Patent Abstract:
An aircraft engine for use in a low-bypass turbofan application has a high pressure turbine having a blade and an engine casing disposed about the blade. A shield is disposed around the casing adjacent to the blade to create an area between the shield and the casing. A gate selectively controls entry of cooling air into the area and may be closed if the engine is maneuvering and may be open if cruising.

Full Description:
RELATED APPLICATION 
       [0001]    This application claims priority to PCT/US2010/029341, filed on Mar. 31, 2010. 
     
    
     BACKGROUND 
       [0002]    Aircraft gas turbine case cooling systems help the efficiency of gas turbine engines by lowering fuel consumption thereof. The systems distribute relatively cool air from an engine compressor to the casing surface of turbine cases causing the casing surface to shrink. Clearance between the case inner diameter and turbine blade tips shrinks to minimize the amount of air that escapes around the blade tip thereby increasing fuel savings to optimize the system. 
         [0003]    Generally, during a cruise condition, compressor air is ducted to manifolds that surround the turbine cases. The manifolds direct the cooler air on a case surface causing case diameter to shrink, closing blade tip-to-case clearances. 
         [0004]    However, at take off or during climbing, the cooling air is shut off causing the cases to grow in diameter. Clearances between the blade tips and the casing are increased and the system is not optimized but blade-to-case interactions are minimized 
       SUMMARY 
       [0005]    According to an exemplary embodiment, a low bypass turbofan gas turbine engine, such as in a fighter jet application, has a high pressure turbine having a blade and an engine casing disposed about the blade. A shield is disposed around the casing adjacent to the blade to create an area between the shield and the casing. A gate selectively controls entry of cooling air into the area and may be closed if the engine is maneuvering and may be open if cruising. 
         [0006]    According to a further exemplary embodiment, a cooling system is disposed in a low-bypass turbofan gas turbine engine, the engine having a high pressure turbine having a blade and an engine casing disposed about the blade. The cooling system has a shield disposed around the casing adjacent to the blade to create an area between the shield and the casing. A gate may selectively controls entry of fan air into the area if disposed about the casing such that the gate is adapted to be closed if the engine is maneuvering and may be open if the engine is cruising. 
         [0007]    According to a further exemplary embodiment, a method of cooling a low-bypass turbofan engine includes the steps of providing a shield around a casing adjacent a high pressure turbine blade in the engine, gating fan air to an area between the shield and the casing to shrink the casing around the blades if the engine is in a cruise mode. 
         [0008]    These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]      FIG. 1  shows a schematic drawing in which a jet engine utilizes a clearance control system that is off. 
           [0010]      FIG. 2  is an embodiment of the schematic embodiment of the jet system of  FIG. 1  in which the air flow is vented through the duct. 
           [0011]      FIG. 3  shows an exploded view of the air cooling system of  FIG. 1 . 
           [0012]      FIG. 3A  shows an expanded view taken along the lines  3 A in  FIG. 3 . 
           [0013]      FIG. 3B  shows a back view of  FIG. 3A . 
           [0014]      FIG. 4  shows a perspective view of the system disclosed herein in cruise condition. 
           [0015]      FIG. 5  shows the system disclosed herein in take-off or maneuver condition. 
           [0016]      FIG. 6  shows the system disclosed herein in steady-state condition. 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       [0017]    Referring to  FIG. 1 , a jet engine  15  used with aircraft that have performance as a priority, e.g., a military fighter aircraft  10  that is used for quick acceleration and deceleration, is schematically shown. Such engines  15  frequently employ high speed maneuvers, in which the engine may be throttled upwardly and downwardly quickly and often. 
         [0018]    Historical active clearance control systems (“ACS” and not shown) do not work with these engines and aircraft  10 . The cooling provided by an ACS cannot keep up with the rapid heat changes in the engine caused by maneuvering. For instance, a pilot (not shown) may need rapid acceleration in one instance that causes the case  20 , and clearance, to expand rapidly. Air directed to the case by an ACS to minimize that clearance may not be delivered in time to cool the case during that maneuver. But cooling caused by the ACS may occur too rapidly as the throttle is pulled back to decelerate the aircraft (and the temperature of the engine) so that blade tip-to-case interference may occur. Such situations are clearly undesirable. Moreover, ACS may be heavy and may limit the aircraft&#39;s ability to maneuver. As a result, engines in this type of aircraft  10  do not have ACS and particularly in the high pressure turbine section  25  of the engine  15  where such tip-to-case in clearance is critical and in which tip-to-case interference is undesirable. 
         [0019]    Referring to  FIGS. 1 and 2 , a portion  17  of an engine  15  is shown. The engine casing  20  encloses high pressure turbine blades  30 , low pressure turbine blades  35  and a plurality of stationary struts  40 . A ducting system  45  directs cooling air (indicated by arrows  50 ) on a continual basis to the case  20  outside the low pressure turbine blades  35  via boss  55 . This cooling air is typically directed from a compressor (not shown) through the ducting system  45  in an area between the case  20  and a nacelle  60 . 
