Patent Publication Number: US-2016230675-A1

Title: Fan flow control valve

Description:
BACKGROUND 
     In modern aircraft environments, turbine exhaust case (TEC) modules are typically air cooled using fan bypass air in a gas turbine engine. Future engines may include, or be associated with, variable and adaptive cycles where the major air streams are varied as a function of operating conditions to maximize/optimize performance and operability. There is a need to vary the flow into the TEC in these engines to provide high levels of air flow for conditions where high cooling is needed, but also provide reduced flow for other conditions since the air supplied to the TEC takes the form of a loss with respect to the engine cycle. 
     A valve may be used to support varying the air streams. Most valves rely on an introduction of a small control area with an abrupt area change which creates a parasitic pressure drop to reduce flow, which means losses are high in this mode. Valve methods which provide a smooth aerodynamic flow path to provide a very low pressure drop, the subject here, would benefit the engine cycle. 
     BRIEF SUMMARY 
     The following presents a simplified summary in order to provide a basic understanding of some aspects of the disclosure. The summary is not an extensive overview of the disclosure. It is neither intended to identify key or critical elements of the disclosure nor to delineate the scope of the disclosure. The following summary merely presents some concepts of the disclosure in a simplified form as a prelude to the description below. 
     Aspects of the disclosure are directed to a system associated with an engine of an aircraft, comprising: a first duct configured to convey a first flow, a second duct configured to convey a second flow that corresponds to a first portion of the first flow, a third duct configured to convey a third flow that corresponds to a second portion of the first flow, and at least one valve configured to control a cross-sectional area associated with at least one of the second duct or the third duct in order to control the ratio of the first portion to the second portion. In some embodiments, the first flow is a fan flow, the second flow is a nozzle flow, and the third flow is a turbine exhaust case flow. In some embodiments, at least a portion of the second flow is used to cool a component associated with a nozzle of the aircraft. In some embodiments, at least a portion of the second flow is exhausted in order to provide forward thrust for the aircraft. In some embodiments, the at least one valve comprises an actuator and a piston. In some embodiments, the valve is coupled to an engine control system, and a state of the valve is controlled based on a command received by the valve from the engine control system. In some embodiments, the at least one valve is located within a fan duct. In some embodiments, the at least one valve is located outside of a fan duct. In some embodiments, the system is configured to provide the third flow through a plurality of turbine exhaust case vanes. In some embodiments, the system is configured to divert the third flow downstream of turbine exhaust case vanes to provide a flow going into a core. 
     Aspects of the disclosure are directed to a system associated with an engine of an aircraft, comprising: a first duct configured to convey a fan bypass flow, a second duct configured to convey a first portion of the fan bypass flow as at least one of a cooling nozzle flow or a thrust flow, a third duct configured to convey a second portion of the fan bypass flow as a turbine exhaust case flow, and at least one valve configured to adaptively control a cross-sectional area associated with at least one of the second duct or the third duct based on a command received by the at least one valve. In some embodiments, the third duct is configured to convey the turbine exhaust case flow to at least one strut associated with a bearing at an output of a turbine. In some embodiments, the at least one valve is located within a fan duct. In some embodiments, the at least one valve is located outside of a fan duct. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The present disclosure is illustrated by way of example and not limited in the accompanying figures in which like reference numerals indicate similar elements. 
         FIG. 1  illustrates a gas turbine engine. 
         FIG. 2A  illustrates a system for modulating a fan flow in terms of a turbine exhaust case (TEC) flow and a nozzle flow. 
         FIG. 2B  illustrates a cross-section of a channel associated with the TEC flow of  FIG. 2A . 
         FIG. 3A  illustrates a system for modulating a fan flow in terms of a TEC flow and a nozzle flow. 
         FIG. 3B  illustrates a cross-section of a channel associated with the TEC flow of  FIG. 3A . 
         FIGS. 4A-4C  illustrate a system for diverting a fan flow downstream of TEC vanes to control a flow going into a core. 
     
    
    
     DETAILED DESCRIPTION 
     It is noted that various connections are set forth between elements in the following description and in the drawings (the contents of which are included in this disclosure by way of reference). It is noted that these connections are general and, unless specified otherwise, may be direct or indirect and that this specification is not intended to be limiting in this respect. A coupling between two or more entities may refer to a direct connection or an indirect connection. An indirect connection may incorporate one or more intervening entities. 
