Abstract:
A system for cooling a gas turbine engine control component comprises a control component, at least one duct in fluid communication with the component and the atmosphere, and a valve assembly located with the duct for modulating airflow to the component. A method for cooling a gas turbine engine control component comprises the steps of: a) providing a system for cooling the component, the system comprising at least one duct in fluid communication with the component and the atmosphere and a valve assembly located with the duct for modulating airflow to the component; b) exposing the system to a first condition where cooling of the component is required; c) opening the valve assembly in response to the first condition; d) exposing the system to a second condition where cooling of the component is not required; and e) closing the valve assembly in response to the second condition.

Description:
CROSS REFERENCE TO RELATED APPLICATION 
       [0001]    This application is a continuation-in-part of U.S. patent application Ser. No. 11/967,904 filed Dec. 31, 2007. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    The technology described herein relates generally to gas turbine engines, and more particularly, to a system and method for cooling engine control components for such engines. 
         [0003]    At least one known gas turbine engine assembly includes a fan assembly that is mounted upstream from a core gas turbine engine. During operation, a portion of the airflow discharged from the fan assembly is channeled downstream to the core gas turbine engine wherein the airflow is further compressed. The compressed airflow is then channeled into a combustor, mixed with fuel, and ignited to generate hot combustion gases. The combustion gases are then channeled to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight. The other portion of the airflow discharged from the fan assembly exits the engine through a fan stream nozzle. 
         [0004]    A variety of different control systems and methods are available for controlling the operation and performance of such aircraft gas turbine engines. One type of control system is commonly referred to as a FADEC (Full Authority Digital Engine Control), which is a generic name for a system for electronically controlling an aviation engine. A FADEC system includes a central computer system (often referred to as the FADEC) which receives information regarding the position of a throttle lever and other engine control devices operated by a pilot, as well as various information from various sensors fitted to the engine, and controls the engine so as to provide the required levels of performance such as engine thrust, fuel economy, etc. Engine control components such as FADEC units often include a number of small wires and electronic components such as computer chips and printed circuit boards. 
         [0005]    In service, gas turbine aircraft engines are subject to a wide range of operating conditions such as high and low altitudes, high and low temperatures, and high and low speed airflows over, around, and through the engine. Even during a single flight, the aircraft, its engine(s), and engine control components may experience low speed, low altitude, and high temperature conditions during taxi, takeoff, and landing operations, as well as high speed, high altitude, and low temperature conditions during the cruise portion of the flight. 
         [0006]    Engine control components such as FADEC units are normally housed within the engine cowling or nacelle, where it is protected from precipitation, dirt, debris, and other contaminants which may pass over, around, or through the engine. However, housing the engine control components in close proximity to the engine itself exposes them to potentially high temperatures generated during normal engine operation. Additionally, because of the electrical connections and components, as well as operating computer components which may be inside the engine control component, the engine control components may themselves generate heat which in a confined space tends to expose the unit to higher than desired operating temperatures. 
         [0007]    To manage the operating temperatures of the engine control components, ventilation is often provided to direct air which is cooler than the components onto the components to carry heat away and maintain the temperature of the component at a satisfactory operating level. However, cooling needs often vary greatly during the course of a flight or operating session. For example, a much greater degree of cooling may be needed on a hot day during ground operations at engine idle power settings than at high altitude during cruise conditions and high power settings. 
         [0008]    Due to the increased weight and complexity which would be required to modulate cooling air flow with conventional systems, such systems normally are configured to operate in a fixed operating condition to provide the maximum level of cooling air volume required during the foreseeable operating environment. However, such a configuration may result in over-cooling or chilling of the engine control components due to greater-than-necessary airflow of low temperature air, thereby inducing thermal fatigue stress and cyclic strain and potentially leading to cracks in sensitive internal elements such as printed circuit boards. 
