Patent Grant
9249687
The subject matter disclosed herein relates to exhaust diffusion for turbine systems.
A gas turbine system may include an exhaust diffuser coupled to a gas turbine engine. The gas turbine engine combusts a fuel to generate hot combustion gases, which flow through a turbine to drive a load and/or compressor. The exhaust diffuser receives the exhaust from the turbine, and gradually reduces the pressure and velocity. Unfortunately, exhaust diffusers often consume a considerable amount of space. For instance, the exhaust diffuser may be as long as the gas turbine engine. Therefore, it may prove beneficial to implement design strategies for reducing the footprint of the exhaust diffuser, and, thus, the overall footprint of the gas turbine system.
Certain embodiments commensurate in scope with the originally claimed invention are summarized below. These embodiments are not intended to limit the scope of the claimed invention, but rather these embodiments are intended only to provide a brief summary of possible forms of the invention. Indeed, the invention may encompass a variety of forms that may be similar to or different from the embodiments set forth below.
In accordance with a first embodiment, a system includes a gas turbine engine. The gas turbine engine includes a combustion section and a turbine section coupled to the combustion section. The turbine section includes a turbine stage having multiple turbine blades coupled to a rotor, a stationary shroud disposed about the multiple turbine blades, and a clearance between the stationary shroud and each end of the multiple turbine blades. The turbine blades may have a rotating shroud attached to their ends or not. The gas turbine engine includes a diffuser section coupled to the turbine section. The diffuser section includes an outer wall defining an expanding flow path downstream from the multiple turbine blades. The outer wall includes a first wall portion having a first angle relative to a rotational axis of the multiple turbine blades, and the clearance is configured to enable over tip leakage flow to energize a boundary layer along the outer wall.
In accordance with a second embodiment, a system includes a rotary section. The rotary section includes multiple blades coupled to a rotor, a stationary shroud disposed about the multiple blades, and a clearance between the stationary shroud and each end of the multiple blades, wherein the clearance is configured to enable over tip leakage flow. The turbine blades may have a rotating shroud attached to their ends or not. The system also includes a diffuser section that includes an outer wall defining an expanding flow path downstream from the multiple blades. The outer wall includes a first wall portion having a first angle relative to a rotational axis of the multiple blades, and the clearance is configured to enable an increase in the first angle by maintaining the boundary layer along the outer wall with the over tip leakage flow.
In accordance with a third embodiment, a method includes enabling an over tip leakage flow to pass between a stationary shroud and multiple turbine blades of a turbine stage. The method also includes energizing a boundary layer along a wall of a turbine diffuser with the over tip leakage flow.
These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein:
One or more specific embodiments of the present invention will be described below. In an effort to provide a concise description of these embodiments, all features of an actual implementation may not be described in the specification. It should be appreciated that in the development of any such actual implementation, as in any engineering or design project, numerous implementation-specific decisions must be made to achieve the developers' specific goals, such as compliance with system-related and business-related constraints, which may vary from one implementation to another. Moreover, it should be appreciated that such a development effort might be complex and time consuming, but would nevertheless be a routine undertaking of design, fabrication, and manufacture for those of ordinary skill having the benefit of this disclosure.
When introducing elements of various embodiments of the present invention, the articles “a,” “an,” “the,” and “said” are intended to mean that there are one or more of the elements. The terms “comprising,” “including,” and “having” are intended to be inclusive and mean that there may be additional elements other than the listed elements.
The disclosed embodiments are directed to over tip leakage flow in a turbine, such as a gas turbine or steam turbine, to reduce flow separation from along an outer wall of an exhaust diffuser. In general, it may be desirable to minimize the clearance between ends of rotating blades and the surrounding stationary shroud, thereby maximizing the work of the fluid (e.g., steam or hot gases) on the rotating blades. However, some amount of clearance may be provided to reduce the possibility of rub between the blades and the stationary shroud. However, this consideration for clearance does not relate to the fluid flow downstream from the rotating blades. As discussed below, flow separation and other undesirable fluid flow may occur downstream from the rotating blades. The disclosed embodiments specifically adjust the clearance to control an over tip leakage flow, thereby controlling the fluid flow downstream of the blades. For example, the over tip leakage flow that passes between the blade ends of multiple blades and a stationary shroud disposed about the blades energizes a boundary layer along an outer wall of an exhaust diffuser, thereby allowing large angles relative to a rotational axis of the blades to be incorporated into the outer wall of the exhaust diffuser. In other words, the over tip leakage flow increases the flow velocity along the boundary layer, thus, reducing or preventing the separation of the flow from the outer wall of the exhaust diffuser that normally occurs when large angles relative to the rotational axis of the blades are used, while also maintaining the pressure recovery of the exhaust diffuser. Over tip leakage flow, while allowing an increase in the angles in the exhaust diffuser, may also allow the length of the diffuser to be reduced, as well as, the overall length of the turbine system.