     In accordance with various aspects of the disclosure, apparatuses, systems and methods are described for varying a flow to a turbine exhaust case (TEC)  29 . The variation may be obtained without causing a large pressure loss (e.g., a loss in an amount greater than a threshold). 
     Aspects of the disclosure may be applied in connection with a gas turbine engine. For example,  FIG. 1  is a side-sectional illustration of a gas turbine engine  10 . The engine  10  includes a compressor section  12 , a turbine section  14  and one or more engine hot sections. The engine hot sections may include, for example, a first engine hot section  16  configured as a combustor section and a second engine hot section  18  configured as an augmentor section. The compressor section  12 , the first engine hot section  16 , the turbine section  14  and the second engine hot section  18  may be sequentially aligned along an axial centerline  20  between a forward engine airflow inlet  22  and an aft engine airflow exhaust  24 . The second engine hot section  18  involving secondary combustion to augment the engine thrust may include a first (e.g., annular, radial inner) duct case  26 , a second (e.g., annular, radial outer) duct case  28 , and one or more hot section vanes  30  to connect and support the ducts  26  and  28  as well as support the rotor shaft of the engine using a bearing. 
     One skilled in the art would appreciate that, in connection with the design and operation of an engine (e.g., engine  10 ), there may exist at least two flows. A first such flow, which may be referred to as a core flow  40 , may pass through the engine hardware and be subjected to combustion in, e.g., the first engine hot section  16 . A secondary flow, which may be referred to as a bypass flow  50 , bypasses the engine core. A bypass ratio may be established for denoting the ratio between the bypass flow  50  and the core flow  40 . The bypass ratio may be one measure of the efficiency (e.g., the fuel efficiency) of the engine  10 . The bypass flow  50  typically provides cooling air  51  passing through the hot surfaces  32  of the exhaust nozzle and/or flows out of the exhaust  52  to add thrust. 
       FIG. 1  represents one possible configuration for an engine  10 . Aspects of the disclosure may be applied in connection with other engine configurations. 
     TEC modules  29  associated with, e.g., the turbine  14 , such as for example struts associated with bearings at the output of the turbine  14 , may be cooled using air from a given flow (e.g., the bypass flow  50 ) passing through openings  35  in the TEC vanes  30 . In some embodiments, a moving flowpath boundary may be provided to streamline a capture of a feed of air flow to the TEC, thereby changing the flow into the TEC without creating large parasitic pressure losses. 
     Referring to  FIG. 2A , an exemplary system  200  for obtaining such a moving flowpath boundary is shown. The system  200  includes a flow  250 , which may correspond to at least a portion or the entirety of the bypass flow  50  described above in connection with  FIG. 1 . The flow  250  may be referred to as a fan flow. 
     The flow  250  may effectively be split into two flows. A first of these two flows is denoted in  FIG. 2A  via reference character  260 . The flow  260  may be referred to as a nozzle flow. At least a portion of the flow  260  may be used to cool components associated with a nozzle  51  of an aircraft. Alternatively, or additionally, at least a portion of the flow  260  may be dumped/exhausted  52  in order to provide (forward) thrust for the aircraft. In this respect, the flow  260  may be referred to as a thrust flow. 
     A second of these two flows is denoted in  FIG. 2A  via reference character  270 . The flow  270  may be associated with the TEC modules described above, and as such, may be referred to as a TEC flow passing though openings  35  into the TEC vanes  30 . 
     The ratio of (the splitting of the flow  250  into) the flow  260  to the flow  270  may be controlled based on a valve  280 . The valve  280  is shown in  FIG. 2A  as being located within a duct associated with the fan. The valve  280  may include a piston/rod  284  that is driven by an actuator  282 . The valve  280 /rod  284  is shown as being in a first, open position in  FIG. 2A , thereby providing for a large area, and hence, flow  270 . 
