Gas turbine with transition duct

p0001The invention relates to a gas turbine, in particular an aircraft engine, according to the preamble of claim 1.

p0002Gas turbines, such as aircraft engines, have usually over several compressors, more turbines and a combustion chamber. For several compressors are usually at a low pressure compressor or a medium-pressure compressor and a high pressure compressor, wherein the plurality of turbines to a high pressure turbine and low pressure turbine. The gas turbine is flowed through in the axial direction, said low pressure compressor or medium-pressure compressor upstream of the high-pressure compressor and the high pressure turbine upstream of the low pressure turbine is positioned. From the low pressure compressor or medium-pressure compressor, the flow passes into the high-pressure compressor by means of a transition channel between the two compressors. Likewise, such a transition duct is positioned between the high pressure turbine and the low pressure turbine. Another transition duct located in the turbine exhaust case downstream of a low-pressure turbine.

p0003From the prior art it is already known to arrange spaced-apart support ribs in such a transition channels in the circumferential direction of the transition duct. The support ribs serve to carry out, for example, oil lines and sensors as well as the absorption of forces, which is why the support ribs are designed to be relatively thick. Supporting ribs are known from the prior art which, although flow leader but not formed strömungsumlenkend. Further support ribs are known, which have a suction side and a pressure side and assumes the function of a flow diversion. The thickness of the support ribs may be approximately 30% of the chord length of the supporting ribs. Due to the large relative thickness of such support ribs occur at the flow outlet edge of the support ribs on large wedge angle, which cause high flow delays, particularly in the region of the transition duct radially inner limiting duct wall and / or the transition duct radially outer limiting duct wall. This gives rise to secondary flows, flow separation and thus flow losses. Furthermore, a flow downstream of the support ribs positioned gas turbine blades disturbed. A similar problem Tritz particular even when cooled high pressure turbine blades having a relatively large wedge angle at the flow outlet edge.

p0004From the documents <patcit id="pcit0001" dnum="US2292673A"><text>US 2292673 A</text></patcit>. <patcit id="pcit0002" dnum="GB2257736A"><text>GB 2257736 A</text></patcit>. <patcit id="pcit0003" dnum="US3104525A"><text>US 3104525 A</text></patcit> and <patcit id="pcit0004" dnum="US2735612A"><text>US 2735612 A</text></patcit> are known with support ribs for gas turbine transition ducts, at least one duct wall in the region of the flow trailing edges of the support ribs eingezogenist inwardly into the channel.

p0005Proceeding from this, the present invention addresses the problem to provide a gas turbine having a further improved transition duct.

p0006This problem is solved by a gas turbine according to Patent Claim first Here, / is a transition channel radially inwardly delimits channel wall and / or the transition duct radially outwardly retracted limiting duct wall in the region of the flow trailing edges of the support ribs inwardly into the transition duct into. Erfindungegemäß is the retracting range between 70% to 12D% of the chord length of the support ribs.

p0007Hereby the invention, the harmful influence of thickness of the supporting ribs with a large wedge angle can be reduced at the flow outlet edge on the flow within the meaning. Another advantage of the invention is that an improved Anströmqualität for downstream of the support ribs positioned blade rows may be achieved. Furthermore, the overall length of the support ribs, as well as of the transition duct can be shortened. Finds the flow structure of the invention in the field of turbines used, so the noise of the gas turbine can be reduced, since setting a lesser degree of interaction between the support ribs or the turbine blades and the downstream blade row positioned.

p0008the inner channel wall and / or outer duct wall between two adjacent support ribs in the region of the flow trailing edges are preferably / is the same to the side walls of the support ribs fed exclusively adjacent to the inside in the transition passage into it, are in the range of half the pitch between two adjacent supporting ribs / is the inner channel wall and / or the outer channel wall, however not feed.

p0009the inner channel wall and / or the outer duct wall According to an advantageous development of the invention are / is curved outwards adjacent to the flow inlet edge of the support ribs outward from the transition duct out.

