Patent Description:
Known heat transfer devices are disclosed in "<NPL>, and<CIT>. <CIT> discloses a hybrid cooling system for a gas turbine engine.

According to a first aspect of the present invention, there is provided a gas turbine engine as defined by claim <NUM>.

According to a second aspect of the present invention, there is provided a method of cooling a first fluid in a gas turbine engine as defined by claim <NUM>.

<FIG> schematically illustrates an example gas turbine engine <NUM> that includes a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section <NUM> drives air along a bypass flow path B while the compressor section <NUM> draws air in along a core flow path C where air is compressed and communicated to the combustor section <NUM>. In the combustor section <NUM>, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section <NUM> where energy is extracted and utilized to drive the fan section <NUM> and the compressor section <NUM>.

Although the disclosed non-limiting embodiment depicts a geared turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with geared turbofans as the teachings may be applied to other types of traditional turbine engines. For example, the gas turbine engine <NUM> can have a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.

The example gas turbine engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearing systems <NUM>. It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided.

The low speed spool <NUM> generally includes an inner shaft <NUM> that connects a fan <NUM> and a low pressure (or first) compressor section <NUM> to a low pressure (or first) turbine section <NUM>. The inner shaft <NUM> drives the fan <NUM> through a speed change device, such as a geared architecture <NUM>, to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The high-speed spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure (or second) compressor section <NUM> and a high pressure (or second) turbine section <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine central longitudinal axis A.

A combustor <NUM> is arranged between the high pressure compressor <NUM> and the high pressure turbine <NUM>. In one example, the high pressure turbine <NUM> includes at least two stages to provide a double stage high pressure turbine <NUM>. In another example, the high pressure turbine <NUM> includes only a single stage. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.

The example low pressure turbine <NUM> has a pressure ratio that is greater than about <NUM>. The pressure ratio of the example low pressure turbine <NUM> is measured prior to an inlet of the low pressure turbine <NUM> as related to the pressure measured at the outlet of the low pressure turbine <NUM> prior to an exhaust nozzle.

The mid-turbine frame <NUM> further supports bearing systems <NUM> in the turbine section <NUM> as well as setting airflow entering the low pressure turbine <NUM>.

The air in the core flow path C is compressed by the low pressure compressor <NUM> then by the high pressure compressor <NUM> mixed with fuel and ignited in the combustor <NUM> to produce high speed exhaust gases that are then expanded through the high pressure turbine <NUM> and low pressure turbine <NUM>. The mid-turbine frame <NUM> includes vanes <NUM>, which are in the core flow path C and function as an inlet guide vane for the low pressure turbine <NUM>. Utilizing the vane <NUM> of the mid-turbine frame <NUM> as the inlet guide vane for low pressure turbine <NUM> decreases the length of the low pressure turbine <NUM> without increasing the axial length of the mid-turbine frame <NUM>. Reducing or eliminating the number of vanes in the low pressure turbine <NUM> shortens the axial length of the turbine section <NUM>. Thus, the compactness of the gas turbine engine <NUM> is increased and a higher power density may be achieved.

The disclosed gas turbine engine <NUM> in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine <NUM> includes a bypass ratio greater than about six (<NUM>), with an example embodiment being greater than about ten (<NUM>). The example geared architecture <NUM> is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about <NUM>.

In one disclosed embodiment, the gas turbine engine <NUM> includes a bypass ratio greater than about ten (<NUM>:<NUM>) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor <NUM>. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.

A significant amount of thrust is provided by the air in the bypass flow path B due to the high bypass ratio. The fan section <NUM> of the engine <NUM> is designed for a particular flight condition -- typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (about <NUM>,<NUM> metres). The flight condition of <NUM> Mach and <NUM>,<NUM> ft (<NUM>,<NUM>), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of poundmass (Ibm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point.

In another non-limiting embodiment the low fan pressure ratio is less than about <NUM>.

