Patent Description:
A gas turbine engine typically includes a compressor section, a turbine section, and a combustion section disposed therebetween. The compressor section includes multiple stages of rotating compressor blades and stationary compressor vanes. The combustion section typically includes a plurality of combustors. The turbine section includes multiple stages of rotating turbine blades and stationary turbine vanes. Turbine blades and turbine vanes often operate in a high temperature environment and are internally cooled. Gas turbines having these features are disclosed i. in <CIT>, <CIT>, <CIT>, <CIT>, <CIT> and <CIT>.

According to another aspect of the invention, there is provided a turbine blade as set forth in claim <NUM>.

Advantageous aspects of the invention are defined in the dependent claims.

Before any embodiments of the invention are explained in detail, it is to be understood that the invention is not limited in its application to the details of construction and the arrangement of components set forth in this description or illustrated in the following drawings, but is only limited by the literal language of the appended claims. The invention is capable of other embodiments and of being practiced or of being carried out in various ways, and is only limited by the literal language of the appended claims.

Various technologies that pertain to systems and methods will now be described with reference to the drawings, where like reference numerals represent like elements throughout. The drawings discussed below, and the various embodiments used to describe the principles of the present disclosure in this patent document are by way of illustration only and should not be construed in any way to limit the scope of the invention, which is only limited by the literal language of the appended claims. Those skilled in the art will understand that the principles of the present disclosure may be implemented in any suitably arranged apparatus. It is to be understood that functionality that is described as being carried out by certain system elements may be performed by multiple elements. Similarly, for instance, an element may be configured to perform functionality that is described as being carried out by multiple elements. The numerous innovative teachings of the present application will be described with reference to exemplary non-limiting embodiments.

Also, it should be understood that the words or phrases used herein should be construed broadly, unless expressly limited in some examples. For example, the terms "including", "having", and "comprising", as well as derivatives thereof, mean inclusion without limitation. Further, the term "and/or" as used herein refers to and encompasses any and all possible combinations of one or more of the associated listed items. The term "or" is inclusive, meaning and/or, unless the context clearly indicates otherwise. The phrases "associated with" and "associated therewith" as well as derivatives thereof, may mean to include, be included within, interconnect with, contain, be contained within, connect to or with, couple to or with, be communicable with, cooperate with, interleave, juxtapose, be proximate to, be bound to or with, have, have a property of, or the like. Furthermore, while multiple embodiments or constructions may be described herein, any features, methods, steps, components, etc. described with regard to one embodiment are equally applicable to other embodiments absent a specific statement to the contrary.

Also, although the terms "first", "second", "third" and so forth may be used herein to refer to various elements, information, functions, or acts, these elements, information, functions, or acts should not be limited by these terms. Rather these numeral adjectives are used to distinguish different elements, information, functions or acts from each other. For example, a first element, information, function, or act could be termed a second element, information, function, or act, and, similarly, a second element, information, function, or act could be termed a first element, information, function, or act, without departing from the scope of the present invention, which is only limited by the literal language of the appended claims.

Also, in the description, the terms "axial" or "axially" refer to a direction along a longitudinal axis of a gas turbine engine. The terms "radial" or "radially" refer to a direction perpendicular to the longitudinal axis of the gas turbine engine. The terms "downstream" or "aft" refer to a direction along a flow direction. The terms "upstream" or "forward" refer to a direction against the flow direction.

In addition, the term "adjacent to" may mean that an element is relatively near to but not in contact with a further element or that the element is in contact with the further portion, unless the context clearly indicates otherwise. Further, the phrase "based on" is intended to mean "based, at least in part, on" unless explicitly stated otherwise. Terms "about" or "substantially" or like terms are intended to cover variations in a value that are within normal industry manufacturing tolerances for that dimension. If no industry standard is available, a variation of twenty percent would fall within the meaning of these terms unless otherwise stated.

<FIG> illustrates an example of a gas turbine engine <NUM> including a compressor section <NUM>, a combustion section <NUM>, and a turbine section <NUM> arranged along a central axis <NUM>. The compressor section <NUM> includes a plurality of compressor stages <NUM> with each compressor stage <NUM> including a set of stationary compressor vane <NUM> or adjustable guide vanes and a set of rotating compressor blade <NUM>. A rotor <NUM> supports the rotating compressor blade <NUM> for rotation about the central axis <NUM> during operation. In some constructions, a single one-piece rotor <NUM> extends the length of the gas turbine engine <NUM> and is supported for rotation by a bearing at either end. In other constructions, the rotor <NUM> is assembled from several separate spools that are attached to one another or may include multiple disk sections that are attached via a bolt or plurality of bolts.