The gas turbine engine 118 includes one or more fuel nozzles 160 located inside a combustor section 162. In certain embodiments, the gas turbine engine 118 may include multiple combustors 120 disposed in an annular arrangement within the combustor section 162. Further, each combustor 120 may include multiple fuel nozzles 160 attached to or near the head end of each combustor 120 in an annular or other arrangement.
Air enters through the air intake section 163 and is compressed by the compressor 132. The compressed air from the compressor 132 is then directed into the combustor section 162 where the compressed air is mixed with fuel. The mixture of compressed air and fuel is generally burned within the combustor section 162 to generate high-temperature, high-pressure combustion gases, which are used to generate torque within the turbine section 130. As noted above, multiple combustors 120 may be annularly disposed within the combustor section 162. Each combustor 120 includes a transition piece 172 that directs the hot combustion gases from the combustor 120 to the turbine section 130. In particular, each transition piece 172 generally defines a hot gas path from the combustor 120 to a nozzle assembly of the turbine section 130, included within a first stage 174 of the turbine 130.
As depicted, the turbine section 130 includes three separate stages 174, 176, and 178. Each stage 174, 176, and 178 includes a plurality of blades 180 coupled to a rotor wheel 182 rotatably attached to a shaft 184. Each stage 174, 176, and 178 also includes a nozzle assembly 186 disposed directly upstream of each set of blades 180. The nozzle assemblies 186 direct the hot combustion gases toward the blades 180 where the hot combustion gases apply motive forces to the blades 180 to rotate the blades 180, thereby turning the shaft 184. The hot combustion gases flow through each of the stages 174, 176, and 178 applying motive forces to the blades 180 within each stage 174, 176, and 178. The hot combustion gases may then exit the gas turbine section 130 through an exhaust diffuser section 188. The exhaust diffuser section 188 functions by reducing the velocity of fluid flow through the diffuser section 188, while also increasing the static pressure to increase the work produced by the gas turbine engine 118. As illustrated, the exhaust diffuser section 188 has a length 190, which is a portion of an overall length 192 of the gas turbine engine 118. The disclosed engine 118 provides over tip leakage flow from the turbine section 180 into the exhaust diffuser section 188 to energize the boundary layer in the exhaust diffuser section 188, thereby enabling a reduction in length 190.
In the illustrated embodiment, the last stage 178 includes a clearance 194 between ends of the plurality of blades 180 and a stationary shroud 196 disposed about the plurality of blades 180. The clearance 194 allows an over tip leakage flow to energize the boundary layer between an outer wall 198 of the exhaust diffuser section 188 and the flow of the hot combustion gases, thereby allowing the use of large angles in the diffuser section 188 and the shortening of the length 190 of the diffuser section 188 relative to the total length 192 of the gas turbine engine 118. In certain embodiments incorporating over tip leakage flow, the length 190 of the diffuser section 188 may range from approximately 25 to 50 percent, 30 to 45 percent, or 35 to 40 percent of the total length 192 of the gas turbine engine 118. For example, the length 190 of the diffuser section 188 may account for 30, 35, 40, 45, or 50 percent, or any percent therebetween of the total length 192 of the gas turbine engine 118.
