Rotor blade for a gas turbine and gas turbine

Abstract

Rotor blade for use in a high-speed low-pressure turbine of a gas turbine, comprising a blade root element (12) and a jet deflection section (16) which connects to the blade root element (12) in a radial direction (RR), wherein respective centers of gravity (24, 24') of blade cross-sectional areas (20) of the jet deflection section (16) lie on a common threading axis (26), wherein the threading axis (26) extends from a first center of gravity (24) of a first blade cross-sectional area (20) adjoining the blade foot element (12) within a cone (28) whose cone apex lies in the first center of gravity (24) and whose cone height (KH) extends orthogonally to the plane of the blade cross-sectional area (20), wherein the cone height (KH) in the installed state of the rotor blade (10) on the gas turbine extends in the radial direction (RR) of the gas turbine, and wherein the threading axis (26) in the installed state of the rotor blade (10) on the gas turbine is inclined with respect to the radial direction (RR) in the gas flow direction in the axial direction (AR) and / or in the circumferential direction of the gas turbine, characterized in that the opening angle (β) of the cone (28) is greater than 0° and less than or equal to 4°.

Description

[0001] The invention relates to a turbine blade for a gas turbine, in particular an aircraft turbine, comprising a blade root element and a jet deflection section that extends radially to the blade root element, wherein the respective centers of gravity of the blade cross-sectional areas of the jet deflection section lie on a common axis of rotation. The jet deflection section is typically a blade with a convex suction side and a concave pressure side, which are connected to each other via a leading and trailing edge. The blade cross-sectional areas are defined as sections through the jet deflection section with cylindrical surfaces arranged concentrically around the engine axis of the gas turbine.

[0002] Exemplary turbine blades for a gas turbine are known from documents DE 43 44 189 C1, DE 602 16 360 T2, JP 4 299 301 B2 and GB 260 411 A.

[0003] Directional terms such as "axial", "radial" and "circumferential" are generally to be understood as referring to the machine axis of the gas turbine, unless the context explicitly or implicitly indicates otherwise.

[0004] In the development of aircraft gas turbines, it was recognized that decoupling the two components, the low-pressure turbine and the fan, which are usually mounted on a single shaft, allows for optimized operation. In such aircraft gas turbines, the low-pressure turbine and the fan are connected by a reduction gearbox, with the low-pressure turbine rotating approximately three times faster than the fan. These low-pressure turbines are also known as high-speed low-pressure turbines.

[0005] A gas turbine with a low-pressure turbine coupled to the fan of the aircraft gas turbine by means of a reduction gear is known from document EP 2 915 962 A1.

[0006] Due to the altered, and in particular higher, rotational speed of the low-pressure turbine compared to conventional aircraft (gas turbines), the turbine blades, which are attached to a rotor of a turbine stage, are subjected to higher centrifugal forces. In addition to the increased centrifugal forces, the forces of the oncoming hot gas flow must also be taken into account.

[0007] In conventional aircraft gas turbines, it is known to tilt the threading axis of blade cross-sectional areas by up to 5 degrees relative to a purely radial direction (relative to the center of the gas turbine), with optimal tilts in the range of about 2.5 to 4.5° for the slower rotating conventional low-pressure turbines.

[0008] It has been shown that the use of conventional rotor blades in aircraft gas turbines with high-speed, low-pressure turbines is problematic in that local stress peaks occur in the rotor blades, resulting in uneven blade utilization. Overall, such rotor blades, when used in high-speed rotors of a turbine stage, result in unfavorable stress profiles.

[0009] In contrast, the object of the invention is to provide a turbine blade for a gas turbine which overcomes the aforementioned disadvantages and is particularly suitable for use in high-speed low-pressure turbines.

[0010] To solve this problem, a guide vane is proposed in which the threading axis extends from a first centroid of a first vane cross-sectional area adjoining the vane foot element within a cone whose cone apex lies in the first centroid and whose cone height extends orthogonally to the plane of the vane cross-sectional areas, wherein the opening angle of the cone is greater than 0° and less than 4°, preferably greater than or equal to 0.5° and less than or equal to 2°.