         [0009]    Accordingly, there remains a need for a system and method for cooling gas turbine engine control components as needed while maintaining a satisfactory temperature range for operation of the components. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0010]    In one aspect, a system for cooling a gas turbine engine control component is described. The system comprises a control component, at least one duct in fluid communication with the component and the atmosphere, and a valve assembly located with the duct for modulating airflow to the component. 
         [0011]    In another aspect, a method for cooling a gas turbine engine control component is described. The method comprises the steps of: a) providing a system for cooling the component, the system comprising at least one duct in fluid communication with the component and the atmosphere and a valve assembly located with the duct for modulating airflow to the component; b) exposing the system to a first condition where cooling of the component is required; c) opening the valve assembly in response to the first condition; d) exposing the system to a second condition where cooling of the component is not required; and e) closing the valve assembly in response to the second condition. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0012]      FIG. 1  is a cross-sectional illustration of an exemplary gas turbine engine assembly; 
           [0013]      FIG. 2  is a partial cross-sectional elevational view of an exemplary gas turbine engine, illustrating an exemplary embodiment of an engine control component cooling system installed on the engine with a valve shown in an open position; 
           [0014]      FIG. 3  is an enlarged cross-sectional view of the valve assembly shown in  FIG. 2 ; 
           [0015]      FIG. 4  is an enlarged cross-sectional view taken along line  4 - 4  of  FIG. 2 ; 
           [0016]      FIG. 5  is a partial cross-sectional elevational view similar to  FIG. 2 , but illustrating the cooling system of  FIG. 2  with the valve in a closed position; 
           [0017]      FIG. 6  is an enlarged cross-sectional view of the valve assembly shown in  FIG. 5 ; 
           [0018]      FIG. 7  is an enlarged cross-sectional view taken along line  7 - 7  of  FIG. 7 ; 
           [0019]      FIG. 8  is a view similar to  FIG. 7  of an alternative embodiment of a valve assembly; 
           [0020]      FIG. 9  is a temperature vs. time plot of the temperatures experienced by an engine control component with a conventional cooling system; and 
           [0021]      FIG. 10  is a temperature vs. time plot of the temperatures experienced by an engine control component utilizing an exemplary embodiment of the cooling system described herein. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0022]    Exemplary embodiments of the present invention will now be described with regard to the accompanying drawing figures, in which like numerals refer to like elements throughout the drawing figures. 
         [0023]      FIG. 1  is a cross-sectional schematic illustration of an exemplary gas turbine engine assembly  10  having a longitudinal axis  30 . Gas turbine engine assembly  10  includes a fan assembly  14  and a core gas turbine engine  15 . Core gas turbine engine  15  includes a high pressure compressor  18 , a combustor  20 , and a high pressure turbine  22 . In the exemplary embodiment, gas turbine engine assembly  10  also includes a low pressure turbine  24 , and a multi-stage booster compressor  16 , and a splitter  17  that substantially circumscribes booster  16 . 
         [0024]    Fan assembly  14  includes an array of fan blades extending radially outward from a rotor disk, the forward portion of which is enclosed by a streamlined spinner. Gas turbine engine assembly  10  has an intake side  11  and an exhaust side  13 . Fan assembly  14 , booster  16 , and turbine  24  are coupled together by a first rotor shaft  28 , and compressor  18  and turbine  22  are coupled together by a second rotor shaft  26 . 
         [0025]    In operation, incoming air  42  flows through fan assembly  14  and a first portion of the airflow is channeled through booster  16  and onward through internal flowpath  50  of core gas turbine engine  15 . The compressed air that is discharged from booster  16  is channeled through compressor  18  wherein the airflow is further compressed and delivered to combustor  20 . Hot products of combustion (not shown in  FIG. 1 ) from combustor  20  are utilized to drive turbines  22  and  24 , and turbine  24  is utilized to drive fan assembly  14  and booster  16  by way of shaft  28 . Air and combustion products flowing through internal flowpath  50  exit the core gas turbine engine  15  at the trailing edge  38  of the core cowl  36  as shown at  44 . Gas turbine engine assembly  10  is operable at a range of operating conditions between design operating conditions and off-design operating conditions. 
