whitney Canada Corp. Rotary airfoil and method of forming a rotary blade
US10539157B2 (en) 2015-04-08 2020-01-21 Horton, Inc. Fan blade surface features
US10662975B2 (en) 2015-04-08 2020-05-26 Horton, Inc. Fan blade surface features
CN110637151A (zh) * 2017-10-31 2019-12-31 三菱重工发动机和增压器株式会社 涡轮动叶片、涡轮增压器以及涡轮动叶片的制造方法
CN110637151B (zh) * 2017-10-31 2021-09-07 三菱重工发动机和增压器株式会社 涡轮动叶片、涡轮增压器以及涡轮动叶片的制造方法
US10605087B2 (en) * 2017-12-14 2020-03-31 United Technologies Corporation CMC component with flowpath surface ribs
CN110873075A (zh) * 2018-08-31 2020-03-10 赛峰航空助推器股份有限公司 用于涡轮机的压缩机的具有突起的叶片
CN110873075B (zh) * 2018-08-31 2023-09-26 赛峰航空助推器股份有限公司 用于涡轮机的压缩机的具有突起的叶片
US20240254882A1 (en) * 2021-02-02 2024-08-01 Ge Avio S.R.L. Turbine engine with reduced cross flow airfoils
US12421853B2 (en) * 2021-02-02 2025-09-23 Ge Avio S.R.L. Turbine engine with reduced cross flow airfoils

Also Published As

Publication number Publication date
GB1366924A (en) 1974-09-18
FR2114693A5 (Direct) 1972-06-30

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