EP2299124A1 - Aube de rotor pour un compresseur axial - Google Patents

Aube de rotor pour un compresseur axial Download PDF

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Publication number
EP2299124A1
EP2299124A1 EP09011392A EP09011392A EP2299124A1 EP 2299124 A1 EP2299124 A1 EP 2299124A1 EP 09011392 A EP09011392 A EP 09011392A EP 09011392 A EP09011392 A EP 09011392A EP 2299124 A1 EP2299124 A1 EP 2299124A1
Authority
EP
European Patent Office
Prior art keywords
blade
chord
profiles
blade tip
compressor
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Withdrawn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
EP09011392A
Other languages
German (de)
English (en)
Inventor
Christian Dr. Cornelius
Georg Kröger
Eberhard Nicke
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.)
Deutsches Zentrum fuer Luft und Raumfahrt eV
Siemens AG
Original Assignee
Deutsches Zentrum fuer Luft und Raumfahrt eV
Siemens AG
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 Deutsches Zentrum fuer Luft und Raumfahrt eV, Siemens AG filed Critical Deutsches Zentrum fuer Luft und Raumfahrt eV
Priority to EP09011392A priority Critical patent/EP2299124A1/fr
Priority to EP10743094.4A priority patent/EP2473743B1/fr
Priority to US13/393,264 priority patent/US8911215B2/en
Priority to JP2012527268A priority patent/JP5678066B2/ja
Priority to CN201080039406.0A priority patent/CN102483072B/zh
Priority to HUE10743094A priority patent/HUE025789T2/en
Priority to PCT/EP2010/061580 priority patent/WO2011026714A1/fr
Priority to RU2012112930/06A priority patent/RU2534190C2/ru
Priority to ES10743094.4T priority patent/ES2548254T3/es
Publication of EP2299124A1 publication Critical patent/EP2299124A1/fr
Withdrawn legal-status Critical Current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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • 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
    • F05D2250/711—Shape curved convex
    • 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
    • F05D2250/712—Shape curved concave
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10S—TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10S416/00—Fluid reaction surfaces, i.e. impellers
    • Y10S416/02—Formulas of curves
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10S—TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10S416/00—Fluid reaction surfaces, i.e. impellers
    • Y10S416/05—Variable camber or chord length

Definitions

  • the invention relates to a compressor blade for an axial compressor according to the features of the preamble of claim 1.
  • Compressor blades for axial compressors are known from the prior art in a large scale.
  • the EP 0 991 866 B1 a compressor blade having a profile whose suction side contour has a radius of curvature smaller than half the length of the chord on a suction side intersection with a reference line perpendicular to the chord at 5% of the chord length. This is to be achieved that after a comparatively short distance of the flow around the blade on the suction side, the maximum velocity is reached and the location of the envelope of flow from laminar to turbulent coincides with the location of the maximum velocity, whereby this profile has a particularly large work area in the It efficiently compresses the gas flow.
  • Radialspaltmanne is, for example, from SU 1 751 430-A1 known to form the blade tip of blades of an axial compressor S shape.
  • the skeleton line of the profile is formed by two mutually opposite circular arcs, which merge into one another at a turning point.
  • the inflection point is in the range between 5% and 15% of the relative chord length.
  • the leading edge portion is rotated toward the suction side of the airfoil, whereby the forward, ie upstream, portion of the profile has a reverse curvature as compared to the aft, ie, downstream, portion of the airfoil.
  • the object of the invention is to provide a compressor blade with a blade tip, which has particularly low leakage currents and radial gap losses during operation in a turbomachine.
  • a compressor rotor blade for an axial compressor with a curved blade, which comprises a pressure side wall and a suction side wall, which in each case from a common leading edge to a common trailing edge and on the other hand to form a span from a mounting side end of the blade to a
  • the airfoil has a profile with a suction side contour and a pressure side contour, an at least partially curved skeleton line, and a straight chord, which contours, skeleton line, and chord each extend from a leading edge point located on the leading edge to an airfoil extend on the trailing edge arranged rear edge point, wherein that at least some of the skeleton lines of the blade point profile have at least two inflection points.