         [0020]    Referring now to  FIGS. 1 and 2 , exemplary clearance control system  65  (“CCS”) for the high pressure turbine blades  30 , or other areas of the engine  15 , is shown. The CCS  65  includes a heat shield  70 , an actuation valve  75 , and a finger seal  80 , or other means of conventionally constraining the heat shield to a cylindrical case, such as a band clamp (not shown).  FIG. 1  shows the actuation valve  75  closed thereby causing a flow of cooling air  85  not to pass between the heat shield  70  and the case  20  thereby allowing the case to expand and minimize a probability of tip-to-case interference. Such a condition is used if said aircraft  10  is maneuvering.  FIG. 2  shows the actuation valve  75  open thereby causing a flow of cooling air  85  from an engine fan (not shown) to pass between the heat shield  70  and the case  20  thereby causing the case  20  to shrink and improve fuel consumption. Such a condition is used if said aircraft  10  is cruising or in steady state as will be discussed herein. 
         [0021]    Referring now also to  FIGS. 3 ,  3 A, and  3 B, the heat shield  70  is a piece of annular sheet metal that is contoured radially from its inlet end  90  to its outlet end  95  a distance from the casing to allow a proper amount of air  85  into a space  100  between the heat shield  70  about the case  20  adjacent to the high pressure turbine blades  30 . 
         [0022]    The inlet end  90  has a vertically-oriented face  105  (though other orientations are contemplated herein) that has a plurality of openings  110  that are roughly rectangular having curved sides  115  as the heat shield  70  is designed to enclose the case  20 . On that face  105 , the heat shield  70  has one or more slots  120  for cooperating with an annular strap  125  as will be discussed herein. The strap  125  and the face  105  and its openings  110  form the valve (or gate)  75 . 
         [0023]    The face  105  on its back portion  130  (see  FIG. 3B ) thereof has annular L-shaped flanges  135  that form races  140  for holding the flat annular strap  125  against the back portion  130 . The strap  125  has a plurality of spaced slots  145  that ape the shape of the openings  110  are designed to be in register, partially in register and out of register with the openings  110  in the face  105  to meter air  85  in the space  100 . 
         [0024]    The heat shield  70  has a bottom flange  145  which is designed to be in register with the casing  20 . A finger seal  150  (see  FIGS. 1 and 2 ) is attached to the bottom flange  145  by conventional means and is disposed against the case  20  and against the flange  145  to prevent the air  85  from entering the area  100  closed by the heat shield if not desired. The finger seal  150 , is one embodiment and it should be apparent to those skilled in the art that the forward heat shield can be attached by other means, including a band clamp (not shown). 
         [0025]    Referring to  FIG. 3A , the face  105  of the heat shield may have an electro mechanical device  155  that engages a boss  160  in the slot  120  to move the strap radially or about an axis  165  of the engine  15 . This electromechanical device  155 , such as a solenoid or the like) is attached to a controller  170 , as will be discussed herein, via a rod  175  attaching to the tab  175  attached to the strap  125 . The strap is placed within the races  140  within the back  130  of face  105  and is controlled by the electromechanical device  155  to move the strap  125  slots  145  into an out of registry with the openings  110  in the face  105  of the heat shield  70 . One may also recognize that the strap may be rotated by a remote linkage (not shown) or the like. 
         [0026]    The heat shield has several openings  180  therein to allow the boss  55  that extends from the duct system  50  to pass therethrough to provide a cooling air to the low pressure turbine blades  35  of the engine  15 . 
         [0027]    Referring now to  FIG. 4  and  FIGS. 1-3 , the operation of the heat shield is described. If the air craft is maneuvering, the strap  125  is rotated in its races  140  so that the slots  145  in the strap  125  do not align with the openings  110  in the face  105 . Air  85  cannot enter the space  100  and the case  20  is not cooled. Clearance between the blade  30  and the case  20  is allowed to grow thereby minimizing a possibility of tip-to-case interference. 
         [0028]    Referring now to  FIG. 5  and  FIGS. 1-3 , the operation of the heat shield  70  is described. If the aircraft  10  is in a steady state, e.g., where it is neither cruising nor maneuvering but cooling is somewhat effecting and maneuvering is possible, the strap  125  is rotated in its races  140  so that the slots  145  in the strap  125  align partially with the openings  110  in the face  105 . Some air  85  enters the space  100  and the case  20  is cooled a degree. Clearance between the blade  30  and the case  20  is being controlled to a degree thereby starting to minimize fuel consumption. 
         [0029]    Referring now to  FIG. 6  and  FIGS. 1-3 , the operation of the heat shield is described. If the air craft is cruising, e.g., where maneuvering is not anticipated, the strap  125  is rotated in its races  140  so that the slots  145  in the strap  125  align with the openings  110  in the face  105 . Air  85  enters the space  100  and the case  20  is cooled to minimize tip clearance and to minimize fuel consumption. 
         [0030]    This simple, light-weight CCS may provide a fuel efficiency benefit, in the range of 0.5%-1.0% TSFC (thrust specific fuel consumption). 
         [0031]    Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments. 
         [0032]    The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.

Technology Classification (CPC): 5