The invention relates to a compressor, in particular to a high pressure compressor, of a gas turbine, in particular of a gas turbine aircraft engine, in accordance with one aspect.
Gas turbine aircraft engines are equipped with at least one compressor, at least one combustion chamber as well as at least one turbine. From the praxis, gas turbine aircraft engines have been known with two compressors, namely one low pressure compressor and one high pressure compressor, as well as two turbines, namely one high pressure turbine and one low pressure turbine.
Moreover, gas turbine aircraft engines have been known with three compressors, namely one low pressure compressor, one intermediate pressure compressor and one high pressure compressor, as well as three turbines, namely one high pressure turbine, one intermediate pressure turbine and one low pressure turbine.
A compressor of a gas turbine aircraft engine, for example the high pressure compressor, has several stages, with each stage being formed by a rotor-side blade ring and a stator-side guide vane ring. The stator-side guide vane rings are formed by several guide vane segments, with each guide vane segment being composed of several individual blades.
In the case of compressors known from the praxis, all individual blades of a guide vane segment are permanently connected to each other radially on the outside as well as radially on the inside through soldering so that the respective guide vane has an overall rigid design. In the praxis it has shown that vibration tears can form in the guide vane segments of such guide vane rings.
Starting from this, the invention at hand is based on the problem of creating a novel compressor of a gas turbine.
This problem is solved by a compressor in the sense of claim 1.
In accordance with the invention, within each guide vane segment of at least one guide vane ring, adjacent individual blades are permanently attached to each other at opposite surfaces located radially on the exterior while not being connected to each other at opposite surfaces located radially on the interior.
In accordance with the invention at hand, it is suggested to permanently connect the adjacent individual blades of each guide vane segment exclusively at opposite surfaces located radially on the exterior within the area of at least one guide vane ring of a compressor of a gas turbine aircraft engine but to leave them unconnected at opposite surfaces located radially on the interior. By way of this, each guide vane segment will retain a certain flexibility, thereby minimizing the danger of vibration tears forming in the guide vane segments. The life of the guide vane segments can be extended thereby. This will ultimately lead to cost reductions.
Preferred continued developments of the invention will result from the subclaims and from the subsequent description. Embodiments will be explained in detail by way of the drawings without being limited thereto.
In accordance with the invention, within the guide vane segment 10, the adjacent individual blades 11, 12, 13 and 14 are permanently attached to each other at opposite surfaces located radially on the exterior while not being connected to each other at opposite surfaces located radially on the interior. The adjacent individual blades 11 and 12, 12 and 13 as well as 13 and 14 are thus permanently connected to each other in the area of the exterior shrouding band 16 at opposite surfaces, but in the area of the interior shrouding band 15 they are unconnected at the surfaces located opposite each other. The connection of the individual blades 11, 12, 13 and 14 in the area of the exterior shrouding band 16 of the same preferably occurs through soldering.
As can be seen from
Due to the fact that within the guide vane segment 10 the individual blades 11, 12, 13 and 14 are unconnected in the area of the interior shrouding band 15, the guide vane segment 10 retains a certain flexibility, thereby minimizing the danger of vibration tears forming within the guide vane segment 10.
Number | Date | Country | Kind |
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10 2007 009 134.8 | Feb 2007 | DE | national |
This application is a U.S. National Phase application submitted under 35 U.S.C. §371 of Patent Cooperation Treaty application serial no. PCT/DE2008/000277, filed Feb. 14, 2008, and entitled COMPRESSOR OF A GAS TURBINE, which application claims priority to German patent application serial no. DE 10 2007 009 134.8, filed Feb. 24, 2007, and entitled VERDICHTER EINER GASTURBINE, the specifications of which are incorporated herein by reference in their entireties.
Filing Document | Filing Date | Country | Kind | 371c Date |
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PCT/DE2008/000277 | 2/14/2008 | WO | 00 | 8/24/2009 |