  • the invention is based on the finding that losses in the radial gap can be reduced if one for the losses also responsible crevice vertebrae is influenced accordingly.
  • the gap vortex which is generated and driven by the gap mass flow, compared to a conventional airfoil tip profile, now later, ie at a downstream point, arise.
  • the thus resulting relative to the conventional profile gap vortex can be explained by a lower load on the improved profile at the front edge.
  • a stronger local impulse for generating the crevice vertebra should now be generated, in which case its fluidic support should decrease considerably more than in the conventional profile.
  • the skeleton lines In order to produce the desired gap vortex, at least some of the skeleton lines, preferably all the skeleton lines of the blade tip-side profiles, have at least two inflection points. Due to the presence of two inflection points in the skeleton line and the use of a conventional thickness distribution, the blade tip-side profiles, and also the suction side contour and the pressure side contour have a rather unusual for the expert eye kink, which is hereinafter referred to as profile bend with respect to the relevant profile.
  • the profile crease itself causes in its place a local increase in the split mass flow, which drives the crevice vortex more than before, as desired, and expels it from the suction side of the airfoil.
  • the mass flow density in the radial gap drops considerably more than when using previous profilings on the blade tip. Overall, this results in a reduced gap mass flow, compared with the conventional profilings. Due to the suction-side contour of the profile bend, the gap vortex develops along a line which also has a bend downstream of the bend of the suction side contour. The early kinking of the crevice vortex coincides with the sharp increase in the mass flow density in the radial gap to its maximum and the subsequent drop of the same together. The gap vortex line is after her kink at a larger angle from the suction side wall than in the conventional profile of the case.
  • the profiling according to the invention causes less radial gap losses and less blocking of the flow field at the outlet of the blade row.
  • the first of the two turning points in the case of perpendicular projection on the chord on this one first projection point, which is removed from the leading edge point between 10% and 30% of the length of the chord.
  • the second of the two inflection points in the case of perpendicular projection onto the chord on this one second projection point, which is located from the leading edge point between 30% and 50% of the length of the chord.
  • the two turning points are at least 3% of the length of the profile chord apart.
  • the skeleton lines of the profiles comprise a front section, which extends in each case from the leading edge point to an end point of the front section, the projection point of which projects perpendicularly onto the chord from the leading edge point is between 2% and 10% of the length of the chord, with at least some of the front portions, preferably all of the front portions of the blade tip side profiles having a radius of curvature greater than 100 times the chord.
  • the front portions of the skeleton line of blade tip side profiles respectively correspond to a straight line, or at least almost.
  • the profile in the respective front section is symmetrical-virtually without buckling-which means that even the local velocity distribution around the blade tip-side leading edge region of the blade leaves virtually no pressure potential from the pressure side to the suction side. Since the pressure potential between the pressure side and the suction side in the leading edge region is considered as the cause of the crevice vortex and thus as a cause for the gap losses, this relief of the leading edge region causes a weakening and a delayed, ie downstream occurrence of the crevice vortex.
  • the suction side contour and the pressure side contour of blade tip side profiles in the front portion of the skeleton line are symmetrical or in a wedge shape with almost straight contour sections on the suction side and pressure side.
  • each front section has an angle of attack with respect to an incoming gas flow, wherein in addition to or instead of the almost straight front skeleton line section at least some of the angles, but preferably all angles of attack of the blade tip side profiles are smaller than the angle of attack of the other profiles of the airfoil.
  • the angle of attack of the front skeleton line section blade tip-side profiles are less than 5 °, preferably even equal to 0 °.
  • leading edge points preferably all the leading edge points of the blade tip-side profiles
  • leading edge points of the blade tip-side profiles can be arranged further upstream than the leading edge points of the remaining profiles of the blade leaf.
