EP2299124A1 - Rotor blade for an axial compressor - Google Patents

Abstract

The vane has a vane blade with a suction side wall, which is extended towards a vane blade tip by forming span width by an attached vane blade end. The blade comprises a profile (28) and a blade tip side profile (30) with a suction side contour (42) and a pressure side contour (40), and a partially curved skeleton line (32), and a linear chord (34) for vane blade height. The contours, the skeleton line and the chord are extended from a front edge point (24) to a rear edge point. The skeleton line of the blade tip side profile comprises two inflection points (36, 38).

Description

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. For example, 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.

Furthermore, it is known that so-called radial gap losses occur at the blade tips of compressor rotor blades. In this case, part of the pressure gain during operation of the axial compressor is lost in that a leakage flow is established across the blade tip from a pressure side of the blade to a suction side of the blade. In order to reduce this leakage flow, it is known that a radial gap formed between the blade blade tips and an annular wall of the compressor channel opposite this is always to be kept as small as possible. Nevertheless, minimum sizes of gaps must be adhered to in order to avoid tarnishing of blade tips on the annular wall. This applies especially for transient operating conditions in which thermally induced strains of both the channel wall and blades are not yet completed.

In addition, it was often the case that the previous profiling of blade tips was adapted only to the particular inflow conditions in the area of the annular wall. The actual profiling, however, did not take into account the actual three-dimensional flow effects at the blade tip. Conventionally designed airfoil profilings were therefore not optimally adapted to the complex flow conditions in the area of the blade tip. As a result, there is a considerable potential for improvement in particular with compressor wings with a small span and large relative gap heights (in terms of span).

Because modern, like from the EP 0 991 866 B1 known Turbomachinery Blades have now reached a very high aerodynamic efficiency, arises with the tendency to ever higher profile loads an increasing proportion of the total losses caused by these radial gap losses that occur in the outer wall near the annulus area. A reduction of these significant losses thus causes a significant improvement in the efficiency of turbomachinery and axial compressors.

To reduce these Radialspaltverluste 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. This reduces secondary flow losses and flow non-uniformities at the exit of subsonic compressor blades due to the reduction in the pressure gradient. In particular, while the pressure gradient in front and middle area in the passages between the blades are reduced. According to the SU 1 751 430-A1 For example, 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.

Regardless of existing solutions, there is still a great deal of interest in reducing radial gap losses of turbomachinery in order to further increase the efficiency of these machines.

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.

This object is achieved with 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 For each span along the span, 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. According to the invention, 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. Contrary to the previous general effort to weaken the crevice vertebrae as a whole, according to the invention, 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. Overall, this leads to low flow losses in the radial gap. 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. In the downstream zone behind the bend in the suction side contour, 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. As a result, from now on, the gap vortex with increasing distance from the suction side away than in the conventional profiling. The larger angle is due to the larger gradient of mass flow density of the slit flow in both ascent and waste. Overall, 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.

By reducing the radial gap losses achieved, the efficiency of the blading and thus also the efficiency of a turbomachine equipped with the compressor blade can be substantially improved.

Advantageous embodiments are specified in the subclaims.

Preferred dimensions of 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. At the same time, 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. In particular, with such arranged inflection points, the advantages associated with the invention occur to a particularly great extent. The two turning points are at least 3% of the length of the profile chord apart.

According to a further preferred embodiment of the invention, 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. In other words, the front portions of the skeleton line of blade tip side profiles respectively correspond to a straight line, or at least almost. Accordingly, 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. Preferably, 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.

According to a further advantageous embodiment, 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. Preferably, the angle of attack of the front skeleton line section blade tip-side profiles are less than 5 °, preferably even equal to 0 °. It can thus be said that the leading edge region of the blade tip, in contradiction to the solution according to SU 1751430 A1 , is screwed into the flow, which also ensures that a pressure potential between the pressure and suction side in the leading edge area shovel tip side is avoided. This also prevents the generation of the gap vortex in the leading edge region.

As an alternative or in addition to the proposed further developments, preferably at least some of the leading edge points, preferably all the 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. In other words, the 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.

Preferably, 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.

Accordingly, in principle, the invention relates to a modified airfoil tip of compressor rotor blades arranged in a rim for axial compressors.

According to a further advantageous embodiment, 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. As a result, the exit metal angles of shovel point-side profiles smaller than the outlet metal angle of profiles at the level of half the span or in the region of the mounting side, ie hub side airfoil end. Preferably, 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. Overall, 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 ideegemäß to a particularly simple extent.

Conveniently, the blade tip is freestanding.

If 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. This measure ensures a particularly large impulse for the formation of the crevice vortex. In order to keep the radial gap losses as low as possible, it is provided that the energy supply for the gap vortex particularly fast, d. H. on a particularly short length, particularly decreases. For this purpose, it is provided that the relevant profiles are selected so that in a subsequent to the maximum location suction side portion of Saugseitenenkontur with a maximum length of 15% of the length of the chord, a gradient of the speed sets, the slope is maximum. As a result, 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.

