US20010014285A1 - Rotor blade an axial-flow engine - Google Patents

Rotor blade an axial-flow engine Download PDF

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Publication number
US20010014285A1
US20010014285A1 US09/424,481 US42448100A US2001014285A1 US 20010014285 A1 US20010014285 A1 US 20010014285A1 US 42448100 A US42448100 A US 42448100A US 2001014285 A1 US2001014285 A1 US 2001014285A1
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Prior art keywords
blade
edge
airfoil
afflux
rotor blade
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US09/424,481
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US6358003B2 (en
Inventor
Stefan Schlechtriem
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Rolls Royce Deutschland Ltd and Co KG
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Rolls Royce Deutschland Ltd and Co KG
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Assigned to BMW ROLLS-ROYCE GMBH reassignment BMW ROLLS-ROYCE GMBH ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SCHLECHTRIEM, STEFAN
Publication of US20010014285A1 publication Critical patent/US20010014285A1/en
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Publication of US6358003B2 publication Critical patent/US6358003B2/en
Assigned to ROLLS-ROYCE DEUTSCHLAND GMBH reassignment ROLLS-ROYCE DEUTSCHLAND GMBH CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: BMW ROLLS-ROYCE GMBH
Assigned to ROLLS-ROYCE DEUTSCHLAND LTD & CO KG reassignment ROLLS-ROYCE DEUTSCHLAND LTD & CO KG ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: ROLLS-ROYCE DEUTSCHLAND GMBH
Assigned to ROLLS-ROYCE DEUTSCHLAND GMBH reassignment ROLLS-ROYCE DEUTSCHLAND GMBH CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: BMW ROLLS-ROYCE GMBH
Assigned to ROLLS-ROYCE DEUTSCHLAND LTD & CO KG reassignment ROLLS-ROYCE DEUTSCHLAND LTD & CO KG CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: ROLLS-ROYCE DEUTSCHLAND GMBH
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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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
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • 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
    • 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
    • F05D2250/71—Shape curved
    • F05D2250/713—Shape curved inflexed
    • 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