  • leading edge of the profiles for blade tips is preceded by an extension of the profile to the front - in the upstream direction - compared to the rest of the leading edge. This has the consequence that no radial pressure gradient can act in the leading edge region of the blade tip, so that it can not come to a potential between the pressure side and suction side even with the radial pressure distribution.
  • only the skeleton lines of the profiles present in the area of the blade tip have two points of inflection, wherein the blade tip area comprises a maximum of 20% of the span width of the blade tip.
  • the remaining area of the airfoil, from a mounting-side airfoil end to a blade height of at least 80% of the span, may be profiled in a conventional manner.
  • the invention relates to a modified airfoil tip of compressor rotor blades arranged in a rim for axial compressors.
  • the skeleton lines comprise a rear portion which extends from a starting point of the rear portion to the trailing edge point, wherein the rear portion of at least some, preferably all blade tip side skeleton lines has a greater curvature than the rear portions of skeleton lines of the rest Profiles of the airfoil.
  • the section starting point of the rear skeleton line section when projected perpendicularly onto the chord, predetermines a projection point located on the chord, which is at most 60% of the length of the chord from the leading edge point.
  • the trailing edge is therefore more curved in the blade tip area than in the remaining area of the blade.
  • the increased curvature leads to a larger work conversion in the preferably rear 40% of the airfoil, so that overall the load of the airfoil is displaced to the rear.
  • This embodiment can serve as a balance of relief on the leading edge to achieve despite the relief of the blade tip side profile in the front region of the chord still a high work transformation.
  • the flow of the following guide blade in the outer annular wall region can thus also be improved by reducing the blockage in the blade tip region of the compressor blade. This reduces the local misfire of the downstream vanes.
  • Preferred dimensions are at least some, preferably all of the blade tip side profiles in the "Aft-Loaded Design” and the rest, d. H. not blade tip-side profiles in the "front-loaded design” designed.
  • the gap vortex responsible for the gap losses can be influenced very efficiently, although the suction side contour and the pressure side contour have at least three successive curved sections with alternating signs, which adjoin adjacent curved sections in each case a turning point.
  • This can be achieved with a suitable thickness distribution, which is applied in a conventional manner perpendicular and symmetrical, ie on both sides in equal parts on the skeleton line.
  • Such measures lead on the suction side to concave contour sections and on the pressure side to convex contour sections, with which the slit vertebrae can be influenced Economics to a particularly simple extent.
  • the blade tip is freestanding.
  • a velocity distribution of the gas adjusts along the suction side contour from the leading edge point to the trailing edge point in the case of a flow around a gas
  • at least some, preferably all blade tip-side profiles are selected such that a maximum velocity occurs at a maximum location, the projection point of which projects perpendicularly onto the chord on it from the leading edge point is between 15% and 40% of the length of the chord.
  • the relevant profiles are selected so that in a subsequent to the maximum location suction side portion of Sauguzeenkontur with a maximum length of 15% of the length of the chord, a gradient of the speed sets, the slope is maximum.
  • the gap vortex is severely underserved for its size, which causes it to move away from the surface of the suction side at a larger angle. This leads to particularly low gap losses in an axial compressor whose rotor is equipped with the compressor rotor blades according to the invention.
  • FIG. 9 and FIG. 10 each show a freestanding compressor blade 10 from different perspectives.
  • Their blade leaf 12 comprises a pressure side wall 14 and a suction side wall 16, which on the one hand in each case by a common, flowed by the gas flow leading edge 18 to a common trailing edge 20 and on the other under education a span of one in 9 and FIG. 10 not shown fastening side airfoil end to a blade tip 22 extend.
  • FIG. 9 the perspective is chosen so that the view of the trailing edge 20 of the airfoil 12 falls
  • FIG. 10 the view of the leading edge 18 of the blade 12 falls on the attachment side airfoil end can be provided in a known manner, a platform and arranged thereon a blade root.
