Patent Application
20180023475
Gas turbine engines operate at relatively high temperatures. Engine components that are exposed to the high temperatures may be fabricated from superalloys, may include barrier coatings, and/or may include cooling schemes to reduce component temperature.
A gas turbine engine according to an example of the present disclosure includes a core nacelle, and a core engine disposed in the core nacelle. The core engine has at least a compressor section, a combustor section, and a turbine section, which define a core flow path. There is a bypass duct radially outwards of the core nacelle, and a component is disposed in the core engine. A cooling cavity is disposed outside of the bypass duct. A heat pipe contains a working medium sealed therein. The heat pipe includes a first section configured to accept thermal energy and a second section configured to dissipate the thermal energy in the cooling cavity. A tap from at least one of the compressor section or the bypass duct is configured to provide bleed air to the cooling cavity and dissipate the thermal energy from the second section.
In a further embodiment of any of the foregoing embodiments, the cooling cavity is disposed radially between the core flow path and the core nacelle.
In a further embodiment of any of the foregoing embodiments, the core nacelle defines an outer boundary of the cooling cavity.
In a further embodiment of any of the foregoing embodiments, the tap is from the compressor section.
In a further embodiment of any of the foregoing embodiments, the tap is from the bypass duct.
In a further embodiment of any of the foregoing embodiments, the cooling cavity includes a discharge port that opens into the bypass duct.
In a further embodiment of any of the foregoing embodiments, the core nacelle includes an aft-sloped wall section, and the discharge port is in the aft-sloped wall section.
In a further embodiment of any of the foregoing embodiments, the first section includes at least two legs that are axially separated.
In a further embodiment of any of the foregoing embodiments, the second section is axially offset from the first section.
In a further embodiment of any of the foregoing embodiments, the component is a static component.
In a further embodiment of any of the foregoing embodiments, the cooling cavity is radially outwards of the bypass duct.
In a further embodiment of any of the foregoing embodiments, the bypass duct includes a guide vane, and the heat pipe extends through the guide vane into the cooling cavity.
In a further embodiment of any of the foregoing embodiments, the tap includes an annular flow guide or passage.
A gas turbine engine according to an example of the present disclosure includes a core engine defining a core flow path, and a bypass duct radially outwards of the core engine. There is a component in the core engine and an environmentally-controllable cavity disposed outside of the core flow path and outside of the bypass duct. A heat pipe contains a working medium sealed therein. The heat pipe has a first section configured to accept thermal energy and a second section configured to dissipate the thermal energy in the environmentally-controllable cavity. The environmentally-controllable cavity is operational to adjust an environment therein to variably dissipate the thermal energy from the second section.
A further embodiment of any of the foregoing embodiments includes a controller configured to adjust the environment in the environmentally-controllable cavity based on a temperature of the component.
A further embodiment of any of the foregoing embodiments includes the component is in a final one or two stages of the compressor section.
A gas turbine engine according to an example of the present disclosure includes a core engine that has at least a compressor section, a combustor section, and a turbine section. A component is disposed in the core engine, and a tap configured to provide bleed air from at least one of the compressor section or a location outside of the core engine. A heat pipe contains a working medium sealed therein. The heat pipe includes a first section configured to accept thermal energy and a second section configured to dissipate the thermal energy to the bleed air from the tap.
In a further embodiment of any of the foregoing embodiments, the tap is configured to provide bleed air from the compressor section.
A further embodiment of any of the foregoing embodiments includes a bypass duct outwards of the core engine, and the tap is configured to provide bleed air from the bypass duct.
A further embodiment of any of the foregoing embodiments includes a discharge port in communication with the tap and configured to provide the bleed air into the bypass duct.
The various features and advantages of the present disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.
The fan section 22 drives air along a bypass flow path B in a bypass duct 31 defined within a nacelle 33, 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, the examples herein are not limited to use with two-spool turbofans and may be applied to other types of turbomachinery, including direct drive engine architectures, three-spool engine architectures, and ground-based turbines.
The 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 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 the 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. The flight condition of 0.8 Mach and 35,000 ft, 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.
In the example shown, the core engine 29 is disposed in a core nacelle 29a. The core nacelle 29a defines an inner (radial) boundary of the bypass duct 31. The core nacelle 29a may be contoured to provide a desired geometry of the bypass duct 31. In this regard, the core nacelle 29a is spaced radially outward of the core engine 29 such that there is an intermediate cooling cavity 62 outside of (i.e., not in) the bypass duct 31 and outside of (i.e. not in) the core flow path C. In this example, the cooling cavity 62 is radially between the core flow path C of the engine core 29 and the core nacelle 29a. The cooling cavity 62 spans axially from the compressor section 24 to the turbine section 28. Alternatively, the cooling cavity 62 may span with two of these sections or may span with a single one of these sections. The core nacelle 29a may define an outer (radial) boundary of the cooling cavity 62.
