Abstract:
A midframe portion ( 313 ) of a gas turbine engine ( 310 ) is presented and includes a compressor section with a last stage blade to orient an air flow ( 311 ) at a first angle ( 372 ). The midframe portion ( 313 ) further includes a turbine section with a first stage blade to receive the air flow ( 311 ) oriented at a second angle ( 374 ). The midframe portion ( 313 ) further includes a manifold ( 314 ) to directly couple the air flow ( 311 ) from the compressor section to a combustor head ( 318 ) upstream of the turbine section. The combustor head ( 318 ) introduces an offset angle in the air flow ( 311 ) from the first angle ( 372 ) to the second angle ( 374 ) to discharge the air flow ( 311 ) from the combustor head ( 318 ) at the second angle ( 374 ). While introducing the offset angle, the combustor head ( 318 ) at least maintains or augments the first angle ( 372 ).

Description:
STATEMENT REGARDING FEDERALLY SPONSORED DEVELOPMENT 
       [0001]    Development for this invention was supported in part by Contract No. DE-FC26-05NT42644, awarded by the United States Department of Energy. Accordingly, the United States Government may have certain rights in this invention. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The invention relates to can-annular gas turbine engines, and more specifically, to a midframe portion of a can-annular gas turbine engine. 
       BACKGROUND OF THE INVENTION 
       [0003]    A conventional midframe design for a can-annular gas turbine engine is discussed in U.S. Pat. No. 7,721,547 (“&#39;547 Patent”), assigned to the assignee of the present invention, which is incorporated by reference herein. FIG. 1 of the &#39;547 Patent is reproduced as  FIG. 1  herein, and illustrates a cross-section through a midframe portion  13  of a conventional can-annular gas turbine engine  10 . The major components of the gas turbine engine  10  are a compressor section  12 , a combustion section  16  and a turbine section  48 . A rotor assembly  17  is centrally located and extends through the three sections. In operation, the compressor section  12  receives air through an intake (not shown) and compresses it. The compressed air flow  11  passes from the compressor section  12  to an axial diffuser  14 , after which the air flow  11  enters a chamber  15  within a casing  19 , where the total air flow  11  is separated and enters one of multiple combustor heads  18  of the can-annular combustion section  16  that encircle the rotor assembly  17  in an annular configuration. 
         [0004]    As illustrated in  FIG. 1 , the compressor section  12  includes cylinders  27 , 29  that enclose alternating rows of stationary vanes  23  and rotating blades  25 . The stationary vanes  23  can be affixed to the cylinder  27  while the rotating blades  25  can be mounted to the rotor assembly  17  for rotation with the rotor assembly  17 . The stationary vanes  23  include a last stationary vane  26  and an outlet guide vane  28  positioned adjacent to an outlet of the compressor section  12 . Additionally, the rotating blades  25  include a last stage blade  24  positioned upstream from the last stationary vane  26  and the outlet guide vane  28 . The last stationary vane  26  and outlet guide vane  28  are used to remove an absolute tangential swirl angle (measured in an absolute reference frame with respect to the longitudinal direction) of the air flow  11  coming off the last stage blade  24 . 
         [0005]    As further illustrated in  FIG. 1 , load-bearing struts  30  are provided to support a shaft cover  32  of the rotor assembly  17  at the casing  19  of the combustion section  16 . As appreciated by one of skill in the art, one strut  30  may be provided per each one to four combustor heads  18 . As illustrated in  FIG. 1 , the axial diffuser  14  includes an inner cone  36  and an outer cone  34  and the cross-sectional area between the inner and outer cones  36 , 34  increases in the longitudinal direction  68 , such that the air flow  11  expands and decelerates through the diffuser  14 , thereby converting velocity head into pressure head. As illustrated in  FIG. 1 , the strut  30  is attached between a shaft cover  32  to the outer cone  34  of the axial diffuser  14 , and thus the casing  19  of the combustion section  16  supports the strut  30  at the shaft cover  32 . 
         [0006]    As further illustrated in  FIG. 1 , a rotor-cooling extraction pipe  38  is provided, which extracts compressed air from the chamber  15  and passes the compressed air into a cooler  42 . The cooled air passes from the cooler  42  and through rotor-cooling injection pipes  40  that are positioned within the chamber  15  and direct the cooled air below the shaft cover  32 , to cool the rotating components of the engine. 