p0010Preferred embodiments of the invention result from the dependent claims and the following description. Embodiments of the invention are not limited thereto, are explained in greater detail using the drawing. In which:<dl id="dl0001" compact="compact"><dt>Fig. 1</dt><dd>a highly schematic representation of a gas turbine according to the invention with a transition channel between two compressor stages;</dd><dt>FIG. 2</dt><dd>a cross-section through the gas turbine of the invention <figref idrefs="f0001">Fig. 1</figref> in the region of a hub; and</dd><dt>Fig. 3</dt><dd>a cross-section through the gas turbine of the invention <figref idrefs="f0001">Fig. 1 and 2</figref> in cutting direction AA.</dd></dl>

p0011<figref idrefs="f0001">FIGS. 1-3</figref> show highly schematically a preferred embodiment of a gas turbine according to the invention with a transition duct 10 between a medium-pressure compressor 11 and a low pressure compressor and a high pressure compressor 12 of an aircraft engine, wherein a gas flow through the transition duct 10 from the medium-pressure compressor 11 and low-pressure compressor is fed into the area of ​​the high-pressure compressor 12th <figref idrefs="f0001">Fig. 1</figref> shows a highly schematic, that the medium-pressure compressor 11 is completed in the area of ​​its seen in the flow direction last compressor stage vcn a rotor blade ring. 13 The blade ring 13 is formed of several veneinander circumferentially spaced blades 14th

p0012As especially <figref idrefs="f0001">FIG. 2</figref> It can be seen 10 more 10 spaced circumferentially of the transition duct support ribs 15 are arranged in the region of the transition duct. The support ribs 15 are relatively thick and relatively long and have a small aspect ratio. The Szützrippen 15 have in the region of the flow inlet edge 16 and its flow outlet edge 17 over relatively large wedge angle. The supporting ribs 15 are formed as a flow-conducting but as a non-strömungsumlenkende supporting ribs so that their side walls 18 only have a flow guiding function.

p0013The transition duct 10 is primarily limited by two channel walls, namely a radially inner duct wall 19 and a radially au-Benliegenden channel wall 20. The channel walls 19 and 20 may, in particular <figref idrefs="f0001">Fig. 1 and 3</figref> be removed. <figref idrefs="f0001">Fig. 1 and 3</figref> show the channel walls 19 and 20 on the one hand in a solid line and the other in dashed lines. In the solid lines is known from the prior art contour of the channel walls 19 and 20, wherein the dashed representation on the other hand is the inventive design of the channel walls 19 and 20th

p0014In the preferred embodiment of the <figref idrefs="f0001">FIGS. 1-3</figref> are the transition duct 10 radially inwardly limiting duct wall 19 and 17 of the support ribs 15 pulled the transition duct 10 radially outwardly bounding the duct wall 20 in the region of the flow trailing edge inwardly into the transition channel 10 in, and indeed as <figref idrefs="f0001">Fig. 3</figref> It can be seen exclusively adjacent to the side walls 18 of the support ribs 15. In the range of half the pitch between two adjacent support ribs 15, however, (see <figref idrefs="f0001">Fig. 3</figref>), The inner channel wall 19 and the outer duct wall 20 is not fed to the inside, but rather extend in the range known from the prior art contour. It is therefore in the sense of the present invention, 19 and 20 feed the same, the channel walls in the region of the flow trailing edge 17 of the support ribs 15 only adjacent to the side walls 18 inwardly into the transition duct 10 into it.

p0015As <figref idrefs="f0001">Fig. 1</figref> can be removed, the inner channel wall 19 and the outer channel wall according to the invention are 20 retracted in a range between 70% and 120%, in particular in a range between 80% and 110%, of the chord length of the support ribs 15th So can<figref idrefs="f0001">Fig. 1</figref> be seen that the region in which the channel walls 19 and 20 are drawn into the transition duct 10 into it, also downstream of the flow trailing edge 17 of the supporting ribs 15 extends.

p0016As <figref idrefs="f0001">Fig. 1</figref> still can be seen, differs in terms of the present invention, the contour of the channel walls 19 and 20 not only in the field of flow trailing edge 17 of the support ribs 15 of the contouring known from the prior art from, but also in the region of the flow inlet edge 16 thereof. So can<figref idrefs="f0001">Fig. 1</figref> be seen that the inner channel wall 19 and the outer duct wall 20 are bulged in the region of the flow leading edge 16 of the support ribs 15 to the outside of the transition duct 10 out. The area in which the contour according to the invention by inner duct wall 19 and outer duct wall 20 relative to the contouring known from the prior art out arched the same to the outside of the transition duct 10 extends to about 50% of the chord length of the support ribs 15th

p0017Although with reference to <figref idrefs="f0001">FIGS. 1-3</figref> the present invention has been described herein for non-deflecting supporting ribs 15, it should be noted that the invention can also be used in deflecting support ribs.

p0018The inventive solution found in transitional passages use, preferably in a transition channel between two compressors or between two turbines. The inventive solution may be positioned in a transition channel between a medium-pressure compressor and a low-pressure compressor and a high pressure compressor. Furthermore, the solution according to the invention can be arranged in a transition channel between a high pressure turbine and a low pressure turbine. The inventive solution can be positioned in a transition duct of a turbine exhaust housing downstream of a low-pressure turbine.