Abstract:
An aircraft engine according to an example of the present disclosure includes, among other things, a high pressure turbine having a blade, an engine casing disposed about the blade, a shield disposed around the casing adjacent to the blade and creating an area between the shield and the casing, and a gate disposed along the shield. The gate is rotatable about the engine casing between an opened position and a closed position for selectively controlling entry of cooling air into the area. A method of cooling an engine is also disclosed.

Description:
RELATED APPLICATION 
     This application is a continuation of U.S. patent application Ser. No. 13/635,421, filed on Dec. 19, 2012, which is a National Phase Application of International Application No. PCT/US2010/029341, filed on Mar. 31, 2010. 
    
    
     BACKGROUND 
     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. 
     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. 
     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 
     An aircraft engine for use in a fighter jet according to an example of the present disclosure includes a high pressure turbine having a blade, an engine casing disposed about the blade, a shield disposed around the casing adjacent to the blade and creating an area between the shield and the casing, and a gate disposed along the shield. The gate is rotatable about the engine casing between an opened position and a closed position for selectively controlling entry of cooling air into the area. 
     In a further embodiment of any of the foregoing embodiments, the gate is configured to be partially open between the opened and closed positions when the engine is being operated in a steady state. 
     In a further embodiment of any of the foregoing embodiments, the gate is built into a front of the shield. 
     In a further embodiment of any of the foregoing embodiments, the shield defines an opening. The gate comprises a strap having a slot. The strap is movable relative to the opening such that the slot and the opening may be in register with each other. 
     In a further embodiment of any of the foregoing embodiments, the opening is disposed in a face of the shield. The face extends in a radial direction relative to an axis of the high pressure turbine. 
     In a further embodiment of any of the foregoing embodiments, the face has a race therein for holding the strap. 
     In a further embodiment of any of the foregoing embodiments, the strap is moveable within the race for moving the slot of the strap into and out of register with the opening. 
     In a further embodiment of any of the foregoing embodiments, an outer wall of the shield slopes radially inward from the face relative to the axis. 
     In a further embodiment of any of the foregoing embodiments, the strap is moveable about the axis. 
     In a further embodiment of any of the foregoing embodiments, the opening is a plurality of openings circumferentially distributed about the face, and the slot is a plurality of slots circumferentially distributed about the strap, each of the plurality of slots corresponding to one of the plurality of openings. 
     In a further embodiment of any of the foregoing embodiments, the shield and the strap form an annulus. 
     In a further embodiment of any of the foregoing embodiments, the shield defines a duct opening configured to receive a boss. The boss defines a passage configured to communicate cooling airflow to the high pressure turbine. The boss fluidly separates the passage and the area. 
     A further embodiment of any of the foregoing embodiments includes a controller coupled to an actuator. The controller is operable to cause the actuator to selectively move the gate relative to the shield. 
     In a further embodiment of any of the foregoing embodiments, the gate is configured to be located in the closed position when the engine is maneuvering, and the gate is configured to be located in the opened position when the engine is cruising. 
     A cooling system for an aircraft engine for use in a fighter jet according to an example of the present disclosure includes the aircraft engine having a high pressure turbine having a blade and an engine casing disposed about the blade. The cooling system includes a shield disposed around the casing adjacent to the blade and for creating an area between the shield and the casing, and a gate disposed along the shield. The gate is rotatable about the engine casing between an opened position and a closed position for selectively controlling entry of cooling air into the area, the gate disposed about the casing. 
     In a further embodiment of any of the foregoing embodiments, the gate is adapted to be partially open between the opened and closed positions when the engine is being operated in a steady state. 
     In a further embodiment of any of the foregoing embodiments, the gate is built into a front of the shield. 
     In a further embodiment of any of the foregoing embodiments, the shield defines an opening. The gate comprises an opening and a strap having a slot. The strap is movable relative to the opening such that the slot and the opening are in register with each other. 
     In a further embodiment of any of the foregoing embodiments, the opening is disposed in a front of the shield. 
     In a further embodiment of any of the foregoing embodiments, the front has a race therein for holding the strap, and the strap is moveable within the race for moving the slot of the strap into and out of register with the opening. 
     In a further embodiment of any of the foregoing embodiments, the gate is configured to be located in the closed position when the engine is in a maneuvering mode. The gate is configured to be located in the opened position when the engine is in a cruising mode. 
     A method of cooling an engine used in a fighter jet according to an example of the present disclosure includes providing a shield around a casing adjacent to a high pressure turbine blade in the engine. The shield includes a radially extending face providing a gate adjacent to the face, the gate moveable between an opened position and a closed position, and moving the gate about the engine casing toward the opened position such that cooling air is delivered to an area between the shield and the casing to shrink the casing around the blade. 
     A further embodiment of any of the foregoing embodiments includes moving the gate from the opened position toward the closed position to partially block cooling air from entering the area when operation of the engine changes between a cruise mode and a steady state mode. 
     A further embodiment of any of the foregoing embodiments includes moving the gate to the closed position to fully block cooling air from entering the area when the engine is in a maneuvering mode. 
     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 
         FIG. 1  shows a schematic drawing in which a jet engine utilizes a clearance control system that is off. 
         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. 
         FIG. 3  shows a perspective view of the air cooling system of  FIG. 1 . 
         FIG. 3A  shows an expanded view taken along the lines  3 A in  FIG. 3 . 
         FIG. 3B  shows a back view of  FIG. 3A . 
         FIG. 4  shows a perspective view of the system disclosed herein in a closed condition. 
         FIG. 5  shows the system disclosed herein in an opened condition. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
     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. 
     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. 
     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 . 
     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. 
     Referring now also to  FIGS. 3, 3A, and 3B , 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 . 
     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 . 
     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 complement the shape of the openings  110  and 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 . 
     The heat shield  70  has a bottom flange  245  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  245  by conventional means and is disposed against the case  20  and against the flange  245  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). 
     Referring to  FIG. 3A , the face  105  of the heat shield may have an electro mechanical device  155  that engages a boss 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  176  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  into and 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. 
     The heat shield  70  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 . 
     Referring now to  FIGS. 1 and 3-4 , the operation of the heat shield is described. If the aircraft 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. 
     Referring now to  FIGS. 3 and 3A , 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. 
     Referring now to  FIGS. 2-3 and 5 , the operation of the heat shield is described. If the aircraft 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. 
     This simple, light-weight CCS may provide a fuel efficiency benefit, in the range of 0.5%-1.0% TSFC (thrust specific fuel consumption). 
     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. 
     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.