         [0026]    A second portion of the incoming airflow  42  discharged from fan assembly  14  is channeled through a bypass duct  40  to bypass a portion of the airflow from fan assembly  14  around the core cowl  36  which encloses the core gas turbine engine  15 . More specifically, bypass duct  40  extends between a fan casing or shroud, which forms a first or inner surface  31  of the engine nacelle  32  and the core cowl  36  which has a leading edge formed by splitter  17 . Air flowing through bypass duct  40  exits the trailing edge  34  of the nacelle  32 . Nacelle  32  encloses the major portion of the aircraft engine  10  and is secured to the aircraft by appropriate mounting apparatus. In the embodiment shown in  FIG. 1 , the nacelle  32  is secured to the aircraft wing  12  via mounting pylon  19 . 
         [0027]    Accordingly, a first portion of the airflow from fan assembly  14  is channeled through booster  16  and then into compressor  18  as described above, and a second portion of the airflow from fan assembly  14  is channeled through bypass duct  40  to provide thrust for an aircraft, for example. Splitter  17  divides the incoming airflow into first and second portions. Gas turbine engine assembly  10  also includes a fan frame assembly  46  to provide structural support for fan assembly  14  and is also utilized to couple fan assembly  14  to core gas turbine engine  15 . 
         [0028]    Fan frame assembly  46  includes a plurality of outlet guide vanes that extend substantially radially between a radially outer mounting flange and a radially inner mounting flange and are circumferentially-spaced within bypass duct  40 . Fan frame assembly  46  may also include a plurality of struts that are coupled between a radially outer mounting flange and a radially inner mounting flange. In one embodiment, fan frame assembly  46  is fabricated in arcuate segments in which flanges are coupled to outlet guide vanes and struts. In one embodiment, outlet guide vanes and struts are coupled coaxially within bypass duct  40 . Optionally, outlet guide vanes may be coupled downstream from struts within bypass duct  40 . 
         [0029]    Fan frame assembly  46  is one of various frame and support assemblies of gas turbine engine assembly  10  that are used to facilitate maintaining an orientation of various components within gas turbine engine assembly  10 . More specifically, such frame and support assemblies interconnect stationary components and provide rotor bearing supports. Fan frame assembly  46  is coupled downstream from fan assembly  14  within bypass duct  40  such that outlet guide vanes and struts are circumferentially-spaced around the outlet of fan assembly  14  and extend across the airflow path discharged from fan assembly  14 . 
         [0030]    In the embodiment shown in  FIG. 1 , a control component for gas turbine engine  10  such as a Full Authority Digital Engine Control (FADEC) is located within the engine nacelle  32  between the first or inner surface  31  and the second or outer surface  33 . 
         [0031]      FIG. 2  is a partial cross-sectional elevational view of the forward portion of an exemplary aircraft gas turbine engine, such as depicted in  FIG. 1 , illustrating installation details of the engine control component  60 . As shown in  FIG. 2 , engine control component  60  is mounted in a compartment  62  within the nacelle  32 , which may house other components (not shown) in addition to the engine control component  60 . Compartment  62  is located between the inner and outer surfaces,  31  and  33 , respectively, of the nacelle  32 . Various locations within the nacelle are possible, depending upon installation and construction details of the gas turbine engine  10  and its control and accessory systems. In the embodiment shown, compartment  62  is approximately midway between the leading and trailing edges of the nacelle. 
         [0032]    In the exemplary embodiment depicted in  FIG. 2 , the control component  60  is housed in a casing  61 , which at least partially surrounds control component  60  to isolate it from the other elements and space inside of the compartment  62 . Duct  63  provides fluid communication between an opening  66  on the inner surface  31  and the casing  61 , such that air may flow through duct  63  between the casing  61  surrounding control component  60  and the inner surface  31 . Duct  64  provides fluid communication between an opening  65  on the outer surface  33  and the casing  61 , such that air may flow through duct  64  between the casing  61  surrounding control component and the outer surface  33 . Accordingly, control component  60  may be exposed to airflow traveling through ducts  63  and  64  from inner surface  31  and/or outer surface  33 . 