         [0007]    Another portion of the engine needing cooling is a turn in the transition  20  at an inlet to the turbine section  48 , which typically experiences an especially high heat flux during an operation of the gas turbine engine  10 . In order to cool a rear end  54  of the transition  20  during operation of the gas turbine  10 , a portion  58  of the air flow  11  entering the chamber  15  makes contact with the rear end  54  of the transition  20  proximate the highest heat flux region in order to cool the rear end  54  of the transition  20  using thermal convection. 
         [0008]    FIG. 8 of the &#39;547 Patent is reproduced herein as  FIG. 2 , and illustrates a “trans-vane” transition  20 ′ which improves upon the transition  20  of  FIG. 1 .  FIG. 2  illustrates a top down radial view of the midframe portion  13 ′ of the gas turbine engine  10 ′ including the combustion section  16 ′ and a first stage turbine blade array  49 ′ of the turbine section  48 ′ located downstream from the combustion section  16 ′, with the trans-vane transition  20 ′ located therebetween. The midframe portion  13 ′ of  FIG. 2  includes a compressor section (not shown) similar to the compressor section  12  of  FIG. 1 . A first stage housing encloses the first stage turbine blade array  49 ′ and includes a blade ring  51 ′. An upstream side  53 ′ of the blade ring  51 ′ is preferably adapted to couple to a transition outlet  55 ′. The trans-vane transition  20 ′ includes a transition duct body  60 ′ with an inlet  62 ′ to receive a gas flow exhausted from the combustor section  16 ′ and the outlet  55 ′ to discharge a gas flow toward the first stage blade array  49 ′ with an internal passage  66 ′ therebetween. The outlet  55 ′ is offset from the inlet  62 ′ in the three coordinate directions—in the radial direction (in/out of the figure), the longitudinal direction  68  and the tangential direction  70 ′. The gas flow discharged from the outlet  55 ′ is angled in the tangential direction  70 ′ within an absolute reference frame, relative to the longitudinal direction  68  as depicted by the arrow  72 ′, as required by the first stage turbine blade array  49 ′. A brief discussion will be provided of the absolute and relative reference frames of the midframe portion  13 ′, as well as how the velocity vector of an air flow exiting the compressor and entering the turbine  48 ′ of the gas turbine engine  10 ′ is represented in each of those reference frames.  FIG. 3  illustrates a top down radial view of the last stage blade  24  of the compressor section  12  of the gas turbine engine  10 ′ and the first stage blade  49 ′ of the turbine  48 ′ of the midframe portion  13 ′, separated along a longitudinal axis  75  of the conventional gas turbine engine  10 ′ of  FIG. 2 . An outgoing air flow off the last stage blade  24  is oriented in the (relative) reference frame of the last stage blade  24  along a relative outgoing velocity vector  76 . During an operation of the compressor section  12 , the last stage blade  24  rotates around the longitudinal axis  75  with a blade velocity vector  78  that is oriented perpendicular to the longitudinal axis  75 . In order to determine the velocity vector of the outgoing air flow off the last stage blade  24  in an absolute reference frame, the blade velocity vector  78  is added to the relative outgoing velocity vector  76 , resulting in an absolute outgoing velocity vector  80  that is angled in the tangential direction  70  by an angle  82 , relative to the longitudinal direction  68 . In an exemplary embodiment, the angle  82  is approximately 45 degrees. Accordingly, the absolute outgoing velocity vector  80  of the outgoing air flow off the last stage blade  24  is oriented approximately 45 degrees in the tangential direction  70 , relative to the longitudinal direction  68 . The last stage vanes  26 , 28  of the conventional midframe portion  13 ′ are configured to reduce the angle  82  of the absolute outgoing velocity vector  80  from 45 degrees to approximately 0 degrees, to align the air flow along the longitudinal axis  75 . However, as discussed below, the embodiments of the present invention do not utilize the last stage vanes, and thus utilize the initial angle  82  of the absolute outgoing velocity vector  80  off the last stage blade  24 .  FIG. 3  also illustrates an incoming air flow to the first stage blade  49 ′ of the turbine  48 ′ illustrated in  FIG. 2 . In order to maximize the effectiveness of the turbine  48 ′, the incoming air flow is oriented in the (relative) reference frame of the first stage blade  49 ′ along a relative incoming velocity vector  84 . During an operation of the turbine  48 ′, the first stage blade  49 ′ rotates around the longitudinal axis  75  with a blade velocity  86  that is oriented perpendicular to the longitudinal axis  75 . In order to determine the velocity vector of the incoming air flow in the absolute reference frame, the blade velocity vector  86  is added to the relative incoming velocity vector  84 , resulting in an absolute incoming velocity vector  88  that is angled in the tangential direction  70  by an angle  90 , relative to the longitudinal direction  68 . In an exemplary embodiment, the angle  90  is approximately 70 degrees. Accordingly, the absolute incoming velocity vector  88  of the incoming air flow onto the