The engine 20 includes one or more heat pipes 64, an example of which is shown in an isolated view
The first section 64a is associated with a component 68 and is configured to accept thermal energy from the core engine 29. For instance, the component 68 is in the core engine 29 and may be located in the core gas path C, at least partially bound the core gas path C, or be exposed to hot combustion gases in the core gas path C. In this example, the component 68 is a static turbine vane in the turbine section 28, such as a vane in the second (or high) pressure turbine 54. Alternatively, the component 68 may be an outer air seal, a wall, or other non-rotating component that is subject to high temperatures in the engine 20. In one example, the component 68 is hollow and the first section 64a is located at least partially within the hollow cavity of the component 68 in order to accept and remove thermal energy from the component 68 and/or accept thermal energy transferred from the hot combustion gases in the core gas path C. For vanes or other circumferentially-arranged components, each such component 68 may have a corresponding heat pipe 64 for thermal energy removal. Additionally or alternatively, the first section 64a or a portion thereof may be integrally formed with the component 68. For instance, the hollow cavity or a portion of the hollow cavity of the component 68 may be configured as the first section 64a or a portion of the first section 64a.
The second section 64b is configured to dissipate (reject) the thermal energy in the cooling cavity 62. For instance, in this example, the second section 64b is located at least partially within the cooling cavity 62, i.e., at least a heat transfer surface of the second section 64b is exposed in the cooling cavity 62 or is located in close proximity to the cooling cavity 62 to reject thermal energy thereto.
The engine 20 further includes a tap 70 that serves to facilitate the removal of the thermal energy from the second section 64b of the heat pipe 64. For instance, the tap 70 may include an inlet port, a duct or passage, a flow guide in the cooling cavity 62, the cooling cavity 62 or portion thereof as a duct or passage, or any combination of such features. An example flow guide or passage is shown at 71. As an example, the flow guide or passage 71 may be annular. The flow guide or passage 71 may serve to guide flow of bleed air and/or diffuse the bleed air. In this example, the tap 70 is from the bypass duct 31 at a location forward or upstream of the second section 64b. As an example, the tap 70 may include a valve 70a or similar device that is operable to control the flow of the bleed air from the bypass duct 31 to the cooling cavity 62.
The tap 70 is configured to provide bleed air from the bypass duct 31 to the cooling cavity 62 and second section 64b to facilitate removal of the thermal energy from the second section 64b. For instance, the bleed air flows over at least a portion of the second section 64b to absorb and remove the thermal energy from the second section 64b. A single tap 70 may be used to provide bleed air to a single second section 64b. For instance, there may be a plurality of taps 70 providing bleed air to a corresponding plurality of second sections 64b of heat pipes 64 that are associated with a plurality of components 68 (e.g., a 1:1 ratio of taps to condensation sections). Alternatively, a single tap 70 may provide bleed air to multiple second sections 64b, or a single tap 70 may provide bleed air to all the second sections 64b.
The cooling cavity 62 also includes one or more discharge ports 62a that open into the bypass duct 31. The discharge port 62a may be downstream or aft of the second section 64b. For example, the discharge port 62a may include a flow control valve 62b for controlling flow from the second cavity 62 back into the bypass duct 31. There may be a plurality of discharge ports 62a, such as one discharge port 62a for each tap 70 or for each second section 64b. Alternatively, the discharge port 62a may be annular or may be an arc segment. There may be a greater number or a lesser number of discharge ports 62a than taps 70 or second sections 64b.
The tap 70 provides bleed air to remove thermal energy from the second section 64b; however, the removal of the bleed air from the bypass duct 31 penalizes thrust efficiency. The bleed air receives thermal energy from the second section 64b and increases in temperature and pressure in the closed second cavity 62. The bleed air is then released through the discharge port 62a back into the bypass duct 31 to recover at least a portion of the lost thrust. Thus, heat removal from the component 68, which might otherwise penalize engine cycle efficiency, serves for thrust recovery. Additionally, since the second section 64b is located at least partially within the cooling cavity 62, rather than in the bypass duct 31, the second section 64b does not physically disturb or impede flow in the bypass duct 31. The heat pipe 64 thus provides for a compact heat removal system that avoids physically interfering with the bypass flow.
In a further example, the discharge port 62a is shaped and/or oriented to control the direction of the discharged stream of bleed air into the bypass duct. Streaming the bleed air into the bypass duct 31 in a direction that is locally misaligned with the flow of bypass air in the bypass duct 31 potentially results in flow turbulence and/or loss of energy. Directing the stream of bleed air in a direction that is locally aligned with the flow direction of bypass air in the bypass duct 31 facilitates reduction in turbulence and/or energy loss. As an example, to permit greater local alignment of the discharged bleed air, the discharge port 62a may be located in an aft-sloped wall section 29a1 of the core nacelle 29a.
In this example, there is a tap 470 from the bypass duct 31 at a location forward or upstream of the second section 464b. The tap 70 may include a valve 70a or similar device that is operable to control the flow of the bleed air from the bypass duct 31 to the cooling cavity 462. The cooling cavity 462 also includes one or more discharge ports 462a that open into the bypass duct 31 for thrust recovery as described elsewhere herein.
Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments.
The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.
