Patent Publication Number: US-8113003-B2

Title: Transition with a linear flow path for use in a gas turbine engine

Description:
FIELD OF THE INVENTION 
     This invention is directed generally to gas turbine engines, and more particularly to transition ducts for routing gas flow from combustors to the turbine section of gas turbine engines. 
     BACKGROUND OF THE INVENTION 
     Referring to  FIG. 1 , there is shown a cross-section through a portion of a combustion turbine  10 . The major components of the turbine are a compressor section  12 , a combustion section  14  and a turbine section  16 . A rotor assembly  18  is centrally located and extends through the three sections. The compressor section  12  can include cylinders  20 ,  22  that enclose alternating rows of stationary vanes  24  and rotating blades  26 . The stationary vanes  24  can be affixed to the cylinder  20  while the rotating blades  26  can be mounted to the rotor assembly  18  for rotation with the rotor assembly  18 . 
     The combustion section  14  can include a shell  28  that forms a chamber  30 . Multiple combustors, for example, sixteen combustors (only one combustor  32  of which is shown) can be contained within the combustion section chamber  30  and distributed around a circle in an annular pattern. Fuel  34 , which may be in liquid or gaseous form—such as oil or gas—can enter each combustor  32  and be combined with compressed air introduced into the combustor  32  from the chamber  30 , as indicated by the unnumbered arrows surrounding the combustor  32 . The combined fuel/air mixture can be burned in the combustor  32  and the resulting hot, compressed gas flow  36  can be exhausted to a transition duct  38  attached to the combustor  32  for routing to the turbine section  16 . 
     The turbine section  16  can include a cylindrical housing  40 , including an inner cylinder  42 , can enclose rows of stationary vanes and rotating blades, including vanes  44  and blades  46 . The stationary vanes  44  can be affixed to the inner cylinder  42  and the rotating blades  46  can be affixed to discs that form parts of the rotor assembly  18  in the region of the turbine section  16 . The first row of vanes  44  and the first row of blades  46  near the entry of the turbine section  16  are generally referred to as the first stage vanes and the first stage blades, respectively. 
     Encircling the rotor assembly  18  in the turbine section  16  can be a series of vane platforms  48 , which together with rotor discs  50 , collectively define an inner boundary for a gas flow path  52  through the first stage of the turbine section  16 . Each transition duct  38  in the combustion section  14  can be mounted to the turbine section housing  40  and the vane platforms  48  to discharge the gas flow  30  towards the first stage vanes  44  and first stage blades  46 . 
     In operation, the compressor section  12  receives air through an intake (not shown) and compresses it. The compressed air enters the chamber  30  in the combustion section  14  and is distributed to each of the combustors  32 . In each combustor  32 , the fuel  34  and compressed air is mixed and burned. The hot, compressed gas flow  30  is then routed through the transition duct  38  to the turbine section  16 . In the turbine section  16 , the hot, compressed gas flow is turned by the vanes, such as first stage vane  44  and rotates the blades, such as first stage blade  52 , which in turn drive the rotor assembly  18 . The gas flow is then exhausted from the turbine section  16 . The turbine system  10  can include additional exhaust structure (not shown) downstream of the turbine section  16 . The power thus imparted to the rotor assembly  18  can be used not only to rotate the compressor section blades  26  but also to additionally rotate other machinery, such as an external electric generator or a fan for aircraft propulsion (not shown). 
     For a better understanding of the invention, a coordinate system can be applied to such a turbine system to assist in the description of the relative location of components in the system and movement within the system. The axis of rotation of the rotor assembly  18  extends longitudinally through the compressor section  12 , the combustion section  14  and the turbine section  16  and defines a longitudinal direction. Viewed from the perspective of the general operational flow pattern through the various sections, the turbine components can be described as being located longitudinally upstream or downstream relative to each other. For example, the compressor section  12  is longitudinally upstream of the combustion section  14  and the turbine section  16  is longitudinally downstream of the combustion section  14 . The location of the various components away from the central rotor axis or other longitudinal axis can be described in a radial direction. Thus, for example, the blade  46  extends in a radial direction, or radially, from the rotor disc  50 . Locations further away from a longitudinal axis, such as the central rotor axis, can be described as radially outward or outboard compared to closer locations that are radially inward or inboard. 