The further explanation of the invention is based on the illustrated in the drawing embodiment.

FIG 1

ein erfindungsgemäßes Profil und ein aus dem Stand der Technik bekanntes Profil für eine Verdichterlaufschaufel;

FIGs 2, 3, 6

die Geschwindigkeitsverteilungen entlang der Saugseitenkontur und der Druckseitenkontur des erfindungsgemäßen Profils und des herkömmli- chen Profils aus FIG 1;

FIG 4

die Kontur von Saugseite und Druckseite des erfindungsgemäßen Profils für eine Verdichter- laufschaufel;

FIG 5

den Krümmungsverlauf des erfindungsgemäßen Profils entlang der Saugseite und Druckseite;

FIG 7

die Massenstromdichte des Massenstrom in einem Radialspalt bei Verwendung eines erfindungsge- mäßen Profils für eine freistehende Schaufel- blattspitze;

FIG 8

die Topologie der Spaltwirbeltrajektoren für das erfindungsgemäße Profil und das konventio- nelle Profil und

FIG 9, 10

perspektivische Darstellungen auf die freiste- hende Schaufelblattspitze einer erfindungsge- mäßen Verdichterlaufschaufel.

In detail show:

FIG. 1

an inventive profile and a known from the prior art profile for a compressor blade;

FIGS. 2, 3, 6

the velocity distributions along the suction side contour and the pressure side contour of the profile according to the invention and of the conventional profile FIG. 1 ;

FIG. 4

the contour of the suction side and pressure side of the profile according to the invention for a compressor blade;

FIG. 5

the curvature of the profile according to the invention along the suction side and pressure side;

FIG. 7

the mass flow density of the mass flow in a radial gap when using a profile according to the invention for a free-standing blade tip;

FIG. 8

the topology of the Spaltwirbeltrajektoren for the profile according to the invention and the conventional profile and

9, 10

perspective views of the freestanding blade tip of a compressor blade according to the invention.

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.

In FIG. 9 the perspective is chosen so that the view of the trailing edge 20 of the airfoil 12 falls, in 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. Depending on 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.

Mounted in the rotor of an axial compressor is the orientation of the airfoil 12 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.

As is known, 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. In principle, 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. For each existing along the span blade height exists such a profile. In this respect, 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. In FIG. 9 are printed page contours 40 of three profiles 28, 30 shown in full line. In FIG. 10 a plurality of suction side contours 42 of profiles 28, 30 different blade heights are also shown in solid lines.

This in 9 and 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.

In FIG. 1 two fundamentally different profiles 28, 30 are shown. The first profile 28 shown in dotted line style shows a cross section through the compressor blade 10 according to FIG 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.

In addition to the skeleton line 32, 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. In 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.

Of course, there is also a profile chord for the profile 28 known from the prior art. However, this chord is in for the sake of clarity FIG. 1 not shown.

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. Due to the different signs of the curvatures of the curved sections A, B, C, suction side contour 42 and pressure side contour 40 also have corresponding curved sections: 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 Saugseitenenkonturabschnitt D and convex pressure side contour section E leads to a locally kinking profiling, which is referred to here as a profile kink.

It is provided that 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. Furthermore goes out 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.

Furthermore, it is provided that 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.

Furthermore, goes from the FIG. 1 that 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. In this front portion H of the skeleton line 32, the skeleton line 32 is almost not arched, ie approximately straight. Likewise, 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. In general, in the front section H of blade tip-side profiles 30, a symmetrical course of the suction side contour 42 and pressure side contour 40 is symmetrically desirable.

In 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 speeds are given in Mach numbers, where Mach = 1 means the speed of sound for a given temperature. 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. In 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. The greater the distance between the profile of the suction side velocity distribution 44, 48 and the pressure side velocity distribution 46, 50 for each position of the normalized chord 34, the greater Pressure difference and thus the load at the respective considered location of the chord of the respective considered profile 28, 30th 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.

Due to the self-adjusting speed distributions 44, 46 occurs a higher load in the rear portion G of the blade tip-side profile 30, since the area between suction side velocity distribution 44 and pressure side velocity distribution 46 for a rear profile section from 60% of the chord 34 to 100% of the chord 34 is greater than the corresponding area between the corresponding velocity distributions 48, 50 of the known from the prior art conventional profile 28. Since the conventional Profile 28 is provided for non blade tip portions of the compressor blade 10, thus, along the blade blade height, a change of load from the front portion ("front-loaded design") to the rear portion of the airfoil ("aft-loaded design") occurs. It is characteristic that 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. In particular, 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.

Figures 3 to 8 give a further overview of the effects caused by the profile bend. In 3 and FIG. 6 Again, the Mach number distributions of the conventional profile 28 and the blade tip side profile 30 are represented by the relative chord length. 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 Druckseitenknicks 56 a sharp increase in Machzahldifferenz and thus the pressure potential between Saugseitenenkontur 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.

