Patent Abstract:
An apparatus and method are disclosed for a gas turbine engine including an offtake located within the air flow of the engine. The offtake has an inlet and a louver covering the inlet. The louver has multiple airfoils arranged to direct the air flow into the inlet of the offtake.

Full Description:
BACKGROUND 
       [0001]    Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine onto a multitude of rotating turbine blades. Typically a fan section is utilized to take in ambient air and direct it to different components of the engine for extracting energy and cooling purposes. Some of the fan air is initially directed into the compressor stages, while other portions of the fan air continue through outlet guide vanes and can later be directed into the engine components as needed. 
         [0002]    Gas turbine engines include offtakes in areas of the engine where air is extracted from high-velocity, swirling channels to the internal air system for cooling, sealing or heat management purposes. When the angle of redirection is 90° or higher louvers or other aerodynamic shapes are required to turn the flow effectively. The louvers are typically cascades of equal length, shape and camber angle. 
       BRIEF DESCRIPTION 
       [0003]    In one aspect, embodiments of relate to gas turbine engine comprising an annular fan exhaust section, an engine core at least partially located within the fan exhaust section, a cooling air offtake located in the engine core and having an inlet, a louver located at the inlet and having at least two different size airfoils in spaced axial arrangement. 
         [0004]    In another aspect, embodiments relate to a louver assembly for an off take of a gas turbine engine comprising at least four airfoils in axial arrangement, with none of the airfoils are of the same size. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]    In the drawings: 
           [0006]      FIG. 1  is a schematic cross-sectional diagram of a gas turbine engine for an aircraft. 
           [0007]      FIG. 2  is an enlarged view of a fan exhaust section of the gas turbine engine of  FIG. 1 . 
           [0008]      FIG. 3  is an enlarged view of an inlet with a louver having multiple airfoils for a cooling offtake duct for the gas turbine engine of  FIG. 1 . 
           [0009]      FIG. 4A  is a flow diagram of a conventional louver assembly. 
           [0010]      FIG. 4B  is a flow diagram of an embodiment of the proposed louver assembly. 
       
    
    