     The third coordinate direction—a circumferential direction—can describe the location of a particular component with reference to an imaginary circle around a longitudinal axis, such as the central axis of the rotor assembly  18 . For example, looking longitudinally downstream at an array of turbine blades in a turbine engine, one would see each of the blades extending radially outwardly in several radial directions like hands on a clock. The “clock” position—also referred to as the angular position—of each blade describes its location in the circumferential direction. Thus, a blade in this example extending vertically from the rotor disc can be described as being located at the “12 o&#39;clock” position in the circumferential direction while a blade extending to the right from the rotor disc can be described as being located at the “3 o&#39;clock” position in the circumferential direction, and these two blades can be described as being spaced apart in the circumferential direction. Thus, the radial direction can describe the size of the reference circle and the circumferential direction can describe the angular location on the reference circle. 
     Generally, the longitudinal direction, the radial direction and the circumferential direction are orthogonal to each other. Also, direction does not connote positive or negative. For example, the longitudinal direction can be both upstream and downstream and need not coincide with the central axis of the rotor. The radial direction can be inward and outward, and is not limited to describing circular objects or arrays. The circumferential direction can be clockwise and counter-clockwise, and, like the radial direction, need not be limited to describing circular objects or arrays. 
     Further, depending on the context, the relevant position of two components relative to each other can be described with reference to just one of the coordinate directions. For example, the combustor  32  can be described as radially outboard of the rotating blade  46  because the combustor  32  is located radially further away from the central axis of the rotor assembly  18  than the rotating blade  46  is—even though the combustor  32  is not in the same longitudinal plane of the rotating blade  46 , and in fact, is longitudinally upstream of the rotating blade  46  and may not be circumferentially aligned with a particular rotating blade  46 . 
     The coordinate system can also be referenced to describe movement. For example, gas flow  36  in the transition  38  is shown to flow in the direction of arrow  36 . This gas flow  36  travels both longitudinally downstream from the combustor  32  to the turbine section  16  and radially inward from the combustor  32  to the first stage vanes  44  and blades  46 . 
     In the context of describing movement, such as the flow of a gas, the circumferential direction can also be referred to as the tangential direction. When gas flows in the circumferential direction, a component of the flow direction is tangential to a point on the circular path. At any given point on the circle path, the circumferential flow can have a relatively larger tangential component and a relatively smaller radial component. Since the tangential component predominates, particularly for larger diameter paths, such as around vane and blade arrays in a turbine engine, a circumferential direction and tangential direction can be regarded as substantially the same. 
     Bearing this coordinate system in mind and referring to  FIG. 2 , a transition duct  54  is shown alone as it would be seen when viewed from longitudinally downstream. This particular transition duct  54  is oriented in the 12 o&#39;clock circumferential position and it should be understood that a turbine engine would have additional transition ducts, for example, a total of sixteen, spaced in an annular array. 
     The transition duct  54  can include a transition duct body  56  having an inlet  58  for receiving a gas flow exhausted by an associated combustor (not shown, but see  FIG. 1 ). The transition duct body  56  can include an internal passage  60  from the inlet  58  to an outlet  62  from which the gas flow is discharged towards the turbine section (not shown). Because the combustor is radially outboard of the first stage of the turbine section (see  FIG. 1 ), the transition duct  54  extends radially inwardly from its inlet  58  to its outlet  62 . In  FIG. 2 , this radial direction is depicted by the axis  64 . The transition duct  54  includes a longitudinal bend  66  near the outlet  62  to discharge the gas flow predominantly longitudinally. Because the gas flow in the transition duct  54  is redirected radially inwardly and then longitudinally, the transition duct  54  experiences substantial turning in the radial direction  64 . This radial thrust pushes the outlet region of the transition duct  54  radially outwardly (up in the plane of the page of the figure). To support the transition duct  54  against this bending thrust, the transition duct  54  can be radially supported by various braces (not shown) at its ends, as it well known in the art. It can be seen that the outlet  62  and the inlet  58  are aligned along the circumferential or tangential direction, which is depicted by the axis  68 . 
