EP3392459A1 - Verdichterschaufeln - Google Patents

Verdichterschaufeln Download PDF

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Publication number
EP3392459A1
EP3392459A1 EP17175536.6A EP17175536A EP3392459A1 EP 3392459 A1 EP3392459 A1 EP 3392459A1 EP 17175536 A EP17175536 A EP 17175536A EP 3392459 A1 EP3392459 A1 EP 3392459A1
Authority
EP
European Patent Office
Prior art keywords
aerofoil
free end
machine
tip
vicinity
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Withdrawn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
EP17175536.6A
Other languages
English (en)
French (fr)
Inventor
Christopher Hall
Anastasios Kovanis
Anthony Dickens
James Taylor
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.)
Rolls Royce PLC
Original Assignee
Rolls Royce PLC
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 Rolls Royce PLC filed Critical Rolls Royce PLC
Publication of EP3392459A1 publication Critical patent/EP3392459A1/de
Withdrawn legal-status Critical Current

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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
    • F04D29/324—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
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel
    • 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
    • F04D29/38—Blades
    • F04D29/384—Blades characterised by 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
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/542—Bladed diffusers
    • F04D29/544—Blade shapes
    • 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
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/125—Fluid guiding means, e.g. vanes related to the tip of a stator vane
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade
    • 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