first stage blade  49 ′ of the turbine  48 ′ is oriented approximately 70 degrees in the tangential direction  70 , relative to the longitudinal direction  68 . In contrast with the transition  20 ′ of  FIG. 2 , the transition  20  illustrated in  FIG. 1  discharges a gas flow to the turbine section  48  with an offset in only the radial direction and the longitudinal direction  68 , and thus the gas flow is not angled in the tangential direction relative to the longitudinal direction  68 . Since the first stage turbine blade array  49  of the turbine section  48  requires an incoming gas flow that is angled in the tangential direction relative to the longitudinal direction  68 , the turbine section  48  of  FIG. 1  includes a first stage vane  74 , to introduce an offset in the tangential direction for the gas flow discharged from the transition  20 . However, by implementing the trans-vane design in the transition  20 ′, the gas flow is discharged from the outlet  55 ′ at the necessary angle  90  in the tangential direction  70  relative to the longitudinal direction  68  to accommodate the first stage turbine blade array  49 ′, and thus the first stage vanes  74  are not needed. In the &#39;547 Patent, the inventors made various improvements to the midframe portion of the gas turbine engine, downstream of the combustion section, to enhance the operating efficiency of the gas turbine engine. In the present invention, the present inventors make various improvements to the midframe portion of the gas turbine engine, upstream of the combustion section, to also enhance the operating efficiency and/or cost efficiency of the gas turbine engine. 
       SUMMARY OF THE INVENTION 
       [0009]    The present inventors have recognized that significant improvements in the operating efficiency of a can annular gas turbine engine may be obtained by innovation in the design of the mid-section of the engine. The inventors have recognized that the movement of air from the compressor section to the combustor section in a can-annular gas turbine engine is a generally unstructured, chaotic process. Compressed air produced by the compressor section is directed into the annular chamber  15  and is allowed to find its path of least resistance around various structural obstacles and into one of the respective combustor heads  18 . As a result, the flow experiences turbulence and fluid friction induced pressure losses. The present inventors have recognized that an improved engine mid-section design can minimize such losses, by minimizing swirl reduction, thereby providing improved overall engine performance. 
         [0010]    The air flow  11  experiences aerodynamic loss based on a total angle of rotation while traveling from the compressor section  12  to one of the combustor heads  18 . The last stationary vane  26  and outlet guide vane  28  collectively rotate the air flow  11  by an initial absolute tangential swirl angle to remove the initial absolute tangential swirl angle that is imparted by the rotating blades  24 , such as 45 degrees, to align the air flow  11  in a downstream longitudinal direction  68  upon entering the axial diffuser  14 . In addition to the initial absolute tangential swirl angle rotation, upon exiting the diffuser  14  and entering the chamber  15 , the air flow  11  experiences two 180 degree rotations: a first approximate 180 degree rotation to orient the air flow  11  from an initial downstream longitudinal direction  68  to an upstream longitudinal direction to travel longitudinally backward to a respective combustor head  18 , and a second approximate 180 degree rotation at the combustor head  18  to direct the air flow  11  into an inlet of the combustor head  18 . Accordingly, the air flow  11  undergoes a total rotation of approximately 400 degrees while traveling from the compressor section  12  to one of the combustor heads  18 , and most of this rotation is accomplished in the unstructured environment of the chamber  15 . The present inventors have recognized that the aerodynamic efficiency of the air flow from the compressor section to the respective combustor head can be enhanced by reducing this total rotation of the air and/or controlling the rotation process more precisely. For example, a radial velocity component may be introduced to the air flow at the diffuser outlet, such that the air flow is a mixed-flow (which has combined longitudinal, tangential and radial velocity components) at the diffuser outlet. By introducing the radial velocity component to the air flow at the diffuser outlet, the required total angle of rotation within the chamber  15  will be decreased, and the aerodynamic efficiency of the air flow from the compressor section to the respective combustor head would be enhanced. As discussed above, the last stage vane and outlet guide vane of a conventional can-annular gas turbine engine  10  are provided to eliminate the initial tangential swirl angle of approximately 45 degrees that is imparted by the rotating compressor blades  24 , so that the air flow is directed into the chamber  15  along a downstream longitudinal direction (0 degree tangential swirl angle). The present inventors have recognized that some transition designs, such as the trans-vane design discussed above in  FIG. 2 , require a final absolute tangential swirl angle (with respect to the longitudinal