Abstract:
A method facilitates assembling a gas turbine engine including a compressor and a rotor assembly coupled in axial flow communication downstream from the compressor. The method comprises coupling a bypass system in flow communication with the compressor to channel a portion of flow discharged from the compressor towards the rotor assembly is channeled through the bypass system, and coupling a downstream end of the bypass system within the gas turbine engine such that the flow entering the bypass system flows past the rotor assembly and is discharged downstream from the rotor assembly.

Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to gas turbine engines, and more specifically to methods and apparatus for assembling gas turbine engines. 
   At least some known gas turbine engines used with aircraft include a core engine having, in serial flow arrangement, a compressor which compresses airflow entering the engine, a combustor which burns a mixture of fuel and air, and low and high pressure rotary assemblies which each include a plurality of rotor blades that extract rotational energy from airflow exiting the combustor to generate thrust from the engine. In addition, within at least some known gas turbines some of the work generated by the rotary assemblies is transmitted to an engine accessory gearbox by means of shaft wherein the available work can then be used to drive electrical equipment utilized on the aircraft. 
   As aircraft accessory power demands have increased, there also has been an increased need to run the gas turbine engines at idle speeds that may be higher than other engines not subjected to increased power demands. More specifically, increasing the idle speeds enables the increased power demands to be met without sacrificing compressor stall margins. However, the increased idle speeds may also generate thrust levels for the engine which are higher than desired for both flight idle decent operations and/or during ground idle operations. Over time, continued operation with increased thrust levels during such idle operations may increase maintenance costs and the increased fuel flows may also increase aircraft operating expenses. 
   BRIEF SUMMARY OF THE INVENTION 
   In one aspect, a method for assembling a gas turbine engine including a compressor and a rotor assembly coupled in axial flow communication downstream from the compressor is provided. The method comprises coupling a bypass system in flow communication with the compressor to channel a portion of flow discharged from the compressor towards the rotor assembly is channeled through the bypass system, and coupling a downstream end of the bypass system within the gas turbine engine such that the flow entering the bypass system flows past the rotor assembly and is discharged downstream from the rotor assembly. 
   In another aspect, a rotor assembly for a gas turbine engine including a compressor is provided. The rotor assembly includes a rotor coupled in axial flow communication downstream from the combustor, and a bypass system coupled in flow communication to the compressor for channeling a portion of flow discharged from said compressor around said rotor. 
   In a further aspect, a gas turbine engine is provided. The gas turbine engine includes a compressor, a rotor, and a bypass system. The rotor is coupled downstream from and in axial flow-communication with the compressor. The bypass system is coupled in flow communication to the compressor for channeling a portion of flow discharged from the compressor around the rotor during engine operation. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is schematic illustration of an exemplary gas turbine engine; 
       FIG. 2  is an enlarged cross-sectional schematic view of a portion of the gas turbine engine shown in  FIG. 1 ; and 
       FIG. 3  is a perspective view of a portion of the gas turbine engine shown in  FIG. 2 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
     FIG. 1  is a schematic illustration of a gas turbine engine  10  including a fan assembly  12 , a booster  22 , a high pressure compressor  14 , and a combustor  16 . Engine  10  also includes a high pressure turbine  18 , and a low pressure turbine  20 . Fan assembly  12  includes an array of fan blades  24  extending radially outward from a rotor disc  26 . Engine  10  has an intake side  28  and an exhaust side  30 . Fan assembly  12  and turbine  20  are coupled by a first rotor shaft  31 , and compressor  14  and turbine  18  are coupled by a second rotor shaft  32 . In the exemplary embodiment, high pressure turbine  18  is also coupled to a shaft (not shown) which provides a rotary motive force to drive a driven machine, such as, but, not limited to a gearbox, a transmission, a generator, a fan, or a pump. 
   In operation, air flows through fan assembly  12  and compressed air is supplied to high pressure compressor  14 . The highly compressed air is delivered to combustor  16 . Airflow (not shown in  FIG. 1 ) from combustor  16  drives turbines  18  and  20 , and turbine  20  drives fan assembly  12  by way of shaft  31 . Moreover, the accessory gearbox is also driven by turbine  18 . 
