WO2018219611A1

WO2018219611A1 - Compressor stator vane for axial compressors having a corrugated tip contour

Info

Publication number: WO2018219611A1
Application number: PCT/EP2018/061959
Authority: WO
Inventors: Christian Cornelius; Stephan Klumpp; David Monk
Original assignee: Siemens Aktiengesellschaft
Priority date: 2017-06-01
Filing date: 2018-05-09
Publication date: 2018-12-06

Abstract

In summary, the invention relates to a compressor stator vane (14) for axial compressors (10), comprising a cantilevered airfoil (22) attached to a compressor casing (18) via a blade root, the airfoil (22) comprising an airfoil tip (28) extending from a leading edge (34) to a trailing edge (35). To improve aerodynamic efficiency, a corrugated airfoil tip (28) is suggested, exposing a wave-like shape having at least one wave crest (38). To further reduce parasitic secondary leakage flows, the airfoil tip (28) comprises a first trough (42) and a second wave crest (40). The amplitude of the first crest (38) is larger than the amplitude of the second crest (40). Furthermore, the facing rotor hub endwall (30) may be formed correspondingly. Thus, tip leakage swirl length and position is shifted upstream substantially thereby reducing overall tip losses and minimising blockage of the flow passage which both improves compressor efficiency.

Description

COMPRESSOR STATOR VANE FOR AXIAL COMPRESSORS HAVING A CORRUGATED TIP CONTOUR

The invention relates to a compressor stator vane for axial compressor, comprising a blade root and a cantilevered air^¬ foil attached to said blade root, the airfoil comprising a suction side and a pressure side extending in a spanwise di- rection from said blade root to an airfoil tip and in chord direction from an upstream-sided leading edge to a downstream-sided trailing edge. Further the invention relates to an axial compressor with at least one ring of said compressor stator vanes.

Modern turbo engines and especially their bladings achieve already a very high aerodynamical efficiency. With the ten^¬ dency to continuously increased profile loadings an increas^¬ ing part of overall losses originates in the area close to the endwalls of the flow path due to tip leakage flows. A re^¬ duction of these losses would enable a significant improve^¬ ment of the efficiency of the turbo engine, especially of ax^¬ ial compressors. Up to now the conical flow path geometry of compressors do often not address the problem of tip leakage flow over the rotor blade tips. In detail rather constructive and economi^¬ cal targets are of interest. However, especially airfoils with a rather short span height and relative large radial tip gaps have potential for reducing losses and therefore in^¬ creasing efficiency.

To address this problem it is known that rotor blades with corrugated tips with a constant radial tip gap design to the opposing endwall for reducing the tip could reduce tip losses of said rotor blade.

Therefore the object of the invention is the provision of a compressor stator vane and an axial compressor with increased efficiency .

The object of the invention is achieved by the independent claims. The dependent claims describe advantageous develop- ments and modifications of the invention.

The invention is based on the recognition, that also at can- tilevered compressor stator vanes a significant tip loss oc^¬ curs. To reduce said tip loss a very specific compressor sta- tor vane tip design and a corresponding endwall design which faces the tips of a ring of appropriate compressor stator vanes is proposed with this invention. In detail the inven^¬ tion proposes a compressor stator vane for an axial compres^¬ sor, comprising a blade root and a cantilevered airfoil at- tached to said blade root, the airfoil comprising a suction side and a pressure side extending in a spanwise direction from said blade root to an airfoil tip and in chord direction from an upstream-sided leading edge to an downstream-sided trailing edge, wherein the airfoil tip in chord direction is wave-like, such, that along the chord direction at least one wave crest is provided. With the proposed geometry of the compressor stator vane tip and the correspondingly shaped opposing hub endwall as part of the rotor the tip leakage flow is influenced. Said influence reduces the local aerodynamical losses, which contributes to the increase of the overall ef^¬ ficiency of the compressor.