direction  68 ) to be output from the transition  20 ′ to the first stage blade  49 ′ of the turbine section  48 ′ which is greater than the initial absolute tangential swirl angle generated by the last stage blade of the compressor section. For example, in an exemplary embodiment, a trans-vane design requires a final absolute tangential swirl angle of 70 degrees to the first stage blade  49 ′ of the turbine section  48 ′, in excess of the initial absolute tangential swirl angel of 45 degrees off the last stage blade  24  of the compressor section. The present inventors have recognized that it would be advantageous to maintain and then augment the initial absolute tangential swirl angle off the last stage blade of the compressor section rather than the conventional practice of eliminating it through the last stage compressor vane  26  and outlet guide vane  28  and replacing it through the transition  20 ′ to the first stage blade  49 ′ of the turbine section  48 ′. The inventors have recognized that the initial absolute tangential swirl angle off the last stage blade of the compressor section can be increased by an offset angle through a trans-vane transition so that the absolute tangential swirl angle of the air flow output from the trans-vane transition fulfills the required final absolute tangential swirl angle incident on the first stage blade  49 ′. This offset angle through the transition may be 25 degrees, for example, which is additive to the 45 degree angle provided by the last stage compressor blades  24  to achieve a required 70 degree tangential angle at the first stage turbine blade  49 ′, thereby eliminating the need for the last stage compressor vane  26  and outlet guide vane  28 . By maintaining or supplementing the initial absolute tangential swirl angle of the air flow off the last stage blade of the compressor section, the air flow can be passed from the compressor outlet to a respective combustor head with a substantially reduced total angle of rotation when compared to the conventional design, thereby enhancing the aerodynamic efficiency of the midframe portion of the gas turbine engine. 
         [0011]    The present inventors have also recognized that the aerodynamic efficiency of the air flow  11  from the compressor section  12  to the combustor heads  18  may be enhanced by directly coupling the air flow from respective sections of the compressor-diffuser outlet to each individual combustor head inlet. In the conventional gas turbine, the compressed air flow  11  passes from the compressor section  12  to the axial diffuser  14 , after which the air flow  11  enters the chamber  15  within the casing  19 , where the air flow  11  experiences aerodynamic loss in the process of randomly entering any of the multiple combustor heads  18  within the chamber  15 . Upon exiting the diffuser  14  and entering the chamber  15 , the air flow  11  also experiences aerodynamic losses as a result of making contact with the load-bearing struts  30 , the rotor-cooling injection pipes  40  and a near side  22  of the transition  20 , which are each positioned adjacent to the outlet of the diffuser  14 . By directly coupling the air flow  11  from the compressor-diffuser outlet to a respective combustor head inlet as described herein, the air flow avoids entering the chamber altogether, thereby allowing for the reduction of the aerodynamic losses associated with randomly entering one of the multiple combustor heads  18 . 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0012]    The invention is explained in the following description in view of the drawings that show: 
           [0013]      FIG. 1  is a cross-sectional view of a portion of a conventional turbine engine; 
           [0014]      FIG. 2  is a cross-sectional view of a trans-vane design of a transition of a conventional turbine engine; 
           [0015]      FIG. 3  is a radial view of a last compressor blade and a first turbine blade of the conventional turbine engine of  FIG. 2 ; 
           [0016]      FIG. 4  is a partial longitudinal view of a linear diffuser duct within a turbine engine; 
           [0017]      FIG. 5  is a partial radial view of the linear diffuser duct illustrated in  FIG. 4 ; 
           [0018]      FIG. 6  is a partial radial view of a helical diffuser duct within a turbine engine; 
           [0019]      FIG. 7  is a partial longitudinal view of the helical diffuser duct illustrated in  FIG. 6 ; 
           [0020]      FIG. 8  is a plot of total pressure and static pressure of an air flow versus a longitudinal location of the air flow along a gas turbine engine; 
           [0021]      FIG. 9  is a partial radial view of a helical duct within a turbine engine; and 
           [0022]      FIG. 10  is a partial longitudinal view of the helical duct illustrated in  FIG. 9 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0023]    As discussed above, the inventors of the present invention recognized that an improved midframe portion of the gas turbine engine features initiating a mixed air flow (axial, tangential plus radial flow velocities) from the diffuser outlet. By initiating the mixed-air flow from the diffuser outlet, the air flow passes from the diffuser outlet to the combustor head inlet while undergoing a reduced total angle of rotation when compared to the air flow with the conventional midframe portion. 