Definitions

  • the present disclosure concerns fluid-washed components of axial flow machines, e.g. such as gas turbine engines. More specifically, the disclosure concerns blades and/or vanes used to drive or redirect flow through an axial flow machine.
  • tip leakage flow remains an ongoing efficiency loss in axial compressors.
  • the efficiency is degraded not only by the volume of fluid passing over the tip from the pressure to the suction side of the aerofoil, but also by the subsequent mixing losses when the leakage flow mixes with the mainstream flow downstream of the compressor blade/vane.
  • US 2008/0213098 A1 and corresponding EP 1 953 344 A1 disclose an attempt to resolve tip leakage problems by providing a winglet at the compressor blade tip, characterised by a sharp curvature, i.e. a corner, at the tip compared to the remainder of the blade.
  • the benefits disclosed for such a winglet comprise a reduction in driving pressure difference across the blade tip, a reduction in tip clearance and the shielding of aerodynamic shocks of the flow against interaction with the casing.
  • an axial flow compressor having an axis of rotation and comprising an array of aerofoils angularly spaced about said axis, the aerofoils comprising leading and trailing edges extending in a direction spanning a flow region defined between a radially inner rotor component and a radially outer casing, each aerofoil comprising opposing pressure and suction surfaces extending between the leading and trailing edges and terminating at a free end of the aerofoil, wherein each aerofoil leans towards the pressure surface by an angle of between 10° and 80° in the vicinity of the free end of the aerofoil.
  • the vicinity of the free end of the aerofoil may comprise a minority of the height of the aerofoil towards the free end.
  • the vicinity of the free end of the aerofoil may include the free end of the aerofoil.
  • the vicinity of the free end may be close to the free end.
  • the vicinity of the free end of the aerofoil may comprise up to 20% or 15% of the aerofoil height.
  • the vicinity of the free end of the aerofoil may comprise less than or equal to 12%, 10% or 8% of the aerofoil height.
  • the angle that the aerofoil leans towards the pressure surface by may be the angle between the aerofoil and the radial direction.
  • the angle that the aerofoil leans towards the pressure surface by may be the angle between the free end of the aerofoil and the radial direction.
  • the vicinity of the free end of the aerofoil may comprise greater than 3%, 4% or 5% of the aerofoil height.
  • a specific aerofoil lean over 5-10% of the blade height towards its free end may be used specifically for the purpose of tip leakage reduction.
  • the lean towards the pressure surface may be referred to herein as a negative lean.
  • the lean may represent a relatively aggressive change in orientation of the aerofoil, e.g. a change in orientation of between 20° and 80°.
  • the angle formed at the tip relative to a radial direction e.g. the angle of lean
  • the angle formed at the tip relative to a radial direction may be in the vicinity of 40-60°.
  • the total change in orientation in the vicinity of the free end may be greater than the angle at the tip if the aerofoil has a positive lean over an aerofoil region adjacent the vicinity of the free end.
  • Each aerofoil may lean towards the pressure surface by an angle of at least 10°, 20°, 30° or 40° in the vicinity of, or at, the free end of the aerofoil.
  • Each aerofoil may lean towards the pressure surface by an oblique angle of less than or equal to 75°, 70°, 65°, 60°, 55°, 50° or 45° in the vicinity of, or at, the free end.
  • An aerofoil change in orientation or negative lean in the vicinity of the tip of approximately 50-60° may be used.
  • a change in orientation of the aerofoil towards the tip may be used to trigger separation of the flow over the aerofoil surface in the vicinity of the tip gap. This may reduce the effective flow area for flow over the tip between the pressure and suction surfaces.
  • the lean towards the pressure surface may be over a majority of the chord length, such as all, or substantially all, of the chord length.
  • the lean may be substantially constant over the chord length of the aerofoil or may vary over the chord length, e.g. having a greater lean towards the trailing edge.
  • the aerofoil cross-section profile may remain substantially constant through the height of the aerofoil in the vicinity of the free end.
  • the aerofoil may lean away from a radial direction towards the axis of rotation of the compressor in the vicinity of the free end.
  • the aerofoil may be angled and/or curved towards the pressure surface in the vicinity of the free end.
  • the lean towards the pressure surface may be relative to a radial direction defined with respect to the axis of rotation. Additionally or alternatively, the lean towards the pressure surface in the vicinity of the free end may be defined with respect to the orientation, or average orientation, of the remainder of the aerofoil. Thus the orientation may be defined in absolute/Cartesian or relative terms.
  • the remainder of the aerofoil (e.g. the remainder of the aerofoil height) may lean towards either the pressure or suction surface.
  • the remainder of the aerofoil may comprise a positive lean of less than 40°, 30 °, 20° or 10°.
  • the aerofoil may comprise a negative turning point or height.
  • the lean of the aerofoil towards the pressure side may increase from the turning point/height to the tip.
  • the rate of change of orientation of the aerofoil may increase from the turning point/height to, or towards, the tip.
  • the aerofoil may comprise two turning points.
  • the angular orientation of the aerofoil at the tip may be less than or equal to 45°, 40° or 35° from the orientation of the aerofoil at the turning point/height.
  • the aerofoil may be smoothly contoured/curved from the turning point/height to the free end.
  • any of the angular definitions provided herein with respect to the aerofoil, or a blade or vane may be references to the component as a whole or else a longitudinal section thereof, e.g. from the root to the tip.
  • the aerofoils may be mounted to the compressor rotor, e.g. blades.
  • the aerofoils may be mounted to a compressor drum. Additionally or alternatively the aerofoils may be mounted to the compressor casing/stator, e.g. vanes.
  • a gap may be provided between the free end of the aerofoils and the opposing casing/rotor surface.
  • a blade for an axial flow machine comprising leading and trailing edges extending from a root of the blade to a tip of the blade.
  • the blade comprises opposing major surfaces extending between the leading and trailing edges and terminating at a blade tip.
  • the blade leans towards the pressure surface by an angle of between 10° and 80° in the vicinity of the blade tip.
  • the blade may be an aerofoil.
  • the opposing major surfaces may be a pressure surface and a suction surface.
  • the vicinity of the blade tip may include the blade tip.
  • the vicinity of the blade tip may be a minority of the height of the aerofoil towards the tip.
  • the blade may be a compressor blade.
  • a rotor or rotor drum may be provided comprising one or more of the compressor blades.
  • Each blade may lean towards the pressure surface by an oblique angle of less than or equal to 75°, 70°, 65°, 60°, 55°, 50°, 45°, 40°, 35°, 30°, 25°, 20° or 15° in the vicinity of, or at, the blade tip.
  • a stator vane for an axial flow machine comprising leading and trailing edges extending from a base of the vane to a tip of the vane, the vane comprising opposing major surfaces extending between the leading and trailing edges and terminating at a vane tip, wherein the vane leans towards the pressure surface by an angle of between 10° and 80° in the vicinity of the tip.
  • a stator or casing may be provided comprising one or more of the stator vanes.