     DETAILED DESCRIPTION 
       [0011]    The described embodiments of the present invention are directed to a gas turbine engine have a louver to redirect fan air. For purposes of illustration, embodiments of the present invention will be described with respect to the turbine for an aircraft gas turbine engine. It will be understood, however, that the embodiments of the invention are not so limited and may have general applicability within an engine, including compressors, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications. 
         [0012]    As used herein, the term “forward” or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine relative to the engine centerline. 
         [0013]    Additionally, as used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. 
         [0014]    All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, aft, etc.) are only used for identification purposes to aid the reader&#39;s understanding of the present embodiments, and do not create limitations, particularly as to the position, orientation, or use of the embodiments. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary. 
         [0015]      FIG. 1  is a schematic cross-sectional diagram of a gas turbine engine  10  for an aircraft. The engine  10  has a generally longitudinally extending axis or centerline  12  extending forward  14  to aft  16 . The engine  10  includes, in downstream serial flow relationship, a fan section  18  including a fan  20 , a compressor section  22  including a booster or low pressure (LP) compressor  24  and a high pressure (HP) compressor  26 , a combustion section  28  including a combustor  30 , a turbine section  32  including a HP turbine  34 , and a LP turbine  36 , and an exhaust section  38 . 
         [0016]    The fan section  18  includes a fan casing  40  surrounding the fan  20 . The fan  20  includes a plurality of fan blades  42  disposed radially about the centerline  12 . The fan casing  40  can also surround at least a portion of the fan exhaust section  41 . The HP compressor  26 , the combustor  30 , and the HP turbine  34  form a core  44  of the engine  10 , which generates combustion gases. The core  44  is surrounded by core casing  46 , which can be coupled with the fan casing  40 , so that the core  44  is at least partially located within the fan exhaust section  41 . 
         [0017]    A HP shaft or spool  48  disposed coaxially about the centerline  12  of the engine  10  drivingly connects the HP turbine  34  to the HP compressor  26 . A LP shaft or spool  50 , which is disposed coaxially about the centerline  12  of the engine  10  within the larger diameter annular HP spool  48 , drivingly connects the LP turbine  36  to the LP compressor  24  and fan  20 . 
         [0018]    The LP compressor  24  and the HP compressor  26  respectively include a plurality of compressor stages  52 ,  54 , in which a set of compressor blades  56 ,  58  rotate relative to a corresponding set of static compressor vanes  60 ,  62  (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage  52 ,  54 , multiple compressor blades  56 ,  58  can be provided in a ring and can extend radially outwardly relative to the centerline  12 , from a blade platform to a blade tip, while the corresponding static compressor vanes  60 ,  62  are positioned upstream of and adjacent to the rotating blades  56 ,  58 . It is noted that the number of blades, vanes, and compressor stages shown in  FIG. 1  were selected for illustrative purposes only, and that other numbers are possible. 
         [0019]    The blades  56 ,  58  for a stage of the compressor can be mounted to a disk  59 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 , with each stage having its own disk  59 ,  61 . The vanes  60 ,  62  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
         [0020]    The HP turbine  34  and the LP turbine  36  respectively include a plurality of turbine stages  64 ,  66 , in which a set of turbine blades  68 ,  70  are rotated relative to a corresponding set of static turbine vanes  72 ,  74  (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage  64 ,  66 , multiple turbine vanes  72 ,  74  can be provided in a ring and can extend radially outwardly relative to the centerline  12 , while the corresponding rotating blades  68 ,  70  are positioned downstream of and adjacent to the static turbine vanes  72 ,  74  and can also extend radially outwardly relative to the centerline  12 , from a blade platform to a blade tip. It is noted that the number of blades, vanes, and turbine stages shown in  FIG. 1  were selected for illustrative purposes only, and that other numbers are possible. 
         [0021]    The blades  68 ,  70  for a stage of the turbine can be mounted to a disk  71 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 , with each stage having its own disk  71 ,  73 . The vanes  72 ,  74  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
         [0022]    The portions of the engine  10  mounted to and rotating with either or both of the spools  48 ,  50  are also referred to individually or collectively as a rotor  53 . The stationary portions of the engine  10  including portions mounted to the core casing  46  are also referred to individually or collectively as a stator  63 . 
         [0023]    In operation, the airflow exiting the fan section  18  is split such that a portion of the airflow is channeled into the LP compressor  24 , which then supplies pressurized ambient air  76  to the HP compressor  26 , which further pressurizes the ambient air. The pressurized air  76  from the HP compressor  26  is mixed with fuel in the combustor  30  and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine  34 , which drives the HP compressor  26 . The combustion gases are discharged into the LP turbine  36 , which extracts additional work to drive the LP compressor  24 , and the exhaust gas is ultimately discharged from the engine  10  via the exhaust section  38 . The driving of the LP turbine  36  drives the LP spool  50  to rotate the fan  20  and the LP compressor  24 . 
         [0024]    A remaining portion of the airflow  78  bypasses the LP compressor  24  travelling through the fan exhaust section  41  and exiting the engine assembly  10  through a stationary vane row, and more particularly an outlet guide vane assembly  80 , comprising a plurality of airfoil guide vanes  82 . More specifically, a circumferential row of radially extending airfoil guide vanes  82  are utilized adjacent the fan section  18  to exert some directional control of the airflow  78 . Upon exiting the fan exhaust section  41 , the airflow  78  can be redirected using a cooling air offtake  84  for additional cooling of the engine core  44  and turbine section  32 . 
         [0025]    Some of the ambient air supplied by the fan  20  can bypass the engine core  44  and be used for cooling of portions, especially hot portions, of the engine  10 , and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally the combustor  30  and components downstream of the combustor  30 , especially the turbine section  32 , with the HP turbine  34  being the hottest portion as it is directly downstream of the combustion section  28 . Other sources of cooling fluid can be, but is not limited to, fluid discharged from the LP compressor  24  or the HP compressor  26 . This fluid can be bleed air  77  which can include air drawn from the LP or HP compressors  24 ,  26  that bypasses the combustor  30  as cooling sources for the turbine section  32 . This is a common engine configuration, not meant to be limiting. 
         [0026]      FIG. 2  is an enlarged view of the area near the fan exhaust section  41 . The cooling air offtake  84  comprises a duct  86  having walls  88  that turn through nearly 90° from a primarily radial orientation to a primarily axial orientation. The cooling air offtake  84  includes an inlet  90  located downstream of the outlet guide vane assembly  80 . The inlet  90  includes a louver assembly  92  having a louver  93  comprising at least two different size airfoils  94 ,  96 . While illustrated at a location downstream of the fan exhaust section  41 , the offtake  84  can be located at any appropriate location throughout the engine. 
         [0027]    In an exemplary embodiment illustrated in  FIG. 3  the louver assembly  92  includes four airfoils  94 ,  96 ,  98 ,  100  spaced in an axial arrangement. The inlet  90  has a leading edge  85  with a rounded lip and a trailing edge  87  having a chamfer angle β of at least 20°, but not to exceed 30° measuring from the duct wall  88  towards the trailing edge  87  axially upstream. This feature will allow for higher pressure air bleed and moving the impingement point aft. An excessive angle will result in undesired pressure losses. 
         [0028]    The geometry of airfoil  96 , which will be referred to as the primary airfoil  96 , is outlined in  FIG. 4  described by a chord length C having a length defined as a line from a leading edge  108  to a trailing edge  110  and a height H having a length defined as a line from a radial maximum  112  to a radial minimum  114  relative to the engine centerline. Each airfoil is also described by an angle of attack α measured from a local relative wind direction  116  to a continuous line along the chord length C. For illustrative purposes the dimensions for an initial, third, and fourth airfoil  94 ,  98 ,  100  will be represented by subscripts  1 ,  2 , and  3  respectively. 
         [0029]    The primary airfoil  96  is geometrically larger, both with respect to the chord length C and the height H, than the other three airfoils  94 ,  98 ,  100 . The maximum height H of the primary airfoil  94  is at least 2 times larger than the maximum height H 1  of the initial airfoil  94 . The chord length C is at least 2.5 times larger than the chord length C 1 . The axial arrangement of the airfoils comprises a geometry partially defined by a chord length relationship as follows: 
         [0000]      C&gt;C 3 &gt;C 1 ≧C 4  
 