         [0033]    The elements of the compartment  62 , controller casing  61 , and ducts  63  and  64  are sized, shaped, adapted, and otherwise configured to suit the characteristics of the particular engine application desired. 
         [0034]    As shown in  FIG. 2 , a valve assembly  67  is included to passively modulate airflow through the system and manage the operating temperatures experienced by the engine control component  60 . 
         [0035]      FIGS. 3 and 4  illustrate in greater detail the valve assembly  67  depicted in  FIG. 2 . As shown in  FIGS. 2 ,  3 , and  4 , the valve assembly  67  is in the open position, so that air flows in the direction of the arrows as shown. Accordingly, air flows from opening  65  on the second or outer surface  33  of the nacelle, through duct  64  with the valve assembly  67  in the open position, through casing  61  surrounding control component  60 , through duct  63 , and out through opening  66  on the first or inner surface  31  of the nacelle. This configuration and operating condition would be typical of a low speed, low Mach number, low altitude operating condition such as ground operations or takeoff. Under such conditions, the temperature differences between that of the air available at opening or port  65  and that of the air inside control component  60  are relatively small, and therefore unrestricted air flow through valve assembly  67  is permissible under thermal shock design considerations. 
         [0036]    As shown in  FIGS. 3 and 4 , the valve assembly  67  in the exemplary embodiment shown includes two semicircular valve flaps pivotally mounted with a hinge pin  68 . The valve assembly  67  is installed such that, in the normal aircraft attitude, gravity acts upon the valve flaps to exert a force on them and bias them downward, to an open position, without any springs, solenoids, or other actuators being necessary. The valve flaps include a feature, such as wedges  70 , which prevent the valve flaps from reaching a fully open (i.e., vertical as shown) position where the flaps would contact each other. This feature aids in preventing a flutter situation where dynamic forces of air flowing over the valve flaps could cause them to oscillate and potentially cause vibration and/or premature wear of parts of the valve assembly. A feature which maintains the valve flaps in a slightly closed (not fully open) position also aids in enabling the valve flaps to respond to a change in airflow direction and pressure, such that they respond and move to a closed condition more readily when the airflow reverses direction. The transition point (open/close) is a function of aircraft speed, aircraft angle of attack, and air density. This turning point can be changed by adding to the valve flaps a calibrated weight and changing the position with respect to the turning axis. As will be more fully described below, the valve flaps also include a feature such as apertures  71  to permit a measured, minor amount of airflow through the system when the valve is in a closed position. Also included are one or more valve stops  69  against which the valve flaps rest when in a fully closed condition, to limit their travel and prevent going over-center. The valve stop or stops  69  may also take the form of a complete annular ring, or ring segments, since with the presence of apertures  71  it is not necessary for any particular amount of clearance to be provided at the edges of the valve flaps for the passage of air in the closed position. A bumper or seal  72  may be provided in association with the valve stop or stops  69  to reduce vibration and reduce or prevent air leaks when the valve is closed. For discontinuous valve stops  69 , the bumper or seal  72  is also discontinuous, while for a valve stop in the form of a continuous annular ring the bumper or seal  72  may be either continuous (forming an annular ring) or discontinuous (forming one or more annular segments). The degree of influence the bumper or seal  72  has on air passage, i.e., the degree of sealing capability, is dependent upon the degree of continuity of the seal  72 . Where the bumper or seal  72  is discontinuous, it tends to function more as a bumper than as a seal. 