  • a gas turbine engine comprising any or any combination of a compressor, a blade, a rotor, a vane or a stator/casing according to a preceding aspect of the disclosure.
  • a change in orientation in the region of 20-70° in the final portion of the blade/vane close to the tip e.g. an angle of approximately 50-60° relative to a radial direction, or the remainder of the blade/vane, is generally optimal and that it is not desirable for the aim of the present disclosure to cause a change in direction approaching 90°, akin to a winglet.
  • a gas turbine engine is generally indicated at 10, having a principal and rotational axis 11.
  • the engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate pressure turbine 18, a low-pressure turbine 19 and an exhaust nozzle 20.
  • a nacelle 21 generally surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.
  • the gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust.
  • the intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to the high pressure compressor 15 where further compression takes place.
  • the compressed air from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture combusted.
  • the resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust.
  • the high 17, intermediate 18 and low 19 pressure turbines drive respectively the high pressure compressor 15, intermediate pressure compressor 14 and fan 13, each by a suitable interconnecting shaft.
  • gas turbine engines to which the present disclosure may be applied may have alternative configurations.
  • such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines.
  • the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.
  • the present concept was devised for use within the compressor, i.e. the intermediate 14 or high 15 pressure compressor of engine 10.
  • the disclosure is not limited thereto and may be applied to other axial flow machines that suffer from tip gap flow losses, causing flow efficiency losses.
  • the compressor 14 comprises a compressor drum 14A and a plurality of rows of compressor blades 14B mounted thereto. Each row of blades comprises a radial array of angularly spaced blades.
  • a casing 14C Surrounding the compressor is a casing 14C arranged around the common axis 11 to define an annular flow passage between the compressor drum 14A and casing 14C.
  • the casing 14C comprises a plurality of rows of stator vanes 14D mounted thereto.
  • Each row of vanes 14D comprises a radial array of angularly spaced vanes.
  • Successive compressor stages comprise paired rows of blades 14B and vanes 14C.
  • Blades depend outwardly from the rotor drum 14A towards the casing 14C and terminate at a free end, or tip, with a small clearance from the casing, thereby leaving a tip gap.
  • Vanes 14D depend inwardly from the casing towards the drum 14A and terminate at a free end, or tip, with a small clearance from the drum, thereby leaving a tip gap.
  • the blades 14B and vanes 14D span the majority of the annular flow passage.
  • FIGs. 2-4 there is shown an example of an individual blade 24 according to the disclosure alongside a datum blade 26.
  • the blade 24 is aerofoil-shaped in section over its entire height from its root end 28 to its tip 30.
  • a conventional root formation is not shown but would be provided at the end 28 for mounting the blade on the drum 14A.
  • the blade 24 has a leading edge 32, a trailing edge 34 and opposing convex/pressure 36 and concave/suction 38 faces extending between the leading and trailing edges.
  • the pressure and suction surfaces may be opposingly oriented and the aerofoil curvature (i.e. the sectional profile) may be mirrored.
  • Figs. 2-4 there is shown a change in orientation of the blade 24, whereby the blade is bent/curved towards the pressure side 36 close to the tip 30.
  • the change in orientation may be referred to herein as a lip region 40.
  • the lip region 40 is shown in an exaggerated/schematic manner in Fig. 4 .
  • Figs. 3 and 4 the lip region 40 is shown in the same orientation so that comparison can be draw between the different schematic and three-dimensional views.
  • the end face of the blade lies in a tangential/circumferential plane with respect to the axis of rotation. Therefore the end face is generally perpendicular to a height/radial direction of the blade but obliquely angled relative to the blade (e.g. the pressure and/or suction surfaces thereof) in the lip region 40.
  • the internal angle formed between the end face at the tip and the direction of the blade is thus equal to the angle of lean in the lip region 40. This may result in the end face area being greater than that of the cross-sectional area of the blade.
  • a turning point A at which the blade 24 (e.g. the leading edge 32 or a centreline thereof) starts to lean negatively towards pressure side 36.
  • the point A may be considered to comprise a line, e.g. a chord line, along the blade 24 when viewed from the side or in cross section.
  • the line is of substantially constant height in this example but the relevant region of the blade could vary between the leading 32 and trailing 34 edges within the limits disclosed herein.
  • the blade turns towards the pressure side, whereas the remainder of the blade has a generally neutral or positive lean.
  • the angular orientation of the blade 24 increases gradually from point A towards the tip 30 such that a maximum lean angle is achieved at the tip.
  • Fig. 4 shows a blade height, L, and a height, L1, from point A to the tip 30.
  • the height L1 may thus define the portion of the blade height L that comprises the lip region 40.
  • the angle, ⁇ is defined as the angle formed between the blade 24 or leading edge 32 and a radial direction R with respect to the axis of rotation of the compressor in use.
  • the local negative lean at the tip of the aerofoil increases the angle of attack of the over-tip leakage flow (from pressure to suction side), which increases the blockage in the tip gap, thus reducing the effective tip gap area and resultant leakage flow.
  • the aerodynamic performance of the blade/vane and axial flow compressor as a whole can thus be achieved. Whilst it is not a limitation of aspects of this disclosure, it is feasible that a leakage flow reduction of e.g. 3-7% may be achieved. When considered in conjunction with the mixing losses created as the leakage flow mixes with the mainstream flow through the compressor, it will be appreciated that such aero-efficiency improvements can have a significant impact on performance.
  • the design of the remainder of the blade may be modified slightly to accommodate the lip region 40.
  • a positive lean may be provided in the leading edge 32 and/or blade 24 leading up to the turning point/line A, as can be seen in Fig. 4 .
  • a blade/vane may be made according to the extremes of the depth of the blade/vane in a lateral/circumferential direction, e.g. when viewed front on as shown in Fig. 4 .
  • the depth dimension, S can be defined as shown in Fig. 4 , which may be divided into negative, S1, and positive, S2, components relative to a radial line R1 coinciding with the leading edge 32 at its base/root end 28.
  • the value of S in the tip region 40 may be greater than, equal to, or less than the value of S for the remainder of the blade/vane. Any such relationship may be used in conjunction with any other relationship or geometric feature defined herein as an aspect of the disclosure.
  • the blade/vane may have a more conventional lean of less than 20°, 15° or 10° spread over the remainder of the blade/vane height, as well as a more aggressive lean towards the tip as described herein.
  • the negative lean could also help with the tip rubbing the casing liner since the leaning tip would act more akin to a cutting tool and hence provide a cleaner rub. This may reduce the heat resulting from a tip rubbing against the casing which is linked with the tip cracking.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP17175536.6A 2017-04-18 2017-06-12 Verdichterschaufeln Withdrawn EP3392459A1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GR20170100180 2017-04-18