The example gas turbine engine includes the fan <NUM> that comprises in one non-limiting embodiment less than about <NUM> fan blades. In another non-limiting embodiment, the fan section <NUM> includes less than about <NUM> fan blades. Moreover, in one disclosed embodiment the low pressure turbine <NUM> includes no more than about <NUM> turbine rotors schematically indicated at <NUM>. In another non-limiting example embodiment the low pressure turbine <NUM> includes about <NUM> turbine rotors. A ratio between the number of fan blades and the number of low pressure turbine rotors is between about <NUM> and about <NUM>. The example low pressure turbine <NUM> provides the driving power to rotate the fan section <NUM> and therefore the relationship between the number of turbine rotors <NUM> in the low pressure turbine <NUM> and the number of blades in the fan section <NUM> disclose an example gas turbine engine <NUM> with increased power transfer efficiency. <FIG> illustrates an oscillating heat pipe <NUM> that is employed to transfer heat from the gas turbine engine <NUM>. The oscillating heat pipe <NUM> is a passive heat transfer device that can transport heat over a relatively long distance (for example, more than <NUM> inches, or <NUM> centimetres). The oscillating heat pipe <NUM> can be used with or integrated into existing engine structures, as explained below.

The oscillating heat pipe <NUM> includes a series of meandering, hermetically sealed channels <NUM> having a capillary dimension that define a continuous sealed loop. The oscillating heat pipe <NUM> is filled with a working fluid <NUM> in a bi-phase. In one example, the fluid <NUM> is one of glycol, water, alcohol, refrigerant, or a mixture of any of these fluids. The channels <NUM> may be circular, semi-circular, square, or have any other cross-sectional shape. The hydraulic diameter of the channels <NUM> depends on the fluid <NUM> employed. A pressure difference between the fluid <NUM> in an evaporator <NUM> and the fluid <NUM> in a condenser <NUM> drives the fluid <NUM> (which includes liquid slugs <NUM> and vapor plugs <NUM>) to move through the oscillating heat pump <NUM> and between the evaporator <NUM> to the condenser <NUM> to transport heat.

As shown in <FIG>, the evaporator <NUM> accepts heat from a first fluid X (oil or air) located in or near the engine core. The engine core includes the high pressure compressor <NUM>. The evaporator <NUM> accepts heat from oil of an oil system <NUM> that lubricates the bearing system <NUM>, oil from a gearbox or geared architecture <NUM>, or heated bleed air from the high pressure compressor <NUM> (shown in <FIG>). If heat is transferred from the oil system <NUM>, a small liquid-liquid heat exchanger (not shown) can be used to transfer heat from the oil in the oil system <NUM> to the evaporator <NUM> of the oscillating heat pipe <NUM>.

The condenser <NUM> rejects heat to a second fluid Y located outwardly of the engine core. The condenser <NUM> rejects heat to fuel in a fuel system <NUM>. The condenser <NUM> is located on a surface of a fan exit guide vane <NUM>. The channels <NUM> can be formed in the structure of the fan duct <NUM> or the oscillating heat pipe <NUM> can be located inside the fan duct <NUM>. The fan duct <NUM> has a large surface area to allow for the rejection of heat from the first fluid X to the second fluid Y.

When the liquid slugs <NUM> and the vapor plugs <NUM> enter the evaporator <NUM>, the liquid slugs <NUM> and the vapor plugs <NUM> accept heat from the first fluid X in or near the engine core of the gas turbine engine <NUM>, adding heat Q to the fluid <NUM> and cooling the first fluid X. The increase in temperature and pressure of the fluid <NUM> in the evaporator <NUM> forces the liquid slugs <NUM> and the vapor plugs <NUM> to move towards the condenser <NUM>. In the condenser <NUM>, the fluid <NUM> rejects the heat to the second fluid Y located outwardly of the engine core, and the temperature and the pressure of the fluid <NUM> decreases. The vapor plugs <NUM> in the condenser <NUM> do not collapse due to significant surface tension forces, which leads to a restoration force to initiate oscillations. The fluid <NUM> then travels back to the evaporator <NUM>, completing the cycle.

