Gas turbine compressor

Abstract

A gas turbine compressor has a flow duct wall disposed radially opposite to an airfoil tip and has a circumferential groove having an upstream groove edge and a downstream groove edge, the circumferential groove having a web having a radial cutback. In at least one meridional section through an airfoil-tip-side end face of the web, an axial distance between an upstream beginning of the cutback and the upstream leading edge of the airfoil tip is at least 1% and/or no more than 40% of a chord length and/or an axial distance between the upstream leading edge of the airfoil tip and the downstream groove edge is at least 5% and/or no more than 40% of the chord and/or an axial distance between the upstream leading edge of the airfoil tip and a kink in an airfoil-tip-side upper edge of the web in the cutback is no more than 10% of the chord length and/or a radial distance between the airfoil tip and an airfoil-tip-side upper edge of the web in the cutback is at least 50% and/or no more than 1500% of a radial distance between the airfoil tip and the downstream groove edge radially opposite thereto.

Claims

1. A gas turbine compressor comprising: at least one airfoil tip having an upstream leading edge and a downstream trailing edge; and a flow duct wall disposed radially opposite to the airfoil tip and having a circumferential groove having an upstream groove edge and a downstream groove edge, the circumferential groove having disposed therein at least one web having a radial cutback, wherein, in at least one meridional section through an airfoil-tip-side end face of the web, an axial distance (L.sub.KOZ) between an upstream beginning of the cutback of the web and the upstream leading edge of the airfoil tip is at least 1% and no more than 40% of a chord length between the upstream leading edge and the downstream trailing edge of the airfoil tip or an axial distance (Δ.sub.45) between the upstream leading edge of the airfoil tip and a kink in an airfoil-tip-side upper edge of the web in the cutback is no more than 10% of the chord length between the upstream leading edge and the downstream trailing edge of the airfoil tip.

2. The gas turbine compressor as recited in claim 1 wherein the upstream beginning of the cutback is located axially downstream of the upstream groove edge between the upstream groove edge and the upstream leading edge of the airfoil tip or a downstream end of the cutback is located in an airfoil-tip-proximal half of a radial height of the circumferential groove.

3. The gas turbine compressor as recited in claim 1 wherein in the at least one meridional section, the airfoil-tip-side upper edge of the web has a continuous curvature at the upstream groove edge; or wherein an airfoil-side end face of the web merges axially with the upstream groove edge or merges into a downstream groove flank with a curvature in or opposite to a direction of rotation of the airfoil tip.

4. The gas turbine compressor as recited in claim 3 wherein the airfoil-tip-side upper edge of the web has the continuous curvature at the upstream groove edge up to a beginning of the cutback.

5. The gas turbine compressor as recited in claim 1 wherein the web merges into an upstream or a downstream groove flank of the circumferential groove or, in the at least one meridional section, has a cross-sectional area which is at least 70% of a cross-sectional area of the circumferential groove.

6. The gas turbine compressor as recited in claim 5 wherein the cross-sectional area of the web is at least 75% of the cross-sectional area of the circumferential groove.

7. The gas turbine compressor as recited in claim 1 wherein the circumferential groove extends through the full circumference of the flow duct wall or, in the at least one meridional section, forms an angle (α) of between 60° and 90° with the flow duct wall at the upstream groove edge.

8. The gas turbine compressor as recited in claim 1 wherein an axial distance between the upstream groove edge and the leading edge of the airfoil tip disposed downstream thereof is greater than an axial distance between the downstream groove edge and the leading edge of the airfoil tip disposed upstream thereof; or wherein an axial distance between upstream and downstream groove edges is at least 25% of an axial distance between the upstream leading edge and the downstream trailing edge of the airfoil tip.

9. The gas turbine compressor as recited in claim 1 wherein in at least one section perpendicular to an axis of rotation of the compressor, the web is inclined toward a groove base of the circumferential groove in the direction of rotation of the airfoil tip; or wherein at least three identical or different webs are arranged in the circumferential groove and spaced equidistantly or at varying intervals apart in the circumferential direction.

10. The gas turbine compressor as recited in claim 9 wherein the web is inclined toward the groove base of the circumferential groove in the direction of rotation of the airfoil tip by at least 25° or no more than 65° to a radial direction.

