ROTOR BLADE FOR A TURBOMACHINE, ASSOCIATED TURBINE MODULE, AND USE THEREOF

Abstract

Rotor blade (20) to be arranged in a gas conduit (3) of a turbomachine (1), having a rotor blade airfoil (23), which radially inwardly has a chord length S.sub.i, radially outwardly has a chord length S.sub.a, and in a radial position

r.sub.x inbetween has a chord length S.sub.x, the chord length S in the radial position r.sub.x being at least equal to the chord length S.sub.i radially inwardly (S.sub.i<S.sub.x), and the chord
length S.sub.a radially outwardly corresponding at most 0.9 times the chord length S.sub.x in the radial position r.sub.x inbetween (Sa<0.9 S.sub.x).

Claims

1.-15. (canceled)

16. A rotor blade for arrangement in a gas duct of a turbomachine, wherein the rotor blade comprises a rotor blade airfoil which radially inwardly has a chord length S.sub.i, radially outwardly has a chord length S.sub.a, and, in a radial position r.sub.x inbetween, has a chord length S.sub.x, the chord length S.sub.x in a radial position r.sub.x being at least equal to the chord length S.sub.i radially inwardly (S.sub.i≤S.sub.x), and the chord length S.sub.a radially outwardly corresponds at most to 0.9 times the chord length S.sub.x in the radial position r.sub.x inbetween (S.sub.a≤0.9 S.sub.x).

17. The rotor blade of claim 16, wherein in relation to a rotor blade airfoil height taken from radially inside to radially outside, the radial position r.sub.x with the chord length S.sub.x is at least 20% and at most 50% of the rotor blade airfoil height.

18. The rotor blade of claim 16, wherein the chord length S.sub.i radially inwardly corresponds to at least 0.9 times the chord length S.sub.x radially inbetween (S.sub.i≥0.9 S.sub.x).

19. The rotor blade of claim 16, wherein the chord length S.sub.a radially outwardly corresponds to at least 0.7 times the chord length S, radially inbetween (S.sub.a≥0.7 S.sub.x).

20. The rotor blade of claim 16, wherein a radial progression of a chord length S(r) radially outwardly from the radial position r.sub.x shows a monotonic decrease from S.sub.x to S.sub.a.

21. The rotor blade of claim 16, wherein a radial progression of a chord length S(r) radially inwardly from the radial position r.sub.x shows a monotonic decrease from S.sub.x to S.sub.i.

22. The rotor blade of claim 20, wherein the monotonic decrease is strictly monotonic and follows a constant slope.

23. The rotor blade of claim 21, wherein the monotonic decrease is strictly monotonic and follows a constant slope.

24. The rotor blade of claim 20, wherein the monotonic decrease is strictly monotonic and follows a slope which increases radially inwardly or outwardly away from the radial position r.sub.x.

25. The rotor blade of claim 21, wherein the monotonic decrease is strictly monotonic and follows a slope which increases radially inwardly or outwardly away from the radial position r.sub.x.

26. The rotor blade of claim 16, wherein, in relation to its radial rotor blade airfoil height, the rotor blade airfoil is provided, at least in some section or sections, with a slope toward its suction side, wherein the slope is set in such a way that, during operation, a centrifugal-force bending moment which the centrifugal force brings about on the rotor blade airfoil as a result of the slope is greater than a gas-force bending moment which acts on the rotor blade airfoil as a result of a flow around the rotor blade airfoil.

27. The rotor blade of claim 16, wherein the rotor blade comprises an outer shroud arranged radially outwardly on the rotor blade airfoil, a single sealing fin being arranged radially outwardly on the outer shroud.

28. The rotor blade of claim 16, wherein at least the rotor blade airfoil is made of a high-temperature-resistant material.

29. The rotor blade of claim 28, wherein the high-temperature-resistant material comprises a titanium aluminide.

30. The rotor blade of claim 28, wherein the high-temperature-resistant material comprises a TNM (titanium niobium molybdenum) alloy.

31. The rotor blade of claim 16, wherein the rotor blade airfoil is provided with a coating at least at a leading edge of the rotor blade airfoil.

32. The rotor blade of claim 31, wherein the coating is a multilayer coating.

33. The rotor blade of claim 16, wherein the rotor blade is designed for a high-speed rotor having an An.sup.2 of at least 2000 m.sup.2/s.sup.2.

34. A turbine module for an aircraft engine, wherein the module comprises the rotor blade of claim 16 and is designed to feed a cooling fluid to an outer shroud of the rotor blade, the cooling fluid being fed in from outside the rotor blade.

35. The module of claim 34, wherein the rotor blade is capable of rotating with an An.sup.2 of at least 2000 m.sup.2/s.sup.2.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0033] The invention is explained in greater detail below with reference to an exemplary embodiment, although, within the scope of the additional independent claims, the individual features may also be essential to the invention in some other combination, and, in this case too, no distinction is drawn specifically between the various categories of claims.

[0034] More particularly,

[0035] FIG. 1 shows schematically a turbofan engine in an axial section;

[0036] FIG. 2 shows schematically a rotor blade of the engine according to FIG. 1 in a side view;

[0037] FIG. 3 shows the rotor blade according to FIG. 2 in an axial view.

[0038] FIG. 4 shows the relationship between the chord length S and the radius r;

[0039] FIG. 5 shows the determination of the chord length S on a cross-sectional profile.