     Reference is now made to  FIG. 3 , focusing on a turbine subsection  70  that includes a combustor  72 , a transition duct  74  and first stage vanes  76  and blades  78 .  FIG. 3  shows a view from above of the combustor  72 , the transition duct  74 , a few first stage vanes  76  and a few first stage blades  78 , illustrated schematically. It should be understood that in a turbine, there would be additional first stage vanes spaced apart circumferentially to form an annular array. Similarly, there would be additional first stage blades spaced apart circumferentially to form an annular array around the engine centerline. These additional vanes and blades are not shown in  FIG. 3  to facilitate illustration. A turbine system would typically also include additional combustors and transitions, but a single combustor  72  and transition  74  are shown schematically for purposes of illustration. 
     From this top view, the longitudinal direction can be noted by reference to the axis  80 . The circumferential or tangential direction can be noted by reference to the axis  82 . The radial direction is not illustrated because the radial direction lies into and out of the page of the figure, but would be generally orthogonal to the longitudinal direction and the radial direction. 
     Gas flow, such as hot, compressed gas with perhaps some limited liquid content, is exhausted from the combustor  72  and routed by the transition duct  74  to the first stage vanes  76  and blades  78 . The gas flow as discharged from the exit or outlet  86  of the transition duct  74  generally travels downstream in the longitudinal direction, as indicated by the arrow  84 . There may be some incidental, small-scale radial and circumferential flow components to the discharged gas flow that produce a downstream wake due to edge conditions  86  at the outlet and other factors. The downstream wake can create vibrations in downstream turbine blades. 
     As this longitudinal gas flow  84  discharges from the outlet  86  of the transition duct  74 , the flow passes the first stage vanes  76 . The function of the first stage vanes  76  is to accelerate and turn the predominantly longitudinal flow in the circumferential direction  82  so that the predominant flow direction of the gas flow leaving the trailing edges of vanes  76  is angled in the circumferential or tangential direction relative to the longitudinal direction as shown, for example, by the arrow  88 . This turned flow  88  thus has a longitudinal component and a circumferential component. The flow angle can be substantial, in the range of 40 degrees to 85 degrees measured from the longitudinal axis  80 . By accelerating and angling the gas flow in the circumferential direction  82  relative to the longitudinal direction  80 , the resulting gas flow  88  more effectively imparts its energy to the first row blades  78 , which in turn rotate the associated rotor assembly (not shown). 
     The use of first stage vanes to accelerate and turn the longitudinal gas flow in the circumferential direction present several challenges. The vanes and the associated vane support structure (see  FIG. 1 ) must have high strength characteristics to withstand the forces generated in changing the direction of a extremely hot, high pressure gas flow over a substantial angle in a relatively short distance. The temperature of the gas flow and the heat generated by this turning process also require a vane cooling system. The forces and heat involved diminish material properties that can crack and otherwise damage the vanes and associated support structure. To address these various requirements and operating conditions, the first stage vanes and the associated support structure and cooling systems have developed into a complex system that can be expensive to manufacture, install, and, in the event of damage, repair and replace. Thus, there is a need to accelerate and tangentially turn a gas flow for presentation to a first stage blade array without the complications and related costs and damage risks associated with first stage vanes. 
     SUMMARY OF THE INVENTION 
     This invention is directed to a transition duct for routing gas flow from a combustor to a turbine section of a turbine engine and eliminating damaging stresses created between conventional transitions and row one turbine vanes. The transition duct may have an axis that is generally linear with a generally linear flow path that combines the functions of a transition and row one turbine vanes. In such a configuration, the transition duct channels gases from a combustor basket to a downstream turbine blade assembly and accomplishes the task of redirecting the gases, thereby eliminating the need for row one vanes. The transition duct directs gases into the turbine assembly at the same incidence angle relative to the longitudinal axis of the engine as the row one vanes. However, the transition duct does not include any leading or trailing edges, and the problems inherent with each, that are found in each of the row one vanes. The transition ducts are constructed such that adjacent sides of adjacent ducts are coplanar, which causes the gases to be emitted from each of the transition ducts without an area of decreased fluid flow between adjacent flows. In at least one embodiment of the transition duct, there is no turning of the gases in particular, no radial or circumferential turning of the gases. As a result, there is not a circumferential pressure gradient across the outlet, thereby resulting in reduced excitation and stresses on the row one blades. The nonexistence of gas turning also reduces structural loading on mounts and eliminates aerodynamic losses due to turning of the gas flow. Finally, because there is no uncovered turning, a more uniform flow angle is created over the range of operating conditions. 
