Abstract:
A combustor liner for a gas turbine includes a body having a plurality of angled strips on an outside surface of the combustor liner. The plurality of angled strips are arranged in an array about the outside surface. A space is disposed between each of said plurality of angled strips so as to create vortices in a cooling air flowing in a longitudinal direction across said outside surface of said combustor liner. A method of fabricating a combustor liner includes forming a plurality of angled strips on an outside surface of the combustor liner and arranged in an array about the outside surface, each of the angled strips is disposed so as to be spaced apart to create vortices in a cooling air flowing across the outside surface of the combustor liner.

Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to turbine components and more particularly to a combustor liner. 
   Conventional gas turbine combustors use diffusion (i.e., non-premixed) combustion in which fuel and air enter the combustion chamber separately. The process of mixing and burning produces flame temperatures that can exceed 3900 degrees F. Since conventional combustors and/or transition pieces having liners are generally capable of withstanding for about ten thousand hours (10,000 hrs.), a maximum temperature on the order of about 1500 degrees F., steps to protect the combustor and/or transition piece must be taken. This has typically been done by film-cooling which involves introducing relatively cool compressor air into a plenum formed by the compressor discharge case surrounding the outside of the combustor. In this prior arrangement, the air from the plenum passes through louvers in the combustor liner and then passes as a film over the inner surface of the liner, thereby maintaining combustor liner temperature at an acceptable level. 
   Because diatomic nitrogen rapidly disassociates at temperatures exceeding about 3000° F. (about 1650° C.) and reacts readily with oxygen at such temperatures, the high temperatures of diffusion combustion result in relatively high NOx emissions. One approach to reducing NOx emissions has been to premix the maximum possible amount of compressor air with fuel. The resulting lean premixed combustion produces cooler flame temperatures and thus lower NOx emissions. Although lean premixed combustion is cooler than diffusion combustion, the flame temperature is still too hot for prior conventional combustor components to withstand without some type of active cooling. 
   Furthermore, because the advanced combustors premix the maximum possible amount of air with the fuel for NOx reduction, little or no cooling air is available, making film-cooling of the combustor liner and transition piece impractical. Nevertheless, combustor liners require cooling to maintain material temperatures below limits. In dry low NOx (DLN) emission systems, this cooling can only be supplied as cold side convection. Such cooling must be performed within the acceptable limits of thermal gradients and pressure loss. Thus, means such as thermal barrier coatings in conjunction with “backside” cooling have been utilized to protect the combustor liner and transition piece from destruction by such high heat. Backside cooling involves passing the compressor air over the outer surface of the combustor liner and transition piece prior to premixing the air with the fuel. 
   There are currently three forms of prior art for the convective cooling of combustor chambers. First, a series of longitudinal or axially spaced horizontal turbulators, which appear as straight lines across the surface of the liner, are used in practice to disrupt the thermal boundary layer and provide enhanced heat transfer for cooling. These turbulators are either machined in the metal surface, or applied as tack-welded strips of material to the metal. Second, convective cooling is provided by a series of impingement jets supplied by the external combustor chamber cooling flow sleeve. Typically, it is not possible to provide such impingement cooling over the entire extent of the chamber, and so some mixture of impingement and surface turbulators is employed. Third, an array of surface indentations, also known as dimples or hemispherical concavities, is made in the liner surface to create flow vortices that act to enhance heat transfer. The various known techniques enhance heat transfer but with varying effects on thermal gradients and pressure losses. 
   BRIEF DESCRIPTION OF THE INVENTION 
   Exemplary embodiments of the invention include a combustor liner for a gas turbine that has a body having a plurality of angled strips on an outside surface of the combustor liner. The plurality of angled strips are arranged in an array about the outside surface. A space is disposed between each of said plurality of angled strips so as to create vortices in a cooling air flowing in a longitudinal direction across said outside surface of said combustor liner. 
   Exemplary embodiments of the invention also include a method of fabricating a combustor liner. The method includes forming a plurality of angled strips on an outside surface of the combustor liner and arranged in an array about the outside surface, each of the angled strips is disposed so as to be spaced apart to create vortices in the cooling air flowing across the outside surface of the combustor liner. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a schematic representation of a known gas turbine combustor, 
       FIG. 2  illustrates a top view of an outside surface of a combustor liner. 
       FIG. 3  illustrates an alternative embodiment of the surface of  FIG. 2 . 
       FIG. 4  illustrates an alternative embodiment of the surface of  FIG. 2 . 
       FIG. 5  illustrates an alternative embodiment of the surface of  FIG. 2 . 
       FIG. 6  illustrates an alternative embodiment of the surface of  FIG. 2 . 
       FIG. 7  illustrates an alternative embodiment of the surface of  FIG. 2 . 
       FIG. 8  illustrates the geometry and flow orientation of the surface. 
       FIG. 9  illustrates a cross-section of one of the strips. 
       FIG. 10  is a graph illustrating the Reynolds number versus ratio of the turbulated friction coefficient surface to the smooth surface friction coefficient. 