         [0024]    A midframe design is provided, in which a mixed air flow can be initiated within a diffuser of the midframe portion, downstream of the traditional axial compressor section.  FIG. 4  illustrates a longitudinal cross-sectional view of a midframe portion  313  of a gas turbine engine  310  within a radial-tangential plane, in which a compressor section (not shown) compresses an air flow into an annulus  329  downstream of the compressor section. A plurality of diffuser ducts  314  are positioned in an annular configuration around the rotor assembly (not shown) of the gas turbine engine  310 . The diffuser duct  314  receives compressed air from the annulus  329  and the diffuser duct  314  is directed tangentially in a tangential direction  70  at an angle  373  which is oriented 90 degrees from the radial direction  69 . Accordingly, an air flow  311  from the annulus  329  is respectively directed in a mixed flow direction into each respective diffuser duct  314 , with both a radial and tangential velocity components based on the angle  373  and a longitudinal velocity component along the longitudinal direction (not shown). The diffuser duct  314  is discussed in greater detail below, and includes features which minimize the required total angle of rotation of the mixed air flow  311  as it passes from the compressor section outlet to the combustor  318 , to the transition  320  and eventually to the annulus  331  and the turbine section of the gas turbine engine  310 . The above discussion of the diffuser duct  314  establishes that a mixed-air flow  311  can be established within the diffuser duct  314  of the midframe portion  313  of the gas turbine engine  310 . 
         [0025]      FIG. 5  illustrates a radial view of the diffuser duct  314  of  FIG. 4 , in the longitudinal-tangential plane, where the diffuser duct  314  extends from the annulus  329  at the compressor section outlet, to the respective combustor head  318  inlet of the combustor  316 . The last stage blade (not shown) of the compressor section in the midframe portion  313  of  FIG. 5  is similar to the last stage blade  24  of the midframe portion  13 ′ illustrated in  FIG. 3 , and thus the air flow  311  comes off the last stage blade (not shown) of the midframe portion  313  with an absolute outgoing velocity vector oriented at an angle  372  in the tangential direction  70  with respect to the longitudinal direction  68 . In an exemplary embodiment, the angle  372  is approximately 45 degrees. To accommodate the absolute outgoing velocity vector oriented at the angle  372 , the diffuser duct  314  is a straight duct that is also angled at the angle  372 , to receive the air flow  311 . Also, the first stage blade (not shown) of the turbine in the midframe portion  313  of  FIG. 5  is similar to the first stage blade  49 ′ of the midframe portion  13 ′ illustrated in  FIG. 3 , and thus the air flow  311  is incident on the first stage blade  49 ′ with an absolute incoming velocity vector oriented at an angle  374  in the tangential direction  70  with respect to the longitudinal direction  68 . In an exemplary embodiment, the angle  374  is approximately 70 degrees. To accommodate the absolute incoming velocity vector oriented at the angle  374 , a trans-vane transition  320  is a straight duct that is also angled at the angle  374 , to receive the air flow  311  from the combustor head  318  at the angle  374 . Thus, the air flow  311  experiences an offset in the absolute velocity vector from the angle  372  to the angle  374  at the combustor head  318 . Although the angular offset from the angle  372  to the angle  374  occurs at the combustor head  318 , the angular offset is relatively small, such as 25 degrees, for example, in comparison with the total angle of rotation of the air flow from the compressor outlet to the combustor head inlet in a conventional midframe portion, such as 400 degrees, for example. As illustrated in  FIG. 5 , the diffuser duct  314  is designed to accommodate an air flow from an annulus  329  through a straight tubular duct and to an outlet, while the transition  320  is designed with a reverse design to the diffuser duct  314 , as it accommodates an air flow from an inlet through a straight tubular duct and to an annulus  331  within the turbine section. As illustrated in  FIG. 4 , an injector  325  is positioned to pass a volume of fuel  327  into the combustor head  318 , which is mixed with the air flow  311  and the air-fuel mixture is subsequently ignited. As illustrated in 
         [0026]      FIG. 5 , an outlet of the diffuser duct  314  encloses the inlet of the combustor head  318  inlet, since the outer diameter  342  of the diffuser duct  314  outlet is greater than the outer diameter  344  of the combustor head  318  inlet. 