         [0030]    The spaced axial arrangement includes the initial airfoil  94  nearest the leading edge  85  of the inlet  90 , after which the primary airfoil  96  is located downstream of the initial airfoil  94 , followed in the downstream direction by the third and fourth airfoils  98 ,  100 . The third and fourth airfoils  98 ,  100  are spaced equivalently so that the distance between the duct wall  88  and the fourth airfoil  100  is nearly the same as the distance between the third and fourth airfoils  98 ,  100 . This spacing prevents flow separation between airfoils whilst keeping a Mach number high (See  FIGS. 4A and 4B ) 
         [0031]    The angle of attack a for the third and fourth airfoils  98 ,  100  is different than the angle of attack a for the first and second airfoil  94 ,  96 . In an exemplary embodiment the angle of attack a for the third and fourth airfoils  98 ,  100  is greater than that of the first and second airfoil  94 ,  96 . 
         [0032]    In an exemplary embodiment, the trailing edges  110  of the third and fourth airfoils  98 ,  100  terminate in a line L connecting the trailing edge  110  of the primary airfoil  96  to a point  118  downstream of a trailing edge  87  of the inlet  90 . This geometry causes corresponding chord lengths C 3 , C 4  for the third and fourth airfoil  96 ,  98  become consecutively shorter. This relationship manages to turn effectively the flow whilst reducing any friction losses due to flow contact with the airfoil surface. 
         [0033]    The overall benefit of the current embodiments is seen by the comparison of  FIG. 4A , showing a contemporary louver assembly with equal sized airfoils, with substantially the same angle of attack and equal spacing, as compared to the embodiment of  FIGS. 2-3 . For the conventional louver assembly the flow direction is changed by guiding the airflow using a louver  122  having similar shaped airfoils  124  as depicted in  FIG. 4A . This design can cause airflow separation  126  which is undesirable for effective airflow movement. Increasing the size of the primary airfoil  94  so that the louver assembly  92  comprises at least two different size airfoils  94 ,  96  where the second  96  is geometrically larger than the first  94 . This geometry differentiation causes an acceleration  128  of the flow depicted in  FIG. 4B  allowing for a total engine pressure P t  increase. 
         [0034]    Each of the airfoils  94 ,  96 ,  98 ,  100  in the louver assembly  92  is designed with a purpose, ensuring the effective use of the individual aerodynamic geometry. The initial airfoil  94  is configured to stabilize a boundary layer  130  and contain recirculation  132  in the duct  86 . With a conventional louver assembly  120  the boundary layer  131  is too thick and will induce separation, wherein as seen in  FIG. 4B , the boundary layer  130  by both the initial airfoil  94  and primary airfoil  96  is well defined. The primary airfoil  96  is configured to accelerate  128  the flow to maximum speed  134  without flow separation. The third and fourth airfoils  98 ,  100  are configured to guide the flow from downstream of the primary airfoil  96  in order to prevent separation. 
         [0035]    Thorough CFD (Computational Fluid Dynamics) analyses has been conducted and supports the benefit of the louver assembly  92  as compared to conventional louver assemblies  120 . 2D optimization backed up with a 3D analysis has been carried out with tabulated results following. The pressure recovery is maximized whether considering an area from the fan exhaust section  41  to the HP turbine  34  or from the fan exhaust section  41  to the LP turbine  36  both of which enable a reduction in bled flow. The following table compares a first engine recovery ratio to a second engine recovery ratio where the second engine recovery ration includes the louver assembly  92  in place and the pressure recovery is at least 0.30. The goal is to maintain the highest total pressure (P t ) as possible so as to best move air through the duct to the turbine sections. 
         [0000]    
       