[0011] In this arrangement, the cone height extends essentially along the radial direction of the gas turbine, so that in the proposed design, the thread axis deviates from the radial direction or cone height by a maximum of 2° in the axial or circumferential direction. In other words, the cone apex is located at the intersection of the thread axis with an inner annular space of the turbine. It has been shown that, particularly in high-speed, low-pressure turbines, the relationship between centrifugal force and the force of the oncoming hot gas has changed, so that more uniform blade loading without local stress peaks can be achieved with rotor blades of the described design. The deviation of the rotor blade thread axis from the purely radial direction thus occurs within a very narrow range, which is particularly advantageous for high-speed rotors of a turbine stage.

[0012] Additionally, the cone height extends radially along the gas turbine when the rotor blade is attached.

[0013] Furthermore, it is preferred that the threading axis extends linearly within the cone from the first center of gravity.

[0014] To counteract the forces acting on the hot gas flowing through it, it is proposed that the thread axis, when the rotor blade is installed on the gas turbine, be inclined in the axial direction of the gas turbine relative to the radial direction in the gas flow direction, and / or, to optimize the stress distribution with regard to the higher centrifugal force acting on it, that the thread axis, when the rotor blade is installed on the gas turbine, be inclined in the circumferential direction of the gas turbine relative to the radial direction.

[0015] Since the cover band element forms the radially outer end of the running blade and centrifugal forces also act on it, it is proposed that the centers of gravity of cross-sectional areas of the cover band element are also arranged on the threading axis.

[0016] According to a further aspect, the invention relates to a gas turbine, in particular an aviation gas turbine, comprising at least one turbine stage with a rotor to which several adjacent running blades with one of the above-mentioned features are attached in the circumferential direction.

[0017] It is preferred that at least one turbine stage is part of a low-pressure turbine. It is further proposed that the low-pressure turbine be coupled to the fan of the aircraft gas turbine by means of a reduction gearbox.

[0018] Furthermore, in a gas turbine, it is preferred that the respective threading axes of the rotor blades are inclined equally to a respective straight line running radially through the first center of gravity of the first blade cross-sectional area.

[0019] The invention is described below by way of example and without limitation with reference to the accompanying figures. Fig. Figure 1 shows an example of a turbine blade of an aircraft gas turbine in perspective view. Fig. Figure 2 shows an example of a rotor blade cross-section. Fig. Figure 3 schematically shows in sub-figures A and B the relative arrangement of a threading axis with respect to a cone and an oblique cylinder. Fig. Figure 4 shows two superimposed blade cross-sections at different radial positions along the threading axis of the blade.

[0020] Fig. Figure 1 shows an exemplary perspective view of a turbine blade 10 of an aircraft gas turbine. The turbine blade 10 comprises a blade root element 12, a shroud element 14, and a jet deflection section 16, which is arranged in the radial direction RR between the blade root element 12 and the shroud element 14. The jet deflection section 16 has radially inner and radially outer transition areas 18 in which the cross-sectional area changes, in particular increasing, with a view to a stable, preferably one-piece connection with the blade root element 12 and the shroud element 14, respectively. The turbine blade 10 shown here is purely exemplary in order to describe the general structure of a turbine blade. The turbine blade 10 shown here is of the Fig. Figure 1 does not necessarily show all features of a running blade 10 according to the invention.

