Turbomachine and turbomachine stage

Abstract

A turbomachine stage includes guide vanes and an airfoil platform forming a guide vane cascade, and rotor blades and an airfoil platform forming a rotor blade cascade. Airfoil platforms have cascade regions extending between circumferentially adjacent airfoils, and gap regions which radially and/or axially bound an axial gap extending axially between the guide vane cascade and the rotor blade cascade. A contour of at least one of these gap regions varies in the radial and/or axial direction around the circumference. A maximum extent of this contour in the radial direction toward the spoke-like pattern is circumferentially spaced from an airfoil edge of this cascade by no more than 50% of the cascade pitch, a maximum variation in the radial direction being no more than 50% of the cascade pitch and/or a maximum extent in the axial direction away from the spoke-like pattern is circumferentially spaced from an airfoil edge of this cascade by no more than 50% of the cascade pitch, a maximum variation in the axial direction being no more than 50% of the cascade pitch.

Claims

1. A turbomachine stage comprising: guide vanes and at least one of a radially inner and radially outer guide vane airfoil platform defining a guide vane cascade, the guide vane airfoil platform having a guide vane platform cascade region and a guide vane gap region; and rotor blades and at least one of a radially inner and radially outer rotor blade airfoil platform defining a rotor blade cascade, the rotor blade airfoil platform having a rotor blade platform cascade region and a rotor blade gap region, the rotor blade cascade being adjacent to the guide vane cascade, the guide vane gap region and the rotor blade gap region at least one of radially and axially bounding an axial gap extending axially between the guide vane cascade and the rotor blade cascade, a contour of at least one of the guide vane and rotor blade gap regions varying in at least one of the radial and axial direction around a circumference, at least one of (a) and (b) being true: (a) a maximum extent of the contour in the radial direction toward a spoke-like pattern of the respective guide vane or rotor blade cascade being circumferentially spaced from a respective airfoil edge of the respective guide vane or rotor blade cascade by no more than 50% of the pitch of the respective guide vane or rotor blade cascade, a maximum variation of the contour in the radial direction being no more than 50% of the pitch of the respective guide vane or rotor blade cascade; and (b) a maximum extent of the contour in the axial direction away from the spoke-like pattern is circumferentially spaced from the respective airfoil edge of the respective guide or rotor blade cascade by no more than 50% of the pitch of the respective guide vane or rotor vane cascade, a maximum variation of the contour in the axial direction being no more than 50% of the pitch of the respective guide vane or rotor blade cascade; and whereby a radially opposite contour of the other of the guide vane and the rotor blade gap regions varies identically around the circumference as the contour but with a phase shift, wherein: the contour varies according to R()=R.sub.0+Rsin(.sub.R) and the radially opposite contour varies with the phase shift PS(x) according to R(, x)=R.sub.0+Rsin(.sub.R+PS(x)), wherein the phase shift PS(x) varies with the axial position x, the contour varies according to X()=X.sub.0+Xsin(.sub.x) and the radially opposite contour varies with the phase shift PS(r) according to X(, r)=X.sub.0+Xsin(.sub.x+PS(r)), wherein the phase shift PS(r) varies with the radial position r, wherein [0, 360], R.sub.0, R, X.sub.0, X, .sub.R, .sub.x, .sub.R, .sub.x are constant, R defines a variation in a radial direction, and X defines a variation in an axial direction.

2. The turbomachine stage as recited in claim 1 wherein the contour varies in the radial direction, the maximum extent of the contour in the radial direction toward the spoke-like pattern of the respective guide vane or rotor blade cascade being circumferentially spaced from a respective airfoil edge of the respective guide vane or rotor blade cascade by no more than 25% of the pitch of the respective guide vane or rotor blade cascade.

3. The turbomachine stage as recited in claim 2 wherein the maximum variation in the radial direction is no more than 40% of the pitch of the respective guide vane or rotor blade cascade.

4. The turbomachine stage as recited in claim 1 wherein the contour varies in the radial direction and with the maximum variation in the radial direction being no more than 40% of the pitch of the respective guide vane or rotor blade cascade.