     The transition duct may also be configured to include an outlet with canted side surfaces that is configured to tilt the downstream wake thereby resulting in reduced vibration in downstream row one turbine blades. As such, the outlet reduces vibration of downstream blade that may be caused by the combustor gases exiting the transition duct. 
     The transition duct may be configured to route gas flow in a combustion turbine subsystem that includes a first stage blade array having a plurality of blades extending in a radial direction from a rotor assembly for rotation in a circumferential direction, said circumferential direction having a tangential direction component, the rotor assembly axis defining a longitudinal direction, and at least one combustor located longitudinally upstream of the first stage blade array and located radially outboard of the first stage blade array. The transition duct may be formed from a transition duct body having an internal passage extending between an inlet and an outlet. The outlet may be offset from the inlet in the longitudinal direction. An axis of the transition duct body may be generally linear such that a flow path for gases is generally linear. 
     In at least one embodiment, the inlet may be generally cylindrical and an adjacent midsection of the duct may be generally conically shaped. A throat adjacent to the midsection may have a cross-section with a generally consistent cross-sectional area. The outlet may be formed from a radially outer side generally opposite to a radially inner side, and the radially outer and inner sides may be coupled together with opposed first and second side walls. The radially inner side may be positioned radially outward a distance equivalent to the position of the ID of adjacent turbine blades, and the radially outer side may be positioned radially outward a distance equivalent to the position of the OD of adjacent turbine blades. The first side wall may be canted relative to a radial axis when viewing the outlet longitudinally upstream. The second side wall may also be canted relative to a radial axis when viewing the outlet longitudinally upstream. In one embodiment, the second side wall may be nonparallel to the first side wall of the outlet. The first or second side walls, or both, may be canted between about 20 and about 70 degrees relative to a radial axis when viewing the outlet longitudinally upstream. More particularly, the first or second side walls, or both, may be canted between about 30 and about 60 degrees relative to a radial axis when viewing the outlet longitudinally upstream. 
     In some embodiments, the transition duct body may be generally linear and positioned within a turbine engine such that row one vanes are unnecessary. In particular, the outlet may be offset from the inlet in the tangential direction and positioned such that gases are discharged from the outlet at an angle between the longitudinal direction and the tangential direction. The transition duct body is located between the combustor and the first stage blade array to receive the gas flow from the combustor into the internal passage through the inlet and to discharge the gas flow toward the first stage blade array. 
     During operation, hot combustor gases flow from a combustor into inlets of the transitions. The gases are directed through the internal passages. The position of the transition duct is such that gases are directed through the inlet, into the conical midsection where the flow is accelerated, through the adjacent throat and are expelled out of the outlet. The gases are expelled at a proper orientation relative to the turbine blades such that the gases are directed into the turbine blades in correct orientation without need of row one turbine vanes to alter the flow of the gases. Thus, energy is not lost through use of row one turbine vanes. The canted first and second sides of the outlet distribute the wake across a downstream turbine blade. In particular, the wake is distributed from a pressure side, across a leading edge of the blade, to a suction side, thereby distributing the wake across the entire blade. Such a configuration reduces vibrations and stresses in the downstream, stationary turbine blades. 
     An advantage of this invention is that the transition ducts have generally linear axes that enable gases to be emitted from the ducts in proper alignment relative to the row one turbine blades, thereby eliminating the necessity of row one turbine vanes and the inefficiencies associated with the row one turbine vanes. 
     Another advantage of this invention is that the transition duct eliminates leakages that exist between conventional transitions and turbine vanes because such connection does not exist. 
     Yet another advantage of this invention is that the transition duct eliminates leakage between adjacent turbine vanes at the exit frame because the transition duct eliminates the need for row one turbine vanes. 
     Another advantage of this invention is that the incidence angle at which the transition duct is positioned eliminates uncovered turning of gases exiting the transition, thereby making the flow angles more consistent through the range of operating power levels and enabling more power to be extracted from the first stages of the turbine. 
     Still another advantage of this invention is that the canted sides of the outlet of the transition reduce the trailing wake affect on the downstream turbine blades. 