The compressor section <NUM> is in fluid communication with an inlet section <NUM> to allow the gas turbine engine <NUM> to draw atmospheric air into the compressor section <NUM>. During operation of the gas turbine engine <NUM>, the compressor section <NUM> draws in atmospheric air and compresses that air for delivery to the combustion section <NUM>. The illustrated compressor section <NUM> is an example of one compressor section <NUM> with other arrangements and designs being possible.

In the illustrated construction, the combustion section <NUM> includes a plurality of separate combustors <NUM> that each operate to mix a flow of fuel with the compressed air from the compressor section <NUM> and to combust that air-fuel mixture to produce a flow of high temperature, high pressure combustion gases or exhaust gas <NUM>. Of course, many other arrangements of the combustion section <NUM> are possible.

The turbine section <NUM> includes a plurality of turbine stages <NUM> with each turbine stage <NUM> including a number of stationary turbine vanes <NUM> and a number of rotating turbine blades <NUM>. The turbine stages <NUM> are arranged to receive the exhaust gas <NUM> from the combustion section <NUM> at a turbine inlet <NUM> and expand that gas to convert thermal and pressure energy into rotating or mechanical work. The turbine section <NUM> is connected to the compressor section <NUM> to drive the compressor section <NUM>. For gas turbine engines <NUM> used for power generation or as prime movers, the turbine section <NUM> is also connected to a generator, pump, or other device to be driven. As with the compressor section <NUM>, other designs and arrangements of the turbine section <NUM> are possible.

An exhaust portion <NUM> is positioned downstream of the turbine section <NUM> and is arranged to receive the expanded flow of exhaust gas <NUM> from the final turbine stage <NUM> in the turbine section <NUM>. The exhaust portion <NUM> is arranged to efficiently direct the exhaust gas <NUM> away from the turbine section <NUM> to assure efficient operation of the turbine section <NUM>. Many variations and design differences are possible in the exhaust portion <NUM>. As such, the illustrated exhaust portion <NUM> is but one example of those variations.

A control system <NUM> is coupled to the gas turbine engine <NUM> and operates to monitor various operating parameters and to control various operations of the gas turbine engine <NUM>. In preferred constructions the control system <NUM> is typically micro-processor based and includes memory devices and data storage devices for collecting, analyzing, and storing data. In addition, the control system <NUM> provides output data to various devices including monitors, printers, indicators, and the like that allow users to interface with the control system <NUM> to provide inputs or adjustments. In the example of a power generation system, a user may input a power output set point and the control system <NUM> may adjust the various control inputs to achieve that power output in an efficient manner.

The control system <NUM> can control various operating parameters including, but not limited to variable inlet guide vane positions, fuel flow rates and pressures, engine speed, valve positions, generator load, and generator excitation. Of course, other applications may have fewer or more controllable devices. The control system <NUM> also monitors various parameters to assure that the gas turbine engine <NUM> is operating properly. Some parameters that are monitored may include inlet air temperature, compressor outlet temperature and pressure, combustor outlet temperature, fuel flow rate, generator power output, bearing temperature, and the like. Many of these measurements are displayed for the user and are logged for later review should such a review be necessary.

<FIG> illustrates a perspective view of a turbine blade <NUM>. The turbine blade <NUM> or similar blades may be used in the gas turbine engine <NUM> as the rotating turbine blades <NUM>.

The turbine blade <NUM> has a blade platform <NUM>, a blade airfoil <NUM>, and a blade root <NUM>. The blade root <NUM> extends from a first side of the blade platform <NUM> toward the rotor <NUM> to engage the turbine blade <NUM> with the rotor <NUM>.

The blade airfoil <NUM> extends from a second side of the blade platform <NUM>, which is opposite to the first side, toward a blade tip <NUM>. The blade airfoil <NUM> has a pressure side wall <NUM> and a suction side wall <NUM> that join together at a blade leading edge <NUM> and a blade trailing edge <NUM> with respect to a flow direction of the working fluid <NUM>. A mean camber line <NUM> of the blade airfoil <NUM> is defined from the blade leading edge <NUM> to the blade trailing edge <NUM> passing through a midway points between the pressure side wall <NUM> and the suction side wall <NUM>. The blade airfoil <NUM> is exposed in a stream of working fluid <NUM>. The working fluid <NUM> may include the exhaust gas <NUM> from the combustor <NUM> shown in <FIG>.