The invention proposes also to reduce the radius of the hub endwall locally for reducing the losses. In detail the hub endwall as that axial section of the rotor being opposite lo^¬ cated of the compressor stator vane tip is also wave-like. The hub endwall is wave-like, such, that along the chord di^¬ rection at least one rotor wave trough is provided. Even it was recognized that some beneficial effects occurs also inde^¬ pendently of the stator vane tip geometry, preferably both the stator vane tip geometry and the hub endwall geometry are correspondingly shaped. The provision of said wave-like geometries mean, that the lo^¬ cal gradient of the hub endwall is changed compared to con^¬ ventional designs, which were until present substantially conical or cylindrical. Due to thermal effects and effects with regard to rotating forces during the operation of the compressor and the gas turbine equipped with such a compres^¬ sor, a local, relative displacement between hub endwall and oppositely arranged compressor stator vane tips appear. This leads to a tip gap having a variable size along the chord. These displacements and variable gap size do not significant- ly influence the positive effects of the wave-like geometries so that during operation said misalignment and non-parallel tip gap could be tolerated. Nevertheless the misalignment has to be considered in the whole engine model and is important to determine the final size of the radial gap.

In detail comprehensive investigations have shown that in service a main gap flow structure and a secondary gap flow structure are generated from the vane tip. In that case the term "swirl" has to be understood in that way that said gap refers to the vertical flow structure generated by the leak^¬ age flow over the tip. The main gap swirl is located rather upstream and is stronger in force than the secondary swirl which is located rather downstream with regard to the main flow direction in the flow path of the compressor. Both swirls weaken the main flow and generate local blockages downstream of the airfoil passages close to the rotating hub endwall. While establishing at least one first wave crest lo^¬ cated close to the leading edge of the airfoil the main gap swirl can be weakened which increases the compressor effi^¬ ciency. The new wave-like tip design with at least one wave crest causes the main gap swirl to remain in the passage be^¬ tween neighbored airfoils of the vane ring so that it does not extend into the flow area downstream of the trailing edg- es of the respective airfoils. This limits the blockage and enables an increased free flow area for the main flow.

Additionally it was recognized that the rotating hub endwall influences the swirl behavior different from the static end- wall: The first evidence is the difference in flow structures between the rotating hub endwall and the static endwall. The influence of the rotating hub endwall is known to produce both a primary and secondary gap swirl, which is not seen for rotor tip flow under a static casing. These gap swirls can be further influenced by local hub gradients with the effect to push the gap in a more circumferential direction relative to rotor blade tip swirls.

In a first preferred embodiment the weakening effect of the first wave crest can be further increased, when advantageous^¬ ly a wave trough in the tip profile is established downstream of the first wave crest. The wave trough can determined with regard to a virtual traverse axis which extends straight be^¬ tween the radial most inner point of the leading edge of the airfoil and the radial most inner point of the trailing edge of the airfoil. In a longitudinal section the wave crest (s) and the wave trough are arranged on different sides of said virtual traverse axis. According to a second preferred embodiment the airfoil tip comprises a second wave crest which is located downstream of the first wave crest. Further preferred, the amplitude of the first wave crest is larger than the amplitude of the second wave crest. Due to the appearance of the main gap swirl and the secondary gap swirl besides the main flow through the compressor passages the shape of the waved tip is adapted ac^¬ cordingly. The second wave crest fights the secondary gap swirl so also that remains in the main flow passage between two neighbored airfoils of the same row.

Accordingly to a another preferred embodiment in chord direc^¬ tion as seen from leading edge to trailing edge the rise to the first wave crest is steeper than the descent from the first wave crest to the first wave trough. The steeper rise to the first wave crest leads to an exchange in the pressure ratio and generates a relative strong aerodynamically im^¬ pulse, that influences the main gap swirl significantly over a longer section of the chord, why the descent from the first wave crest to a first wave trough is flatter and longer than the rise to the first wave crest.

In a further preferred embodiment a virtual straight chord with a chord length is determinable, which extends from the leading edge to the trailing edge of the airfoil tip, and wherein the position of the maximum amplitude of the first wave crest is not more than 15% of the chord length away from that location, at which the highest pressure difference be^¬ tween suction side and the pressure side appears. Hence the first wave crest is located at a position of the airfoil tip chord, where the largest tip leakage flow probably appeared. Therefore the preferred feature fights against said largest tip leakage flow at best. In a further preferred embodiment the maximal thickness of the airfoil profile at the airfoil tip is not more than 15% of the chord length away from the maximum amplitude of the first wave crest.