         [0037]      FIGS. 5 ,  6 , and  7  correspond to  FIGS. 2 ,  3 , and  4 , and illustrate comparable views of the same components but with  FIGS. 5 ,  6 , and  7  illustrating the reversed airflow (as indicated by the arrows) and the valve assembly  67  in a closed condition. The valve flaps are against the valve stops  69 , and the apertures  71  may be sized and adapted to provide a measured amount of airflow through the substantially closed valve assembly to maintain the engine control component  60  at a desired temperature during valve-closed operation. Apertures  71  are optional, and if desired could be omitted if a total obstruction to airflow through the system were desired. Accordingly, air flows (begins to flow and closes valve, and/or continues to flow in a minor amount through apertures  71 ) from opening  66  on the first or inner surface  31  of the nacelle, through duct  63 , through casing  61  surrounding control component  60 , through duct  64  with the valve assembly  67  in the closed position, and out through opening  65  on the second or outer surface  33  of the nacelle  32 . This configuration and operating condition would be typical of a high speed, high Mach number, high altitude operating condition such as cruise or other high speed flight. Under such conditions, the temperature of the air available to flow through the system is comparatively cool, therefore the amount of cooling air required to keep the control component  60  under normal operating temperature is relatively small. 
         [0038]    Operating the system described herein with minimal or no airflow (mass flow) through the system during aircraft cruise conditions also reduces drag experienced by the aircraft, thereby reducing fuel consumption (often in terms of Specific Fuel Consumption (SFC)). 
         [0039]      FIG. 8  illustrates a variation of the valve configuration described above. In the embodiment of  FIG. 8 , the valve flaps are sized so as to be smaller in diameter when fully closed than the diameter of the duct. This creates a gap  72 , which functions similarly to the apertures  71  described above in terms of permitting a measured, comparatively small amount of airflow through and past the valve assembly  67 . In contrast to the embodiment described above, the valve stop or stops  69  preferably are discrete small elements or ring segments and do not form a complete annular ring, since with the presence of gap  72  it is necessary for clearance to be provided at the edges of the valve flaps for the passage of air in the closed position. 
         [0040]    Other placements of the valve assembly  67  in either duct  63  or duct  64  may be chosen, depending upon the characteristics of the particular application. However, orienting the valve assembly  67  so as to take advantage of gravity as the biasing force to bias the valve to an open position eliminates the need for springs, solenoids, or other actuators being necessary so as to maintain a simple, reliable, lightweight design which operates automatically or passively in response to operating conditions without manual or system intervention. 
         [0041]    The ducts, casing, and valve assembly are sized and configured to provide particular flow characteristics needed to ensure proper temperature operation of the engine control component.  FIGS. 9 and 10  are graphical illustrations of internal operating temperatures experienced by a FADEC unit for one particular engine application during the course of an operating session. The maximum temperature difference, denoted by the reference identifier T in  FIGS. 9 and 10 , is approximately 60 degrees Fahrenheit in  FIG. 9  when operating with a conventional fixed ducting system where the airflow is not regulated and approximately 25 degrees Fahrenheit in  FIG. 10  when a system as described herein is employed. 
         [0042]    In the exemplary embodiment, the ducts and casing may be a fiberglass material, a graphite material, a carbon material, a ceramic material, an aromatic polyamid material such as KEVLAR, a thin metallic material such as, but not limited to, titanium, aluminum, and/or a Metal Matrix Composite (MMC) material, and/or mixtures thereof. Any suitable thermosetting polymeric resin can be used in forming the ducts and casing, for example, vinyl ester resin, polyester resins, acrylic resins, epoxy resins, polyurethane resins, bismalimide resin, and mixtures thereof. Materials should be selected based on the temperature, vibration, and flexure considerations to which the components may be exposed. 
         [0043]    Likewise, the elements of the valve assembly may be constructed from a wide variety of suitable materials, including metals and plastics, particularly those of relatively light weight. 
         [0044]    While much of the foregoing discussion has focused on engine control components (such as FADEC units), the systems and method described herein could be applied to other compartments requiring cooling and ventilation under similar conditions and circumstances. 
         [0045]    This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to make and use the invention. 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.