  • This invention relates to a rotor blade airfoil of an axial-flow turbomachine, more particularly for a fan or a compressor high-pressure stage of a gas turbine system, having means in the tip area to improve supersonic performance.
  • a rotor blade airfoil of an axial-flow turbomachine more particularly for a fan or a compressor high-pressure stage of a gas turbine system, having means in the tip area to improve supersonic performance.
  • Associated state of the art is described in U.S. Pat. No. 3,989,406 and U.S. Pat. No. 4,012,172.
  • the first-cited patent specification deals with the supersonic performance of rotor blade airfoils generally, while the second illustrates a forward swept airfoil, i.e.
  • the present invention also relates to the sweep of rotor blade airfoils, more particularly of the first compressor high-pressure stages or of a fan of gas turbines or aircraft gas turbine engines. Owing to the high peripheral speeds of said components and the practically zero-swirl inlet flow to the blade, supersonic regions occur in the blade tip area with mach numbers in excess of 1.4. These for several reasons negatively affect the performance of this component. Firstly, the efficiency decreases with increasing blade span, owing to growing shock losses, much more severely than with subsonic rotors. Also, the interaction of duct shock with blade tip swirl negatively affects the stability of the flow, because large blockage regions occur in that area whose nonlinear growth ultimately determines the stability limit of the compressor.
  • the arrangement of the present invention provides means to remedy said problems.
  • Reference numeral 1 indicates the blade root and reference numeral 2 the airfoil of the rotor blade of an axial-flow turbomachine, more particularly for a fan or compressor high-pressure stage of a gas turbine system.
  • the axis of rotation of the rotor omitted on the drawing, accordingly runs parallel to the direction of flow 3 and extends (far) below the blade root 1 .
  • airfoil 2 has a root region 2 a and a tip region 2 b , the latter substantially extending along the outer 15% to 25%, more particularly along about 20% of the blade span h.
  • the latter by amount is the difference (R a ⁇ R i ), where R a is the outer radius measured from the axis of rotation and R i the inner radius of the airfoil 2 .
  • Shown in the tip region 2 b is both the afflux edge 4 a and the efflux edge 4 b of a conventionally designed airfoil 2 , indicated by dashed lines a′ and b′.
  • a conventionally shaped airfoil of this type is impaired by the disadvantages cited above.
  • the airfoil 2 of the present invention is in the tip region 2 b provided with a forward-back sweep at least on the blade afflux edge 4 a as shown on the drawing, i.e. in this tip region 2 b the blade afflux edge 4 a extends, unlike in a conventional design, first counter to the direction of flow 3 and, upon reaching a radius point U, retreats in the direction of flow 3 in a manner more pronounced than with a conventionally designed blade afflux edge a′.
  • an angle ⁇ of a 90° order of magnitude is sought in the uppermost region, i.e. at the blade tip, between the afflux edge 4 a and the case contour indicated by line 5 of a case (omitted on the drawing) surrounding the rotor.
  • the blade efflux edge 4 b has a shape essentially similar to that of the blade afflux edge 4 a , i.e. it reflects the forward-back sweep, as a comparison of the solid line 4 b with the line b′ representing the conventional design will show.
  • the back sweep can be made more pronounced than the forward sweep, so as to allow for the mach number distribution.
  • the distance, not indicated on the drawing, by which the blade afflux edge 4 a in the tip region is shifted back compared with the conventional design in accordance with line a′ is larger than the distance, again not indicated, by which the blade afflux edge 4 a in the radius point U area is shifted forward relative to the conventional design in accordance with line a′.
  • this may look different on account of the two-dimensional representation, whereas a rotor blade airfoil is obviously a three-dimensional shape.
  • the sweep can also be designed such that the topmost airfoil section of a conventional rotor blade airfoil is retained. This will allow the blading in a previously existing case of an axial-flow turbomachine also to be renewed in a case with contoured walls, without having to adjust rub rings.
  • the forward-back sweep of the present invention primarily reduces the local afflux mach number and so reduces shock losses. In the process, the forward-back sweep does not induce greater radial velocities that would be the cause of additional losses.
  • a forward-back sweep also provides structural mechanical advantages, considering that the thrust, elastic axis and gravity centers will, when compared with the straight forward or back sweep, not change at all across blade span h until about 80% of the blade span, and only slightly from 80% blade span to the blade tip. Additional bending moments in the hub area, such as arising with straight forward or back sweep, are therefore avoided.
  • the modified modes of vibration in the upper region of the airfoil 2 are above the second E.O., so that the arrangement of the present invention also causes no restrictions from the structural dynamic aspect. It is apparent that a plurality of especially design features other than those described herein may be incorporated in the present embodiment without departing from the inventive concept.

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

Abstract

This invention relates to a rotor blade airfoil of an axial-flow turbomachine, more particularly for a fan or a compressor high-pressure stage of a gas turbine system, having means in the tip area to improve supersonic performance. In accordance with the present invention, the blade afflux edge has a forward-back sweep. The tip region of the blade efflux edge can be designed essentially similar to the blade afflux edge, with the tip region extending across the outer 15% to 25% of the blade span.

