Abstract:
A gas turbine engine comprises a main fan that delivers air into a bypass duct and into a core engine. A heat exchanger is positioned within the bypass duct and receives a fluid to be cooled from a component associated with the gas turbine engine. A heat exchanger fan is positioned to draw air across the heat exchanger and a control for the heat exchanger fan. The control is programmed to stop operation of the fan during certain conditions, and to drive the heat exchanger fan under other conditions. A method of forming a heat exchanger is also disclosed.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This application relates to the placement of a heat exchanger in a gas turbine engine bypass duct where a fan is controlled to selectively draw air across the heat exchanger. 
         [0002]    Gas turbine engines are known and typically include a fan delivering air into a bypass duct as propulsion air and into a compressor as core airflow. The air is compressed in the compressor and delivered into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving them to rotate. 
         [0003]    As known, there are a number of accessories associated with a gas turbine engine. Various fluids associated with the accessories require cooling. Thus, it is known to place heat exchangers into various locations in a gas turbine engine. 
         [0004]    One such location is in the bypass duct, such that a fluid within the heat exchanger is cooled by the bypass air. However, the pressure ratio delivered by the fan is becoming lower with the recent advent of a gear reduction driving the fan rotor at slower speeds. In addition, during the course of operation of a gas turbine engine associated with an aircraft, the amount of air driven through the bypass duct will vary. Further, the cooling challenges on the heat exchanger will vary. As an example, at takeoff conditions, the cooling load will tend to be greater than it will be at cruise conditions. 
         [0005]    Known gas turbine engines have heat exchanger which are sized for the highest heat load challenge. 
       SUMMARY OF THE INVENTION 
       [0006]    In a featured embodiment, a gas turbine engine comprises a main fan that delivers air into a bypass duct and into a core engine. A heat exchanger is positioned within the bypass duct and receives a fluid to be cooled from a component associated with the gas turbine engine. A heat exchanger fan is positioned to draw air across the heat exchanger and a control for the heat exchanger fan. The control is programmed to stop operation of the fan during certain conditions, and to drive the heat exchanger fan under other conditions. 
         [0007]    In another embodiment according to the previous embodiment, the control commands a motor to drive the heat exchanger fan under high heat load conditions. 
         [0008]    In another embodiment according to any of the previous embodiments, the high heat load conditions at least include takeoff of an associated aircraft. 
         [0009]    In another embodiment according to any of the previous embodiments, the control commands the motor to turn off the heat exchanger fan under low heat load conditions. 
         [0010]    In another embodiment according to any of the previous embodiments, the low heat load conditions include a cruise condition of an associated aircraft. 
         [0011]    In another embodiment according to any of the previous embodiments, the motor is positioned within a space of a fairing such that it is out of a path of air downstream of the heat exchanger fan. 
         [0012]    In another embodiment according to any of the previous embodiments, the heat exchanger is formed through an additive manufacturing process. 
         [0013]    In another embodiment according to any of the previous embodiments, air downstream of the heat exchanger fan is delivered to mix back into a bypass airflow path. 
         [0014]    In another embodiment according to any of the previous embodiments, a shaft connects a rotor of the heat exchanger fan to the motor. 
         [0015]    In another embodiment according to any of the previous embodiments, the control commands the motor to turn off the heat exchanger fan under low heat load conditions. 
         [0016]    In another embodiment according to any of the previous embodiments, the low heat load conditions include a cruise condition of an associated aircraft. 
         [0017]    In another embodiment according to any of the previous embodiments, the motor is positioned within a space of a fairing such that it is out of a path of air downstream of the heat exchanger fan. 
         [0018]    In another embodiment according to any of the previous embodiments, a shaft connects a rotor of the heat exchanger fan to the motor. 
         [0019]    In another embodiment according to any of the previous embodiments, the heat exchanger is formed through an additive manufacturing process. 
         [0020]    In another embodiment according to any of the previous embodiments, air downstream of the heat exchanger fan is delivered to mix back into a bypass airflow path. 
