Gas turbine with transition duct

The invention relates to a flow structure for a gas turbine, in particular for an aircraft engine, according to the preamble of patent claim 1 and 12 respectively.

Gas turbines, such as aircraft engines, have usually over several compressors, more turbines and a combustion chamber. In the multi-compressor is usually 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 low-pressure compressor or intermediate-pressure compressor, the flow enters the high-pressure compressor by 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 ..

From 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 occurs in particular even when cooled high pressure turbine blades having a relatively large wedge angle at the flow outlet edge.

Proceeding from this, the present invention addresses the problem of creating a new flow structure for a gas turbine.

This problem is solved by a flow structure for a gas turbine according to Patent Claim first According to the invention are / is drawn inwards in the transition duct into a transition channel radially inner limiting duct wall and / or the transition duct radially outer limiting duct wall in the flow trailing edge of the support ribs.

In terms of the present invention accordingly namely the transition duct radially inner limiting duct wall and / or the transition duct radially outer limiting duct wall, adjacent to the region of the flow trailing edge of the support ribs on the side walls of the channel wall of the transition duct, fed them inwardly into the transition duct into , The transition duct or flow channel is thus fed exclusively near the flow outlet edge of the relatively thick support ribs, resulting in an asymmetric contouring of the transition duct radially inner limiting duct wall and the transition duct radially outer limiting duct wall. Hereby 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.

the 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 duct 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.

the 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.

A flow structure of the invention for a flow channel of a high-pressure turbine is defined in the independent claim 12th

Preferred 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 flow structure of the invention for a gas turbine, namely a flow structure for a transition channel between two compressor stages;</dd><dt>FIG. 2</dt><dd>a cross-section through the inventive flow structure of Figure 1 in the region of a hub. and</dd><dt>Fig. 3</dt><dd>a cross-section through the inventive flow structure of FIG. 1 and 2 in the cutting direction AA.</dd></dl>

FIG. 1 to 3 show highly schematically a preferred embodiment of a flow structure of the invention for a gas turbine in a transition duct 10 between a medium-pressure compressor 11 and a low pressure compressor and a high pressure compressor 12 a formed as a gas turbine aircraft engine, wherein a gas flow or by the transition duct 10 from the medium-pressure compressor 11 . low-pressure compressor is fed into the area of ​​high pressure compressor 12.. FIG. 1 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 of a rotor blade ring. 13 The blade ring 13 is formed of a plurality of spaced apart in the circumferential direction blades 14th

In particular in FIG. 2 can be removed, disposed more 10 spaced circumferentially of the transition duct support ribs 15 in the region of the transition duct 10. The support ribs 15 are relatively thick and relatively long and have a small aspect ratio. The support ribs 15 are equipped 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.

The transition duct 10 is primarily limited by two channel walls, namely a radially inner duct wall 19 and a radially outer channel wall 20. The channel walls 19 and 20 of FIGS. 1 and 3 in particular can be removed. Fig. 1 and 3 show the channel walls 19 and 20 on the one hand in a solid line and the other in dashed lines. In the solid line guide is known from the prior art contour of the channel walls 19 and 20, wherein the dotted presentation, however, is the inventive embodiment of the channel walls 19 and 20 for the provision of the flow structure of the invention.

In the preferred embodiment of Figs. 1 to 3 which the transition duct 10 radially inwardly limiting duct wall 19 and the drawn in the transition duct 10 radially outwardly bounding the duct wall 20 in the region of the flow trailing edge 17 of the support ribs 15 inwardly into the transition channel 10 in, and indeed as shown in can be removed. 3, only 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 FIG. 3), the inner channel wall 19 and the outer duct wall 20 is not drawn inwards, but run rather in the 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.

As can be one taken from Fig., The inner channel wall 19 and the outer duct wall 20 are drafted in a range between 70% and 120%, in particular in a range between 80% and 110%, of the chord length of the support ribs 15. Thus, Fig. 1 it can 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.

As shown in Fig. 1 also can be removed, giving way in the sense of the present invention, the contour of the channel walls 19 and 20 not only in the region of the flow trailing edge 17 of the support ribs 15 of the contouring known from the prior art from, but rather also in the region of the flow inlet edge 16 thereof. Thus, Fig. 1 it can 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

Although the present invention, for non-deflecting supporting ribs 15 described with reference to FIGS. 1 to 3, it should be noted that the invention can also be used in deflecting support ribs.

Furthermore, the invention is not limited to the use in the field of positioned in transition channels supporting ribs, but rather, it can also be used to deflecting turbine blades with large wedge angles of the flow trailing edge. So it is in the sense of the present invention to utilize the flow structure of the invention in the field of high-pressure turbine blades with large wedge angles of the flow trailing edge.

Find the flow structure of the invention in transition channels use, so is the same preferably positioned in a transition channel between two compressors or between two turbines. Thus, the flow structure of the invention may be positioned in a transitional channel between a medium-pressure compressor and a low pressure compressor and a high pressure compressor. Furthermore, the flow structure of the invention can be arranged in a transition channel between a high pressure turbine and a low pressure turbine. Also, the flow structure of the invention may be positioned in a transition duct of a turbine exit housing downstream of a low pressure turbine.