Description

  • This invention relates to a rotor blade airfoil of an axial-flow turbomachine, more particularly for a fan or a compressor high-pressure stage of a gas turbine system, having means in the tip area to improve supersonic performance. Associated state of the art is described in U.S. Pat. No. 3,989,406 and U.S. Pat. No. 4,012,172. The first-cited patent specification deals with the supersonic performance of rotor blade airfoils generally, while the second illustrates a forward swept airfoil, i.e. an airfoil on which the leading edge, starting from the blade root, first arches counter to the direction of flow, so that in a lateral view, the airfoil area or blade chord first continuously increases until, upon reaching a radius point, it arches in the direction of flow toward the blade tip area, where in said lateral view, the airfoil area or chord of the blade again decreases. [0001]
  • The present invention also relates to the sweep of rotor blade airfoils, more particularly of the first compressor high-pressure stages or of a fan of gas turbines or aircraft gas turbine engines. Owing to the high peripheral speeds of said components and the practically zero-swirl inlet flow to the blade, supersonic regions occur in the blade tip area with mach numbers in excess of 1.4. These for several reasons negatively affect the performance of this component. Firstly, the efficiency decreases with increasing blade span, owing to growing shock losses, much more severely than with subsonic rotors. Also, the interaction of duct shock with blade tip swirl negatively affects the stability of the flow, because large blockage regions occur in that area whose nonlinear growth ultimately determines the stability limit of the compressor. [0002]
  • In a broad aspect, the arrangement of the present invention provides means to remedy said problems. [0003]
  • It is a particular object of the present invention to provide a solution to said problems by providing a forward-back-swept blade leading edge. Further objects and advantages of the present invention will become apparent from the subclaims. [0004]
  • The invention is described more fully by means of a preferred embodiment in the light of a single drawing showing a rotor blade airfoil in greatly simplified lateral view. [0005]
  • [0006] Reference numeral 1 indicates the blade root and reference numeral 2 the airfoil of the rotor blade of an axial-flow turbomachine, more particularly for a fan or compressor high-pressure stage of a gas turbine system. Air is impinging on airfoil 2 in the direction of arrowhead 3 (=direction of flow 3) , i.e. the left-hand leading edge of airfoil 2 is its afflux edge 4 a and the right-hand trailing edge is its efflux edge 4 b. The axis of rotation of the rotor, omitted on the drawing, accordingly runs parallel to the direction of flow 3 and extends (far) below the blade root 1.
  • As usual, [0007] airfoil 2 has a root region 2 a and a tip region 2 b, the latter substantially extending along the outer 15% to 25%, more particularly along about 20% of the blade span h. The latter by amount is the difference (Ra−Ri), where Ra is the outer radius measured from the axis of rotation and Ri the inner radius of the airfoil 2. Shown in the tip region 2 b is both the afflux edge 4 a and the efflux edge 4 b of a conventionally designed airfoil 2, indicated by dashed lines a′ and b′. A conventionally shaped airfoil of this type, however, is impaired by the disadvantages cited above.
  • To remedy these disadvantages, the [0008] airfoil 2 of the present invention is in the tip region 2 b provided with a forward-back sweep at least on the blade afflux edge 4 a as shown on the drawing, i.e. in this tip region 2 b the blade afflux edge 4 a extends, unlike in a conventional design, first counter to the direction of flow 3 and, upon reaching a radius point U, retreats in the direction of flow 3 in a manner more pronounced than with a conventionally designed blade afflux edge a′. For mechanical reasons, an angle α of a 90° order of magnitude is sought in the uppermost region, i.e. at the blade tip, between the afflux edge 4 a and the case contour indicated by line 5 of a case (omitted on the drawing) surrounding the rotor.
  • In the [0009] tip region 2 b, the blade efflux edge 4 b has a shape essentially similar to that of the blade afflux edge 4 a, i.e. it reflects the forward-back sweep, as a comparison of the solid line 4 b with the line b′ representing the conventional design will show.
  • In this arrangement, the back sweep can be made more pronounced than the forward sweep, so as to allow for the mach number distribution. This means that the distance, not indicated on the drawing, by which the blade afflux edge [0010] 4 a in the tip region is shifted back compared with the conventional design in accordance with line a′, is larger than the distance, again not indicated, by which the blade afflux edge 4 a in the radius point U area is shifted forward relative to the conventional design in accordance with line a′. On the drawing, this may look different on account of the two-dimensional representation, whereas a rotor blade airfoil is obviously a three-dimensional shape.
  • The sweep can also be designed such that the topmost airfoil section of a conventional rotor blade airfoil is retained. This will allow the blading in a previously existing case of an axial-flow turbomachine also to be renewed in a case with contoured walls, without having to adjust rub rings. [0011]
  • The forward-back sweep of the present invention primarily reduces the local afflux mach number and so reduces shock losses. In the process, the forward-back sweep does not induce greater radial velocities that would be the cause of additional losses. [0012]
  • A forward-back sweep also provides structural mechanical advantages, considering that the thrust, elastic axis and gravity centers will, when compared with the straight forward or back sweep, not change at all across blade span h until about 80% of the blade span, and only slightly from 80% blade span to the blade tip. Additional bending moments in the hub area, such as arising with straight forward or back sweep, are therefore avoided. The modified modes of vibration in the upper region of the [0013] airfoil 2 are above the second E.O., so that the arrangement of the present invention also causes no restrictions from the structural dynamic aspect. It is apparent that a plurality of especially design features other than those described herein may be incorporated in the present embodiment without departing from the inventive concept.
    • LIST OF REFERENCES
  • [0014] 1 Blade root
  • [0015] 2 Airfoil
  • [0016] 2 a Root region of 2
  • [0017] 2 b Tip region of 2
  • [0018] 3 Direction of flow
  • [0019] 4 a Blade afflux edge
  • [0020] 4 b Blade efflux edge
  • [0021] 5 Case contour
  • R[0022] a Outer radius of 2
  • R[0023] i Inner radius of 2
  • W Radius point [0024]
  • a′ Conventional contour of [0025] 4 a in 2 b
  • b′ Conventional contour of [0026] 4 b in 2 b
  • h Blade span [0027]
  • α Angle between [0028] 4 a and 5