     Also superimposed in  FIG. 2A  for reference purposes is the rod  284  in a second position, as indicated by the reference character  284 ′. When the valve  280 /rod  284  is in this second position  284 ′, the valve  280  is closed and a larger portion of the flow  250  is directed/provided to the flow  260  (relative to the flow  270 ). 
     The actuator  282  may be driven by, or respond to commands from, an engine control system (not shown). The engine control system may include logic to determine a state/position for the valve  280 , and hence, the actuator  282 . 
     Regardless of the state/position of the valve  280 , the channels/ducts associated with the flows  260  and  270  may include relatively smooth, aerodynamic surfaces. Accordingly, loss (e.g., a pressure drop) that is associated with the flow  260  or the flow  270  may be small/minimal. 
     Referring to  FIG. 2B , a cross-section associated with the flow  270  of the system  200  is shown. In particular, a TEC module  290  (e.g., a strut) is shown as proximate to a valve moving body  292 , wherein the body  292  may correspond to the valve  280  of the system  200  opening of closing the opening  35  into the TEC vane  30 . 
     Referring to  FIG. 3A , an exemplary system  300  for obtaining a moving flowpath boundary is shown. The system  300  includes many of the flows and components/devices, and many of the same characteristics, as described above in connection with the system  200  of  FIG. 2A , and so, a complete re-description is omitted for the sake of brevity. In contrast to the system  200 , in the system  300  the valve  280  is shown as being located outside of the fan duct. All other things being equal, the valve  280  may operate at cooler temperatures in the system  300  relative to the system  200 . Easier access to the valve  280  for, e.g., maintenance or inspection activities may be obtained using the system  300  relative to the system  200 . In this case the opening  35  to the vane  30  is raised above the fan duct flow path  250  to enable closing of the valve with a horizontal translating valve  200 . 
     Referring to  FIG. 3B , a cross-section associated with the flow  270  of the system  300  is shown. A translating ring  390  is included in proximity to the vane opening  35  wherein flow  270  into the vane  30  is controlled. Flow to the nozzle  260  is allowed to pass around the valve through passage  395 . 
     In conjunction with the systems and flows described above in connection with  FIGS. 2A-2B and 3A-3B , the modulation/metering of the flows  250 - 270  may be performed to control flow through one or more TEC vanes  30 . In contrast thereto, the embodiment of  FIGS. 4A-4C  may be used to divert flow downstream of the TEC vanes, to control/provide a flow combining with the core  40 . The flow can combine with the core  40  in close proximity to the TEC to participate in combustion of the augmentor  18 , or combine farther downstream as with the cooling flow  51  thereby changing the engine cycle bypass ratio. 
     As shown in  FIG. 4A , one or more passages  402  may be selectively provided in a translating ring  400  to accommodate the flow  250 . In  FIG. 4B , one of the passages  402  is shown in an open state/position, allowing for traversal of the flow  250  therethrough. In contrast to  FIG. 4B , in  FIG. 4C  the passage  402  is substantially closed, such that the flow  250  is generally precluded from flowing through the passage  402 . 
     In view of the foregoing, aspects of the disclosure may be used to modulate one or more flows by controlling (e.g., modifying) a cross-section/area of one or more channels/ducts associated with the flows that are always aerodynamically smooth to reduce parasitic pressure losses. A valve may be used to provide such control. By utilizing arrangements such as those described herein, aerodynamic efficiency may be enhanced/increased. 
     Technical effects and benefits of this disclosure include providing a variable flow into a TEC  29  or the core  40 . Such a flow may be provided with minimal, parasitic pressure losses, thereby maintaining engine performance/efficiency. Aspects of the disclosure may be applied in connection with so-called adaptive engines to facilitate a dynamic alteration of one or more engine parameters. For example, if maximum thrust is desirable then a TEC flow may be reduced, whereas if it desirable to increase cooling to the TEC then the TEC flow may be increased. 
     Aspects of the disclosure have been described in terms of illustrative embodiments thereof. Numerous other embodiments, modifications, and variations within the scope and spirit of the appended claims will occur to persons of ordinary skill in the art from a review of this disclosure. For example, one of ordinary skill in the art will appreciate that the steps described in conjunction with the illustrative figures may be performed in other than the recited order, and that one or more steps illustrated may be optional in accordance with aspects of the disclosure.