5. The turbomachine stage as recited in claim 1 wherein the contour varies in the axial direction, the maximum extent of the contour in the axial direction away from the spoke-like pattern being circumferentially spaced from the respective airfoil edge of the respective guide or rotor blade cascade by no more than 25%.

6. The turbomachine stage as recited in claim 5 wherein the maximum variation in the axial direction is no more than 40% of the pitch of the respective guide vane or rotor blade cascade.

7. The turbomachine stage as recited in claim 1 wherein the contour varies in the axial direction and with the maximum variation in the axial direction being no more than 40% of the pitch of the respective guide vane or rotor blade cascade.

8. The turbomachine stage as recited in claim 1 wherein the respective maximum extent is disposed in the pressure-side region of an airfoil leading edge or in the suction-side region of an airfoil trailing edge.

9. The turbomachine stage as recited in claim 1 wherein the contour varies radially and axially.

10. The turbomachine stage as recited in claim 1 wherein the contour varies radially but is constant in the axial direction.

11. The turbomachine stage as recited in claim 1 wherein the respective guide vane platform or rotor blade platform gap region whose contour merges smoothly into the respective guide vane or rotor blade cascade region.

12. The turbomachine stage as recited in claim 1 wherein the contour varies periodically.

13. A turbomachine comprising at least one turbomachine stage as recited in claim 1.

14. A gas turbine comprising at least one turbomachine stage as recited in claim 1.

15. An aircraft engine gas turbine comprising at least one turbomachine stage as recited in claim 1.

16. A compressor stage comprising the turbomachine stage as recited in claim 1.

17. A turbine stage comprising the turbomachine stage as recited in claim 1.

18. The turbomachine stage as recited in claim 1 wherein the contour varies axially.

19. The turbomachine stage as recited in claim 1 wherein the phase shift PS(r) varies linearly with the radial position r according to PS(x)=.sub.xr.

20. The turbomachine stage as recited in claim 1 wherein the phase shift PS(x) varies with the axial position x according to PS(x)=.sub.Rx.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) Further features and advantages will become apparent from the dependent claims and the exemplary embodiments. To this end, the drawings show, partly in schematic form, in:

(2) FIG. 1: a developed view of a portion of a gas turbine stage according to the present invention including a guide vane cascade and a rotor blade cascade having radially inner airfoil platforms whose gap region contour varies in the axial direction around the circumference;

(3) FIG. 2: an axial elevation view of a cascade of guide vanes or rotor blades of a gas turbine stage according to the present invention including radially inner airfoil platforms whose gap region contour varies in the radial direction around the circumference; and

(4) FIG. 3: a meridional section through a gas turbine stage according to the present invention having an annular-flange-like platform extension.

DETAILED DESCRIPTION

(5) FIG. 1 shows a developed view of a portion of a gas turbine stage according to the present invention, including a stationary cascade of guide vanes 1 and, opposite thereto, a rotating cascade of rotor blades 2. The rotation is indicated by a filled vertical arrow, the flow of working fluid is indicated by an empty arrow in the region of the guide vane cascade. This configuration is merely exemplary for purposes of illustration. The present invention may be used equally in turbine and compressor stages, where the guide vane cascade is disposed upstream and/or downstream of the rotor blade cascade.

(6) Integrally formed with the airfoils 1, 2 are radially inner airfoil platforms, which are shown from above in FIG. 1. Each airfoil may either have a separate airfoil platform, or several or all of the airfoils of a cascade may be connected to, in particular integrally formed with, the same airfoil platform which, in accordance with the present invention, may then be imagined as being divided into separate airfoil platforms associated with the individual airfoils. Therefore, FIG. 1 does not show any airfoil platform boundaries in the circumferential direction (vertically in FIG. 1).

(7) A cascade region 10.1 of the guide vane platforms and a cascade region 20.1 of the rotor blade platforms extend axially between the respective leading edge (left in FIG. 1) and the respective trailing edge (right in FIG. 1), said cascade regions being hatched from top left to bottom right in FIG. 1.