Further the virtual straight chord possesses a chord length of about 100% and the location of 0% chord coincides with the leading edge and the location of 100% chord coincides with the trailing edge. Comprehensive investigations have shown that advantageously the location of the maximum amplitude of the first wave crest is arranged in the section between 1% chord and 30% chord and/or the location of the maximum amplitude of the second wave crest is arranged in the section be^¬ tween 70% chord and 80% chord. Interestingly it was found that for said double-crest structure the first crest is placed further away from the leading edge than for a single crest structure. For a single crest structure the maximum am^¬ plitude of the first wave crest between 1% - 10% chord where^¬ as for a double crest design the first wave crest is between 10% - 30% chord.

During said investigation it was recognized that to be most efficient the maximum amplitude of the first wave crest amounts between 2% of the chord length and 6% of the chord length and/or the maximum amplitude of the second wave crest amounts between 0.5% of the chord length and 3% of the chord length, related to said virtual traverse axis.

With regard to an axial compressor it comprises at least one ring of compressor stator vanes according to the invention and a rotor, whose surface pointing radially outwardly. Said rotor surface comprises a section, which as hub endwall faces the tips of the respective compressor stator vanes under es^¬ tablishing radial gaps. In cold status (at ambient tempera- tures) the hub contour is selected such, that it is arranged rotation-symmetrically and shaped in a corresponding way as described above for the compressor stator vane. This means in example that the hub contour comprises a rotor wave crest at a location where the compressor stator vane tip comprises a wave trough and wherein the hub contour comprises a rise where the compressor stator vane tip comprises a descent.

Due to the recognition that the tip losses relatively in- creases with short span of airfoils, the application of the before mentioned features are in particular advantageously when these are implemented in the rear stages of the axial compressors with subsonic flow velocities. Embodiments of the invention are now described, by the way of example only, with reference to the accompanying drawings, of which :

Figure 1 shows a longitudinal section through the ax- compressor of a gas turbine,

Figure 2 shows a compressor stator vane according

the detail of figure 1 and its opposite ar^¬ ranged hub endwall as part of a rotor,

Figure 3 shows in a radial view onto the compressor stator vanes of an axial compressor the swirl structures in the main flow generated by cantilevered stator vanes,

Figure 4 shows the overall pressure downstream of the compressor stator vanes and Figure 5 shows schematically the vane tip geometry in comparison to a normalized virtual traverse axis . In all drawings similar or identical elements are provided with the same reference numbers. The illustration in the drawings is only in schematic form.

Figure 1 shows in a longitudinal section an axial compressor 10 according to the invention comprising two compressor stages 11. Each compressor stage 11 comprises a row of compressor rotor blades 12 and a row of compressor stator vanes 14 from each only one is shown. The compressor rotor blade 12 is attached to the rotor 16, which is rotatable about a machine axis 17. The individual compressor rotor blades 12 could be attached to compressor rotor disks in a conventional way. Al^¬ so in a conventional way the compressor stator vanes 14 are attached accordingly to a compressor casing 18. Due to the schematic structure of the drawings the details of the at- tachments and especially the roots are not shown. Both the compressor rotor blades 12 and the compressor stator vanes 14 are embodied with cantilevered airfoils 20, 22, which are lo^¬ cated within the air flow path 23 of the axial compressor 10. The airfoil 20 of the compressor rotor blade 12 comprises a tip 24 which faces the compressor casing 18 under establishing a tip gap 26. Simultaneously and in an analogues way the airfoils 22 of the compressor stator vanes 14 comprise a tip 28 which faces the rotor hub endwall 30 while establishing a tip gap 32. The tip gaps 26, 32 are embodied as radial clear- ances. Details with regard to the radial gap 32 of the com^¬ pressor stator vane 14 are shown in figure 2.

