Methods and systems for controlling thermal differential in turbine systems

Method and system are provided for controlling a thermal differential within a turbine rotor for use with a turbine system. A thermal barrier coating is applied to a surface of the turbine rotor. The surface is proximate to a wheel rim of the turbine rotor.

BACKGROUND OF THE INVENTION

The subject matter disclosed herein relates generally to turbine systems and, more particularly, to methods and systems for controlling thermal differential in turbine systems.

At least some known turbine systems include a compressor that compresses air channeled downstream through the turbine system. During startup operations and/or shutdown of at least some known turbine systems, at least some portions of such compressors and/or rotors may be subjected to high temperatures, thermal gradients, high stresses, and/or vibrations. For example at least some known compressors include a plurality of stages that increasingly compress and, consequently, increase a temperature of air channeled therethrough. Such differences in airflow temperature may result in thermal gradients developing within the compressor. Moreover, at least some surfaces are exposed to the flow path more than other portions of the compressor. As a result, other thermal gradients may develop within compressor. Such thermal gradients and/or thermal stresses may cause uneven expansion, bending, and/or other stresses within the compressor that, at times, could lead to undesired slipping and/or rubbing of compressor components.

To improve an efficiency and/or a power output of at least some turbine systems, a compressor discharge temperature, a firing temperature, and/or a core flow rate may be increased. However, increasing any of the aforementioned operational characteristics could increase the likelihood of thermal gradients and/or thermal stresses being developed within the compressor and/or the rotor.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, a method is provided for controlling a thermal differential within a turbine rotor for use with a turbine system. The method includes selecting a surface of the turbine rotor proximate to a wheel rim of the turbine rotor and applying a thermal barrier coating to the surface of the turbine rotor.

In another aspect, a turbine rotor is provided for use with a turbine system. The turbine rotor includes a surface proximate to a wheel rim of the turbine rotor and a thermal barrier coating applied to the surface.

In yet another aspect, a turbine system is provided. The turbine system includes a turbine rotor including a surface proximate to a wheel rim of the turbine rotor. A thermal barrier coating is applied to the surface of the turbine rotor.

DETAILED DESCRIPTION OF THE INVENTION

The subject matter described herein relates generally to turbine systems and, more particularly, to methods and systems for controlling a thermal differential within a turbine rotor. In one embodiment, a thermal barrier coating is applied to at least one surface of the turbine rotor, wherein the surface is selected based on stress levels, transient thermal gradients, and/or surrounding heat transfer environment. Applying the thermal barrier coating to such surfaces facilitates controlling transient thermal stresses within the turbine rotor and, thus, facilitates increasing a useful operating life of the turbine system. The turbine rotor surface may be disposed within a compressor and/or a turbine section of the turbine system.

FIG. 1is a schematic illustration of an exemplary turbine system100. In the exemplary embodiment, turbine system100includes, coupled in serial flow arrangement, a compressor104, a combustor assembly106, and a turbine108that is rotatably coupled to compressor104via a rotor shaft110.

During operation, in the exemplary embodiment, ambient air is channeled through an air inlet towards compressor104. The ambient air is then compressed by compressor104prior to being directed towards combustor assembly106. In the exemplary embodiment, compressed air within combustor assembly106is mixed with fuel, and the resulting fuel-air mixture is then ignited within combustor assembly106to generate combustion gases that are directed towards turbine108. In the exemplary embodiment, turbine108extracts rotational energy from the combustion gases and rotates rotor shaft110to drive compressor104. Moreover, in the exemplary embodiment, turbine system100drives a load112, such as a generator, coupled to rotor shaft110. In the exemplary embodiment, load112is positioned downstream of turbine system100. Alternatively, load112may be positioned upstream of turbine system100.

FIG. 2is a cross-sectional illustration of a portion of a turbine rotor or, more specifically, compressor104. In the exemplary embodiment, compressor104includes a plurality of stages230that include at least a first stage232and a second stage234that is positioned downstream from first stage232. In the exemplary embodiment, each stage230includes at least one rotor disc240that is formed with a disc web242bounded circumferentially by a wheel rim244.

In the exemplary embodiment, at least some portions of airflow220channeled through compressor104is diverted through at least one slot260defined within compressor104. For example, in the exemplary embodiment, compressor104includes a circumferential slot262that is oriented to channel a portion224of airflow220in a circumferential direction, and a radial slot264that is oriented to channel a portion222of airflow220in a radial direction. More specifically, in the exemplary embodiment, airflow portion224is directed through circumferential slot262, and airflow portion222is directed through radial slot264. In the exemplary embodiment, slots262and/or264facilitate cooling portions of compressor104that are proximate to wheel rim244and/or combustor assembly106.

In the exemplary embodiment, compressor104includes a rim aft face270that is downstream from stages230and discs240. More specifically, in the exemplary embodiment, rim aft face270is between wheels240and combustor assembly106. As such, in the exemplary embodiment, rim aft face270provides a buffer area between compressor104and combustor assembly106, to facilitate shielding relatively hot air-fuel mixtures ignited in combustor assembly106.