(8) The cascade regions merge axially into respective gap regions 10.2T and 20.2L beyond the respective airfoil leading or trailing edges, said gap regions being hatched from bottom left to top right in FIG. 1. Gap regions 10.2T and 20.2L each have substantially the shape of a radial shoulder whose circumferential surface facing toward the spoke-like pattern of the respective cascade and whose end face facing toward the respective other airfoil cascade radially and axially bound a radially inner axial gap between the rotor blade cascade and the guide vane cascade.

(9) As can be seen in the elevation or developed view of FIG. 1, the contour of trailing-edge gap region 10.2T of the guide vane cascade, and more particularly its end face facing the rotor blade cascade, varies in the axial direction around the circumference; i.e., in the vertical direction in FIG. 1. That is, the generating lines of the end face extending from the axis of rotation of the turbomachine to the peripheral edge of the radial shoulder have different axial positions, so that the end face has a maximum axial extent at certain circumferential positions and a minimum axial extent at other circumferential positions in a direction away from the guide blade cascade; i.e., toward the right in FIG. 1. In the exemplary embodiment, the axial positions of the generating lines vary sinusoidally with an amplitude of A.sub.max, 1T/2, which results in a maximum axial variation of A.sub.max, 1T, which is illustrated in FIG. 1.

(10) The maximum axial extent of trailing-edge gap region 10.2T of the guide vane cascade is located near a guide blade trailing edge. In this regard, FIG. 1 shows a region .sub.1T which is 50% of guide vane pitch 1 (.sub.1T=0.5 1). It can be seen that the maximum axial extent of trailing-edge gap region 10.2T of the guide vane cascade is located in this region .sub.1T around the guide vane trailing edge; i.e., at a distance of no more than 50% of guide vane pitch 1.

(11) Analogously, the contour of leading-edge gap region 20.2L of the rotor blade cascade, and more particularly its end face facing the upstream guide vane cascade, varies in the axial direction around the circumference, so that the end face has a maximum axial extent at certain circumferential positions and a minimum axial extent at other circumferential positions in a direction away from the rotor vane cascade; i.e., toward the left in FIG. 1. In the exemplary embodiment, the axial positions of the generating lines vary sinusoidally with an amplitude of A.sub.max, 2L/2, which results in a maximum axial variation of A.sub.max, 2L, which is illustrated in FIG. 1.

(12) The maximum axial extent of leading-edge gap region 20.2L of the rotor blade cascade is located near a rotor vane leading edge. In this regard, FIG. 1 shows a region .sub.2L which is 50% of rotor vane pitch 2 (.sub.2L=0.52). It can be seen that the maximum axial extent of leading-edge gap region 20.2L of the rotor blade cascade is located in this region .sub.2L around the rotor blade leading edge; i.e., at a distance of no more than 50% of rotor vane pitch 2.

(13) The generating lines may be perpendicular to the axis of rotation of the turbomachine, or inclined thereto at the same angle or at an angle that varies in the circumferential direction. In the exemplary embodiment, the generating lines are perpendicular to the axis of rotation.

(14) FIG. 2 shows an axial elevation view of a cascade of guide vanes or rotor blades of a gas turbine stage according to the present invention having radially inner airfoil platforms whose gap region contour varies in the radial direction around the circumference. In particular, this gas turbine stage may be the one described hereinabove with reference to FIG. 1, so that a radial undulation is combined with an axial undulation. Therefore, in the following, reference is made to the above description and only the aspects of the radial undulation will be described. It is equally possible to provide only an axial undulation, as described hereinabove with reference to FIG. 1, or only a radial undulation, such as will be described hereinafter.