<FIG> illustrates a portion of the perspective view of the turbine blade <NUM> shown in <FIG> that better illustrates the blade tip <NUM>. The blade airfoil <NUM> has a tip cap surface <NUM> which is a surface at an end of the blade airfoil <NUM> facing the blade tip <NUM>. The blade airfoil <NUM> has a first plurality of cooling holes <NUM> that are formed at the tip cap surface <NUM> and pass through the tip cap surface <NUM>. The first plurality of cooling holes <NUM> are in flow connection with an interior of the blade airfoil <NUM>. According to the invention, the blade airfoil <NUM> has an offset surface <NUM> that is offset a non-zero distance from the tip cap surface <NUM> toward the blade platform <NUM>. The offset surface <NUM> is disposed at a region that is closer to the blade leading edge <NUM> than the blade trailing edge <NUM>. The offset surface <NUM> may be parallel to the tip cap surface <NUM>. In other constructions that are not part of the invention, the blade airfoil <NUM> may not have the offset surface <NUM> such that the tip cap surface <NUM> extends from the blade leading edge <NUM> to the blade trailing edge <NUM> and extends between the pressure side wall <NUM> and the suction side wall <NUM> at the end of the blade airfoil <NUM> facing the blade tip <NUM>.

The blade tip <NUM> include a so-called "squealer tip". The squealer tip is defined by a squealer tip wall <NUM> that extends along a portion of the pressure side wall <NUM> and a portion of the suction side wall <NUM> from the tip cap surface <NUM> to the blade tip <NUM> and from blade leading edge <NUM> toward the blade trailing edge <NUM>. The squealer tip wall <NUM> includes a pressure side squealer tip wall <NUM> and a suction side squealer tip wall <NUM>. The pressure side squealer tip wall <NUM> extends along a portion of the pressure side wall <NUM>. The suction side squealer tip wall <NUM> extends along a portion of the suction side wall <NUM>. In the construction illustrated in <FIG>, the pressure side squealer tip wall <NUM> extends along the pressure side wall <NUM> from the blade leading edge <NUM> to a location before the blade trailing edge <NUM>. The suction side squealer tip wall <NUM> extends along the suction side wall <NUM> from the blade leading edge <NUM> to the blade trailing edge <NUM>. In other constructions, the pressure side squealer tip wall <NUM> may extends along the pressure side wall <NUM> from the blade leading edge <NUM> to the blade trailing edge <NUM> and/or the suction side squealer tip wall <NUM> may extends along the suction side wall <NUM> from the blade leading edge <NUM> to a location before the blade trailing edge <NUM>.

The blade airfoil <NUM> has a second plurality of cooling holes <NUM> that are formed at the squealer tip wall <NUM> and pass through the squealer tip wall <NUM>. The second plurality of cooling holes <NUM> are arranged at the pressure side squealer tip wall <NUM> and pass through the pressure side squealer tip wall <NUM> and are arranged at the suction side squealer tip wall <NUM> and pass through the suction side squealer tip wall <NUM>. The second plurality of cooling holes <NUM> are in flow connection with the interior of the blade airfoil <NUM>.

A chamfered surface <NUM> is formed as a part of the squealer tip wall <NUM>. In the construction illustrated in <FIG>, the portion of the squealer tip wall <NUM> that is adjacent to the blade trailing edge <NUM> is chamfered to form the chamfered surface <NUM>. As used herein "adjacent" means that the chamfered surface <NUM> begins at the blade trailing edge <NUM> or within <NUM>% of a length of the mean camber line <NUM> from the blade trailing edge <NUM>. The chamfered surface <NUM> may extend along the squealer tip wall <NUM> from the blade trailing edge <NUM> toward the blade leading edge <NUM> for a distance between <NUM>-<NUM>% of the length of the mean camber line <NUM>. The length of the mean camber line <NUM> is defined as the curved length of the mean camber line <NUM> from the blade trailing edge <NUM> to the blade leading edge <NUM>. The chamfered surface <NUM> may extend from the blade tip <NUM> toward the blade platform <NUM> for a distance between <NUM> - <NUM>% of a height of the blade airfoil <NUM>. The height of the blade airfoil <NUM> is defined from the blade platform <NUM> to the blade tip <NUM>. The chamfered surface <NUM> may have any desired dimensions and orientations to meet design requirements of the gas turbine engine <NUM>.

