Patent Application
20170307311
This application relates to a heat exchanger for providing cooling air in a gas turbine engine.
Gas turbine engines are known and typically include a fan delivering air into a bypass duct as propulsion air, and further providing air into a core housing. Air in the core housing passes into a compressor where it is compressed, and then into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving them to rotate.
As is known, turbine components see very high temperatures and thus cooling air has been typically provided to those components. Historically, the fan and a low pressure compressor have rotated as a single unit along with a fan drive turbine. However, more recently, a gear reduction has been placed between the fan rotor and the fan drive turbine. This allows the fan rotor to rotate at slower speeds and the fan drive turbine to rotate at faster speeds. This raises the challenges on the turbine components and requires more efficient provision of the cooling air.
At the same time, the overall pressure ratio provided by the compressor has increased. Historically, the air to cool the turbine components has been tapped from a location downstream of a highest pressure location on the compressor. However, with the increase in overall pressure ratio, this air has become hotter.
The heat exchangers for cooling this air are thus subject to extreme challenges.
In a featured embodiment, a heat exchanger system for use in a gas turbine engine has a plurality of circumferentially spaced heat exchangers. The spaced heat exchangers are formed of a nickel alloy material including more than 50-percent by volume gamma-prime intermetallic phase material.
In another embodiment according to the previous embodiment, the heat exchangers are formed of elongated members having fins on an outer surface.
In another embodiment according to any of the previous embodiments, the elongated members are tubes.
In another embodiment according to any of the previous embodiments, the elongated members extend radially outwardly to an elbow which takes air radially outwardly to the elbow and a second elongated member returns air radially inwardly into a housing for the engine.
In another embodiment according to any of the previous embodiments, there are a plurality of axially spaced heat exchangers.
In another featured embodiment, a gas turbine engine has a compressor section, a combustor section, and a turbine section. A core housing contains the compressor section, the combustor and the turbine section. A first conduit taps hot compressed air to be cooled and passes the air to a heat exchanger. The air is cooled in the heat exchanger and returned to a return conduit. The return conduit passes the cooled air to the turbine section. The heat exchanger includes a plurality of circumferentially spaced heat exchangers. The circumferentially spaced heat exchangers are formed of a nickel alloy material including more than 50-percent by volume gamma-prime intermetallic phase material.
In another embodiment according to the previous embodiment, the heat exchangers are formed of elongated members having fins on an outer surface.
In another embodiment according to any of the previous embodiments, the elongated members are tubes.
In another embodiment according to any of the previous embodiments, the elongated members extend radially outwardly to an elbow which takes air radially outwardly to the elbow and a second elongated member returns air radially inwardly into a housing for the engine.
In another embodiment according to any of the previous embodiments, there are a plurality of axially spaced heat exchangers.
In another embodiment according to any of the previous embodiments, the heat exchanger is positioned in a bypass duct outwardly of the core housing.
In another embodiment according to any of the previous embodiments, the heat exchanger is positioned forwardly of a pivot point for a pivoting portion of the core housing, and the exchanger being positioned radially outwardly of a fixed inner structure.
In another embodiment according to any of the previous embodiments, the heat exchanger is positioned within the core housing.
In another embodiment according to any of the previous embodiments, a pivoting door selectively allows bypass air to pass over the heat exchanger for cooling the heat exchanger.
In another embodiment according to any of the previous embodiments, a valve selectively controls the flow of the compressed air to the heat exchanger.
In another embodiment according to any of the previous embodiments, a valve selectively controls the flow of the compressed air to the heat exchanger.
In another embodiment according to any of the previous embodiments, a duct for controlling the flow of air downstream of the heat exchanger is positioned upstream of a fan nozzle plane of the gas turbine engine.
In another embodiment according to any of the previous embodiments, a ramp causes a lower pressure downstream of the ramp to facilitate flow of the bypass air over the heat exchanger and into an exhaust.
In another embodiment according to any of the previous embodiments, a duct for controlling the flow of air downstream of the heat exchanger is positioned downstream of a nozzle plane.
In another embodiment according to any of the previous embodiments, the return conduit passing into a strut and radially inwardly to pass to the turbine section.
These and other features may be best understood from the following drawings and specification.
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 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.
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.
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.
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).
Preferably there are a plurality of circumferentially spaced conduits 110, 112 and struts 114.
The air reaches an elbow 124 and then returns inwardly through another tube 126 which may be provided with fins 128 and also trip strips, if desired. That air returns to the conduit 112.
As shown in
In embodiments, the heat exchanger tubes 118 and 126, and optionally the fins 120 and 128 and trip strips 122 may be formed of a super alloyed material typically utilized for turbine components. In particular, a cast nickel alloy material including more than 50-percent by volume gamma-prime (Y′). Intermetallic phase material may be utilized as the Y′ material. The intermetallic phase material may be Ni3AL or Ni3TI as examples.
The use of this alloy, which has been typically reserved for use in the turbine, allows the heat exchanger to survive much higher temperatures than with typical heat exchangers utilized in gas turbine engines. As such, the challenges mentioned above can be addressed.
Cooling air passes over the heat exchanger 142 and through a duct 144, which may also be selectively closed by control 145. Air is tapped through a valve 146 from the hot location, as in the
The
Although embodiments of this invention have 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.