[0021] Fig. Figure 2 shows an exemplary cross-sectional area 20 of a running blade 10, for example in the area of ​​section line II-II of the Fig. 1. The blade cross-sectional area 20 is preferably obtained as a section through the jet deflection section 16 with a cylindrical surface arranged concentrically around the machine axis of the gas turbine (not shown here). For simplicity, this surface can also be represented flat in a plane. The cross-sectional area 20 can be the same over the entire length of the jet deflection section 16. However, it is also conceivable that the outer contour 22, and thus also the cross-sectional area 20, changes from a first cross-sectional area, which may be located approximately in the region of the dashed line S1 (i.e., above the transition region 18), to a radially outermost cross-sectional area, which may be located approximately in the region of the dashed line Sn, depending on the desired aerodynamic effect of the rotor blade 10 or the jet deflection section 16. In particular, the cross-sectional area can, preferably monotonically, decrease radially outwards.This has advantages with regard to the structural mechanical requirements of the rotor blades 10.

[0022] The cross-sectional area 20 of the rotor blade has a center of gravity 24, which is qualitatively indicated in the figure. The position of the center of gravity 24 is purely illustrative and not mathematically or geometrically precise. A [missing information] extends through this center of gravity 24. Fig. 2. A threading axis (not shown) runs along which all imaginary cross-sectional areas are arranged or threaded onto one another, with the threading axis passing through the centroids of all cross-sectional areas. All these centroids lie on a straight line.

[0023] Fig. 3 shows in the Fig. 3A and Fig. 3B Geometric relationships between the threading axis 26 and the radial direction RR for a rotor blade and its cross-sectional areas. According to Fig. 3A is located at line S1 of the first cross-sectional area of ​​the beam deflection section 16 ( Fig. 1) At the center of gravity 24, a point of intersection between the radial direction RR and the threading axis 26 is defined. The threading axis 26 is inclined relative to the radial direction RR and thus deviates from it. In the illustrated example, the threading axis 26 lies within a cone 28, the apex of which is located at the center of gravity 24 and the height of which KH coincides with the radial direction RR. The cone 28 has an opening angle β, which lies in a range from greater than 0° to 4°. The threading axis 26, which preferably has a linear progression starting from the centroid 24 of the first cross-sectional area at the radial level S1, lies within the lateral surface of the cone 28. The threading axis 26 can thus run inclined in the circumferential direction and / or in the axial direction of the gas turbine to the radial direction RR, wherein the deviation from the radial direction is no more than 2°, which corresponds to half the opening angle β of the cone 28.

[0024] In an embodiment not necessarily according to the invention (the present invention is defined by means of a cone and not by means of a cylinder), the threading axis 26 can run within an inclined cylinder 30, wherein the inclined cylinder has a cylinder height ZH that coincides with the radial direction RR and whose cylinder axis ZA is formed at an angle φ of greater than 0° to 2° to the radial direction RR or cylinder height ZH. The diameter of the cylinder 30 can, for example, be selected depending on the size of a specific cross-sectional area of ​​the rotor blade; for example, the diameter of the cylinder could be half the thickness of a cross-sectional area, wherein the thickness of the cross-sectional area can be measured, for example, in a thickness direction DR perpendicular to the axial direction AR of the gas turbine and intersecting the center of gravity 24, as is done in Fig. Figure 4 indicates that the thickness direction DR can also be understood as tangential to the circumferential direction of the gas turbine. However, this is only a simplified example. The diameter of the cylinder 30 can also be selected depending on other dimensions or parameters of the rotor blade. In particular, the diameter can also be set to any value that is not necessarily related to the geometric dimensions of the rotor blade. It is particularly preferred that the cylinder be inclined axially in the direction of hot gas flow. This allows the bending moment resulting at the connection point of the blade or jet deflection section with the blade root element, which arises from centrifugal forces and pressure forces of the flow, to be kept as small as possible.