       FIG. 11  is a graph illustrating the Reynolds number versus the ratio of the turbulated heat transfer coefficient (“HTC”) to the smooth surface HTC. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   Referring to  FIG. 1 , a can-annular revers-flow combustor  10  is illustrated. The combustor  10  generates the gases needed to drive the rotary motion of a turbine by combusting air and fuel within a confined space and discharging the resulting combustion gases through a stationary row of vanes. In operation, discharge air  11  from a compressor (compressed to a pressure on the order of about 250-400 lb/sq-in) reverses direction as it passes over the outside of the combustors (one shown at  14 ) and again as it enters the combustor en route to the turbine (fist stage indicated at  16 ). Compressed air and fuel are burned in the combustion chamber  18 , producing gases with a temperature of about 1500° C. or about 2730° F. These combustion gases flow at a high velocity into turbine section  16  via transition piece  20 . The transition piece  20  connects to a combustor liner  24  at connector  22 , but in some applications, a discrete connector segment may be located between the transition piece  20  and the combustor liner. The combustor liner  24  and the transition piece  20  have an outside surface  26  that the discharge air  11  flows over, which cools the combustor liner  24 . 
   In particular, there is an annular flow of the discharge air  11  that is convectively processed over the outside surface  26  (cold side) of liner  24 . In an exemplary embodiment, the discharge air flows through a flow sleeve  28 , which forms an annular gap  30  so that the flow velocities can be sufficiently high to produce high heat transfer coefficients. The flow sleeve  28 , which is located at both the combustor liner  24  and the transition piece  20 , may be two separate sleeves connected together. The flow sleeve  28  has a series of holes, slots, or other openings (not shown) that allow the discharge air  11  to move into the flow sleeve  28  in sufficient quantity without incurring a large pressure drop. 
     FIGS. 2-7  illustrate alternative embodiments of patterned chevron and broken chevron surfaces that are machined or fabricated on the outside surface  26  (cold side) of the combustor liner  24 . In particular, the figures illustrate the general geometry and flow orientation of such surfaces. It is understood that  FIGS. 2-7  represent just a portion of the various embodiments that are encompassed by the angled, segmented strips  40 . The various embodiments improve upon the cooling augmentation, and in particular convection cooling, that may be obtained on industrial gas turbine combustor liners and transition pieces. 
   In particular, each of the embodiments illustrate the formation of an array of projecting strips  40  in the specific form of chevrons, i.e., v-shape, or broken chevron strips, which have the base of V-shape removed and may also include a first side of the V-shape offset from the second side of the V-shape so as to create staggered angled strips. The chevron surfaces formed with V-shaped strips in either inline or staggered arrays serve to disrupt the boundary layer flow over the cooling surface, but also to create important secondary flow vortices along the angled strips. These secondary flows add to the heat transfer enhancement and also interact in the regions between chevrons to further mix and disrupt the flows. The edges or ends of these strips also create local shedding vortices. The broken chevron arrays remove the base of the chevrons, thus leaving staggered and oriented strips which act as do the chevrons, but further create additional enhanced flows and heat transfer at the added broken edges. 
   Referring to  FIGS. 2-9  and in particular,  FIGS. 8 and 9 , the strips  40  are segmented so that there are spaces  42  and  44  between the v-shaped strips and the angled strips. Space  42  is the horizontal distance between each of the v-shaped strips. Space  44  is the horizontal distance of the space created when the base of the chevron is removed. There is also a space P, which is in the longitudinal distance between each strip  40 . In an exemplary embodiment, spaces  42  and  44  are the same dimension, but it is not required that spaces  42  and  44  be the same dimension. The spaces  42  and  44  create distinct edges in which the discharge air  11  interacts with the edges to create vortices to turbulate the flow. It is noted that while the air flows mainly in the longitudinal direction, the spaces  42  and  44  are formed laterally next to each strip  40 . 
   The v-shaped strips  40  may be staggered so as to create an offset  46 , which is defined by the pitch of the strips  40 . Each strip  40  is located so as to have the space P between each strip  40 . In an exemplary embodiment, the offset  46  is half the distance P between any two strips  40  in the longitudinal direction. The offset  46  can range from about 0.3 of the space P to about 0.7 of the space P. 
   The different angles A of the strips  40  create a different magnitude of local secondary flow or secondary flow vortices. The amount of the angle A of the strips  40  includes angles of about 30° to about 60° from a horizontal line. Each strip  40  has a length L that is about 1 cm to about 2 cm. One parameter is defined as the ratio between the distance P between the strips to the height H of the strips (P/H). In an exemplary embodiment, the ratio P/H is about 6 to 14. This ratio determines the preferred spacing of the strips in a row of such features so that the flow may reattach to the surface  26  between each strip for best heat transfer. Another parameter is the spacing  42  between rows of surface strips. This spacing must be large enough to allow the generation of flow vortices freely from the edges of the strips. In an exemplary embodiment, this spacing  42  is from 5 to 10 times the height H. There is no upper limit to this spacing  42  as he adjacent strip edges will continue to generate vortices even if the spacing is very large. However, spacing  42  is related to the overall surface area augmentation achieved by these enhancement geometries in that a closer spacing means more features and more added surface area 
   For a large combustor chamber of about 14 to 16 inch (35.5 to 40.6 cm) in diameter, an exemplary embodiment of the height H of the strips  40  is between 0.020 inches (0.051 cm) and 0.120 inches (0.305 cm). Spacing  42  and height H of the strips  40  may vary along the surface array as desired to achieve some tailoring of cooling augmentation. Moreover, by utilizing the strips in the segmented chevron and segmented, broken chevron patterns, the surface area that the discharge air  11  (coolant) interacts with increases by up to approximately 25% as compared to using no strips on the outside of the liner. 