Claims (3)

What is claimed is:
1. Rotor blade airfoil (2) of an axial-flow turbomachine, more particularly for a fan or a compressor high-pressure stage of a gas turbine system, having means in the tip region (2 b) for improving supersonic performance, characterized by a forward-back sweep of the blade afflux edge 4 a):
2. Rotor blade airfoil of
claim 1
,
characterized in that the tip region (2 b) of the blade efflux edge (4 b) is designed essentially similar to that of the blade afflux edge (4 a).
3. Rotor blade of
claim 1
 or
2
, characterized in that the tip region (2 b) extends across the outer 15% to 25% of the blade span (h).
US09/424,481 1998-03-23 1999-03-23 Rotor blade an axial-flow engine Expired - Lifetime US6358003B2 (en)

Applications Claiming Priority (4)

Application Number Priority Date Filing Date Title
DE19812624 1998-03-23
DE19812624.7 1998-03-23
DE19812624A DE19812624A1 (en) 1998-03-23 1998-03-23 Rotor blade of an axial flow machine
PCT/EP1999/001981 WO1999049185A1 (en) 1998-03-23 1999-03-23 Rotor blade of an axial-flow engine

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US20010014285A1 true US20010014285A1 (en) 2001-08-16
US6358003B2 US6358003B2 (en) 2002-03-19

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US (1) US6358003B2 (en)
EP (1) EP0990090B1 (en)
DE (2) DE19812624A1 (en)
WO (1) WO1999049185A1 (en)

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2969230A1 (en) * 2010-12-15 2012-06-22 Snecma COMPRESSOR BLADE WITH IMPROVED STACKING LAW
CN103946110A (en) * 2011-09-29 2014-07-23 斯奈克玛 Blade for a fan of a turbomachine, notably of the unducted fan type, corresponding fan and corresponding turbomachine
CN104583604A (en) * 2012-07-12 2015-04-29 斯奈克玛 Turbomachine vane having an airfoil designed to provide improved aerodynamic and mechanical properties
CN105332948A (en) * 2015-10-23 2016-02-17 上海交通大学 Improved compressor blade and achieving method thereof
US9695695B2 (en) 2012-01-30 2017-07-04 Snecma Turbojet fan blade
CN107023513A (en) * 2017-06-16 2017-08-08 广东美的制冷设备有限公司 Axial-flow windwheel and air conditioner
CN113167122A (en) * 2018-12-12 2021-07-23 索拉透平公司 Modal response tuned turbine blade