Figure 2 shows in a longitudinal sectional view only a single stator vane 14 of a row of compressor stator vanes according to figure 1. As displayed in figure 2, the airfoil tip 28 is in general not cylindrically or conically but rather wave^¬ like in chord direction. According to the main flow direction 36 of the air to be compressed from the compressor 10 the airfoil 22 comprises a leading edge 34, at which the air flow approaches the airfoil 20 at first. After flowing around the pressure side and suction side of the airfoil 20, the com^¬ pressed air leaves the airfoil 20 at the trailing edge 35. A chord of the airfoil can be determined according to the defi- nition of the book "Stationare Gasturbinen" (Author: Christof Lechner, Jorg Seume, published in Spriner-Verlag Belin, 2003, 1. edition , page 310) .

The chord extends from the leading edge 34 to the trailing edge 35 of the profiled airfoil 22. According to a first ex^¬ emplary embodiment of the invention and as seen in the direc^¬ tion from the leading edge 34 to the trailing edge 35 the tip 28 of the airfoil 20 is wave-like, such, that along the chord direction two wave crests 38 and 40 and in between one wave trough 42 are provided. In detail the first wave crest 38 is located upstream of the second wave crest 40 according to the main flow direction 36 and the amplitude of the first wave crest 38 is larger than the amplitude of the second wave crest 40. In further detail as seen from the leading edge 34 towards the trailing edge 35 the rise 44 to the first wave crest 38 is steeper and shorter than the descent 46 from the first wave crest 38 to the first wave trough 42.

Fig 3 shows a sectional view according to cut line III-III of Fig. 2. As can be seen in figure 3, due to the application of the wave-like tip of the airfoil 20, the main gap swirl 48 and the secondary gap swirl 50 remains further upstream in comparison to conventional compressors comprising a conical or cylindrical hub endwall and a corresponding airfoil tip without any wave-like structures. The swirl length and posi^¬ tion has been shifted upstream significantly: both swirls 48, 50 are located within the passage between adjacent airfoils 22 of the dedicated row. This leads to a positive effect onto the tip losses so that the overall pressure downstream of the stator airfoils is increased compared to conventional tip gaps while the areas comprising a lower pressure are reduced. Said reduction leads to a smaller blockage within the flow path (Fig. 4) . The lines 51 of equal pressure represents schematically the pressure distribution, which according to the invention have smaller areas with lowest local pressure compared to conventional compressors.

Figure 5 shows schematically a graph of a second exemplary embodiment of a vane tip geometry as full line 52. According to the secondary exemplary embodiment the tip comprises only a single crest 38 and only a single trough 42. The abscissa represents the normalized virtual traverse axis extending be^¬ tween a leading edge located at 0% chord and a trailing edge at 100% chord. The ordinate represents the radial deviation from the virtual traverse axis extending from the radially inner-most leading edge point to the radially inner-most trailing edge position. The radially inner-most points are determined when the compressor stator vane 14 is assembled in the compressor 10. The vane tip profile starts from the radi^¬ ally most inner point of the leading edge with a rise 44 to the wave crest 38. The rise 44 has a substantially constant gradient. The apex of the wave crest 38 is rather rounded than sharp and slopes downward to the trough 42. The value of the gradient of the descent 46 between the first crest 38 and the trough 42 is considerably smaller than the gradient of the rise 44. Further downstream the remaining vane tip profile extends from a bottom of the trough 42 to the trailing edge at 100% chord to approach there at zero level. In this exemplary embodiment the length of the descent 46 is more than about 50% of the cord length. With regard to the virtual traverse axis x which represents zero level or a zero-radial deviation from the virtual traverse axis constant radius, the value of the wave crest apex has another sign than the value of the wave trough apex.

The relative long descent 46 leads to a smooth pressure re^¬ lief in the main flow which weakens the main tip swirl effec- tively. Of course this style of vane tips is most beneficial when it is used at stator vanes of a compressor comprising, at ambient temperatures, opposite of the respective compres^¬ sor vane tips a correspondingly shaped hub endwall. In summary the invention relates to a compressor stator vane 14, for an axial compressor 10, comprising at blade root and a cantilevered airfoil 20 attached to said blade root, the airfoil comprising a suction side and a pressure side extend^¬ ing in a span wise direction from said blade root to an air- foil tip 28 and in chord direction from an upstream-sided leading edge 34 to a downstream-sided trailing edge 35. To provide a compressor stator vane with increased efficiency, it is proposed that the airfoil tip 28 in chord direction is wave like, such, that along the chord direction at least one wave crest 38 is provided.