In the exemplary embodiment, at least some sections and/or portions of compressor104are subjected and/or exposed to higher temperatures than other sections and/or portions. Moreover, within such sections and/or portions, at least some surfaces of compressor104are exposed to airflow220and/or higher temperatures more than other portions of compressor104. For example, in the exemplary embodiment, disc rim244has more exposure to airflow220than a disc bore and/or disc web242. Moreover, in the exemplary embodiment, compressor104includes a dovetail slot and loading locations250sized and/or oriented to receive an airfoil (not shown) therein. In the exemplary embodiment, slot and loading locations250is exposed to higher temperatures than other portions of compressor104.

Accordingly, in the exemplary embodiment, a thermal barrier coating (TBC)210is applied to at least some high heat transfer and/or high temperature regions of compressor104including, for example, disc rim244to facilitate controlling a thermal differential within compressor104. As such, TBC210facilitates shielding at least some of these surfaces within compressor104from relatively high temperatures and/or heat transfer convention coefficient. In the exemplary embodiment, TBC210is fabricated from a material having a thermal conductivity that is suitable to enable turbine system100to function as described herein. For example, TBC210may be fabricated from, without limitation, ceramics and/or sprayed oxides.

TBC210may be applied to any suitable rotor and/or static structure within turbine system100. Additional surfaces for application of TBC210may be selected based on various factors including, but not limited to, stress levels, transient thermal gradients, and/or a surrounding heat transfer environment. For example, in the exemplary embodiment, TBC210shields the outer surfaces of each stage230of compressor104including disc rim244, slot and loading locations250, slots262and264, and/or rim aft face270from exposure to high temperatures of airflow220.

In the exemplary embodiment, TBC210is applied with a thickness that is suitable to provide a thermal barrier without substantially affecting airflow220within compressor104. For example, in the exemplary embodiment, TBC210has a thickness of less than approximately 0.04 inches. Particularly, TBC210has a thickness of less than approximately 0.02 inches. More particularly, TBC210has a thickness of less than approximately 0.005 inches. Alternatively, TBC210may have a thickness of greater than or equal to approximately 0.04 inches.

During operation, as airflow220is channeled through compressor104, a portion of airflow220is channeled through compressor104directly to combustor assembly106. As airflow220flows downstream and is compressed by compressor104, the temperature of airflow220generally increases. For example, in the exemplary embodiment, airflow220upstream from first stage232may be approximately 700-800° F. while airflow220downstream from second stage234may be approximately 750-850° F. Moreover, in the exemplary embodiment, a portion of airflow220is directed through slots262and264. More specifically, in the exemplary embodiment, airflow portion224flows circumferentially within compressor104, and airflow portion222flows radially outward within compressor104. In the exemplary embodiment, a compressor discharge is directed through rim aft face270towards combustor assembly106. In the exemplary embodiment, the compressor discharge temperature may be greater than approximately 800° F.

FIG. 3is a graph300illustrating an exemplary temperature difference, or a delta temperature310, of a surface protected by TBC210during operation of turbine rotor or, more specifically, compressor104. More specifically, the X-axis is representative of a mission time measured in seconds, and the Y-axis is representative of a difference in temperature between wheel rim244and bore242. As shown inFIG. 3, the mission time includes three stages to the operating cycle of turbine system100: a startup320, a main operating stage330, and a shutdown340.

In the exemplary embodiment, during startup320of turbine system100, compressor104is subjected to a quick increase in temperature. Accordingly, surfaces within compressor104may be subjected to a relatively quick increase in delta temperature310, as is indicated by the spike in delta temperature310. In the exemplary embodiment, TBC210facilitates reducing delta temperature310during startup320as compared to a compressor104without TBC210. As such, TBC210facilitates reducing a transient response rate during startup320of turbine system100.

Until operation of compressor104stabilizes for main operating stage330, in the exemplary embodiment, delta temperature310decreases relatively steadily. In the exemplary embodiment, the operating temperature of compressor104is relatively steady and, thus, during main operating stage330, delta temperature310is relatively constant at approximately 0° F.

In the exemplary embodiment, during shutdown340of turbine system100, compressor104is subjected to a quick decrease in temperature. Accordingly, surfaces within compressor104may be subjected to a relatively quick increase in delta temperature310in a negative direction, as is indicated by the dip in delta temperature310. In the exemplary embodiment, TBC210facilitates decreasing delta temperature310during shutdown340as compared to a compressor104without TBC210. In the exemplary embodiment, delta temperature310increases overtime until compressor104is cooled down. As such, TBC210facilitates reducing a transient response rate during shutdown340of turbine system100.

The exemplary methods and systems described herein facilitate controlling a rotor thermal differential of a turbine rotor and/or turbine system. By reducing a change in temperature difference of the turbine rotor surface, the exemplary methods and systems facilitate decreasing stresses and, therefore, increasing an operating life of the turbine rotor and/or a turbine system. Additionally, the methods and systems described herein may enable the compressor to have high compressor discharge temperatures, firing temperatures, and/or core flow rates, thereby increasing an efficiency of turbine system as well as faster operational cycles. Moreover, controlling such temperature differences of the turbine rotor facilitates reducing thermal gradients, which may result in reducing undesired slipping and/or rubbing of rotor and/or static components.

Exemplary embodiments of methods and systems are described and/or illustrated herein in detail. The exemplary systems and methods are not limited to the specific embodiments described herein, but rather, components of each system and/or steps of each method may be utilized independently and separately from other components and/or method steps described herein. Each component and each method step may also be used in combination with other components and/or method steps.