  • the manner of attachment of the blade root of the compressor blade 10 is designed either dovetail, fir-tree or hammer-shaped.
  • the compressor blade may also be welded to a rotor.
  • the orientation of the airfoil 12 is such that the airfoil 12 extends from the leading edge 18 to the trailing edge 20 in approximately the axial direction of the axial compressor which is in the 9 and FIG. 10 associated coordinate system with the axis X is designated.
  • the radial direction of the axial compressor coincides with the Z axis of the illustrated coordinate system and the tangential direction, ie the circumferential direction with the Y axis.
  • a span of the airfoil 12 is thus detected in the Z-axis direction.
  • compressor blades 10 for axial compressors are designed in such a way that different or even identical profiles are strung together along an unillustrated rectilinear or even slightly curved stacking axis whose enclosed space predetermines the blade 12.
  • each profile has a centroid on the stack axis.
  • a profile is understood in detail to mean an endless polyline which has a suction side contour and a pressure side contour an airfoil.
  • the contours meet on the one hand in a leading edge point and on the other hand in a trailing edge point, which are also part of the profile and lie on the corresponding edge of the airfoil.
  • the profile represents the contour of a cross section through the blade for a particular blade height, wherein the cross section either perpendicular to the radial direction of the axial compressor or even slightly inclined - may be oriented - according to an annular channel contraction.
  • FIG. 9 are printed page contours 40 of three profiles 28, 30 shown in full line.
  • FIG. 10 a plurality of suction side contours 42 of profiles 28, 30 different blade heights are also shown in solid lines.
  • FIG. 10 illustrated curved airfoil 12 has a comparison with the prior art according to the invention modified blade tip region 43, the specific configuration and operation will be described in more detail below in detail.
  • the first profile 28 shown in dotted line style shows a cross section through the compressor blade 10 according to FIG. 10 at an airfoil height of half the span of the airfoil 12.
  • the profile 28 may be a conventional profile known in the art.
  • the profile 30 shown in full line shows a cross section through the compressor blade 10 according to the invention FIG. 10 in the area 43 of the blade tip 22.
  • Each profile 28, 30 according to FIG. 1 has a skeleton line associated with it, for reasons of clarity in FIG. 1 only one skeleton line 32 of the blade tip side profile 30 is shown in dashed line style.
  • the skeleton line 32 begins at a leading edge point 24, terminates at an associated trailing edge point 26, and is always centered between the printing page contour 40 and suction side contour 42. It is also known as profile centerline.
  • profiles are also defined in the prior art with the aid of a straight chord.
  • the chord is a straight line which extends from the leading edge point to the trailing edge point.
  • FIG. 1 only one profile chord 34 for the blade tip-side profile 30 is shown. Since the profile chord 34 is subsequently used for the geometric definition of significant points of the profile 30, its length is normalized to one, wherein in the leading edge point 24, the length of the chord 0% and in the trailing edge point 26, the length of the chord 100%. This also means a relative chord length.
  • the normalized chord 34 is indicated by x / c.
  • This in FIG. 1 represented profile 30 is representative of the radially outermost of the blade tip-side profiles 30.
  • Das in FIG. 1 The conventional profile 28 illustrated on the one hand is representative of the profiles known from the prior art and on the other hand for the remaining profiles of the compressor blade 10.
  • the other profiles 28 are understood to mean those which are not arranged on the blade tip side and thus, for example, in the attachment-side region of the blade 12 or can be arranged centrally between the blade tip 22 and the attachment-side blade end.
  • the transition from the conventional profile 28 to the blade tip-side profile 30 takes place, as FIG. 10 shows, stepless.
  • Characteristic of a compressor blade 10 according to the invention is that the skeleton lines 32 of the blade tip-side profiles 30 have at least two turning points 36, 38. That is, the skeleton line 32 upstream of the foremost inflection point 36 has a first curvature portion A of a first curvature and downstream of the first inflection point 36 to the second inflection point 38 a second curvature portion B of a second curvature. The signs of the first curvature and the second curvature are different. Downstream of the second curvature section B, in the second inflection point 38, a third curvature section C follows, whose curvature in turn has a different sign than that of the second curvature.