         [0021]    In another embodiment according to any of the previous embodiments, the motor is positioned within a space of a fairing such that it is out of a path of air downstream of the heat exchanger fan. 
         [0022]    In another embodiment according to any of the previous embodiments, a shaft connects a rotor of the heat exchanger fan to the motor. 
         [0023]    In another embodiment according to any of the previous embodiments, the heat exchanger is formed through an additive manufacturing process. 
         [0024]    In another embodiment according to any of the previous embodiments, air downstream of the heat exchanger fan is delivered to mix back into a bypass airflow path. 
         [0025]    In another featured embodiment, a method of forming a heat exchanger comprises the steps of determining an available space within a gas turbine engine for the heat exchanger, and forming the heat exchanger to conform to the available space utilizing an additive manufacturing process. 
         [0026]    These and other features may be best understood from the following drawings and specification. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]      FIG. 1  shows a gas turbine engine. 
           [0028]      FIG. 2  shows a portion of the  FIG. 1  gas turbine engine. 
           [0029]      FIG. 3  shows a manufacturing technique for a heat exchanger. 
       
    
    
     DETAILED DESCRIPTION 
       [0030]      FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
         [0031]    The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
         [0032]    The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0033]    The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
         [0034]    The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
         [0035]    A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second). 
         [0036]    A fan rotor  99  is shown schematically in  FIG. 2  delivering air within a nacelle  100  and into a bypass duct  102 . A heat exchanger  104  is positioned within the path of air in the bypass duct  102 . A first fluid line  106  and a second fluid line  108  connect an accessory  110  of the gas turbine engine to flow a fluid through the heat exchanger  104 . As an example, lines  106  and  108  may move oil from an accessory  110  (which may be a bearing, a generator, etc.) through the heat exchanger  104 . At any rate, some operational fluid is cooled within the heat exchanger  104  by having air from the bypass duct  102  pass over it. 
         [0037]    A housing  112  is positioned outwardly of the heat exchanger  104 . As can be appreciated, there is a unique space radially inwardly of the housing  112  which receives the heat exchanger  104 . A fairing  114  is also positioned within the bypass duct  102  and together with the housing  112  forms the space for receiving the heat exchanger  104 . A heat exchanger fan rotor  116  selectively draws air across the heat exchanger  104 . Air downstream of the heat exchanger  104  leaves through an exit  118  to mix with a normal downstream bypass path  119 . A motor  120  for the fan rotor  116  sits within an internal space  122  of the fairing  114 . A control  121  controls the speed of the motor  120 . 
         [0038]    During takeoff, when the heat load on the heat exchanger  104  will be high, the control  121  commands the motor  120  to drive the fan rotor  116 . However, during lower heat load conditions, such as cruise, the control  121  commands the motor  120  to turn the fan rotor  116  off. During intermediate operational points, the fan rotor  116  may be rotated at slower speeds as appropriate. 
         [0039]    A shaft  123  connects the fan rotor  116  to the motor  120  such that the motor  120  can be positioned in the space  122 , where it will not be exposed to heat in the air in the downstream portion  118 . 
         [0040]    The size of the heat exchanger  104  may be reduced relative to the prior art, as the air flow across heat exchanger  104  will be optimized due to the fan  116 . 
         [0041]      FIG. 3  shows another feature. An intermediate heat exchanger  130  is being formed by an additive manufacturing tool  132 . As known, additive manufacturing includes a number of distinct steps which essentially lay down material  134  in layers to form a final component, such as the heat exchanger  104 . With additive manufacturing, it is relatively easy to form a heat exchanger  104  of any unique shape such that the space available in a particular engine is utilized. 
         [0042]    The  FIG. 3  method of forming a heat exchanger comprises the steps of determining an available space within a gas turbine engine for the heat exchanger, and forming a heat exchanger  130  to conform to the available space utilizing an additive manufacturing process  132 / 134 . 
         [0043]    Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.