     FIG. 2  is an enlarged cross-sectional schematic view of a portion of gas turbine engine  10 .  FIG. 3  is a perspective view of the portion of gas turbine engine  10  shown in  FIG. 2 . High pressure turbine  18  includes a plurality of stages  50 , and each stage includes a row of rotor blades  52  and a row of stationary vanes (not shown). 
   A load-bearing annular turbine frame  58  extends downstream from turbine  18 . Frame  58  includes a radially outer structural member or casing  60  that extends circumferentially around turbine  18 , and a radially inner member or hub  62  that is coaxially aligned with respect to casing  60  about an axis of rotation of turbine engine  10 . Hub  62  is radially inward from casing  60  and a plurality of circumferentially spaced apart hollow struts  66  extend radially between casing  60  and hub  62 . 
   A transition duct  80  extends downstream from turbine frame  58 . Specifically, transition duct  80  includes a plurality of panels  82  coupled together circumferentially such that a flow passageway  84  is defined through transition duct  80  between high pressure turbine  18  and low pressure turbine  20 . Accordingly, panels  82  extend generally axially between an upstream end  86  of transition duct  80  and a downstream end  88  of duct  80 . 
   Low pressure turbine  18  includes a plurality of stages  90 , and each stage includes a row of circumferentially-spaced rotor blades  92  and a row of circumferentially-spaced stationary vanes  94 . In the exemplary embodiment, turbine  20  is coupled in axial flow communication to turbine  18  and is substantially concentrically aligned with respect to turbine  18 . A casing  96  extends circumferentially around turbine  20 . More specifically, casing  96  extends downstream from extension duct  80  to a turbine rear frame  100 . Turbine rear frame  100  is annular and extends between casing  96  and a primary exhaust nozzle  102 . 
   A bypass system  110  is coupled in flow communication to compressor  14 , and downstream from compressor  14 , such that a portion of flow discharged from compressor  14  is channeled through bypass system  110 , as is described in more detail below. In the exemplary embodiment, bypass system  110  is coupled between high and low pressure turbines  18  and  20 , respectively, and more specifically, between a pair of circumferentially adjacent struts  66  within transition duct  80 . In an alternative embodiment, bypass system  110  is coupled downstream from struts  66 . In another alternative embodiment, bypass system  110  is coupled upstream from struts  66 . Alternatively, bypass system  110  may be coupled at any location downstream from compressor  14  that enables bypass system  110  to function generally as described herein. 
   In the exemplary embodiment, bypass system  110  includes a plurality of circumferentially-spaced bypass ducts  120  that each extend from an inlet  122  to a discharge outlet  124 . In an alternative embodiment, bypass system  110  includes only one bypass duct  120 . In another alternative embodiment, bypass system  110  includes at least one arcuate plenum that extends between ducts  120  and transition duct  80  such that a portion of flow discharged from compressor  14  is channeled through the plenum prior to being routed through ducts  120 . Accordingly, in such an embodiment, the plenum couples at least two adjacent ducts  120  together in flow communication. 
   In the exemplary embodiment, each bypass duct inlet  122  extends through an opening  126  formed in transition duct  80  along a radially outer boundary of flow passageway  84 . Each discharge outlet  124  is coupled to engine  10  downstream from turbine  20 , and more specifically, to primary exhaust nozzle  102 , such that flow discharged from bypass ducts  120  is directed into the gas flowpath of turbine  20 . In an alternative embodiment, flow discharged from bypass ducts  120  is channeled through at least one cavity (not shown) defined within engine  10  aft of turbine rear frame  100 . In another alternative embodiment, flow discharged from bypass ducts  120  is channeled into a primary bypass stream duct (not shown) extending downstream from fan assembly  12  (shown in  FIG. 1 ). In a further alternative embodiment, flow is discharged to ambient from bypass ducts  120 . Alternatively, flow may be discharged from bypass system  110  at any location downstream from the specific turbine being bypassed, i.e., turbine  20 , that enables bypass system  110  to function generally as described herein. 