11. The gas turbine compressor as recited in claim 1 wherein the airfoil tip is a radially outer tip of a rotor blade, and the flow duct wall is located radially outwardly thereof and opposite thereto or a radially inner tip of a stator vane, and the flow duct wall is located radially inwardly of the stator vane and opposite to the stator vane.

12. The gas turbine compressor as recited in claim 1 wherein an upstream groove flank or a downstream groove flank of the circumferential groove has an axial undercut whose cross-sectional area in the at least one meridional section is less than 10% of a cross-sectional area of the circumferential groove between upstream and downstream groove edges.

13. The gas turbine compressor as recited in claim 1 wherein, in the at least one meridional section through the airfoil-tip-side end face of the web, the axial distance (L.sub.KOZ) between the upstream beginning of the cutback and the upstream leading edge of the airfoil tip is at least 1% and no more than 40% of the chord length between the upstream leading edge and the downstream trailing edge of the airfoil tip.

14. The gas turbine compressor as recited in claim 1 wherein, in the at least one meridional section through the airfoil-tip-side end face of the web, the axial distance between the upstream leading edge of the airfoil tip and the downstream groove edge is at least 5% and no more than 40% of the chord length between the upstream leading edge and the downstream trailing edge of the airfoil tip.

15. The gas turbine compressor as recited in claim 1 wherein, in the at least one meridional section through the airfoil-tip-side end face of the web, the axial distance (Δ.sub.45) between the upstream leading edge of the airfoil tip and the kink in the airfoil-tip-side upper edge of the web in the cutback is no more than 10% of the chord length between the upstream leading edge and the downstream trailing edge of the airfoil tip.

16. The gas turbine compressor as recited in claim 1 wherein, in the at least one meridional section through the airfoil-tip-side end face of the web, a radial distance (H.sub.KOZ) between the airfoil tip and the airfoil-tip-side upper edge of the web in the cutback is at least 50% and no more than 1500% of a radial distance (H.sub.GAP) between the airfoil tip and the downstream groove edge radially opposite thereto.

17. An aircraft engine comprising the gas turbine compressor as recited in claim 1.

18. A method for designing the gas turbine compressor as recited in claim 1, the method comprising: selecting, in the at least one meridional section, the axial distance (L.sub.KOZ) between the upstream beginning of the cutback and the upstream leading edge of the airfoil tip to be at least 1% and no more than 40% of thechord length between the upstream leading edge and the downstream trailing edge of the airfoil tip; or selecting the axial distance (Δ.sub.45) between the upstream leading edge of the airfoil tip and the kink in the airfoil-tip-side upper edge of the web in the cutback to be no more than 10% of the chord length between the upstream leading edge and the downstream trailing edge of the airfoil tip.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) Further advantageous refinements of the present invention will become apparent from the dependent claims and the following description of preferred embodiments.

(2) To this end, the only drawing, FIG. 1, shows, in partially schematic form, a portion of a gas turbine compressor in accordance with one embodiment of the present invention in a meridional section.

DETAILED DESCRIPTION

(3) FIG. 1 is a meridional cross section of a portion of a gas turbine compressor in accordance with one embodiment of the present invention; i.e., of a gas turbine compressor designed in accordance with one embodiment of the present invention. The meridional cross section contains the axis of rotation of the compressor (horizontal in FIG. 1). The vertical direction in FIG. 1 is a radial direction.

(4) The gas turbine compressor includes rotor blades arranged adjacent one another in the circumferential direction (perpendicular to the plane of the drawing of FIG. 1) and having shroudless tips, and a flow duct wall 20 disposed outwardly thereof and opposite thereto and fixed relative to the casing, one rotor blade tip 10 being partially shown in the meridional section of FIG. 1.

(5) The flow duct wall has a circumferential groove formed therein, the circumferential groove having an upstream flank 31 which merges into the flow duct wall at an upstream groove edge 21, a downstream flank 32 which merges into the flow duct wall at a downstream groove edge 22, and a groove base 33 connecting these groove flanks.

(6) The upstream groove flank has an axial undercut whose cross-sectional area in the meridional section is less than 10% of a cross-sectional area of the circumferential groove between its upstream and downstream groove edges. This cross-sectional area of the circumferential groove between its upstream and downstream groove edges is the area which, in the meridional section of FIG. 1, is defined by the groove base, a straight connecting line 24 between the upstream and downstream groove edges, and perpendicular lines through the upstream and downstream groove edges, which are indicated by dot-dash lines in FIG. 1. Accordingly, the cross-sectional area of the undercut is the area between upstream groove flank 31 and the dot-dash line perpendicular to connecting line 24 on the left in FIG. 1.