  • the predominantly convex curved suction side contour 42 has a concave shape in a section D between 35% and 50% of the relative chord length.
  • the mainly concave curved pressure side contour 40 has a portion E which is convex. Contrary to the previous known from the prior art profile shapes for compressor blades of axial compressors this concave Sauguzeenkonturabrough D and convex pressure side contour section E leads to a locally kinking profiling, which is referred to here as a profile kink.
  • the first of the two turning points 36 predestinates on perpendicular to the chord on this a first projection point AP, which is removed from the leading edge point 24 between 10% and 30% of the length of the chord 34 and at the second of the two Turning points 38 in perpendicular projection on the chord 34 on this one second projection point BP pretending, which is removed from the leading edge point 24 between 30% and 50% of the length of the chord 34.
  • FIG. 1 clearly shows that the blade tip-side profile 30 relative to the conventional profile 28 has a forwardly displaced leading edge 18 toward the oncoming gas flow.
  • the vorverlagerte leading edge 18 of the blade tip profile 30 is particularly in the perspective views according to 9 and FIG. 10 recognizable.
  • the skeleton line 32 of blade tip side profiles 30 in a rear portion G has a greater curvature than the rear portions of skeleton lines of the remaining profiles 28 of the airfoil 12.
  • the rear portion G of the skeleton line 32 extends from the section starting point GA up to the trailing edge point 26 of the skeleton line 32, which section start point GA when projecting onto the chord 34 on this one projection point GP, which is removed from the leading edge point 24 a maximum of 60% of the length of the chord 34.
  • the blade tip-side profile 30 comprises a skeleton line 32 with a front portion H.
  • the front portion H of the skeleton line 32 extends from the leading edge point 24 to a projection point HP of the skeleton line 32, which is located at 10% of the length of the chord 34.
  • the projection point HP results from the projection of an end point HE of the front portion H perpendicular to the chord 34.
  • the skeleton line 32 is almost not arched, ie approximately straight.
  • the thickness distribution which is known to be applied perpendicular to the skeleton line 32 on both sides in equal parts, chosen here so that there is a wedge-shaped leading edge region for the blade tip side profiles 30 in principle.
  • a symmetrical course of the suction side contour 42 and pressure side contour 40 is symmetrically desirable.
  • FIG. 2 The velocity distributions along the blade tip side profile 30 and along the conventional profile 28 are contrasted for both the suction side flow and the pressure side flow.
  • Each velocity distribution is along the normalized chord x / c applied.
  • the velocity distribution was detected at that blade height of compressor blades, which is 0.5% of the gap of a radial gap between the blade tip 22 and the surrounding annular wall of the axial compressor of the blade tip 22 away.
  • dashed line style are in FIG. 2 . 3 and FIG. 6 the velocity distributions 48, 50 of a conventional profile 28 for the suction side wall 16 and pressure side wall 14 are shown.
  • the velocity distributions 44, 46 for the suction side wall 16 and pressure side wall 14 of the blade tip-side profile 30 are shown in full line.
  • the lower line represents the velocity distribution for the corresponding pressure side
  • the upper line represents the velocity distribution for the corresponding suction side.
  • the suction side velocity distribution for the blade tip side profile 30 is denoted 44, the pressure side velocity distribution for the blade tip profile 46, the suction side velocity distribution for the conventional profile 28 at 48 and the pressure side velocity distribution for the conventional profile 28 with 50.
  • FIG. 2 shows that with the aid of the present invention modified blade tip portion 43, the blade 12 in the front half, ie in particular on the first 15% of the chord 34 seen from the leading edge point 24, has been relieved.