         [0027]    In addition to the angular offset of the air flow  311  traveling within the longitudinal-tangential plane ( FIG. 5 ) of the midframe portion  313 , the air flow  311  experiences an angular offset in the radial-tangential plane of the midframe portion  313 , and both of these angular offsets are combined to determine the total angle of rotation of the air flow  311  while passing from the compressor outlet to the respective combustor head  318  inlet.  FIG. 4  illustrates the midframe portion  313 , within the radial-tangential plane, in which the air flow  311  passes from the annulus  329  at the compressor outlet (not shown) in the tangential direction  70  oriented at an angle  373  which is 90 degrees with respect to the radial direction  69 . The air flow  311  passes within the diffuser duct  314 , that is similarly angled in the radial-tangential plane at the angle  373 , and enters the combustor head  318  inlet in a mixed-flow direction with combined radial and tangential velocity components. The air flow  311  emerges from the combustor head  318  inlet in a mixed-flow direction with combined radial and tangential velocity components at an angle  375  and passes within the transition  320 , that is similarly angled in the radial-tangential plane at the angle  375 . The air flow  311  subsequently exits the transition  320  at an annulus  331  in a tangential direction  70  that is oriented at the angle  375  which is 90 degrees with respect to the radial direction  69  at the transition  320  exit. Upon entering the annulus  331 , the air flow  311  is directed at the first stage blades of the turbine (not shown), at the appropriate angle  374  in the longitudinal-tangential plane ( FIG. 5 ). The angular offset of the air flow  311  in the radial-tangential plane is based on such factors as the radial height of the compressor section outlet, the radial height of the combustor head  318  and the radial height of the turbine inlet. For purposes of  FIG. 4 , it is presumed that the height of the last stage blade of the compressor section is less than the height of the first stage blade of the turbine within the casing  319 , and thus the air flow  311  enters the diffuser duct  314  at a reduced radial height than the air flow  311  exits the transition  320 . In an exemplary embodiment, the angle  373  may be 90 degrees, while the angle  375  may be 90 degrees, and thus the air flow  311  would undergo an approximate 90 degree rotation in the radial-tangential plane, while traveling from the compressor section outlet and into the combustor head  318  inlet. 
         [0028]    As previously discussed, the integrated diffuser duct  314  provides a substantial reduction in the total angle of rotation of the air flow  311 , as the air flow  311  passes from the compressor section outlet to the combustor head  318  inlet. The total angle of rotation of the air flow  311  includes the angle of rotation of the air flow  311  within the longitudinal-tangential plane ( FIG. 5 ) and the angle of rotation of the air flow  311  within the radial-tangential plane ( FIG. 4 ). As discussed above, in an exemplary embodiment, the angle of rotation of the air flow  311  within the longitudinal-tangential plane ( FIG. 5 ) may be approximately 25 degrees, for example. Also, as discussed above, in an exemplary embodiment, the angle of rotation of the air flow  311  within the radial-tangential plane ( FIG. 4 ) may be approximately 90 degrees, for example. Thus, as a result of using the integrated diffuser duct  314 , the total angle of rotation for the air flow  311  passing from the compressor outlet to the combustor head  318  inlet is approximately 115 degrees. This total angle of rotation is substantially less than the approximate 400 degree total angle of rotation of the air flow passing from the compressor outlet to the combustor head inlet in a conventional midframe portion of the gas turbine engine. Indeed, the diffuser duct  314  enhances the aerodynamic efficiency of the midframe portion  313  of the gas turbine engine  310 . 