(7) A plurality of webs are arranged in the circumferential groove and spaced apart in the circumferential direction (perpendicular to the plane of the drawing of FIG. 1), of which one web 40 is shown in cross-section in the meridional section of FIG. 1.

(8) As explained earlier herein, reference numeral 24 in FIG. 1 denotes a straight connecting line 24 between upstream and downstream groove edges 21, 22. Thus, this line represents a reference curve which extends from the upstream groove edge to the downstream groove edge and whose curvature is infinite.

(9) Reference numeral 23 in FIG. 1 denotes another reference curve which also extends from the upstream groove edge to the downstream groove edge, but whose curvature at the upstream and downstream groove edges is in each case equal to the curvature of the flow duct wall and axially continuously linear therebetween; i.e., linearly interpolates the curvature of flow duct wall 20 between groove edges 21, 22. Thus, this reference curve 23 represents a virtual extension of flow duct contour 20 across the circumferential groove.

(10) In the meridional section of FIG. 1 through an airfoil-tip-side end face or upper edge 43 of web 40, reference curves 23, 24 each represent a corresponding circumferentially extending reference plane 23, 24.

(11) As can be seen in the meridional section of FIG. 1, the airfoil-tip-side end face or upper edge 43 deviates from reference curve or plane 23; i.e., from the virtual extension of the flow duct contour, in a direction away from the airfoil tip and toward the groove base (upward in FIG. 1), starting at a point or circumferential line 41 up to another point or circumferential line 42.

(12) Furthermore, starting at the point or circumferential line 41, the airfoil-side end face or upper edge 43 deviates from reference plane or curve 24 toward the groove base by at least 1% of a maximum radial distance between groove base 33 and groove edge 22 (i.e., the one closer to the airfoil tip).

(13) Thus, the point or circumferential line 41 defines an upstream beginning of a radial cutback 44 of the web.

(14) The airfoil-side end face or upper edge of the web continues flow duct contour 20 to this beginning 41 of cutback 44 with a continuous curvature.

(15) The point or circumferential line 42 defines a downstream end of radial cutback 44, where the airfoil-side end face or upper edge 43 of the web merges into downstream groove flank 32.

(16) In a modification (not shown), the airfoil-side end face or upper edge 43 of the web merges back into reference plane or curve 23. In this case, the point or circumferential line where the airfoil-side end face or upper edge 43 of the web merges back into reference plane or curve 23, or the point or circumferential line beyond which the airfoil-side end face or upper edge of the web once again deviates from the straight reference plane or curve 24 toward groove base 33 by less than 1% of the maximum radial distance between groove base 33 and groove edge 22 (i.e., the one closer to the airfoil tip) represents the downstream end of the radial cutback.

(17) In this modification (not shown), the airfoil-side end face or upper edge of the web may continue the flow duct contour with a continuous curvature from downstream groove edge 22 in an upstream direction (toward the left in FIG. 1) to this end of the cutback, as described and illustrated analogously for the region between upstream groove edge 21 and upstream beginning 41 of the cutback.

(18) Thus, the empty space or the free area between the airfoil-side end face or upper edge 43 of the web and reference plane or curve 23 defines radial cutback 44 with its upstream beginning 41 and its downstream end 42.

(19) As can be seen in the meridional section of FIG. 1, this upstream beginning 41 of cutback 44 is located axially downstream (to the right in FIG. 1) of upstream groove edge 21 between this groove edge 21 and upstream leading edge 11 of airfoil tip 10, and downstream end 42 of cutback 44 is located in an airfoil-tip-proximal half 34 of a radial height 35 of the circumferential groove.

(20) The radial height may be defined as the maximum distance between groove base 33 and groove edge 22 (i.e., the one closer to the airfoil tip) in the radial direction (vertical in FIG. 1) or, as indicated in FIG. 1, the maximum distance 35 between groove base 33 and groove edge 22 (i.e., the one closer to the airfoil tip) in a direction perpendicular to the straight connecting line 24 between the upstream and downstream groove edges.

(21) In the embodiment shown, the radial cutback ends in the radially upper half 34 of downstream groove flank 32, and the web is continuously cut back radially, starting at beginning 41. The term “radially upper half” is used to refer to the portion or region of downstream groove flank 32 that extends in the radial direction or a direction perpendicular to connecting line 24 between the upstream and downstream groove edges over 50% of the maximum distance of downstream groove edge 22 from groove base 33 in this direction.

