US20140227102A1

US20140227102A1 - Rotor blade for a compressor of a turbomachine, compressor, and turbomachine

Info

Publication number: US20140227102A1
Application number: US14/122,864
Authority: US
Inventors: Harald Schoenenborn
Original assignee: MTU Aero Engines AG
Current assignee: MTU Aero Engines AG
Priority date: 2011-06-01
Filing date: 2012-06-01
Publication date: 2014-08-14
Also published as: EP2530330A1; EP2530330B1; WO2012163337A1

Abstract

A rotor blade for a compressor of a turbomachine is disclosed, the rotor blade having an airfoil having a leading edge, a trailing edge, an airfoil tip, as well as a pressure side and a suction side extending therebetween, provision being made for a profile variation on the pressure side, the profile variation extending into the trailing edge and into the airfoil tip. Also disclosed are a compressor and a turbomachine.

Description

The present invention relates to a rotor blade of a turbomachine, and to a compressor having such rotor blades, as well as a turbomachine having such a compressor.

BACKGROUND

When, during the operation of turbomachine compressors, the compressor is operated at other than the design point of the rotor blades, so-called compressor surge may frequently occur, in which the pressure in the compressor drops from the rear to the front due to flow separation from the rotor blades and a backflow formed as a result thereof. In such a case, high stresses occur in the blades, which are attributable to some extent to vibration or flutter during the backflow phase. One known measure for reducing stress is selective mistuning, which, however, is technically very complex. An alternative measure is influencing the steady and unsteady flows in the region of the trailing edge near the casing, because this region is decisive for the stress. German Patent Application DE 10 2008 037 154 A1, for example, proposes to form a plurality of parallelogram-like groves or circumferential grooves in a wall of an annular space opposite the rotor blades, the grooves being positioned in the wall of the annular space such that they are located in the leading edge region of the rotor blades. From German Patent Application DE 2 942 703 A1, it is known to form a circumferential groove in a wall of an annular space opposite the rotor blades.
In addition to these modifications on the stator side, it is also known to modify the blades in order to reduce stresses acting thereon. German Patent Application DE 103 52 253 A1, for example, proposes to improve the vibration characteristics of rotor blades by providing them at the tip with a step leading into the trailing edge. United States Patent Application US 2010/0008785 A1 proposes to improve the vibration characteristics by forming each rotor blade with a pressure-side pocket which extends between the leading edge and the trailing edge and is open toward the airfoil tip.

SUMMARY OF THE INVENTION

It is an object of the present invention to provide a rotor blade for a compressor of a turbomachine, which rotor blade will overcome the aforementioned disadvantages and is subjected to reduced stresses. Another object of the present invention is to provide a compressor having high compressor stability and a turbomachine having high efficiency.
The present invention provides a rotor blade for a compressor of a turbomachine that has an airfoil having a leading edge, a trailing edge, an airfoil tip, as well as a pressure side and a suction side extending therebetween. In accordance with the present invention, the airfoil has a profile variation on the pressure side, the profile variation extending into the trailing edge and into the airfoil tip.
The profile variation is a localized change in the profile of the suction side. Since the profile variation is located in a region bounded by the trailing edge and the airfoil tip, it is formed at the stagnation point of a backflow occurring during compressor surge, whereby, in accordance with the present invention, the backflow during compressor surge is diffused. The backflow is split into a plurality of partial flows and, therefore, cannot be transmitted in concentrated form to the neighboring airfoil. This significantly reduces the stress on the blades, so that the airfoils can be made with a smaller wall thickness and need not be thickened. In addition, this eliminates the need for complex selective mistuning.
In an exemplary embodiment which is easy to manufacture, the profile variation is a recess or depression which is open toward the airfoil tip and toward the trailing edge.
The diffusing effect of the recess or profile variation can be improved when the recess or profile variation begins in a region of a maximum pressure gradient.
In an alternative embodiment reducing the stress on the blades, the profile variation is configured as an undulating profile.
The backflow can be diffused very effectively when the undulating profile is oriented transversely to the direction of flow; i.e., radially to the direction of rotation, so that its troughs and crests each form separate barriers to the backflow, which split the partial flows themselves.
A compressor according to the present invention has at least one row of a plurality of rotor blades according to the present invention. A compressor of this kind is characterized by low stress on the rotor blades, for example, in the event of compressor surge; i.e., formation of abackflow, since such backflow is effectively diffused into a plurality of partial flows at the rotor blades. This firstly increases compressor stability and reduces what is known as “susceptibility to flutter.” Secondly, the rotor blades can be made with optimized wall thickness and thus reduced weight, which results in low compressor weight and high running smoothness.
The compressor stability can be further improved when the profile variations on the pressure sides are enlarged toward the suction sides and merge into set-backs extending through the airfoil tips in a transverse direction and in the direction of rotation and when, in addition, a plurality of counter-contours are formed in a radially opposite side wall. In this way, an annular gap bounded by the airfoil tips and the opposite side wall is radially enlarged in the direction of flow, which makes it possible to reduce the pressure gradient.
Preferably, the set-backs and the counter-contours are symmetrical, in particular mirror-symmetrical, to each other.
In an exemplary embodiment which is effective and easy to manufacture, the set-backs and the counter-contours form a plurality of rectangular enlargements of the annular gap.
In an alternative exemplary embodiment which is effective and easy to manufacture, the set-backs and the counter-contours form a plurality of funnel-shaped enlargements of the annular gap.
A turbomachine according to the present invention has a compressor according to the present invention. A turbomachine of this kind is characterized by low efficiency losses, and thus high efficiency, and by reduced airfoil stresses.
Other advantageous exemplary embodiments of the present invention are the subject matter of further dependent claims.