         [0029]      FIGS. 6-7  illustrate an alternate embodiment of the midframe portion  313 ′ of the gas turbine engine  310 ′, which is similar to the midframe portion  313  of  FIGS. 4 and 5 , with the exception that the diffuser duct  314 ′ has an alternate design than the diffuser duct  314  of  FIGS. 4 and 5 . As previously discussed, the diffuser duct  314  of  FIGS. 4 and 5  features a straight tubular configuration, which directs the air flow  311  to the combustor head  318  inlet at the angle  372  in the tangential direction  70  relative to the longitudinal direction  68 , after which the air flow  311  undergoes an angular offset at the combustor head  318  inlet to the angle  374  in the tangential direction  70  relative to the longitudinal direction  68 , before passing through to the transition  320 . Thus, the angular offset of the air flow  311  from the angle  372  to the angle  374  occurs at the combustor head  318  inlet. In contrast with the diffuser duct  314  of  FIGS. 4 and 5 , the diffuser duct  314 ′ of  FIGS. 6-7  takes a helical shape rather than a straight tubular configuration, where the inlet of the diffuser duct  314 ′ at the annulus  329  is aligned at the angle  372  in the tangential direction  70  relative to the longitudinal direction  68 , while the outlet of the diffuser duct  314 ′ at the combustor head  318  inlet is aligned at the angle  374  in the tangential direction  70  relative to the longitudinal direction  68 . Thus, the angular offset of the air flow  311  from the angle  372  to the angle  374  occurs over the length of the helical shape of the diffuser duct  314 ′ between the inlet at the annulus  329  and the outlet at the combustor head  318  inlet. Since the required angular offset from the angle  372  to the angle  374  occurs over the length of the diffuser duct  314 ′, the angular offset need not occur at the combustor head  318  inlet. Accordingly, the air flow  311  passing from the outlet of the diffuser duct  314 ′ into the combustor head  318  inlet experiences minimal angular offset. The outlet of the diffuser duct  314 ′ in  FIGS. 6-7  is attached to the combustor head  318  inlet such that the face  321 ′ of the diffuser duct  314 ′ outlet is aligned parallel with the face  323  of the combustor head  318  inlet. As with the midframe portion  313  illustrated in  FIGS. 4 and 5 , the outer diameter  342  of the diffuser duct  314 ′ outlet is greater than the outer diameter  344  of the combustor head  318  inlet, such that the diffuser duct  314 ′ outlet encloses the combustor head  318  inlet. As illustrated in  FIG. 7 , since the diffuser duct  314 ′ takes a helical shape in the radial-tangential plane, the diffuser duct  314 ′ may rise to a peak radial height that is greater than a peak radial height of the diffuser duct  314  with the straight tubular configuration illustrated in  FIG. 4 . Thus, in order to accommodate the greater peak radial height of the diffuser duct  314 ′, the midframe portion  313 ′ may feature a larger casing  319 ′ than the casing  319  of  FIG. 4 , such that the diffuser duct  314 ′ has adequate radial space within the casing  319 ′ to pass the mixed-air flow  311  from the compressor outlet to the combustor head  318  inlet. Based on the larger casing  319 ′ of the midframe portion  313 ′ and the smaller casing  319  of the midframe portion  313 , the manufacturing cost efficiency for the midframe portion  313  may be greater than the midframe portion  313 ′. However, as previously discussed, the diffuser duct  314  takes on a straight tubular form which directs the angular offset from the angle  372  to the angle  374  at the outlet of the diffuser duct  314 , while the diffuser duct  314 ′ takes on a helical form which directs the angular offset from the angle  372  to the angle  374  along the length of the diffuser duct  314 ′, and thus the aerodynamic efficiency of the midframe portion  313 ′ may be greater than the midframe portion  313 . 