   In the exemplary embodiment, each bypass duct  120  includes a flow control device  140  housed therein. More specifically, in the exemplary embodiment, each flow control device  140  is a butterfly valve that is rotatably coupled within each duct  120 . In an alternative embodiment, each flow control device  140  is a flapper valve that is actuator-controlled. In a further alternative embodiment, each flow control device  140  is a poppet valve that is biased in a closed position. Alternatively, flow control device  140  is any type of flow control mechanism that enables flow control device  140  to function as described herein. 
   In each embodiment, flow control device  140  is resistant to high operating temperatures and is selectably positionable between an open position and a closed position to control an amount of flow entering bypass system  110 . Specifically, in the closed position, flow control device  140  substantially seals bypass duct inlet  122  such that bypass flow is prevented from entering system duct  120 . In contrast, when flow control device  140  is opened, a portion of flow discharged from compressor  14 , or in the exemplary embodiment, turbine  18 , is channeled into bypass system  110  and routed around turbine  20 . In one embodiment, when each control device  140  is opened, approximately 10% of flow discharged from turbine  18  is channeled around turbine  20  through bypass system  110  during pre-selected engine operational periods. 
   In the exemplary embodiment, flow control device  140  is electrically coupled to an engine control system which automatically controls the position of flow control device  140 . In one embodiment, the engine control system is a full authority digital electronic control system (FADEC) commercially available from Lockheed Martin Control Systems, Johnson City, N.Y. The engine control system alters the position of flow control devices  140  to control operation of bypass system  110 . 
   To facilitate cooling flow control device  140 , bypass system  110  is also coupled in flow communication to a cooling source. Specifically in the exemplary embodiment, each duct  120  includes a plurality of cooling openings  160  which enable cooling fluid to be channeled into each duct  120  to facilitate reducing an operating temperature of each flow control device  140 . More specifically, in the exemplary embodiment, openings  160  enable a continuous purge flow of cooling fluid to be channeled into ducts  120 . For example, in one embodiment, compressor discharge air is channeled through openings  160 . In another embodiment, interstage compressor air is channeled through openings  160 . 
   During idle engine operating speeds, and more specifically, during flight idle decent operating conditions and ground idle operating conditions, increased power demands may require engine  10  to operate at an idle speed that is higher than idle speeds of other known gas turbine engines. The increased idle speed enables engine  10  to satisfy the increased power demands while maintaining compressor stall margins. During such engine operating conditions, flow control devices  140  are opened such that a portion of flow discharged from compressor  14  is channeled through bypass system  110 . In the exemplary embodiment, because the flow enters bypass system  110  aft of turbine  18 , flow through high pressure turbine  18  is not disrupted. 
   Accordingly, during operation of bypass system  110 , turbine  18  can continue to operate at an increased operational speed necessary to meet the power demands, without an increased amount of thrust being generated. More specifically, because less flow is channeled through turbine  20  during operation of bypass system  110 , bypass system  110  facilitates reducing an amount of thrust generated from engine  10  in comparison to operating periods when turbine  18  is operated at the same operational speed while bypass system  110  is non-operational. Accordingly, during ground idle operations, because less thrust is generated from engine  10  during operation of bypass system  110 , maintenance of aircraft braking systems, for example, is facilitated to be reduced, as less braking is necessary during such engine operational periods. Moreover, during aircraft flight operations, operation of bypass system  110  facilitates reduced flight idle thrusts during decent operations. As such, bypass system  110  facilitates improving short range fuel burn while maintaining adequate compressor stall margin during high power extraction operating conditions. 
   The above-described frame is cost-effective and highly reliable. The frame includes a bypass system coupled to a transition duct extending between the high and low pressure turbines. The bypass system enables a portion of flow discharged from the high pressure turbine to be channeled around the low pressure turbine during pre-selected engine operational periods. Because a portion of the flow is bypassed around the low pressure turbine, less thrust is generated from the engine during the pre-selected engine operational periods. As a result, the bypass system overcomes known manufacturing gas turbine operating limitations during high power extraction operations in a cost-effective and reliable manner, while maintaining compressor stall margin. 
   Exemplary embodiments of turbine frames are described above in detail. The frames are not limited to the specific embodiments described herein, but rather, components of each assembly may be utilized independently and separately from other components described herein. For example, each bypass system component can also be used in combination with other turbine frame components. Furthermore, each bypass system component may also be used with other gas turbine engine configurations. 
   While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.