Patent Publication Number: US-2020300164-A1

Title: Turbine engine configuration including pylon mounted heat exchanger

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application claims priority to U.S. Provisional Application No. 62/822315 filed on Mar. 22, 2019. 
    
    
     TECHNICAL FIELD 
     The present disclosure relates generally to turbine engine heat exchanger configurations, and particularly to a pylon mounted heat exchanger for a gas turbine engine. 
     BACKGROUND 
     Gas turbine engines, such as those utilized in commercial and military aircraft, include a compressor section that compresses air, a combustor section in which the compressed air is mixed with a fuel and ignited, and a turbine section across which the resultant combustion products are expanded. The expansion of the combustion products drives the turbine section to rotate. As the turbine section is connected to the compressor section via a shaft, the rotation of the turbine section further drives the compressor section to rotate. In some examples, a fan is also connected to the shaft and is driven to rotate via rotation of the turbine as well. 
     Within the gas turbine engine, multiple fluids such as air or oil, are actively cooled and provided to other engine components to cool and/or lubricate the other engine components. To achieve the active cooling, the gas turbine engines include heat exchangers within engine core or the engine nacelle. 
     SUMMARY OF THE INVENTION 
     In one exemplary embodiment an aircraft includes at least one gas turbine engine connected to a wing via an engine pylon, the gas turbine engine including an engine core having a compressor section, a combustor section, and a turbine section and at least one actively cooled engine system, a fan disposed fore of the engine core and rotatably connected to the engine core via a shaft, the engine pylon being disposed within a bifurcation and intersecting a fan stream of the fan and including a pylon space, and at least one heat exchanger disposed within the pylon space, the heat exchanger including a first inlet connect to the at least one actively cooled engine system and a first outlet connected to the at least one actively cooled engine system. 
     In another example of the above described aircraft, the at least one heat exchanger includes at most three heat exchangers. 
     In another example of any of the above described aircraft, the at least one heat exchanger includes at least one air-air heat exchanger. 
     In another example of any of the above described aircraft, the at least one heat exchanger includes at least one air-oil heat exchanger. 
     In another example of any of the above described aircraft, the at least one heat exchanger includes at least one air-air heat exchanger. 
     In another example of any of the above described aircraft, the at least one heat exchanger includes a second inlet disposed in the fan stream and configured to ingest fan stream air. 
     In another example of any of the above described aircraft, the at least one heat exchanger includes a second outlet configured to exhaust spent fan stream air to the fan stream. 
     Another example of any of the above described aircraft further includes an electric generator connected to said engine core such that rotation of the electric generator is driven by the shaft and an electric compressor driven by the electric generator. 
     In another example of any of the above described aircraft, a fluid outlet of the electric compressor is fluidly connected to an aircraft cabin environmental cooling system (ECS). 
     An exemplary method for providing additional heat capacity in a gas turbine engine includes removing a cooling fluid from at least one engine system within a gas turbine engine nacelle, providing the removed cooling fluid to a heat exchanger disposed within a pylon space of a pylon connecting the gas turbine engine nacelle to an aircraft wing, cooling the cooling fluid within the heat exchanger, and returning the cooling fluid to the at least one engine system. 
     In another example of the above method for providing additional heat capacity in a gas turbine engine, removing the cooling fluid from at least one engine system and returning the cooling fluid to the at least one engine system comprise removing the cooling fluid from a single engine system in the at least one engine system and returning the cooling fluid to the single engine system in the at least one engine system. 
     In another example of any of the above methods for providing additional heat capacity in a gas turbine engine, removing the cooling fluid from at least one engine system and returning the cooling fluid to the at least one engine system comprise removing the cooling fluid from a first engine system in the at least one engine system and returning the cooling fluid to a second engine system in the at least one engine system, wherein the first engine system and the second engine system are part of a single coolant cycle. 
     In another example of any of the above methods for providing additional heat capacity in a gas turbine engine, providing the removed cooling fluid to a heat exchanger comprises providing the cooling fluid to three heat exchangers. 
     In another example of any of the above methods for providing additional heat capacity in a gas turbine engine, cooling the cooling fluid within the heat exchanger comprises ingesting fan stream air along a cooling flowpath within the heat exchanger and passing the cooling fluid through the heat exchanger along a circuitous route. 
     In another example of any of the above methods for providing additional heat capacity in a gas turbine engine, the heat exchanger is an air-air heat exchanger. 
     In another example of any of the above methods for providing additional heat capacity in a gas turbine engine, the heat exchanger is an air-oil heat exchanger. 
     Another example of any of the above methods for providing additional heat capacity in a gas turbine engine further includes providing compressed air to a cabin environmental cooling system (ECS) from an electric compressor. 
     Another example of any of the above methods for providing additional heat capacity in a gas turbine engine further includes powering the electric compressor via an electrical generator, wherein the electric generator is driven by a shaft within the gas turbine engine nacelle. 
     In one exemplary embodiment a fluid cooling system for a gas turbine engine includes a fan section defining a fan stream, an engine core connected to the fan section and including at least one engine system interior to the engine core, at least a first heat exchanger disposed in the fan stream and exterior to the engine core, the first heat exchanger being configured to receive a cooling fluid from the at least one engine system and configured to return the cooling fluid to the at least one engine system. 
     In another example of the above described fluid cooling system for a gas turbine engine the at least a first heat exchanger comprises a plurality of no more than three heat exchangers. 
     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  illustrates a high level schematic view of an exemplary imaging system. 
         FIG. 2  schematically illustrates an exemplary engine mounted to an aircraft wing via a mounting pylon. 
         FIG. 3  schematically illustrates an alternative example engine mounted to an aircraft wing via a mounting pylon. 
     
