ROTOR BLADE FOR A TURBOMACHINE

Abstract

The present invention relates to a rotor blade (20) for arrangement in a gas duct (2) of a turbomachine (1), having a rotor blade airfoil (23), which, viewed in a tangential section, has a blade airfoil profile (24) with a leading edge radius RVK and a rotor blade airfoil thickness d, wherein the blade airfoil profile (24) is thickened, at least in sections, specifically the blade airfoil thickness d is specified, in relation to the front edge radius RVK, such that (2d/Rvk2)−d≤5.5.

Claims

1.-13. (canceled)

14. A rotor blade for arrangement in a gas duct of a turbomachine, wherein the rotor blade comprises a rotor blade airfoil which, when viewed in a tangential section, has a blade airfoil profile comprising a front edge radius R.sub.VK, and a blade airfoil thickness d, the blade airfoil profile being thickened at least in sections, and the blade airfoil thickness d being set at a ratio to the front edge radius R.sub.VK in such a manner that
(2.Math.d/R.sub.VK.sup.2)−d≤5.5.

15. The rotor blade of claim 14, wherein the blade airfoil thickness d is set at a ratio to the front edge radius R.sub.VK in such a manner that 0.6≤(2.Math.d/R.sub.VK.sup.2)−d≤5.0.

16. The rotor blade of claim 14, wherein the front edge radius R.sub.VK ranges from 0.2 mm to 1.8 mm.

17. The rotor blade of claim 14, wherein the front edge radius R.sub.VK ranges from 0.6 mm to 1.4 mm.

18. The rotor blade of claim 14, wherein the front edge radius R.sub.VK ranges from 0.7 mm to 1.4 mm.

19. The rotor blade of claim 14, wherein the blade airfoil thickness d in a front edge region is from 0.5 mm to 5 mm.

20. The rotor blade of claim 14, wherein the blade airfoil thickness d in a front edge region is from 1.5 mm to 4 mm.

21. The rotor blade of claim 14, wherein the blade airfoil thickness d in a front edge region is from 2 mm to 2.5 mm.

22. The rotor blade of claim 16, wherein the blade airfoil thickness d in a front edge region is from 0.5 mm to 5 mm.

23. The rotor blade of claim 14, wherein a front edge region in which the blade airfoil profile has the front edge radius R.sub.VK reaches a spacing A, measured along a skeletal line of the blade airfoil profile from a front edge thereof, of from 0.1 mm to 0.3 mm.

24. The rotor blade of claim 14, wherein, in terms of a rotor blade airfoil height H measured from radially inside to radially outside, the blade airfoil profile is thickened at radial positions of at least from 60% to 90% of the rotor blade airfoil height H.

25. The rotor blade of claim 14, wherein, in terms of a chord length l measured on the blade airfoil profile, the blade airfoil thickness d is set at a ratio to the front edge radius R.sub.VK and the chord length l in such a manner that d/(R.sub.VK.Math.l)≤7.2.

26. The rotor blade of claim 25, wherein the blade airfoil thickness d is set at a ratio to the front edge radius R.sub.VK and the chord length l in such a manner than 4.0≤d/(R.sub.VK.Math.l)≤7.0.

27. The rotor blade of claim 25, wherein, in terms of a spacing A that is measured along a skeletal line of the blade airfoil profile from a front edge thereof, the blade airfoil thickness d, at least at spacings between 1.0 mm and 2.0 mm, is set at a ratio to the front edge radius R.sub.VK and the chord length l in such a manner that d/(R.sub.VK.Math.l)≤7.2.

28. The rotor blade of claim 14, wherein at least the rotor blade airfoil is made from a brittle material.

29. The rotor blade of claim 14, wherein at least the rotor blade airfoil is made from a high-temperature-resistant material.

30. The rotor blade of claim 14, wherein the rotor blade airfoil is present as a solid profile.

31. The rotor blade of claim 14, wherein the rotor blade is configured for a high-speed rotor having an An.sup.2 of at least 2000 m.sup.2/s.sup.2.

32. A turbine module for an aircraft engine, wherein the module comprises the rotor blade of claim 14.

33. An aircraft engine, wherein the aircraft engine comprises the turbine module of claim 32.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0028] The invention will be explained in more detail hereunder by means of an exemplary embodiment, wherein the individual features in the context of the coordinate claims may also be relevant to the invention in any other combination and, furthermore, no distinction is made in detail between the different categories of claims.

[0029] In the drawings:

[0030] FIG. 1 shows a fanjet in an axial section;

[0031] FIG. 2 shows a blade airfoil profile of a rotor blade designed according to the invention with a comparative profile;

[0032] FIG. 3 shows a detailed view of the blade airfoil profile according to FIG. 2 for illustrating the thickening on the front edge;

[0033] FIG. 4 shows a diagram for illustrating the ratio between the front edge radius and the blade airfoil thickness, plotted over the spacing from the front edge;

[0034] FIG. 5 shows the determination of the chord length on a blade airfoil profile; and

[0035] FIG. 6 shows a ratio between the blade airfoil thickness, the chord length and the front edge radius, plotted over the spacing from the front edge.