         [0030]    The integrated diffuser duct embodiments of  FIGS. 4-7  are used to reduce a dynamic pressure of the air flow and simultaneously increase a static pressure of the air flow, as the air flow passes through the diffuser duct.  FIG. 8  illustrates a graph of the total pressure  502  and the static pressure  504  of the air flow  311  within the diffuser duct of  FIGS. 4 and 5 , as well as an intermediate pressure  506  outside the diffuser duct in the casing  319 . As appreciated by one of skill in the art, the total pressure  502  of an air flow is the sum of the static pressure  504  of the air flow and a dynamic pressure of the air flow. Thus, the dynamic pressure of the air flow  311  can be determined by the difference between the total pressure  502  and the static pressure  504  of the air flow  311 .  FIG. 8  depicts the various pressures at different locations throughout the midframe portion  313  of the gas turbine engine  310 , including the last stage blade of the compressor  508 , the combustor head inlet  510 , the combustor head outlet  512 , and the trans-vane transition outlet  514 . As illustrated in  FIG. 8 , as the air flow  311  passes through the diffuser duct  314  between the compressor last stage blade  508  and the combustor head inlet  510 , the air flow  311  decelerates, thereby reducing the dynamic pressure, and thus increasing the static pressure  504 . The diffuser duct  314  gradually reduces the dynamic pressure of the air flow  311 , and thus effectively converts most or all of the dynamic pressure of the air flow  311  into static pressure  504 . In contrast, the conventional midframe portion  13  of  FIG. 1  discharges the air flow  11  into the chamber  15 , where the air flow  11  suddenly loses a large amount of dynamic pressure which may not be converted back to dynamic pressure and thus is lost in the process of discharging the air flow  11  into the chamber  15 . Upon entering the combustor head inlet  510 , the fuel-air mixture in the combustor head  318  is ignited, which accelerates the air flow  311  through the combustor outlet  512  and through the outlet of the transition  514 , thereby increasing the dynamic pressure, and thus decreasing the static pressure  504 . Throughout the operation of the midframe portion  313 , a pressure within the casing  319  is set at the intermediate pressure  506 , which is less than the static pressure  504  at each location throughout the midframe portion  313 . As appreciated by one of skill in the art, stagnant air may collect at the interfaces between the diffuser duct  314  and the combustor head  318 , and the combustor head  318  and the transition  320 . By adjusting the intermediate pressure  506  within the casing  319  to be less than the static pressure  504  of the air flow  311  within the diffuser duct  314 , the combustor head  318  or the transition  320 , a leakage of air flow will pass across these interfaces, to discharge any stagnant air within the interfaces, and thus prevent hot, stagnant air from building up at these interfaces. 
         [0031]      FIGS. 9-10  illustrate an alternate embodiment of the midframe portion  313 ″ of the gas turbine engine  310 ″, which is similar to the midframe portion  313 ′ of the gas turbine engine  310 ′ depicted in  FIGS. 6-7 , with the mixed-air flow  311  being passed through a manifold  314 ″, which does not act as a diffuser, but instead accelerates a velocity of the mixed-air flow  311  from the annulus  329  at the compressor outlet to the combustor head  318  inlet. Unlike the diffuser duct  314 ′ of  FIGS. 6-7 , the outer diameter  342 ″ of the manifold  314 ″ is less than the outer diameter  344 ″ of the combustor head  318  inlet, such that the manifold  314 ″ outlet is positioned within the inlet of the combustor head  318 . In contrast, in the above embodiments of  FIGS. 4-7 , the outer diameter  342  of the diffuser duct is greater than the outer diameter  344  of the combustor head  318  inlet, so that the diffuser duct outlet encloses the combustor head  318  inlet. As with the midframe portions  313 ,  313 ′ of  FIGS. 4 and 7 , the midframe portion  313 ″ of  FIG. 9  includes an injector  325  to pass a volume of fuel  327  into the combustor head  318 . The fuel  327  may be a reactive fuel, such as hydrogen, for example, which exhibits high flame speeds, and thus requires a high incoming velocity of the mixed-air flow  311  into the combustor head  318 , to avoid flashback. By sizing the outer diameter  342 ″ of the manifold  314 ″ to be less than the outer diameter  344 ″ of the combustor head  318  inlet, and positioning the manifold  314 ″ outlet within the combustor head  318  inlet, the velocity of the mixed-air flow  311  into the combustor head  318  inlet is increased, to the high incoming velocity threshold of the combustor head  318 , to avoid flashback when reactive fuel  327  is passed into the combustor head  318  through the injector  325 . 
         [0032]    While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.