Patent Grant
8534993
1. Technical Field
The disclosure generally relates to gas turbine engines.
2. Description of the Related Art
A typical gas turbine engine incorporates a compressor section and a turbine section, each of which includes rotatable blades and stationary vanes. Within a surrounding engine casing, the radial outermost tips of the blades are positioned in close proximity to outer air seals. Outer air seals are parts of shroud assemblies mounted within the engine casing. Each outer air seal typically incorporates multiple segments that are annularly arranged within the engine casing, with the inner diameter surfaces of the segments being located closest to the blade tips.
Gas turbine engines and related systems involving blade outer air seals are provided. In this regard, an exemplary embodiment of a blade outer air seal assembly for a gas turbine engine, the engine having a longitudinal axis and rotatable blades, each of the blades having a blade tip, the blade outer air seal assembly comprising: an annular arrangement of outer air seal segments, each of the segments having ends, the segments being positioned in an end-to-end orientation such that each adjacent pair of the segments forms an intersegment gap therebetween, each intersegment gap being angularly offset with respect to a longitudinal axis of the gas turbine engine.
An exemplary embodiment of a gas turbine engine comprises: a compressor; a combustion section; a turbine operative to drive the compressor responsive to energy imparted thereto by the combustion section, the turbine having a rotatable set of blades, the compressor and the turbine being oriented along a longitudinal axis; and a blade outer air seal assembly positioned radially outboard of the blades, the outer air seal assembly having an annular arrangement of outer air seal segments with intersegment gaps being located between the segments, each intersegment gap being angularly offset with respect to the longitudinal axis.
An exemplary embodiment of a blade outer air seal segment for a set of rotatable blades comprises: a blade arrival end; and a blade departure end; each of the blade arrival end and the blade departure end being angularly offset with respect to a longitudinal axis about which the blades rotate.
Other systems, methods, features and/or advantages of this disclosure will be or may become apparent to one with skill in the art upon examination of the following drawings and detailed description. It is intended that all such additional systems, methods, features and/or advantages be included within this description and be within the scope of the present disclosure.
Many aspects of the disclosure can be better understood with reference to the following drawings. The components in the drawings are not necessarily to scale. Moreover, in the drawings, like reference numerals designate corresponding parts throughout the several views.
Gas turbine engines and related systems involving blade outer air seals are provided, several exemplary embodiments of which will be described in detail. In some embodiments, the ends of the outer air seal segments are angularly offset with respect to a longitudinal axis of the gas turbine in which the segments are mounted. In some of these embodiments, the ends of two adjacent segments are shaped to correspond to the mean camber line of the blades at the blade tips. In this manner, a pressure differential between the suction side and the pressure side of a blade as that blade crosses the adjacent ends of the segments tends to be stabilized. In particular, the location of the highest pressure differential during blade passage may tend to wander less along the gap formed between the adjacent segments and/or the rate of hot gas ingestion into the gap may be reduced. Notably, stabilizing of the transient nature of the pressure differential as each blade crosses the gap may allow for a decrease in overall cooling air applied to cool the segments. This may be the case because the region of highest hot gas ingestion along a segment, which corresponds to at least one of a highest temperature of hot gas and a highest volume of hot gas, may be relatively stationary. Thus, increased cooling air can be specifically directed to those regions and less cooling air can be directed to others.
Referring now in more detail to the drawings,
A portion of engine 100 is depicted in greater detail in the schematic diagram of
As shown in
Attachment of the outer air seal to the mounting ring in the embodiment of
With respect to the annular configuration of the outer air seal, outer air seal 125 is formed of multiple arcuate segments, portions of two of which are depicted schematically in
Portions defining the intersegment gap include a blade departure end 152 of segment 140 and a blade arrival end 154 of segment 142. As shown in
The aforementioned configuration may tend to reduce hot gas ingestion and corresponding distress exhibited by the ends of the segments. Notably, the advancing suction side of each rotating blade (e.g., side 170 of blade 112) tends to promote a radial inboard-directed flow of cooling air (depicted by the solid arrow) from the intersegment gap. In contrast, the retreating pressure side of each rotating blade (e.g., side 172 of blade 112) tends to promote a radial outboard-directed ingestion flow of hot gas (depicted by the dashed arrow) into the intersegment gap. By providing an angular offset of the intersegment gap, as defined by the ends of the outer air seal segments, radial penetration of hot gas along the intersegment gap may be reduced. This characteristic may be attributable to a reduction in the length of the intersegment gap over which the instantaneous axial pressure gradient occurs.
In other embodiments, various angular offsets other than those directly corresponding to the blade chord can be used. By way of example, angular offsets of between approximately 5° and approximately 70°, preferably between approximately 20° and approximately 60°, and most preferably between approximately 30° and approximately 45°, can be used. Notably, passage of an intersegment gap by the leading and trailing edges of a blade may occur separately in some embodiments.
Another aspect of the embodiment of
In contrast to the embodiment of
In this embodiment, blade passage region 230 of the gap exhibits a shape that generally corresponds to the mean camber line of the blade at the blade tip (i.e., a line defined by points equidistant from the suction side and pressure side surfaces of the blade tip). The leading and trailing edge regions, which are axially located fore and aft, respectively, of the blade passage region, continue the curvature of the blade passage region. In other embodiments, various other types of curvature can be used for forming an intersegment gap. By way of example, an intermediate portion of the gap (e.g., that portion of the gap located adjacent to the blade tips) can exhibit a shape that generally corresponds to the mean camber line of the blades, while the portions of the gap in the vicinity of the leading and trailing edges can be oriented generally axially. Such a shape may tend to reduce hot gas ingestion, particularly at the leading edge of the gap as the gap shape would not match the airflow direction coming off of the tips of the passing blades.
It should be noted that the angular offset of blade arrival end 154 of segment 142 is depicted in
It should be emphasized that the above-described embodiments are merely possible examples of implementations set forth for a clear understanding of the principles of this disclosure. Many variations and modifications may be made to the above-described embodiments without departing substantially from the spirit and principles of the disclosure. All such modifications and variations are intended to be included herein within the scope of this disclosure and protected by the accompanying claims.
The U.S. Government may have an interest in the subject matter of this disclosure as provided for by the terms of contract number N00019-02-C-3003, awarded by the United States Navy, and contract number F33615-03-D-2345 DO-0009, awarded by the United States Air Force.
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| Number | Date | Country | |
|---|---|---|---|
| 20130213057 A1 | Aug 2013 | US |