In the construction illustrated in <FIG>, the chamfered surface <NUM> is formed as a part of the suction side squealer tip wall <NUM>. A portion of the suction side squealer tip wall <NUM> that is adjacent to the blade trailing edge <NUM> is chamfered to form the chamfered surface <NUM>. The chamfered surface <NUM> extends along the suction side squealer tip wall <NUM> from the blade trailing edge <NUM> toward the blade leading edge <NUM> for the distance between <NUM>-<NUM>% of the length of the mean camber line <NUM>. The chamfered surface <NUM> extends from the blade tip <NUM> on the suction side squealer tip wall <NUM> toward the blade platform <NUM> for the distance between <NUM> - <NUM>% of the height of the blade airfoil <NUM>. In other constructions, the chamfered surface <NUM> may be formed as a part of the suction side squealer tip wall <NUM> and a part of the pressure side squealer tip wall <NUM> that are adjacent to the blade trailing edge <NUM>. In yet other constructions, the chamfered surface <NUM> may be formed as a part of the squealer tip wall <NUM> adjacent to the blade trailing edge <NUM> of a blade airfoil <NUM> having the tip cap surface <NUM> extending from the blade leading edge <NUM> to the blade trailing edge <NUM> without the offset surface <NUM>.

According to the invention, a thermal barrier coating <NUM> is applied to the chamfered surface <NUM>. In other constructions that are not part of the invention, the chamfered surface <NUM> may not be applied with the thermal barrier coating <NUM>.

In operation, with reference to <FIG> and <FIG>, cooling flow exits the blade airfoil <NUM> from the interior of the blade airfoil <NUM> through the first plurality of cooling holes <NUM> disposed at the tip cap surface <NUM> and through the second plurality of cooling holes <NUM> disposed at the squealer tip wall <NUM>. The tip cap surface <NUM> is stepped radially up from the offset surface <NUM> so that the cooling flow exits the blade airfoil <NUM> at a location that is closer to the blade tip <NUM>. Cooling to the blade tip <NUM> is thus improved. The chamfered surface <NUM> at the region of the blade trailing edge <NUM> of the squealer tip wall <NUM> reduces metal temperature of the blade airfoil <NUM> at this region. The chamfered surface <NUM> is coated with the thermal barrier coating <NUM>. The arrangement of the chamfered surface <NUM> with the thermal barrier coating <NUM> reduces the degradation and distress at the trailing edge <NUM> of the squealer tip wall <NUM>. Durability of the turbine blade <NUM> is thus improved.

Although an exemplary embodiment of the present disclosure has been described in detail, those skilled in the art will understand that various changes, substitutions, variations, and improvements disclosed herein may be made without departing from the invention, which is solely defined in the claims.

Claim 1:
A turbine blade (<NUM>) comprising:
a blade platform (<NUM>);
a blade airfoil (<NUM>) that extends from the blade platform (<NUM>) toward a blade tip (<NUM>), the blade airfoil (<NUM>) having a pressure side wall (<NUM>) and a suction side wall (<NUM>) joined at a blade leading edge (<NUM>) and a blade trailing edge (<NUM>);
a tip cap surface (<NUM>) defined at an end of the blade airfoil (<NUM>) facing the blade tip (<NUM>);
a squealer tip wall (<NUM>) that extends along a portion of the pressure side wall (<NUM>) and a portion of the suction side wall (<NUM>) from the tip cap surface (<NUM>) to the blade tip (<NUM>) and from the blade leading edge (<NUM>) toward the blade trailing edge (<NUM>); and
a chamfered surface (<NUM>) formed as a part of the squealer tip wall (<NUM>) at a region that is adjacent to the blade trailing edge (<NUM>),
characterized in that
a thermal barrier coating (<NUM>) is applied to the chamfered surface (<NUM>)
and/or
in that the turbine blade comprises an offset surface (<NUM>) that is offset a non-zero distance from the tip cap surface (<NUM>) toward the blade platform (<NUM>).