  • the profile shape of the blade 12 is selected blade tip side so that the speed increase is achieved to a maximum speed in a maximum location at about 20% of the length of the chord 34 in the shortest possible chord section. Furthermore, a comparatively large decrease in the speed of the suction-side gas flow in a profile chord section that is as short as possible is desired in the subsequent 15% of the chord 34 following the maximum location.
  • this speed course along the suction side wall 16 causes a gap vortex responsible for the gap losses is generated with comparatively more energy, but the comparatively low energy is further supplied to this by the large deceleration after reaching the maximum speed, which weakens him all the more , Overall, this leads to reduced radial gap losses.
  • FIG. 4 describes the blade tip-side profile 30 in the un-graded m'-theta coordinate system.
  • the lower picture, FIG. 5 represents a curvature 52 of the suction side contour 42 and a curvature 54 of the pressure side contour 40 above the m 'coordinate. It is clear to see that in the region of a Druckurgiknicks 56 a sharp increase in Machieredifferenz and thus the pressure potential between Saugateenkontur 42 and pressure side contour 40 is formed.
  • FIG. 7 shows the mass flow density of the mass flow, which flows orthogonal to the chord 34 through the radial gap, based on the considered local area.
  • the mass flow density for a conventional profile 28 is indicated at 58, that for the blade tip-side profile 30 at 60.
  • For the blade tip-side profile 30 is a clear relationship between the increase of the pressure potential and the increase in the mass flow density in the radial gap recognize.
  • the mass flow density in the radial gap also reaches its global maximum shortly after the described profile break.
  • the global maximum of the mass flow density for the blade tip-side profile 30 is higher than in the conventional case.
  • the drop in the mass flow density in the radial gap to its maximum is also greater than in the conventional profiling 28th
  • FIG. 8 shows the topology of the Spaltwirbeltrajektoren (slit vortex lines) for the two profiles 28, 30.
  • the gap vortex line for the conventional profile 28 is designated 62, the gap vortex line for the blade tip profile with 64. Relative to the leading edge 18 of the gap vortex occurs at the blade tip side profile 30 much later - Based on the relative chord length of the profile concerned - and then kinks from the suction side wall 16 with a larger angle than in the conventional profiling 28.
  • the early kinking of the crevasse vortex coincides with the sharp increase in mass flow density to its maximum and the subsequent decrease of the same together.
  • the larger angle is due to the larger gradient in both the increase and decrease in mass flow density.
  • the relative to the conventional profile 28 relatively late emergence of the crevice vertebra can be explained by the low load on the improved profile 30 at the front edge 18.
  • the formation of the gap vortex is delayed. This is followed in the region of the suction-side profile bend by a strong increase in the gap mass flow, which drives the gap vortex and expels it from the suction side wall 16 of the blade tip-side profile 30. In the zone downstream of the suction-side profile bend, the mass flow density in the radial gap drops considerably more than in conventional profiling 28. Overall, this results in a lower gap mass flow.
  • the gap vortex line bends after the suction-side profile bend at a higher angle from the suction side wall 16 than in the conventional profiling 28 is the case. From now on, it runs away from the suction sidewall 16 at a greater distance than in the conventional profiling 28.
  • the split flow in the modified profiling 30 thus causes less losses and less blockage of the flow field at the outlet of the blade row.
  • the load is increased by a higher curvature of the profile 30 in the rear 40% of the chord 34.
  • the invention thus relates to a compressor blade 10 for axially flowed compressor preferably stationary gas turbines.
  • the invention provides that in order to reduce radial gap losses, the skeleton line 32 of the blade tip-side profiles 30 of the blade 12 of the compressor blade 10 have at least two inflection points 36, 38. Due to the presence of two inflection points 36, 38, a suction side contour section D which is concave and for the pressure side contour 40 a pressure side contour section results for the suction side contour 42 in the section of 35% to 50% E, which is convex. With the aid of this geometry, it is possible to generate lower-loss splitter vortices in order to increase the overall efficiency of an axial compressor equipped with these compressor rotor blades 10.