         
               
               
               
               
               
             
               
               
               
               
               
             
           
               
                   
               
               
                   
                   
                 Total 
                 Static 
                   
               
               
                   
                   
                 Pressure 
                 Pressure 
                 Total 
               
               
                 Recovery = (Pt − P s13 )/ 
                 Recov- 
                 (Stage 13) 
                 (Stage 13) 
                 Pressure 
               
               
                 (P t13  − P s13 ) 
                 ery 
                 P t13   
                 P s13   
                 P t   
               
               
                   
               
             
             
               
                   
               
             
          
           
               
                 One Engine with conventional 
                 0.131 
                 8.07 
                 6.82 
                 6.99 
               
               
                 louver design 
               
               
                 Second Engine with proposed 
                 0.362 
                 8.369 
                 7.097 
                 7.557 
               
               
                 louver assembly through the 
               
               
                 HP turbine 
               
               
                 Second Engine with proposed 
                 0.349 
                 8.369 
                 7.097 
                 7.541 
               
               
                 louver assembly through the 
               
               
                 LP turbine 
               
               
                   
               
             
          
         
       
     
         [0036]    Benefits to increasing the pressure recovery and reducing the mass flow include allowing for the duct flow to be reduced while maintain power. As the room for designing pipes is typically constrained, the introduction of this approach enables that duct pipes to be designed with more flexibility. 
         [0037]    It should be appreciated that application of the disclosed design is not limited to turbine engines with fan and booster sections, but is applicable to turbojets and turbo engines as well. 
         [0038]    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 languages of the claims.

Technology Classification (CPC): 5