Publications (1)

Publication Number Publication Date
EP3392459A1 true EP3392459A1 (de) 2018-10-24

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ID=59054024

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EP17175536.6A Withdrawn EP3392459A1 (de) 2017-04-18 2017-06-12 Verdichterschaufeln

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US (1) US20180298912A1 (de)
EP (1) EP3392459A1 (de)

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3059353B1 (fr) * 2016-11-29 2019-05-17 Safran Aircraft Engines Aube directrice de sortie pour turbomachine d'aeronef, comprenant une zone coudee de passage de lubrifiant presentant une conception amelioree
US10982549B2 (en) * 2017-04-17 2021-04-20 General Electric Company Stator vanes including curved trailing edges
EP3421725A1 (de) * 2017-06-26 2019-01-02 Siemens Aktiengesellschaft Kompressorschaufel
USD911512S1 (en) * 2018-01-31 2021-02-23 Carrier Corporation Axial flow fan
US10947851B2 (en) * 2018-12-19 2021-03-16 Raytheon Technologies Corporation Local pressure side blade tip lean
BE1028118B1 (fr) * 2020-03-02 2021-09-27 Safran Aero Boosters Aube pour compresseur de turbomachine

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB946794A (en) * 1961-03-06 1964-01-15 Colchester Woods Improvements in and relating to axial flow fans or compressors
EP1953344A1 (de) 2007-02-05 2008-08-06 Siemens Aktiengesellschaft Turbinenschaufel
US20120243975A1 (en) * 2011-03-25 2012-09-27 Andrew Breeze-Stringfellow Compressor airfoil with tip dihedral
WO2014109959A1 (en) * 2013-01-08 2014-07-17 United Technologies Corporation Gas turbine engine rotor blade
EP2990602A1 (de) * 2014-08-27 2016-03-02 Pratt & Whitney Canada Corp. Verdichterschaufel und verfahren zur herstellung einer schaufel

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB946794A (en) * 1961-03-06 1964-01-15 Colchester Woods Improvements in and relating to axial flow fans or compressors
EP1953344A1 (de) 2007-02-05 2008-08-06 Siemens Aktiengesellschaft Turbinenschaufel
US20080213098A1 (en) 2007-02-05 2008-09-04 Matthias Neef Free-standing turbine blade
US20120243975A1 (en) * 2011-03-25 2012-09-27 Andrew Breeze-Stringfellow Compressor airfoil with tip dihedral
WO2014109959A1 (en) * 2013-01-08 2014-07-17 United Technologies Corporation Gas turbine engine rotor blade
EP2990602A1 (de) * 2014-08-27 2016-03-02 Pratt & Whitney Canada Corp. Verdichterschaufel und verfahren zur herstellung einer schaufel

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