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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)
EP09011392A 2009-09-04 2009-09-04 Aube de rotor pour un compresseur axial Withdrawn EP2299124A1 (fr)

Priority Applications (9)

Application Number Priority Date Filing Date Title
EP09011392A EP2299124A1 (fr) 2009-09-04 2009-09-04 Aube de rotor pour un compresseur axial
EP10743094.4A EP2473743B1 (fr) 2009-09-04 2010-08-10 Aube mobile de compresseur pour un compresseur axial
US13/393,264 US8911215B2 (en) 2009-09-04 2010-08-10 Compressor blade for an axial compressor
JP2012527268A JP5678066B2 (ja) 2009-09-04 2010-08-10 軸流コンプレッサー用のコンプレッサーブレード
CN201080039406.0A CN102483072B (zh) 2009-09-04 2010-08-10 用于轴流式压缩机的压缩机转子叶片
HUE10743094A HUE025789T2 (en) 2009-09-04 2010-08-10 Compressor blade for axial compressor
PCT/EP2010/061580 WO2011026714A1 (fr) 2009-09-04 2010-08-10 Aube mobile de compresseur pour un compresseur axial
RU2012112930/06A RU2534190C2 (ru) 2009-09-04 2010-08-10 Компрессорная рабочая лопатка для осевого компрессора
ES10743094.4T ES2548254T3 (es) 2009-09-04 2010-08-10 Álabe móvil de compresor para un compresor axial

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP09011392A EP2299124A1 (fr) 2009-09-04 2009-09-04 Aube de rotor pour un compresseur axial

Publications (1)

Publication Number Publication Date
EP2299124A1 true EP2299124A1 (fr) 2011-03-23

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EP09011392A Withdrawn EP2299124A1 (fr) 2009-09-04 2009-09-04 Aube de rotor pour un compresseur axial
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EP3112590A1 (fr) * 2015-07-01 2017-01-04 General Electric Company Buse bombée de commande de flux secondaire et performance optimale de diffuseur
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FR2991373A1 (fr) * 2012-05-31 2013-12-06 Snecma Aube de soufflante pour turboreacteur d'avion a profil cambre en sections de pied
WO2013178914A1 (fr) * 2012-05-31 2013-12-05 Snecma Aube de soufflante pour turboreacteur d'avion a profil cambre en sections de pied
US11333164B2 (en) 2012-05-31 2022-05-17 Safran Aircraft Engines Airplane turbojet fan blade of cambered profile in its root sections
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WO2014154997A1 (fr) * 2013-03-28 2014-10-02 Turbomeca Diffuseur à ailettes d'un compresseur radial ou mixte
RU2651905C2 (ru) * 2013-03-28 2018-04-24 Тюрбомека Лопаточный диффузор радиального или диагонального компрессора
EP3088663A1 (fr) * 2015-04-28 2016-11-02 Siemens Aktiengesellschaft Procédé de profilage d'une aube
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EP3112590A1 (fr) * 2015-07-01 2017-01-04 General Electric Company Buse bombée de commande de flux secondaire et performance optimale de diffuseur
US10323528B2 (en) 2015-07-01 2019-06-18 General Electric Company Bulged nozzle for control of secondary flow and optimal diffuser performance
EP3470627A1 (fr) * 2017-10-12 2019-04-17 United Technologies Corporation Surface portante de moteur à turbine à gaz
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WO2011026714A1 (fr) 2011-03-10
JP5678066B2 (ja) 2015-02-25
HUE025789T2 (en) 2016-05-30
EP2473743A1 (fr) 2012-07-11
RU2534190C2 (ru) 2014-11-27
JP2013503999A (ja) 2013-02-04
US8911215B2 (en) 2014-12-16
CN102483072B (zh) 2015-04-08
RU2012112930A (ru) 2013-10-10
CN102483072A (zh) 2012-05-30
ES2548254T3 (es) 2015-10-15
EP2473743B1 (fr) 2015-07-29
US20120230834A1 (en) 2012-09-13

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