PREFERRED EMBODIMENT OF THE INVENTION

[0040] FIG. 1 shows a turbomachine 1 in a schematic view, specifically a turbofan engine. The turbomachine 1 is subdivided functionally into 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. In this case, both the compressor 1a and the turbine 1c are composed of a plurality of stages, each stage being composed of a guide vane ring and a rotor blade ring. In relation to the flow around them in the gas duct 2, the rotor blade ring is arranged downstream of the guide vane ring in each stage. During operation, the rotor blades rotate about the longitudinal axis 3. The fan 4 is coupled via a transmission 5, and the rotor blade rings of the low-pressure turbine module 1cb rotate faster than the fan 4 during operation.

[0041] FIG. 2 shows a rotor blade 20 in a schematic side view, namely a rotor blade 20 of a rotor blade ring of the turbine 1c, specifically of the low-pressure turbine module 1cb. The rotor blade has a blade root 21, which has no further relevance in the present case, and an inner platform 22 radially to the outside of it. The airfoil 23 extends radially outward from the inner platform 22. Arranged at the radially outer end of the airfoil 23 is an outer shroud 24, which has exactly one sealing fin 24.1. This is advantageous with regard to the weight and hence the edge load, cf. the introduction to the description for more detail.

[0042] In relation to the flow around it in the hot-gas duct, the airfoil 23 has a leading edge 23a, a trailing edge 23b, and two side faces 23c,d, which each connect the leading edge 23a and the trailing edge 23b to one another. One of the side faces 23c,d forms the suction side of the rotor blade 20, the other the pressure side. At the leading edge 23a, the rotor blade 20 is provided with a coating 25 for protection against impact damage, said coating being composed of a metallic layer and a ceramic layer arranged thereon (the layers are not shown in detail). From the illustration according to FIG. 2, it can furthermore be seen that the schematically shown chord length S decreases radially outwardly away from a radial position r.sub.x, which likewise reduces the edge load. The chord length S remains constant or even decreases slightly inwardly from the radial position r.sub.x, cf. FIG. 4.

[0043] FIG. 3 shows the rotor blade airfoil 23 schematically in an axial view, which illustrates the slope of the rotor blade airfoil 23. In the illustration, the suction side 41 is on the left of the rotor blade airfoil 23, and the pressure side 42 is on the right. The rotor blade airfoil 23 slopes toward the suction side 41, specifically radially in the center with respect to the rotor blade airfoil height 45. Radially on the inside and radially on the outside, the slope is less steep, and the rotor blade airfoil 23 can also run into the hub or the casing without any slope at all. In this context, the slope toward the suction side 41 is set in such a way that the centrifugal-force bending moment 46 acting on the rotor blade airfoil 23 during operation is greater than the gas-force bending moment 47. As a result, the rotor blade airfoil 23 is bent toward the pressure side 42, which reduces the load there and thus the susceptibility to impact at the leading edge 23a, cf. also the introduction to the description.

[0044] FIG. 4 illustrates the relationship between the chord length S and the radius r, given as a percentage of the radial rotor blade airfoil height. Radially inwardly, the airfoil has the chord length S.sub.i and, radially outwardly, it has the chord length S.sub.a. In a radial position r.sub.x inbetween, it has the chord length S.sub.x (which in the present case represents a maximum). The radial position r.sub.x is between 20% and 50% of the radial rotor blade airfoil height.

[0045] In the radially outer section 46, that is to say radially outwardly from the radial position r.sub.x, the chord length S decreases. This reduces the edge load and thus increases the impact tolerance in this region. Radially outwardly, the chord length S.sub.a is 0.7 to 0.9 times the chord length S.

[0046] In the radially inner section 47, the chord length S does not decrease radially outwardly. It may either be constant (not illustrated) or, as shown in FIG. 4, it may even slightly increase outwardly, that is to say decrease inwardly away from the radial position r. The chord length S.sub.i radially inwardly is 0.9 to 1 times the chord length S. The inventors have observed that overall there are nevertheless no losses in robustness, cf. the introduction to the description for more detail. On the other hand, limiting the chord lengths radially inwardly permits an axially more compact construction, which may be advantageous, for example, with regard to weight and efficiency.

[0047] FIG. 5 illustrates the airfoil 23 in a tangential section. The chord length S is taken along a connecting tangent 50, which is placed against the profile on the pressure side and has a contact point 51.1 axially at the front and a contact point 51.2 axially at the rear on the profile. The chord length S is then taken between two further tangents 52.1, 52.2, which are each perpendicular to the connecting tangent 50, tangent 52.1 having a contact point 53.1 axially at the front and tangent 52.2 having a contact point 53.2 axially at the rear.

TABLE-US-00001 LIST OF REFERENCE SIGNS turbomachine  1 compressor  1a combustion chamber  1b turbine  1c high-pressure turbine module  1ca low-pressure turbine module  1cb gas duct  2 longitudinal axis  3 fan  4 transmission  5 rotor blade 20 blade root 21 inner platform 22 airfoil 23 leading edge 23a trailing edge 23b side faces 23c, d outer shroud 24 sealing fin 24.1 coating 25 chord length 26 profile surface 27 suction side 41 pressure side 42 rotor blade airfoil height 45 outer section 46 inner section 47 centrifugal-force bending moment 48 gas-force bending moment 49 connecting tangent 50 front contact point 51.1 rear contact point 51.2 further tangents 52.1, 52.2 contact points 53.1, 53.2