BRIEF DESCRIPTION OF THE DRAWINGS

Preferred exemplary embodiments of the present invention are described in greater detail with reference to highly simplified schematic drawings, in which:

FIG. 1 a shows a first exemplary embodiment of a rotor blade according to the present invention;

FIG. 1 b is a detail showing a pressure-side profile from FIG. 1 a;

FIG. 2 illustrates a second exemplary embodiment of the rotor blade according to the present invention;

FIG. 3 shows a first exemplary embodiment of an enlargement of a radial gap according to the present invention;

FIG. 4 shows a second exemplary embodiment of an enlargement of a radial gap according to the present invention;

FIG. 5 shows a first exemplary embodiment of a contour variation on the stator side; and

FIG. 6 shows a second exemplary embodiment of a contour variation on the stator side.

DETAILED DESCRIPTION

FIG. 1 is a developed top view of a blade row 1 of a compressor of an aircraft engine. Blade row 1 is composed of a plurality of rotor blades arranged side-by-side in the circumferential direction or direction of rotation y. For the sake of clarity, only one airfoil 2 is shown.
Airfoil 2 has a leading edge 4 and a trailing edge 6 as viewed in the direction of flow x, as well as a pressure side 8 and a suction side 10 extending therebetween. In addition, airfoil 2 has a tip 12 facing the observer and extending in the longitudinal direction of the airfoil, the airfoil tip constituting a radial outer boundary of suction side 8 and pressure side 10.
In accordance with the present invention, a profile variation 14 is formed in a pressure-side surface region between trailing edge 6 and airfoil tip 12. As shown in FIG. 1 b, this profile variation is located in the region of a stagnation point 16 forming in the case of a flow separation and causes backflow 18 to be diffused into a plurality of partial flows 18 a, 18 b oriented in different directions.
In the exemplary embodiment shown, profile variation 14 takes the form of a recess which is open toward trailing edge 6 and toward airfoil tip 12. The profile variation forms a region of reduced cross-section of airfoil 2 and has a flow impingement surface 22 which is set back from an adjacent surface portion 20. Preferably, the profile variation extends in a direction transverse to airfoil 2 up to the center thereof. The profile variation is bounded by an upstream curved vertical surface 24 and by downstream trailing edge 6, as viewed in the direction of flow x. In the radial direction, profile variation 14 is bounded by a radially inner curved longitudinal surface 26 extending in the longitudinal direction of the airfoil and by outer airfoil tip 12.
In an exemplary embodiment shown in FIG. 2, the profile variation 14 on the pressure side takes the form of an undulating profile including a plurality of peaks or crests 28 and a plurality of depressions or troughs 30. Peaks 28 each extend beyond the pressure side profile and form a localized thickening of airfoil 2. Depressions 30 cause a localized reduction in cross-section of airfoil 2. They extend transversely to the direction of flow x, and thus virtually radially to the axis or rotation, thereby forming a plurality of barriers backflow 18 and its partial flows 18 a, 18 b. It is preferred for the undulating profile to begin with a depression 30 as viewed in the direction of flow x. Undulating profile 14 may have a constant amplitude and a constant wavelength. However, the amplitude and/or the wavelength may also vary. It is conceivable, for example, to provide a large amplitude at stagnation point 16 and to decrease this amplitude in the upstream and downstream directions in such a way that undulating profile 14 tapers off to both sides in, as it were, a damped manner. In this connection, the wavelength may remain unchanged. It is possible, in particular, to use only troughs.
In both exemplary embodiments, backflow 18 is diffused into a plurality of partial flows 18 a, 18 b oriented in different directions, so that a pressure perturbation induced by the movement of the airfoil is transmitted to the neighboring airfoil in diffused thrill. At the same time, the profile variations 14 on the pressure side energize a low-energy boundary layer near the airfoil, thereby delaying flow separation. For optimum energization, profile variation 14 begins in the region of a maximum pressure gradient 32 between the low-energy boundary layer and a high-energy layer of flow distant from the airfoil, as viewed in the direction of flow x. To this end, as shown in FIG. 1 b, vertical surface 24 is disposed in the region of a maximum pressure gradient 32.
FIGS. 3 and 4 show exemplary embodiments of an enlargement 34 according to the present invention of a radial gap 36 for influencing the steady flow in the region of the trailing edges 6 of the rotor blades, the enlargement being bounded by the airfoil tips 12 of a row of rotor blades and an opposite side wall portion 38 on the stator side. To this end, a set-back 40 extending front pressure side 8 to suction side 10 is formed in the airfoil approximately in the region of the single-side profile variation 14 shown in FIGS. 1 a, 1 b and 2, so that airfoil tips 12 are radially depressed in the downstream region. On the stator side, a number of counter-contours 42 equal to the number of rotor blades are formed in side wall portion 38. Preferably, counter-contours 42 are configured symmetrically to set-backs 40 and in combination therewith form the enlargements 34 of the radial gap. In the exemplary embodiments shown, these enlargements are preferably rectangular in shape (FIG. 3) or funnel-shaped (FIG. 4).
FIGS. 5 and 6 show exemplary embodiments of a stator-side contour variation 44 according to the present invention for influencing the steady flow in the region of the trailing edges 6 of a blade row 1. To this end, a plurality of grooves 46 are formed in a side wall portion approximately radially opposite the trailing edges 6 of the rotor blades or airfoils 2. In the exemplary embodiment shown in FIG. 5, these grooves have a parallelogram-like shape and are evenly spaced from each other in the circumferential direction y. In the exemplary embodiment shown in FIG. 6, grooves 46 are configured as parallel circumferential grooves, which are preferably evenly spaced from each other in the direction of flow x. However, as illustrated by upstream groove 46 a, grooves 46 may also have different axial distances, as viewed in the direction of low x. A second blade 102 of the blade row is shown schematically.
Disclosed is a rotor blade for a compressor of a turbomachine, the rotor blade having an airfoil having a leading edge, a trailing edge, an airfoil tip, as well as a pressure side and a suction side extending therebetween, provision being made for a profile variation on the pressure side, the profile variation extending into the trailing edge and into the airfoil tip. Also disclosed are a compressor and a turbomachine.