Claims

Patent claims:

1. Compressor stator vane (14) for an axial compressor

(10) ,

comprising a blade root and a cantilevered airfoil (22) attached to said blade root, the airfoil comprising a suction side and a pressure side extending in a spanwise direction from said blade root to an airfoil tip (28) and in chord direction from an upstream-sided leading edge (34) to an downstream-sided trailing edge (35), characterized in that

the airfoil tip (28) in chord direction is wave-like, such, that along the chord direction at least one wave crest (38, 40) is provided.

2. Compressor stator vane (14) according to claim 1,

wherein the airfoil tip (28) in chord direction compris^¬ es a wave trough (42), which is located downstream of said at least one wave crest (35) .

3. Compressor stator vane (14) according to claim 2,

wherein the airfoil tip (28) in chord direction compris^¬ es a second wave crest (40), which is located downstream of said wave trough (42) .

4. Compressor stator vane (14) according to claim 1, 2 or 3,

wherein the first wave crest (38) is located upstream of the second wave crest (40) and the maximum amplitude of the first wave crest (38) is larger than the maximum am^¬ plitude of the second wave crest (40) . Compressor stator vane (14) according to one of the pre ceding claims,

wherein in chord direction as seen from leading edge (34) to trailing edge (35) the rise (44) to the first wave crest (38) is steeper than the descent (46) from the first wave crest (38) to the first wave trough (42)

Compressor stator vane (14) according to one of the preceding claims,

wherein a virtual straight chord with an chord length is determinable, which extends from the leading edge (34) to the trailing edge (35) of the airfoil tip (28) , wherein the position of the maximum amplitude of the first wave crest is not more than 15% of the chord length away of that location, at which the highest pres^¬ sure difference between the suction side and the pres^¬ sure side appears.

Compressor stator vane (14) according to claim 6, wherein a maximal thickness of the airfoil profile at the airfoil tip is not more than 15% of the chord length away from the maximum amplitude of the first wave crest (38) .

Compressor stator vane (14) according to one of the preceding claims,

wherein a resp. the virtual straight chord possesses a chord length of 100 % and the location of 0% chord coin^¬ cidences with the leading edge (34) and the location of 100% chord coincides with the trailing edge (35) , wherein the location of the maximum amplitude of the first wave crest (38) is arranged in the section between 1% cord and 30% chord and/or the location of the maximum amplitude of the second wave crest (40) is arranged in the section between 70% cord and 80% chord.

Compressor stator vane (14) according to one of the preceding claims,

wherein the maximum amplitude of the first wave crest (38) amounts between 2 % of the cord length and 6 % of the chord length and/or

the maximum amplitude of the second wave crest (40) amounts between 0,5 % of the chord length and 3 % of the chord length, related to a virtual traverse axis which extends between the radial most inner point of the lead^¬ ing edge of the airfoil and the radial most inner point of the trailing edge of the airfoil.

Axial compressor (10) with at least one ring of compres^¬ sor stator vanes (14) comprising a rotor (16), whose surface pointing radially outwardly comprises a section, which as hub endwall (30) faces the tips (28) of com^¬ pressor stator vanes (14) arranged within one row under establishing radial tip gaps (32),

characterized in

the hub endwall (30) is in main flow direction wave^¬ like, such, that along the chord direction at least one rotor wave trough is provided.

Axial Compressor (10) according to claim 10,

wherein the compressor stator vanes are embodied accord^¬ ingly to compressor stator vanes of the claims 1 to 10, and wherein

that in cold status of the compressor (10) the hub end- wall (30) is selected that such it is arranged rotation- symmetrically and correspondingly to the wave form of the tips (28) of the airfoils (22)

Axial compressor (10) according to claim 10 or 11, wherein during operation the airfoil tips (28) and the hub endwall (30) are misaligned with regard to their contoured profiles.