    
    
     DETAILED DESCRIPTION 
       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 . The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , and also 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. 
     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. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects, 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 a 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  may be 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. 
     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 the low pressure compressor, or aft of the combustor section  26  or even aft of turbine section  28 , and fan  42  may be positioned forward or aft of the location of gear system  48 . 
     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 and less than about 5:1. 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. 
     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). 
     Existing gas turbine engines have moved toward accommodating higher electrical aircraft demands within the aircraft and accessory loads (i.e. electrical loads external to the engine  20 ) as well as propulsion loads (i.e. electrical loads within the engine  20 ). Increased power demands require the incorporation of additional generators, motors and other electrical components within the engine  20 . Incorporation of these additional electrical components causes the generation of parasitic heat loads within the engine  20  that must be cooled in order to prevent overheating. In addition, physical space within the engine  20  is at a premium, and inclusion of the electrical components associated with higher electrical aircraft and aircraft engines can stress the space limitations. 
     One system that has the potential to use more electrical power, and also has a significant impact on aircraft engine performance is an Environmental Cooling System (ECS). The ECS is an air system that provides conditioned breathing air to the cabin for aircraft crew and passengers. In existing aircraft, the ECS is comprised of ducts that bleed compressed air off the engine compressor and then direct the air to a series of coolers (e.g. air-to-air coolers and vapor cycle coolers) that condition the air for delivery to an aircraft cabin. 
     Due to the high temperatures of compressor bleed air, existing systems positon one or more of the coolers for the ECS in a space within a mount pylon connecting the engine  20  to the aircraft body. The mount pylon is located within a bifurcation in the fan stream. Due to the position in the fan stream fan air can be ingested into the embedded cooler within the mount pylon as a cooling source for the bleed air prior to providing the bleed air to the cabin ECS system. In order to facilitate the removal of fluid from the engine core, and the return of the fluid to the engine core, additional integration between the engine and the aircraft is necessary to allow the engine to be removed and repaired. To facilitate this, the fluid lines can include quick-disconnect couplings that allow for attachment and separation of the two systems. 
     With continued reference to  FIG. 1 ,  FIG. 2  schematically illustrates an exemplary engine  120  mounted to an aircraft wing  110  via a mounting pylon  130 . In the more electric aircraft architecture, air for the ECS is acquired via an electrically driven compressor mounted within the aircraft and bleed air from the turbine engine  120  is not required to drive air to the ECS. In the example architecture, one or more electric motors  140  drive an electric compressor  142  which provides compressed air to the ECS within the aircraft. The electric motors  140  receive generated electricity from a generator  141  within the engine  120 . The air that is compressed by the electric compressor  142  is sourced from any ambient air source, and is not sourced from within the engine core. The utilization of the electric motor  140  creates a new heat load that is not present in conventional aircraft. 
     Due to the utilization of the electric compressor  142 , ambient air can be obtained and compressed for utilization in the ECS, and it is not necessary to pre-cool the air, as would be required when pulling air from a compressor bleed. This in turn allows the cooler that is positioned in the pylon space in conventional systems to be replaced with an engine system cooler  150 . The engine system cooler  150  is positioned in the pylon space  160 , and includes a first fluid inlet  152  and a first fluid outlet  154 . The first fluid inlet  152  connects an internal passage  151  within the cooler  150  to an engine system  121  such that fluid from the engine system  121  can be drawn out of the engine  120  and provided to the internal passage  151 . The first fluid outlet  154  returns the fluid from the internal passage  151  to the engine  120 . In some cases, the fluid is returned to the same engine system  121  that it originated from. In alternative cases, the fluid is returned to a distinct engine system  121  as part of a larger coolant cycle. 