During cruise conditions, the temperature in first fluid X in or near the engine core of the gas turbine engine <NUM> can exceed <NUM>°F (<NUM>), and the temperature of second fluid Y outwardly of the engine core can be approximately -<NUM>°F (-<NUM>). The oscillating heat pipe <NUM> can transfer this heat from the first fluid X to the second fluid Y.

The oscillating heat pipe <NUM> also includes a filling valve <NUM> that allows the fluid <NUM> to be added to the oscillating heat pipe <NUM> and evacuates any air in the channels <NUM> with a vacuum. Once the filling valve <NUM> is closed, the oscillating heat pipe <NUM> is sealed. Alternatively, the filling valve <NUM> may be connected to a pressure regulating system <NUM>. By modulating the pressure of the working fluid <NUM> in the oscillating heat pipe <NUM> with the pressure regulating system <NUM>, the performance can be controlled to compensate for changes in temperature in the evaporator <NUM> and the condenser <NUM>.

By employing the oscillating heat pipe <NUM> to reject heat from the first fluid X to the second fluid Y, it is possible to reduce or eliminate some of the existing heat exchangers of the gas turbine engine <NUM>. As the oscillating heat pipe <NUM> can be incorporated into existing structure of the gas turbine engine <NUM>, there is little or no weight or pressure drop added to the gas turbine engine <NUM>.

The oscillating heat pipe <NUM> provides many benefits. The oscillating heat pipe <NUM> has high effective thermal conductivity, provides fast thermal response and has a low cost. The oscillating heat pipe <NUM> can be made of any material, which allows a coefficient of thermal expansion of the components of the oscillating heat pipe <NUM> to match a coefficient of thermal expansion of the components of the gas turbine engine <NUM>. The oscillating heat pipe <NUM> also has a reduced weight and a reduced volume compared to prior heat exchangers. Additionally, the oscillating heat pipe <NUM> is scalable, orientation-independent, and has a highg tolerance. The oscillating heat pipe <NUM> also functions if temporal and spatial pressure fluctuations are present in the system. A pump is also not needed to move the fluid through the channels <NUM>. Finally, the oscillating heat pipe <NUM> can eliminate any combined weight and pressure drop penalties on the total engine fuel burn associated with current air-oil coolers by about <NUM>%.

Although a gas turbine engine <NUM> with geared architecture <NUM> is described, the oscillating heat pipe <NUM> can be employed in a gas turbine engine without geared architecture.

Claim 1:
A gas turbine engine (<NUM>) comprising:
a high pressure compressor (<NUM>);
an oscillating heat pipe (<NUM>) including an evaporator (<NUM>), a condenser (<NUM>), and a plurality of channels (<NUM>) that define a circuit of a continuous loop through which a working fluid (<NUM>) flows, wherein the condenser (<NUM>) is located on a surface of a fan exit guide vane (<NUM>);
a first fluid (X) that is at least one of oil from an oil system (<NUM>), oil from a gearbox (<NUM>), or bleed air from a high pressure compressor (<NUM>), wherein the first fluid (X) rejects heat to the working fluid (<NUM>) in the evaporator (<NUM>); and
a second fluid (Y) that is at least one of fuel from a fuel system (<NUM>) and air in a fan duct (<NUM>), wherein the working fluid (<NUM>) rejects heat to the fuel in the fuel system (<NUM>) or the air in the fan duct (<NUM>) in the condenser (<NUM>);
wherein, when the second fluid (Y) is air, the fan exit guide vane (<NUM>) is located in the fan duct (<NUM>), and the condenser (<NUM>) rejects heat to the air passing by fan exit guide vanes (<NUM>); and
wherein, when the second fluid (Y) is fuel from a fuel system (<NUM>), the condenser (<NUM>) rejects heat to the fuel in the fuel system (<NUM>).