LIST OF REFERENCE NUMERALS

1 blade row
2, 102 airfoil
4 leading edge
6 trailing edge
8 pressure side
10 suction side
12 airfoil tip
14 profile variation
16 stagnation point
18 backflow
18 a, b partial flows
20 surface portion
22 flow impingement surface
24 vertical surface
26 longitudinal surface
28 peak/crest
30 depression/trough
32 pressure gradient
34 enlargement
36 radial gap
38 side wall portion
40 set-back
42 counter-contour
44 contour variation
46 groove
46 a upstream groove
x longitudinal direction of flow/engine
y circumferential direction/direction of rotation
z vertical direction/radial direction

Claims

1-11. (canceled)

12. A rotor blade for a compressor of a turbomachine, comprising:

an airfoil having a leading edge, a trailing edge, and an airfoil tip (12), the airfoil having a pressure side and a suction side extending therebetween, and a profile variation on the pressure side, the profile variation extending into the trailing edge and into the airfoil tip.

13. The rotor blade as recited in claim 12 wherein the profile variation forms a recess open toward the airfoil tip and toward the trailing edge.

14. The rotor blade as recited in claim 13 wherein the profile variation begins in the region of a maximum pressure gradient.

15. The rotor blade as recited in claim 12 wherein the profile variation forms an undulating profile.

16. The rotor blade as recited in claim 15 wherein the undulating profile is oriented transversely to the direction of flow of the airfoil.

17. A compressor comprising at least one blade row including a plurality of rotor blades as recited in claim 12.

18. The compressor as recited in claim 17 further comprising a side wall radially opposite the rotor blades, the profile variations on the pressure sides being enlarged toward the suction sides and each merging into a set-back extending through the airfoil tips in the direction of rotation, and a plurality of counter-contours are formed in the side wall.

19. The compressor as recited in claim 18 wherein the set-backs and the counter-contours are symmetrical to each other.

20. The compressor as recited in claim 18 wherein the set-backs and the counter-contours form rectangular enlargements of an annular gap between the side wall and the rotor blades.

21. The compressor as recited in claim 18 wherein the set-backs and the counter-contours form funnel-shaped enlargements of an annular gap between the side wall and the rotor blades.

22. A turbomachine comprising the compressor as recited in claim 17.