     Another advantage of this invention is the uniform circumferential pressure gradient at the transition outlet reduces the potential vibration of downstream turbine blades caused by pressure gradients developed in the transition. The transition eliminates the abrupt pressure changes associated with radially aligned transition sides of other transition designs. Eliminating the abrupt pressure changes eliminates the vibrations created by these changes on the turbine blades as the blades rotate about the rotational axis and encounter multiple pressure changes arising from each transition upon each revolution. 
     Yet another advantage of this invention is that the transition eliminates the need for row one turbine vanes and thus eliminates the leading and trailing edges, and the associated problems, including the difficulties of cooling the leading and trailing edges, and the gas blockage caused by the existence of the row one turbine vanes. 
     Another advantage of this invention is that in an assembly of transition ducts in which the transition ducts are positioned adjacent one another and extend radially outward around a centerline of a turbine engine, the flow paths of the transition ducts are parallel downstream of throats within each duct and offset such that the flow from each transition duct is tangential to a circular configuration of transitions. 
     These and other embodiments are described in more detail below. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The accompanying drawings, which are incorporated in and form a part of the specification, illustrate embodiments of the presently disclosed invention and, together with the description, disclose the principles of the invention. 
         FIG. 1  is a cross-sectional view of a portion of a prior turbine engine. 
         FIG. 2  is an upstream longitudinal view of a prior transition duct. 
         FIG. 3  is a schematic radial view of a combustor, transition duct and first stage vanes and blades of a prior turbine engine. 
         FIG. 4  is a longitudinal upstream view of a circular array of transition ducts embodying aspects of the invention. 
         FIG. 5  is a upstream longitudinal view of a circular array of transition ducts embodying aspects of the invention. 
         FIG. 6  is a side view of a transition duct. 
         FIG. 7  is a top view of a circular array of transition ducts. 
         FIG. 8  is an end view of the transition duct of  FIG. 6 . 
         FIG. 9  is a partial perspective view of two transition ducts. 
     
    
    
     DETAILED DESCRIPTION OF EMBODIMENTS OF THE INVENTION 
     As shown in  FIG. 4-9 , this invention is directed to a transition duct  94  for routing gas flow from a combustor to a turbine section of a turbine engine. The transition duct  94  may have an axis  130  that is generally linear. In such a configuration, the transition duct  94  channels gases from a combustor basket to a downstream turbine blade assembly and accomplishes the task of redirecting the gases, which has been accomplished in conventional systems with row one vanes. Thus, the transition duct  94  eliminates the need for row one vanes. The transition duct  94  may also be configured to include an outlet  100  with canted side surfaces  112 ,  114  that is configured to reduce the effect of the transition wake thereby resulting in reduced vibration in downstream turbine blades. As such, the outlet  100  reduces inefficiencies caused by the combustor gases exiting the transition duct  94 . 
     As shown in  FIGS. 4 ,  5  and  7 , the transition ducts  94  may be positioned in an annular array  90 , as shown without surrounding turbine components in an elevation as viewed from longitudinally downstream in a turbine. Each transition duct  94  can include a transition body  96  having an inlet  98  and an outlet  100  and an internal passage  102  between the inlet  98  and the outlet  100  for routing a gas flow through the transition duct  94  from the inlet  98  to the outlet  100 . The array  90  is shown illustrating an arrangement for use in a combustion turbine engine having 16 combustors (not shown). However, the number of transition ducts  94  and their annular arrangement can be varied for use with more or less combustors. 
     As shown in  FIG. 4-6 , the transition duct  94  may include an outlet  100  formed from a radially outer side  108  generally opposite to a radially inner side  110  and configured to match the row one blade annulus. The radially outer side  108  may be positioned radially outward a distance equal to the OD of an adjacent row one turbine blade. The radially inner side  110  may be positioned radially outward a distance equal to the ID of an adjacent row one turbine blade. The radially outer and inner sides  108 ,  110  may be coupled together with opposed first and second side walls  112 ,  114 . The outlet  100  may be offset from the inlet  98  in the longitudinal direction. The term “offset” as used herein and in the claims means that the outlet is spaced from the inlet as measured along the coordinate direction(s) identified. The outlet  100  may also be offset from the inlet  98  in a tangential direction  106 , as shown in  FIG. 4 . The outlet  100  may also be configured such that the outlet  100  is generally orthogonal to a longitudinal axis  136  of the turbine engine such that the transition duct  94  does not interfere with the row one turbine blades, as shown in  FIG. 7 . 