     FIG. 9  also shows that at the base of the strip there is a radius R, which helps reduce the stress at each of the strips. In addition, in an exemplary embodiment, the strip will have a flat top  50  so as to allow better turbulence of the flow of the discharge air. It will be appreciated that the radius and flat top are typical of machined strips and that other fabrication methods are contemplated and may result in rounded tops or very small radii at the base. 
   The spaces  42 ,  44  and offset  46  provide additional disturbance to increase the turbulation of the flow. The increase in turbulation allows the flow of the discharge air  11  to be stirred up at the outside surface  26  of the liner, which brings fresh discharge air down to the outside surface of the liner. In other words, discharge air that is further away from the surface and thus, is cooler, is brought to the liner surface, thereby allowing the cooler air to enhance the heat transfer rate of the surface. It is noted that if the strips  40  are too close together, the air flow may actually be isolated away from the surface to be cooled. In addition, if the strips  40  are too small, the surface features will no longer create substantial secondary flows along the direction of the strips  40 . 
   The strips  40  are oriented in a specified manner so that the bulk discharge air  11  flows across the strips  40  in the longitudinal direction. The flow of air generates numerous and well placed edge vortices in the flow immediately adjacent to the surface to be cooled. It is noted that the flow in the longitudinal direction can be from either the top to the bottom or the bottom to the top direction, shown as the double arrow. While the arrow shows the flow as flowing from either the top to the bottom or the bottom to the top, the change in direction indicates alternate orientations of the strips  40  and not a change in the direction of the main flow between the combustor liner and the flow sleeve, or between the transition piece and the flow sleeve. These shaped and oriented strips also perform the function of adding substantial surface area for heat flux capability, as well as distributing a more uniform cooling augmentation over the entire surface. 
   The method of forming the strips  40  on the outside surface  26  may be accomplished through casting, machining, brazing, welding, or specific deposition techniques such as laser consolidation, etc., which allow the strips  40  to be formed on the liner  24  after the liner  24  has been formed. Thus, the strips  40  may be integrated as part of the liner  24  or may be added to the liner  24  after the liner has been formed. If the strips  40  are integrated as part of the liner  24  (e.g., machining, casting, etc.), then the strips  40  are fully engaged with liner  24  and there is no interface at the liner with the strips. Thus, there is full thermal contact with the strips  40  that are integrated with the liner  24 , which will improve the heat transfer as the discharge air  11  passes over the strips  40 . Alternatively, the strips may be applied to the surface  26  and then bonded in such a way as to provide a seamless interface with the outside surface  26  of the liner. 
   The strips provide higher thermal enhancement as compared to conventional methods in the prior art, and specifically so at the Reynolds number, e.g., 500,000 to 1,000,000, typical of the combustor cooling passage. Recent information has shown that the conventional practice of using transverse turbulators on the liner cooling surface results in a reduced heat transfer augmentation factor as the cooling flow Reynolds number is increased to the very high levels used for low emissions combustors, eg. Reynolds numbers on the order of 500,000 to 1,000,000. The exemplary embodiments discussed herein provide for an increase in the heat transfer coefficient augmentation factor by 15 to 25%, plus an additional 10 to 20% surface area, for a total increase of 20 to 40% heat flux or cooling capability. The strips  40  also have the unexpected benefit of attaining their augmentation levels almost immediately from the location of the surface treatments, as well as holding constant levels along the surface, both characteristics that are not found in the conventional turbulated surfaces. 
     FIGS. 10 and 11  show the advantages of using the segmented strips. In  FIG. 10 , the horizontal axis represents the Reynolds number and the vertical axis represents the ratio of the turbulated friction coefficient surface to the smooth suede friction coefficient. This figure indicates how the turbulation is greater at all of the various Reynolds numbers or flow volumes (and especially the high Reynolds numbers) for both the broken angled strips  40  and the v-shaped strips  40  as compared to the smooth surface, which has no strips, and the transverse turbulators, which are horizontal strips across the surface. In  FIG. 11 , the horizontal axis represents the Reynolds number and the vertical axis represents the ratio of the turbulated heat transfer coefficient (“HTC”) to the smooth surface HTC. This figure indicates how the heat transfer coefficient is greater at all of the various Reynolds numbers or flow volumes (and especially the high Reynolds numbers) for both the broken angled strips  40  and the v-shaped strips  40  as compared to the transverse turbulators, which are horizontal strips across the surface. 
   In addition, while the invention has been described with reference to exemplary embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims. Moreover, the use of the terms first, second, etc. do not denote any order or importance, but rather the terms fir second, etc. are used to distinguish one element from another.