     The cooler  150  also includes a second inlet  156  configured to receive an airflow  160  from a fan section  170  of the engine  120 . The second inlet  156  provides the relatively cold fan air to a second internal passage  153  (or set of second internal passages  153 ) within the cooler  150 . A second outlet  158  dumps spent air from the cooler  150  into a bypass flow and allows the spent air to exit the engine  120  structure. 
     In some examples, the cooler  150  is an air-air cooler, and the fluid extracted from the engine  120  is a compressed air from a compressor bleed. By way of example, a cooled cooling air system could utilize this configuration. To provide cooled engine air from a compressor bleed to a turbine on board injection system, a compressor on board injection system, or to any other engine system utilizing cooled cooling air. 
     In other examples, the cooler  150  is an air-oil cooler with the first inlet  152  withdrawing an oil based lubricant/coolant, cooling the oil based lubricant/coolant using the cooler  150 , and returning the cooled oil based lubricant/coolant to the engine  120  via the first outlet  154 . 
     With continued reference to  FIG. 2 , and with like numerals indicating like elements,  FIG. 3  schematically illustrates an alternative example engine  220  mounted to an aircraft wing  210  via a mounting pylon  230 . Air for the ECS is acquired via an electrically driven compressor mounted within the aircraft and bleed air from the turbine engine  220  is not required to drive air to the ECS. In the example architecture, one or more electric motors  240  drive an electric compressor  242  which provides compressed air to the ECS within the aircraft. The air that is compressed by the electric compressor  242  is sourced from any ambient air source, and is not sourced from within the engine core. The utilization of the electric motor  240  creates a new heat load that is not present in conventional aircraft. 
     Disposed within the pylon space  260  are three heat exchangers  250 ,  250 ′,  250 ″. In some examples, each of the heat exchangers  250 ,  250 ′,  250 ″ is connected to the same engine system  221 . In alternative examples, each heat exchanger  250 ,  250 ′,  250 ″ is connected to a distinct engine system  221 . The heat exchangers  250 ,  250 ′,  250 ″ each include internal configurations substantially similar to that described above with regards to the heat exchanger  150  and illustrated in  FIG. 1 , and can be air-air heat exchangers, air-oil heat exchangers, or any combination of the two. 
     Positioning an engine system cooler, such as the engine system coolers  150 ,  250  in the pylon space  160 ,  260  provides for an improved heat exchanger efficiency, because being positioned in the engine nacelle bifurcation allows the heat exchanger  150 ,  250  efficiency to benefit from a relatively high source pressure obtained from the ram effect of having the entire inlet  156 ,  256  duct perimeter be entirely in the fan stream  161 . In contrast, engine-mounted heat exchangers that are flush with the engine nacelle include only the outer edge of the inlet duct in the fan stream  161 . By taking advantage of the available ram effect, a higher-efficiency, smaller-sized heat exchanger can be used in the pylon space  160  as compared to heat exchangers that are packaged interior to the engine nacelle. 
     Further, including the engine system cooler  150 ,  250  in the pylon space  160 ,  260  improves the nacelle packaging efficiency by removing bulky components from the engine nacelle. Heat exchangers mounted within the engine nacelle are among the physically largest components that fit between the engine and nacelle and inclusion of the heat exchanger within the nacelle may prevent the achievement of the lowest-drag or “ideal’ nacelle flowpath. This affect is especially prominent for a high heat-load applications. By taking advantage of the pylon packaging instead of under the engine nacelle it may be possible to design more efficient nacelle lines. 
     Finally, moving the heat exchangers  150 ,  250  to the pylon space  160 ,  260  provides for additional heat capacity within the engine. For especially-high heat load applications there is insufficient space within an engine nacelle to package the required-size heat exchanger(s). By placing the heat exchanger  150 ,  250  within the pylon space  160 ,  260  this limitation is overcome via the additional space as well as the potential for a reduced-size heat-exchanger due to utilization of improved ram-air effect. 
     It is further understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. 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.