The diffuser section 188 includes greater angles to take advantage of the over tip leakage flow 212. The diffuser section 188 includes the outer wall 198 and a strut 200 disposed radially across the diffuser section 188. The outer wall 198 defines an expanding flow path downstream from the plurality of blades 180. The outer wall 198 includes a first wall portion 214 and a second wall portion 216 downstream of the first wall portion 214. The first wall portion 214 includes a first angle 218 relative to the rotational axis 210 of the plurality of blades 180, as indicated by line 211 parallel to axis 210. In certain embodiments, the first angle 218 may range between approximately 16 to 40 degrees, 20 to 40 degrees, 20 to 30 degrees, 18 to 28 degrees, or 21 to 23 degrees. For example, the first angle 218 may be approximately 16, 18, 20, 22, or 24 degrees, or any angle therebetween. The over tip leakage flow 212 through the clearance 194 enables the increase in the first angle 218 by maintaining the boundary layer along the outer wall 198. Similarly, the second wall portion 216 includes a second angle 220 relative to the rotational axis 210 of the plurality of blades 180, as indicated by line 211 parallel to axis 210. In certain embodiments, the second angle 220 may range between approximately 6 to 12 degrees or 7 to 9 degrees. For example, the second angle 220 may be approximately 6, 8, or 10 degrees, or any angle therebetween. In some embodiments, the first angle 218 may range between approximately 20 to 24 degrees, and the second angle may range between approximately 6 to 12 degrees. The over tip leakage flow 212 may function to energize the boundary layer primarily along the first wall portion 214 at the angle 218, or also along the second wall portion 216 at the angle 220. In either case, the over tip leakage flow enables an increase in the average angle of the diffuser section 188, thereby providing more aggressive diffusion over a shorter distance by virtue of the energized boundary layer.
Incorporating the first angle 218 with the measurements above normally would cause an excessive adverse pressure gradient within the diffuser section 188 causing early flow separation from along the outer wall 198 resulting in poorer performance by the diffuser section 188. However, the over tip leakage flow 212 energizes the boundary layer and reduces or prevents the early flow separation from the outer wall 198 at least along the first wall portion 214. The over tip leakage flow 212 allows the use of a large first angle 218 within the diffuser section 188 and a shortening of the length 190 of the diffuser section 188 relative to the total length 192 of the gas turbine engine 118, while still maintaining diameters 222 and 224 of diffuser section inlet and outlet, respectively. In addition, shortening the length 190 of the diffuser section 188 creates a higher diffusion area ratio per unit length of the diffuser section 188, while maintaining a total diffusion area of the diffuser section 188 for diffusion recovery. As a result, the large first angle 218, in conjunction with the over tip leakage flow 212, allows for at least the same or improved pressure recovery and diffuser performance in a shorter turbine section 188. In certain embodiments, reduction in the length 190 of the diffuser section 188 may range from 30 to 60 percent. As a result, the length 190 of the diffuser section 188 may be at least less than approximately 15 percent of the total length 192 of the gas turbine engine 118.
However, providing some clearance 194 reduces the amount of separation along the boundary layer.
Increasing the clearance 194 imparts even greater momentum and energy to the exhaust flow 240 (e.g., swirl and radial momentum) of the combustion gases.
Plot 256 illustrates, in the absence of clearance 194, a gradual increase in pressure recovery initially along the axial length 190 of the diffuser section 188. As the flow of the combustion gases encounter the leading edge of the strut 200, represented by dashed line 260, the amount of pressure recovery sharply decreases due to flow interaction with the strut 200, but recovers and gradually increases as the flow approaches the trailing edge of the strut 200, represented by dashed line 262, as shown in plot 256. After the strut 200, the pressure recovery gradually increases along the rest of the axial length 190 of the diffuser section 188.
Plot 258 illustrates, in the presence of clearance 194, similar to plot 256, an increase in pressure recovery, but at a greater rate, initially along the axial length 190 of the diffuser section 188. Also, similarly, as the flow of the combustion gases encounter the leading edge 260 of the strut 200, the amount of pressure recovery decreases due to flow interaction with the strut 200, but only slightly, then recovers and increases an upper level of pressure recovery as the flow approaches the trailing edge 262 of the strut 200, as shown in plot 258. After the strut 200, the pressure recovery remains at the upper level of pressure recovery along the rest of the axial length 190 of the diffuser section 188. The graph 200 illustrates that in the presence of clearance 194, as shown in plot 258, pressure recovery occurs at a greater rate and reaches the maximum obtainable pressure recovery sooner along the axial length 190 of the diffuser section 188 than in the absence of clearance 194, as shown in plot 256. As a result of this earlier and greater pressure recovery due to clearance 194, which allows over tip leakage flow 212, large angles may be used in the diffuser section 188 allowing the shortening of the diffuser section 188 in relation to the gas turbine engine 118.