(15) In the axial elevation view of FIG. 2, there are shown bolded edges of airfoils 3 of a cascade of a gas turbine. The edges may be leading or trailing edges of either guide vanes or rotor blades, so that FIG. 2 may be an elevation view looking in the direction of the flow or one looking in a direction opposite to the flow direction. Thus, FIG. 2 is a compact representation of different aspects. For example, it can be considered to be an elevation view looking at rotor blades in the direction of the flow. In this case, the bolded edges correspond to leading edges, airfoils 3 correspond to, for example, rotor blades 2 of FIG. 1, and radial inner gap region 30.2 corresponds to their leading-edge gap region 20.2L, the view showing mainly the pressure sides of the respective rotor blades. FIG. 2 can equally be considered to be an elevation view looking at guide vanes in a direction opposite to the flow direction and showing only the suctions sides of the respective guide vanes. In this case, the bolded edges correspond to trailing edges, airfoils 3 correspond to, for example, guide vanes 1 of FIG. 1, and radial inner gap region 30.2 corresponds to their trailing-edge gap region 10.2T.

(16) The circumferential surface of gap region 30.2 has alternating maximum and minimum extents extending radially outward; i.e., in a direction toward the spoke-like pattern (upwardly in FIG. 2). It can be seen that the circumferential surface of gap region 30.2 varies sinusoidally in the circumferential direction with an amplitude R.sub.max, 3/2. The maximum radial extent of gap region 30.2 is located near the edge of airfoil 3. In this regard, FIG. 2 shows a region which is 50% of guide vane pitch (=0.5). It can be seen that the maximum radial extent of gap region 30.2 in a direction toward the spoke-like pattern (upwardly in FIG. 2) is located in this region around the airfoil edge; i.e., at a distance of no more than 50% of cascade pitch .

(17) If FIG. 2 is considered, for example, to be an elevation view looking at rotor blades in the direction of the flow, then the maximum radial extent of leading-edge gap region 30.2 is located (only just) in the pressure-side half of the segment between two successive rotor blade leading edges. If FIG. 2 is considered, for example, to be an elevation view looking at guide vanes in a direction opposite to the flow direction, then the maximum radial extent of trailing-edge gap region 30.2 is located (only just) in the suction-side half of the segment between two successive guide vane trailing edges.

(18) In addition or as an alternative to a radial-shoulder-shaped gap region 20.2L, the inner rotor blade platforms may have a rotor blade platform extension 20.3, which is preferably shaped like an annular flange, as illustrated in FIG. 3. The radially outer circumferential surface (the upper in FIG. 3) and/or the radially inner circumferential surface (the lower in FIG. 3) of this annular flange 20.3 may have a radial undulation along its entire length or part thereof, such as described hereinabove with reference to gap region 20.2L. However, for the sake of clarity, no such radial undulation is shown in FIG. 3. In particular, annular flange 20.3 may have a constant wall thickness when the two circumferential surfaces vary in parallel around the circumference.

(19) Additionally or alternatively, the axial end face of annular flange 20.3 and/or the radially inner facing circumferential surface of gap region 10.2T may have an undulation.

(20) In addition or as an alternative to radial-shoulder-shaped gap region 20.2L and/or a rotor blade platform extension 20.3, gap regions of radially outer rotor blade platforms and/or gap regions of radially outer guide vane platforms receiving the radially outer rotor blade platforms may have a radial and/or axial undulation (not shown). In a modification (also not shown), leading-edge gap regions of guide vanes and/or trailing-edge regions of rotor blades may additionally or alternatively also have a radial and/or axial undulation, such as described hereinabove with reference to FIGS. 1, 2.

LIST OF REFERENCE NUMERALS

(21) 1 guide vane 2 rotor blade 3 guide vane or rotor blade (edge) 10.1/20.1 cascade region of the radially inner airfoil platform of the rotor blade or guide vane cascade 10.2T trailing-edge gap region of the radially inner airfoil platform of the guide vane cascade 20.2L leading-edge gap region of the radially inner airfoil platform of the rotor blade cascade 30.2 gap region of a radially inner airfoil platform of a (rotor blade or guide vane) cascade 20.3 platform extension (wing, gap region) of the rotor blade platform A.sub.max, 1T maximum axial variation/extent of the trailing-edge gap region A.sub.max, 2L maximum axial variation/extent of the leading-edge gap region R.sub.max, 3 maximum radial variation/extent () cascade pitch (of the rotor blade or guide vane cascade)