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TWI256444B (en) * 2004-05-06 2006-06-11 Sunonwealth Electr Mach Ind Co Air outlet structure for an axial-flow fan
DE502004010281D1 (en) * 2004-06-02 2009-12-03 Rolls Royce Deutschland Compressor blade, especially for the fan of aircraft engines
CN101044213B (en) * 2004-10-12 2014-04-02 卢米尼克斯股份有限公司 Methods for forming dyed microspheres and populations of dyed microspheres
JP4863162B2 (en) * 2006-05-26 2012-01-25 株式会社Ｉｈｉ Fan blade of turbofan engine
DE102006026968A1 (en) 2006-06-09 2008-01-24 Rolls-Royce Deutschland Ltd & Co Kg Turbomachine with rotors of high specific energy output
JP4911344B2 (en) * 2006-07-04 2012-04-04 株式会社Ｉｈｉ Turbofan engine
GB0701866D0 (en) 2007-01-31 2007-03-14 Rolls Royce Plc Tone noise reduction in turbomachines
US20100303604A1 (en) * 2009-05-27 2010-12-02 Dresser-Rand Company System and method to reduce acoustic signature using profiled stage design
US8668446B2 (en) 2010-08-31 2014-03-11 General Electric Company Supersonic compressor rotor and method of assembling same
US9022730B2 (en) 2010-10-08 2015-05-05 General Electric Company Supersonic compressor startup support system
US8864454B2 (en) 2010-10-28 2014-10-21 General Electric Company System and method of assembling a supersonic compressor system including a supersonic compressor rotor and a compressor assembly
US8657571B2 (en) 2010-12-21 2014-02-25 General Electric Company Supersonic compressor rotor and methods for assembling same
US8827640B2 (en) 2011-03-01 2014-09-09 General Electric Company System and methods of assembling a supersonic compressor rotor including a radial flow channel
US8770929B2 (en) 2011-05-27 2014-07-08 General Electric Company Supersonic compressor rotor and method of compressing a fluid
US8550770B2 (en) 2011-05-27 2013-10-08 General Electric Company Supersonic compressor startup support system
US9102397B2 (en) * 2011-12-20 2015-08-11 General Electric Company Airfoils including tip profile for noise reduction and method for fabricating same
FR2999151B1 (en) * 2012-12-07 2017-01-27 Snecma PROPELLER BLADE FOR TURBOMACHINE
FR3021706B1 (en) * 2014-05-28 2020-05-15 Safran Aircraft Engines AIRCRAFT TURBOPROPELLER COMPRISING TWO COAXIAL PROPELLERS.

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Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2969230A1 (en) * 2010-12-15 2012-06-22 Snecma COMPRESSOR BLADE WITH IMPROVED STACKING LAW
EP2652336B1 (en) * 2010-12-15 2019-05-01 Safran Aircraft Engines Turbomachine blade with improved stacking line
CN103946110A (en) * 2011-09-29 2014-07-23 斯奈克玛 Blade for a fan of a turbomachine, notably of the unducted fan type, corresponding fan and corresponding turbomachine
JP2015501247A (en) * 2011-09-29 2015-01-15 スネクマ Especially for ducted fan turbomachine fan blades, corresponding fans, and corresponding turbomachines
RU2606787C2 (en) * 2011-09-29 2017-01-10 Снекма Turbine machine propeller blade, particularly fan without fairing, and corresponding propeller and turbo machine
US9630704B2 (en) 2011-09-29 2017-04-25 Snecma Blade for a fan of a turbomachine, notably of the unducted fan type, corresponding fan and corresponding turbomachine
US9695695B2 (en) 2012-01-30 2017-07-04 Snecma Turbojet fan blade
CN104583604A (en) * 2012-07-12 2015-04-29 斯奈克玛 Turbomachine vane having an airfoil designed to provide improved aerodynamic and mechanical properties
CN105332948A (en) * 2015-10-23 2016-02-17 上海交通大学 Improved compressor blade and achieving method thereof
CN107023513A (en) * 2017-06-16 2017-08-08 广东美的制冷设备有限公司 Axial-flow windwheel and air conditioner
CN113167122A (en) * 2018-12-12 2021-07-23 索拉透平公司 Modal response tuned turbine blade

Also Published As

Publication number Publication date
DE59907613D1 (en) 2003-12-11
WO1999049185A1 (en) 1999-09-30
EP0990090A1 (en) 2000-04-05
DE19812624A1 (en) 1999-09-30
US6358003B2 (en) 2002-03-19
EP0990090B1 (en) 2003-11-05

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