By relieving the blade tip 22 in the leading edge region, 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. Overall, 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. In order to achieve a high work conversion despite the relief of the profile 30 in the front half of the chord 34, the load is increased by a higher curvature of the profile 30 in the rear 40% of the chord 34.

Particularly preferred is the embodiment in which the interaction of the displacement of the load from front to back with the particular curvature distribution of the new profile 30 accounts for about 20% of the chord 34.

Overall, 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.

Claims (17)

A compressor blade (10) for an axial compressor, comprising a curved airfoil (12) comprising a pressure sidewall (14) and a suction sidewall (16) extending from a common leading edge (18) to a common trailing edge (20) on the other hand, to form a span from a fastener side airfoil end to a blade tip (22),
wherein, for each blade height present along the span, the airfoil (12) A profile (28, 30) having a suction side contour (42) and a pressure side contour (40), An at least partially curved skeleton line (32) and • a straight chord (34) which contours (40, 42), skeleton line (32) and chord (34) each extend from a leading edge point (24) to a trailing edge point (26),
characterized in that
at least some of the skeleton lines (32) of the blade tip-side profiles (30) have at least two inflection points (36, 38). A compressor blade (10) according to claim 1,
in which the first of the two turning points (36) when projecting perpendicularly onto the chord (34) on this a first projection point (AP), which is from the leading edge point (24) between 10% and 30% of the length of the chord (34) removed and
in which the second of the two inflection points (38) when projecting perpendicularly onto the chord (34) on this a second projection point (BP), which is from the leading edge point (24) between 30% and 50% of the length of the chord (34) removed , A compressor blade (10) according to claim 1 or 2, wherein the skeleton lines (32) include a front portion (H) extending from the leading edge point (24) to a section end point (HE) whose projection point (HP) is projected perpendicularly the chord (34) is removed from the leading edge point (24) by between 2% and 10% of the length of the chord (34),
wherein at least some of the front portions (H) of the blade tip side profiles (30) have a radius of curvature greater than 100 times the chord (34). A compressor blade (10) according to claim 3,
wherein each front section (H) has an angle of incidence with respect to an incoming gas flow, wherein at least some of the angles of attack of the blade tip side profiles (30) are smaller than the angles of incidence of the remaining profiles (28) of the airfoil (12). A compressor blade (10) according to claim 4,
in which the angle of attack of the front portion (H) blade tip side profile (30) is less than 5 °. A compressor blade (10) according to any one of claims 3 to 5,
the suction side contour (42) and pressure side contour (40) of blade tip-side profiles (30) in the front portion (H) of the skeleton line (32) are formed symmetrically. A compressor blade (10) according to any one of claims 1 to 6,
wherein at least some of the leading edge points (24) of the blade tip side profiles (30) are located further upstream than the leading edge points (24) of the remaining profiles (28) of the airfoil (12). A compressor blade (10) according to any one of claims 1 to 7,
in which exclusively the skeleton lines (32) of the profiles (30) present in the area of the blade tip (22) have two turning points (36, 38). A compressor blade (10) according to any one of claims 1 to 8,
wherein the skeleton lines (32) include a rear portion (G) extending from a section start point (GA) and to the trailing edge point (26),
wherein the rear portion (G) of at least some of the blade tip side skeleton lines (32) has a greater curvature than the rear portions of skeleton lines (32) of the remaining profiles of the airfoil (12). A compressor blade (10) according to claim 9,
in which the section start point (GA) in the case of perpendicular projection onto the chord (34) defines a projection point (GP) arranged on the chord (34) which is at most 60% of the length of the chord (34) removed from the leading edge point (24). A compressor blade (10) according to any one of claims 1 to 10,
in which the suction side contour (42) and the pressure side contour (40) of blade tip-side profiles (30) each have at least two inflection points. A compressor blade (10) according to any one of claims 1 to 11,
in which the blade tip (22) is free-standing. A compressor blade (10) according to any one of claims 1 to 12,
in which at least some of the blade tip-side profiles (30) are designed in "Aft-Loaded Design" and the other profiles (28) in the "front-loaded design". A compressor blade (10) according to any one of claims 1 to 13,
wherein the airfoil tip side comprises a region (43) of at most 20% of the span from the airfoil tip (22). A compressor blade (10) according to any one of claims 1 to 14,
wherein a velocity distribution (44) of the gas is established along the suction side contour (42) from the leading edge point (24) to the trailing edge point (26) when flowing around with a gas,
wherein at least some of the blade tip-side profiles (30) are selected so that at a maximum location a maximum velocity occurs whose projection point at a vertical projection on the chord (34) from the leading edge point (24) between 15% and 40% of the length of the chord (34) is removed. A compressor blade (10) according to claim 15,
in which the relevant profiles (30) are selected such that in a subsequent to the maximum location suction side portion of the Saugseitenenkontur (42) with a maximum length of 15% of the length of the chord (34) adjusts a gradient of speed, the maximum gradient is. Axial compressor with a rotor,
on the outer circumference of which at least one blade ring with compressor blades (10) according to one of claims 1 to 16 is formed.

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