Plot 278 illustrates that, in the absence of clearance 194, the axial velocity slightly decreases as the flow of the combustion gases expands in the radial direction toward the outer wall 198 until the flow expansion proceeds to a point 277 where the expansion results in the sudden and significant loss of axial velocity in the flow of the combustion gases. This sudden loss of axial velocity occurs due to the stalling of the flow of the combustion gases, as a result of the large angles within the diffuser section 188. The low velocity region 279 near the outer wall 198 represents significant flow separation from the outer wall 198. Plot 280 illustrates, in the presence of clearance 194, a slight decrease in axial velocity as the flow of the combustion gases expands in the radial direction. However, as shown in plot 280, the flow of the combustion gases maintains axial velocity, in the presence of over tip leakage flow 212 due to clearance 194, as the flow expands in the radial direction toward the outer wall 198. Thus, the plot 280 does not exhibit the low velocity region 279. Plot 280 illustrates the imparting of momentum and energy to the flow of the combustion gases to maintain the boundary layer (e.g., prevent the stalling of the flow and separation along the boundary layer) along the outer wall 198 of the diffuser section 188. Thus, the over tip leakage flow 212 enables increased of the outer wall 198, while substantially preventing flow separation.
Plot 296 illustrates that, in the absence of clearance 194, the radial velocity slightly increases as the flow of the combustion gases expands in the radial direction toward the outer wall 198 until the flow expansion proceeds to a point 297 where the expansion results in the steady loss of radial velocity in the flow of the combustion gases. The loss of radial velocity, as with the loss of the axial velocity, occurs due to the stalling of the flow of the combustion gases, as a result of the large angles within the diffuser section 188. Plot 298 illustrates, in the presence of over tip leakage flow 212 from clearance 194, a sharp and significant increase in radial velocity occurs as the flow of the combustion gases expands toward the outer wall 198. The radial velocity even continues to increase during expansion, as shown in plot 298, past the point 297 in expansion where in plot 296 the radial velocity decreased. Plot 298 illustrates that the over tip leakage flow 212 imparts a significant amount of energy and momentum to the flow of the combustion gases to increase the radial flow velocity to substantially reduce or eliminate flow separation along the outer wall 198 of the diffuser section 188 in the presence of large angles.
As mentioned above, the blade ends 204 of the plurality of blades 180 may include shrouded ends 205.
As mentioned above, over tip leakage flow 212 may be used in a steam turbine system.
In the illustrated embodiment, a last stage 322 includes clearance, as generally indicated by arrow 324, between the blade ends 204 of the plurality of blades 80 and a shroud 326 disposed about the plurality of blades 180. In certain embodiments, the distance of clearance 324 may range from between approximately 100 to 250 mils. The clearance allows over tip leakage flow 212, as described above, and, thus, allowing the use of large angles in the diffuser section 312 and the shortening of the diffuser section 312 relative to the total length of the steam turbine engine 306. The length of the diffuser section 312 may range from approximately 20 to 60 percent, or any percent therebetween of the total length of the steam turbine engine 306.
In certain embodiments, a method of operating a turbine system may include enabling over tip leakage flow 212 to energize a boundary layer and to prevent flow separation downstream from a turbine, e.g., in a diffuser section 188. For example, the method may include enabling over tip leakage flow 212 to pass between the stationary shroud 196 and the plurality of turbine blades 180 of turbine stage 178. The method also includes energizing the boundary layer along the wall 198 of the turbine diffuser 188 with the over tip leakage flow 212. The method may further include radially expanding the flow from the plurality of turbine blades 180 in a downstream direction through the first portion 214 of wall 198 having an angle at least greater than or equal to approximately 16 degrees, wherein the energizing maintains the boundary layer along the first portion 214. In some embodiments, the angle may be at least greater than or equal to approximately 20 degrees. The method, additionally, may include radially expanding the flow from the first portion 214 of wall 198 to the second portion 216 of the wall 198 having an angle at least greater than or equal to approximately 6 degrees. Also, the method may include diffusing an exhaust flow from the turbine stage through the turbine diffuser 188 over length 190 that is at least less than approximately 15 percent of the total length 192 of turbine engine 118 having the turbine stage 178 and the turbine diffuser 188.
Technical effects of the disclosed embodiments include providing large angles in the diffuser section 188 of a turbine system. Also, providing clearance 194 allows the over tip leakage flow 212 to energize and provide momentum to the flow during radial expansion through the diffuser section 188 to prevent the separation of the flow from the wall 198 that normally occurs with large angles. Using the large angles, in conjunction with the over tip leakage flow 212, allows the length of the diffuser section 188, as well as the total length of the turbine system to be reduced while at least maintaining, if not improving, performance. By shortening the lengths of the diffuser section 188 and turbine system the foot prints of each may be reduced.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.
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