PREFERRED EMBODIMENT OF THE INVENTION

[0036] FIG. 1 shows a turbomachine 1 in a schematic view, specifically a fanjet. The turbomachine 1 in functional terms is made up of a compressor 1a, a combustion chamber 1b and a turbine 1c, the latter having a high-pressure turbine module 1ca and a low-pressure turbine module 1cb. Here, the compressor 1a as well as the turbine 1c are in each case constructed from a plurality of stages, each stage being composed of a guide vane assembly and a rotor blade assembly. In terms of the circulating flow in the gas duct 2, one stage of the rotor blade assembly is disposed downstream of each guide vane assembly. The rotor blades in operation rotate about the longitudinal axis 3. The fan 4 is coupled by way of a gearbox 5; the rotor blade assemblies of the low-pressure turbine module 1cb rotate in operation faster than the fan 4. The reference sign 20 in an exemplary manner identifies some rotor vanes.

[0037] FIG. 2 shows a rotor blade airfoil 23 for a rotor blade 20 of the turbine 1c, specifically of the low-pressure turbine module 1cb, in a tangential section. In relation to FIG. 1, the section plane thus lies so as to be perpendicular to the drawing plane and is horizontal. FIG. 2 thus shows the blade airfoil profile 24 which is defined so as to extend from a front edge 25 up to a rear edge 26 between a suction side 27 and a pressure side 28.

[0038] A profile which is optimized solely in aerodynamic terms is identified by the solid lines in FIG. 2. As has been set forth in detail in the introductory description, this can however be mechanically disadvantageous in terms of the structure, in particular having an insufficient tolerance in terms of impact. For comparison, a blade airfoil profile 24 that has been thickened according to the invention is plotted using the dashed lines, the blade airfoil thickness d being increased in particular in a region of the front edge 25, this improving the impact resistance.

[0039] FIG. 3 shows a region on the front edge 25 in detail. The blade airfoil thickness d is measured perpendicularly to the skeletal line 30. The blade airfoil profile 24 at the front edge 25 has front edge radius R.sub.VK (a corresponding arc is illustrated by dotted lines). Presently, a front edge region 35, across which the rotor blade airfoil 23 has the front edge radius R.sub.VK, extends across approximately 0.15 mm. In order for the front edge radius R.sub.VK to be determined, a circle or arc, respectively, the center of the latter being on the skeletal line 30, can be fitted into the front edge region 35 at a plurality of points, for example by the “best fit” method.

[0040] In the rotor blade airfoil 23, the thickening is set according to the invention such that

[00004]$\begin{array}{cc}\frac{2\times d}{R_{VK}^2}-d& \left(\mathrm{Formula}1\right)\end{array}$

is between 0.6 and 5.5. The thickening achieved therewith is advantageous not only in mechanical-structural but also aerodynamical terms; cf. the introductory description in detail. The thickening is provided in particular in a region proximal to the front edge. In terms of the radial extent of the rotor blade airfoil 23, thus the rotor blade airfoil height H (cf. FIG. 1), said rotor blade airfoil 23 can be correspondingly be optimized above all in a radially outer portion.

[0041] FIG. 4 illustrates the ratio according to the preceding formula for different spacings A from the front edge 25 (cf. FIG. 3), said spacings A being plotted in millimeters here. The curve 40 herein shows the values measured on the blade airfoil profile 24 optimized according to the invention; for comparison, the curves 41 show values measured on different profiles which are in each case optimized solely in aerodynamic terms.

[0042] FIG. 5 shows how the chord length l is determined on a blade airfoil profile 24. Said chord length l is measured along a connecting tangent 50 which on the pressure side is placed on the blade airfoil profile 24 and on the latter has an axially front contact point 51.1 and an axially rear contact point 51.2. The chord length l is then measured between two further tangents 52.1, 52.2 which lie in each case so as to be perpendicular to the connecting tangent 50, wherein the tangent 52.1 axially at the front has a contact point 53.1 and the tangent 52.2 axially at the rear has a contact point 53.2.

[0043] FIG. 6 illustrates a ratio between the blade airfoil thickness d, the front edge radius R.sub.VK and the chord length l set according to the invention, again plotted over the spacing A measured from the front edge 25 along the skeletal line 30. The curve 60 here is obtained by means of a blade airfoil profile 24 optimized in structural as well as aerodynamic terms, whereas the curves 61 for comparison reflect profiles that are optimized solely in aerodynamic terms.

LIST OF REFERENCE SIGNS

[0044] Turbomachine [0045] 1a Compressor [0046] 1b Combustion chamber [0047] 1c Turbine [0048] 1ca High-pressure turbine module [0049] 1cb Low-pressure turbine module [0050] 2 Gas duct [0051] 3 Longitudinal axis [0052] 4 Fan [0053] 5 Gearbox [0054] 20 Rotor blade [0055] 23 Rotor blade airfoil [0056] 24 Blade airfoil profile [0057] 25 Front edge [0058] 26 Rear edge [0059] 27 Suction side [0060] 28 Pressure side [0061] 30 Skeletal line [0062] 35 Front edge region [0063] 40 Curve (optimized according to the invention) [0064] 41 Curve (optimized solely in aerodynamic terms) [0065] 50 Connecting tangent [0066] 51.1 Axially front contact point [0067] 51.2 Axially rear contact point [0068] 52.1, 52.2 Further tangents [0069] 53.2 Axially rear contact point [0070] 60 Curve (optimized according to the invention) [0071] 61 Curve (optimized solely in aerodynamic terms) [0072] A Spacings [0073] d Blade airfoil thickness [0074] H Rotor blade airfoil height [0075] L Skeletal line length [0076] l Chord length [0077] R.sub.VK Front edge radius