     The transition duct  94  may be configured to direct gases along a generally linear flow path along the transition axis  130 . In one embodiment, the transition duct  94  may have a generally cylindrical inlet  98  adjacent to a conical midsection  132 . The conical midsection  132  may be positioned between the inlet  98  and the throat  134 . The conical midsection  132  may include an ever decreasing cross-sectional area until the conical midsection  132  joins an adjacent throat  134 . The conical midsection  132  accelerates the flow of gases before the gases are directed into the row one turbine blades  140 , as shown in  FIG. 6 . Accelerating the flow of gases before the gases strike the row one turbine blades increases the efficiency of the turbine engine. The throat  134  may have any appropriate cross-section. In at least one embodiment, the throat  134 , as shown in  FIG. 8 , may have a cross-section with two opposing, generally linear sides and two opposing, non-linear sides. The cross-sectional area of the throat  134  may be less than a cross-sectional area of the conical midsection  132 . 
     As shown in  FIG. 9 , the transition ducts  94  may be formed from first and second opposing side walls  112 ,  114 . A first side wall  112  of a first transition duct  94  may be positioned such that an inner surface of the first side wall is coplanar with an inner surface of the second side wail  114  of an adjacent transition duct  94 . As such, the gas flows through each transition duct  94  are generally parallel to each other and immediately adjacent to each other without an area of decreased fluid flow between adjacent flows. Instead, the gas flows emitted from each of the transition ducts  94  are parallel and touching each other. 
     As shown in  FIG. 4 , the first side wall  112  may be canted relative to a radial axis  104  when viewing the outlet longitudinally upstream. The second side wall  114  may be canted relative to the radial axis  104  when viewing the outlet longitudinally upstream. In one embodiment, the first and second side walls  112 ,  114  may be canted between about 20 and about 70 degrees relative to the radial axis  104  when viewing the outlet longitudinally upstream. More particularly, the first and second side walls  112 ,  114  may be canted between about 30 and about 60 degrees relative to the radial axis  104  when viewing the outlet longitudinally upstream. In one embodiment, as shown in  FIG. 4 , the second side wall  114  may be nonparallel to the first side wall  112  of the outlet  100 . 
     The first and second side walls  112 ,  114  may be canted as shown in  FIGS. 4 and 5  to reduce the affects of the pressure differential between high pressure regions, denoted by the plus sign  116 , and the low pressure regions, denoted by the minus sign  118 . The high and low pressure regions  116 ,  118  exist within the same transition but in different portions of the cross-section. Such is the case because as the hot combustor gases flow quickly and accelerate through the transition  94 . 
     Inclusion of the canted first and second sides  112 ,  114  in the outlet  100  facilitates an increased incidence angle  142 , which is the angle shown in  FIG. 7  between an axis orthogonal to the longitudinal axis  136  and the linear flow path at the outlet  100  of the transition  94 . A higher incidence angle, which is an angle at which the discharge gas flow path is moving further way from alignment with the longitudinal axis  136 , facilitates positioning the transition duct  94  at improved angles of discharge of the combustor gases to downstream turbine blades. 
     During operation, hot combustor gases flow from a combustor into inlets  98  of the transitions  94 . The gases are directed through the internal passages  102 . The position of the transition duct  94  is such that gases are directed through the inlet  98 , the conical midsection  132 , and the adjacent throat  134  and are expelled out of the outlet  100 . The gases are expelled at a proper orientation relative to the turbine blades such that the gases are directed into the turbine blades in correct orientation without need of row one turbine vanes to alter the flow of the gases. Thus, energy is not lost through use of row one turbine vanes. In transition ducts  94  with linear flow paths, the gases are exhausted through the outlets  100 . The canted first and second sides  112 ,  114  of the outlet  100  distribute the wake across a downstream turbine blade. In particular, the wake is distributed from a pressure side, across a leading edge of the blade, to a suction side, thereby distributing the wake across the entire blade. Such a configuration reduces vibrations and stresses in the downstream, stationary turbine blades. 
     The foregoing is provided for purposes of illustrating, explaining, and describing embodiments of this invention. Modifications and adaptations to these embodiments will be apparent to those skilled in the art and may be made without departing from the scope or spirit of this invention.