[0025] Fig. Figure 4 shows, by way of example, two superimposed contours of blade cross-sectional areas at the level of lines S1 and Sn respectively. Fig. 1. As can be seen from this illustration, the centroid 24' of the cross-sectional area at height Sn (dashed contour) differs from the centroid 24 of the cross-sectional area 20 at height S1. The undrawn threading axis 26 passes through these two centroids 24 and 24' and is therefore opposite to a radial direction that is in Fig. 4 runs orthogonally to the plane of the paper through the center of gravity 24, inclined, whereby the angle enclosed between the radial direction and the threading axis is greater than 0° and less than or equal to 2°. In Fig. 4 furthermore, the boundary defined by a cone 28 or an inclined cylinder 30 is evident through the circle 32, within which the centers of gravity of all cross-sectional areas of the impeller are arranged.

[0026] It should be noted that the angles and inclinations shown are exaggerated for clarity and do not correspond to the preferred angle ranges of greater than 0° to a maximum of 2°. The chosen representations serve to illustrate the principle of the geometric relationships without achieving a dimensionally accurate or true-to-scale representation.

[0027] The presented design of a turbine blade, whose centers of gravity of the cross-sectional areas are arranged along a threading axis that is only slightly inclined to the radial direction, enables optimized blade utilization and loading without local stress peaks for a high-speed, low-pressure turbine of an aircraft gas turbine. The chosen deviation of the threading axis from the radial direction results in a more uniform load distribution, and the inventors have recognized that, in high-speed, low-pressure turbines, deviations from the radial direction (or a purely radial threading) must be strictly limited. Reference symbol list 10 Running shovel 12 Shovel foot element 14 Cover band element 16 Beam deflection section 18 Transition area 20 cross-sectional area 22 Outer contour 24 Focus 26 Threading axis 28 pins 30 cylinders 32 district β Opening angle cone φ Angle Cylinder axis AR axial direction DR Thickness Direction KH cone height RR Radial direction ZA cylinder axle ZH cylinder height

Claims

[1] Rotor blade for use in a high-speed low-pressure turbine of a gas turbine, comprising a blade root element (12) and a jet deflection section (16) which connects to the blade root element (12) in a radial direction (RR), wherein respective centers of gravity (24, 24') of blade cross-sectional areas (20) of the jet deflection section (16) lie on a common threading axis (26), wherein the threading axis (26) extends from a first center of gravity (24) of a first blade cross-sectional area (20) adjoining the blade foot element (12) within a cone (28) whose cone apex lies in the first center of gravity (24) and whose cone height (KH) extends orthogonally to the plane of the blade cross-sectional area (20), wherein the cone height (KH) in the installed state of the rotor blade (10) on the gas turbine extends in the radial direction (RR) of the gas turbine, and wherein the threading axis (26) in the installed state of the rotor blade (10) on the gas turbine is inclined with respect to the radial direction (RR) in the gas flow direction in the axial direction (AR) and / or in the circumferential direction of the gas turbine, characterized by , that the opening angle (β) of the cone (28) is greater than 0° and less than or equal to 4°. [2] Guide vane according to claim 1, characterized by that the gas turbine is an aircraft gas turbine. [3] Guide vane according to any one of the preceding claims, characterized by , that the threading axis (26) extends linearly within the cone (28) starting from the first center of gravity (24). [4] Guide vane according to claim 2 or 3, characterized by , that the opening angle (β) of the cone (28) is greater than or equal to 0.5° and less than or equal to 2°. [5] Guide vane according to any one of the preceding claims, characterized by, that the centers of gravity of cross-sectional areas of a cover band element (14) are arranged on the threading axis (26). [6] Gas turbine, in particular aviation gas turbine, comprising at least one turbine stage with a rotor on which several adjacent running blades (10) are attached in the circumferential direction according to one of the preceding claims. [7] Gas turbine according to claim 6, wherein the at least one turbine stage is part of a low-pressure turbine. [8] Gas turbine according to claim 7, characterized by , that the low-pressure turbine is coupled to the fan of the aircraft gas turbine by means of a reduction gearbox. [9] Gas turbine according to any one of claims 6 to 8, characterized by , that the respective threading axes (26) of the running blades (10) are inclined equally to a respective straight line running radially through the first center of gravity (24) of the first blade cross-sectional area (20).

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