GUIDE SHAFTER SEGMENT WITH SUPPORT SECTION RIBBON

Description

[0001] The present invention relates to a guide vane segment for a gas turbine, in particular an aircraft gas turbine, with at least three guide vanes, a radially inner shroud and a radially outer shroud, wherein the at least one guide vane extends between the inner shroud and the outer shroud, and wherein the inner shroud and / or the outer shroud has or have a respective support section extending along the circumferential direction with a respective axial end face.

[0002] Guide vanes with projections on the radially outer deck strip are known from document WO2007 / 093851A2.

[0003] US Patent 2,605,997 discloses a fastening for guide vanes for axial compressors and axial turbines, wherein individual guide vanes have radial inner and outer projections which are to be inserted into corresponding recesses or grooves of an inner and an outer ring in order to fasten the individual guide vanes in the radial direction and in the circumferential direction.

[0004] 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.

[0005] The area of ​​the guide bands is also commonly used to clamp a guide vane segment in a machining device in order to machine the guide vanes as desired, for example, to grind them. The forces exerted by the clamping and the machining process are absorbed via the support section or the end face supported in the machining device.

[0006] Due to the forces acting during the machining of a guide vane segment, undesirable cracks can occur in the support section because very high compressive forces act on the support section.

[0007] The object underlying the invention is seen as being to provide a guide vane segment in which cracking can be avoided through machining.

[0008] This problem is solved by a guide vane segment with the features of claim 1 and a gas turbine with the features of claim 9. Advantageous embodiments with expedient further developments are specified in the dependent claims.

[0009] A guide vane segment for a gas turbine, particularly an aircraft gas turbine, is proposed, comprising at least three guide vanes, a radially inner shroud, and a radially outer shroud, wherein the at least one guide vane extends between the inner and outer shrouds, and wherein the inner and / or outer shrouds each have a support section extending along the circumferential direction with a respective axial end face. It is provided that the support section on the respective inner and / or outer shroud has at least one radially extending support section rib, which is connected to the support section, in particular is integrally formed as a single piece, and which has a rib end face merging into the end face of the support section.

[0010] The support section rib and its end face allow for better distribution and absorption of the forces occurring during machining of the guide vane segment, thus counteracting crack formation. Furthermore, the support section rib also stiffens the support section, which is typically rather thin in the radial direction.

[0011] The axial end face, which adjoins the (respective) rib end face and which may be arranged axially at the front, axially at the rear, radially outside or radially inside, may have a maximum radial extent of 5mm, in particular 2mm.

[0012] The rib end face and the end face of the support section can be designed to be flat against each other. This results in a flat and consistently enlarged support surface formed by the end face and the rib end face. By increasing the total available area for force transmission, a better force distribution within the support section is achieved, thus counteracting crack formation.

[0013] The rib end face and the end face on the radially inner cover strip and / or the radially outer cover strip, together with a support surface on the other of the inner and outer cover strips (which support surface may comprise the other end face or rib end face and end face), form a support surface of the guide vane segment lying within a common plane. This support surface allows the guide vane segment to be supported axially on a flat surface. In other words, the end face(s) and the rib end face(s) form bearing and / or contact surfaces by means of which the guide vane segment, particularly when clamped in a machining device for machining, can be supported on a flat surface that runs parallel or substantially parallel to the circumferential and radial directions, especially with surface contact between the support surface and the flat bearing surface.The support surface therefore preferably forms an axial interface of the guide vane segment.

[0014] The guide vane segment can have more than three guide vanes.

[0015] In a further development, the guide vane segment includes both such an axially forward (upstream) and such an axially rear (downstream) support surface, both of which include one or more rib end faces on the inner shroud and / or on the radially outer shroud.

[0016] The end face of a support section rib can correspond to approximately 2% to 15%, preferably approximately 6% to 12%, of the end face of the support section. In other words, the available support area is increased by approximately 2% to 15% by providing a support section rib.

[0017] The at least one support section rib can be located at the center of the support section with respect to its circumferential extent. Alternatively, the guide vane segment can have two or more support section ribs distributed along the support section, particularly in a regularly spaced manner. The arrangement of support section ribs can be tailored to the design of the guide vane segment, especially its circumferential extent and the dimensions of the support section. This ensures that the additional rib end face is provided at suitable locations to distribute the acting forces as effectively as possible.

[0018] The at least one support section rib, in particular all support section ribs, can be formed on the side of the respective cover band that faces away from a gas channel. This ensures that the additional support section rib(s) does not impair the flow in a gas turbine.