(22) In the embodiment of FIG. 1, web 40 merges into the upstream and downstream flanks 31, 32 of the circumferential groove, and thus extends axially through the groove.

(23) As explained earlier herein, the airfoil-tip-side end face or upper edge of the web has, at upstream groove edge 21, the same curvature as flow duct contour 20 and smoothly continues this curvature to beginning 41 of cutback 44.

(24) In the embodiment of FIG. 1, web 40 has a cross-sectional area (shown hatched in FIG. 1) which is at least 75% of the cross-sectional area of the circumferential groove in this meridional section, which is defined by groove flanks 31, 32, groove base 33, and connecting line 24 between the two groove edges 21, 22.

(25) In the embodiment of FIG. 1, the circumferential groove forms an angle α of between 60° and 90° with flow duct wall 20 at upstream groove edge 21.

(26) In the embodiment of FIG. 1, an axial distance between upstream groove edge 21 and leading edge 11 of airfoil tip 10 disposed downstream thereof (to the right in FIG. 1) is greater than an axial distance between downstream groove edge 22 and leading edge 11 disposed upstream thereof.

(27) In the embodiment of FIG. 1, an axial distance between upstream and downstream groove edges 21, 22 is at least 25% of an axial distance between upstream leading edge 11 and a downstream trailing edge 12 of airfoil tip 10.

(28) S.sub.AX schematically indicates an axial chord length of airfoil tip 10. This axial chord length may be equal to the axial distance between leading and trailing edges 11, 12 of airfoil tip 10 or also to the length of the chord line or camber line thereof.

(29) An axial distance L.sub.KOZ between upstream beginning 41 of cutback 44 and upstream leading edge 11 of the airfoil tip is between 1% and 40%, preferably between 2% and 15%, of the so-defined chord length S.

(30) An axial distance L.sub.OL between upstream leading edge 11 of the airfoil tip and downstream groove edge 22 is between 5% and 40%, preferably between 10% and 30%, of chord length S.sub.AX.

(31) An axial distance Δ.sub.45 between upstream leading edge 11 of the airfoil tip and a kink 45 in the airfoil-tip-side end face or upper edge 43 of the web in the cutback is no more than 10%, preferably no more than 5%, of chord length S.sub.AX.

(32) A radial distance between airfoil tip 10 and the airfoil-tip-side end face or upper edge 43 of the web in cutback 44 is between 50% and 1500%, preferably between 100% and 1000%, of a radial distance H.sub.GAP between airfoil tip 10 and the downstream groove edge 22 radially opposite thereto. The minimum radial distance H.sub.KOZ between airfoil tip 10 and the airfoil-tip-side end face or upper edge 43 is exemplarily indicated in FIG. 1. Likewise, a maximum distance or mean distance at leading edge 11 may also be taken as a basis.

(33) Although the above is a description of exemplary embodiments, it should be noted that many modifications are possible. It should also be appreciated that the exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration in any way. Rather, the foregoing detailed description provides those skilled in the art with a convenient road map for implementing at least one exemplary embodiment, it being understood that various changes may be made in the function and arrangement of elements described without departing from the scope of protection as set forth in the appended claims or derived from combinations of features equivalent thereto.

LIST OF REFERENCE NUMERALS

(34) 10 airfoil tip 11 leading edge 12 trailing edge 20 flow duct contour 21 upstream groove edge 22 downstream groove edge 23 reference plane/curve 24 straight reference plane/curve 31 upstream groove flank 32 downstream groove flank 33 groove base 34 airfoil-tip-proximal half of the circumferential groove 35 radial height of the circumferential groove 40 web 41 upstream beginning of the cutback 42 downstream end of the cutback 43 airfoil-tip-side end face/upper edge 44 cutback 45 kink α angle H.sub.KOZ radial distance between airfoil tip and airfoil-tip-side end face/upper edge H.sub.GAP radial distance between airfoil tip and downstream groove edge L.sub.KOZ axial distance between beginning of cutback and airfoil tip leading edge L.sub.OL axial distance between airfoil tip leading edge and downstream groove edge S.sub.AX axial chord length Δ.sub.45 axial distance between kink and airfoil tip leading edge