[0019] The above problem is also solved by a gas turbine, in particular an aviation gas turbine, with at least one guide vane ring, located on the compressor or turbine side, and formed from several guide vane segments arranged side by side in the circumferential direction, as described above. With regard to the gas turbine having a guide vane ring with such guide vane segments, it should be noted that the support section ribs have no role or function during the operation of a gas turbine. The weight increase of a guide vane segment caused by the support section ribs can be considered negligible, especially considering that cracking and thus costly repairs or replacement of guide vane segments can be avoided.

[0020] The invention is described below by way of example and without limitation, with reference to the accompanying figures. Fig. 1 A simplified schematic representation of a basic diagram of an aircraft gas turbine. Fig. 2 a known guide vane segment with support section on the radially inner cover band in simplified and schematic front and side view; Fig. 3 An example of a guide vane segment with support section and support section rib in simplified and schematic front and side view;

[0021] Fig. 1Figure 10 schematically and in a simplified manner shows an aircraft gas turbine 10, illustrated purely as an example of a turbofan engine. The gas turbine 10 comprises a fan 12, which is surrounded by a suggested casing 14. A compressor 16, housed in a suggested inner casing 18 and capable of being single-stage or multi-stage, is connected to the fan 12 in the axial direction AR of the gas turbine 10. The combustion chamber 20 is connected to the compressor 16. Hot exhaust gas flowing from the combustion chamber then passes through the subsequent turbine 22, which can be single-stage or multi-stage. In this example, the turbine 22 comprises a high-pressure turbine 24 and a low-pressure turbine 26. A hollow shaft 28 connects the high-pressure turbine 24 to the compressor 16, in particular a high-pressure compressor 29, so that they are driven or rotated together.An internal shaft 30, located radially RR of the turbine, connects the low-pressure turbine 26 to the fan 12 and to a low-pressure compressor 32, so that these are driven and rotated together. A thrust nozzle 33, shown here only, is connected to the turbine 22.

[0022] In the illustrated example of an aircraft gas turbine 10, a turbine intermediate casing 34 is arranged between the high-pressure turbine 24 and the low-pressure turbine 26, and is positioned around the shafts 28, 30. Hot exhaust gases from the high-pressure turbine 24 flow through the turbine intermediate casing 34 in its radially outer region 36. The hot exhaust gas then enters an annular space 38 of the low-pressure turbine 26. The rotor blade assemblies 27 of the compressors 28, 32 and the turbines 24, 26 are shown as examples. For clarity, the guide vane assemblies 31, which are typically present, are shown only as examples for the compressor 32.

[0023] The following description of an embodiment of the invention relates in particular to guide vane segments that form a guide vane ring for the turbines 24, 26 or the compressors 28, 32.

[0024] Fig. 2 Figure 1 shows a simplified and schematic perspective view of a known guide vane segment 40. The guide vane segment 40 has a radially outer cover band 42 and a radially inner cover band 44. In the example shown, three guide vanes 46 extend between the cover bands 42 and 44. The radially inner cover band 44 has a support section 48 extending along the circumferential direction UR. The support section 48 projects in the axial direction AR beyond an edge section 50 of the inner cover band 44. An end face 52 of the support section 48 is typically oriented substantially orthogonal to the axial direction AR and extends in a plane spanned by the circumferential direction UR and the radial direction RR.

[0025] For the sake of completeness, it should also be noted that radially inside the inner cover band 44, fastening sections are visible, which serve to connect the guide vane segment 40 to other structural components of the gas turbine, not shown. Radially outside, fastening structures are also shown on the outer cover band 42, which serve to secure the guide vane segment 40 within the gas turbine. Furthermore, respective cooling lines are shown on both cover bands 42 and 44.

[0026] If a guide vane segment 40 is machined using a machining device not shown here, for example by grinding, the guide vane segment 40 can be clamped in the machining device in the area of ​​the inner cover band 44. The support section 48 rests with its end face 52 against a corresponding counterpart of the machining device and is subjected to pressure for clamping.

[0027] With simultaneous reference to the various representations of the Fig. 3 An example of a guide vane segment 40 according to the invention is described below. The guide vane segment 40 essentially has the features described above with reference to the Fig. 2 The described structure is therefore... Fig. 3 also the same reference symbols as in the Fig. 2 They are shown, but without describing them again.

[0028] In the guide vane segment 40, a support section rib 62 is formed on the support section 48. The support section rib 62 is connected to the support section 48, in particular integrally or as a single piece. The support section rib 62 has a rib end face 64, which also extends substantially orthogonally to the axial direction AR and lies in a plane spanned by the circumferential direction UR and the radial direction RR.

[0029] The rib end face 64 and the end face 52 of the support section 48 together form a continuous or entire support surface. By providing at least one support section rib 62, the total available support area can be increased. The additional rib end face provided by the support section rib 62 has an area that corresponds to approximately 2% to 15%, preferably approximately 6% to 12%, of the area of ​​the end face 50 of the support section 48. Assuming a known guide vane segment 40 according to Fig. 2 In a case where the end face 52 of the support section has an area corresponding to 100%, the entire support surface formed by the end face 52 and the rib end face 64 has an area corresponding to 1.02 to 1.15 times the area of ​​the end face 52.

[0030] As can be seen from the depictions of the Fig. 3As can be seen, the end face 50 and the rib end face 64 can be designed to be planar relative to each other, i.e., they lie in the same plane. By increasing the bearing surface, the forces acting on the guide vane segment 40 can be distributed more effectively when it is clamped into the machining device, thus preventing cracking in the support section 48.

[0031] Regarding the Fig. 3In the upper drawing, which depicts the guide vane segment 40 in the RR-UR plane, it should be noted that the support section rib 62 can be arranged circumferentially at the midpoint of the length of the support section 48. The provision of one support section rib 62 is merely an example, and two or more support section ribs can also be provided. For illustrative purposes only, two further support section ribs 62a are indicated by dashed lines in Fig. 5. A guide vane element 40 could thus have only the support section rib 62, only the support section ribs 62a, or all three support section ribs 62, 62a. Naturally, a guide vane segment that extends over a larger area in the circumferential direction UR and may also have more guide vanes can also have more support section ribs. Accordingly, the total available support area can be further increased by providing several support section ribs 62, 62a. Reference symbol list

[0032] 10 Aircraft gas turbine 12 Fan 14 Casing 16 Compressor 18 Inner casing 20 Combustion chamber 22 Turbine 24 High-pressure turbine 26 Low-pressure turbine 28 Hollow shaft 29 High-pressure compressor 30 Shaft 31 Guide vane ring 32 Low-pressure compressor 33 Thrust nozzle 34 Turbine intermediate casing 36 Radial outer area 38 Annular space 40 Guide vane segment 42 Radial outer shroud 44 Radial inner shroud 46 Guide vane 48 Support section 50 Edge section 52 End face 62, 62a Support section rib 64 Rib end face AR Axial direction RR Radial direction U Circumferential direction

Claims

1. Guide vane segment (40) for a gas turbine (10), in particular an aircraft gas turbine, comprising at least three guide vanes (46), a radially inner shroud (44), and a radially outer shroud (42), the at least one guide vane (46) extending between the inner shroud (44) and the outer shroud (42), and the inner shroud (44) and / or the outer shroud (42) having a relevant support portion (48) extending along a circumferential direction (UR) and having a relevant axial end face (52), the support portion (48) having at least one support portion rib (62) extending in a radial direction (RR) on the relevant inner shroud (44) and / or on the relevant outer shroud (42), which rib is connected to the support portion (48), in particular is integrally formed as a single piece, and which has a rib end face (64) which transitions into the end face (52) of the support portion (48), characterized in that the rib end face (64) and the end face (52) on the inner shroud (44) and / or on the outer shroud (42), with a support surface on the other of the inner and outer shroud (42), form a support surface of the guide vane segment which lies within a common plane and by means of which the guide vane segment can be supported in an axial direction on a planar surface.

2. Guide vane segment according to claim 1, characterized in that the axial end face (52) which adjoins the rib end face (64) has a maximum radial extension of 5 mm, in particular 2 mm.

3. Guide vane segment according to claim 1 or claim 2, characterized in that the rib end face (64) and the end face (52) of the support portion (48) are planar relative to each other.

4. Guide vane segment according to any of the preceding claims, characterized in that the rib end face (64) of a support portion rib (62) corresponds to approximately 2% to 15%, preferably approximately 6% to 12%, of the end face (52) of the support portion (48).

5. Guide vane segment according to any of the preceding claims, characterized in that the guide vane segment has more than three guide vanes (46).

6. Guide vane segment according to any of the preceding claims, characterized in that the at least one support portion rib (62) is formed in the center of the support portion (48) with respect to the extent of the support portion (48) in the circumferential direction (UR).

7. Guide vane segment according to any of the preceding claims, characterized in that it has two or more support portion ribs (62, 62a) which are distributed along the support portion (48), in particular regularly distributed.

8. Guide vane segment according to any of the preceding claims, characterized in that the at least one support portion rib (62), in particular all support portion ribs (62, 62a), are formed on the side of the relevant shroud (42, 44) that faces away from a gas duct.

9. Gas turbine, in particular an aircraft gas turbine (10), comprising at least one guide vane ring which is provided on the compressor side or on the turbine side, and which is formed from a plurality of guide vane segments (40) according to any of the preceding claims, which are arranged next to one another in the circumferential direction (UR).